Beechcraft Be350/350C

KING AIR 350/350C (Model B300/B300C) PRO LINE 21 PILOT TRAINING MANUAL “The best safety device in any aircraft is a well-trained crew.”™

KING AIR 350/350C (Model B300/B300C) PRO LINE 21 PILOT TRAINING MANUAL REVISION 0.2

REVISION 0.2 FlightSafety International, Inc. Marine Air Terminal, LaGuardia Airport Flushing, New York 11371 (718) 565-4100 www.FlightSafety.com

F O R T R A I N I N G P U R P O S E S O N LY

NOTICE The material contained in this training manual is based on information obtained from the aircraft manufacturer’s Pilot Manuals and Maintenance Manuals. It is to be used for familiarization and training purposes only. At the time of printing it contained then-current information. In the event of conflict between data provided herein and that in publications issued by the manufacturer or the FAA, that of the manufacturer or the FAA shall take precedence. We at FlightSafety want you to have the best training possible. We welcome any suggestions you might have for improving this manual or any other aspect of our training program.

F O R T R A I N I N G P U R P O S E S O N LY

Courses for the King Air 300/350C are taught at the following FlightSafety Learning Center:

FlightSafety International Wichita Hawker Beechcraft Learning Center 9720 E. Central Avenue Wichita, KS 67206 Phone: (316) 612-5300 Toll-Free: (800) 488-3747 Fax: (316) 612-5399

Copyright © 2011 by FlightSafety International, Inc. All rights reserved. Printed in the United States of America.

INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES LIST OF EFFECTIVE PAGES Dates of issue for original and changed pages are: Revision............... 0 .............. August 2008 Revision............... .01 ......... October 2009 Revision............... 0.2 .........February 2011

THIS PUBLICATION CONSISTS OF THE FOLLOWING: Page No.

*Revision No.

Cover ...................................................... 0.2 i – vi ........................................................ 0.2 1-i – 1-vi.................................................. 0.2 1-1 – 1-32 ............................................... 0.2 2-i – 2-iv.................................................. 0.2 2-1 – 2-32 ............................................... 0.2 3-i – 3-iv.................................................. 0.2 3-1 – 3-8 ................................................. 0.2 4-i – 4-iv.................................................. 0.2 4-1 – 4-12 ............................................... 0.2 5-i – 5-iv.................................................. 0.2 5-1 – 5-26 ............................................... 0.2 6-i – 6-ii................................................... 0.2 7-i – 7-iv.................................................. 0.2 7-1 – 7-54 ............................................. 0.2 8-i – 8-iv ............................................... 0.2 8-1 – 8-10 ............................................... 0.2 9-i – 9-iv.................................................. 0.2 9-1 – 9-10 ............................................... 0.2 10-i – 10-iv.............................................. 0.2 10-1 – 10-26 ........................................... 0.2 11-i – 11-iv.............................................. 0.2 11-1 – 11-24 ........................................... 0.2 12-i – 12-iv.............................................. 0.2 12-1 – 12-18 ........................................... 0.2

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13-i – 13-ii............................................... 0.2 14-i – 14-iv.............................................. 0.2 14-1 – 14-30 ........................................... 0.2 15-i – 15-iv.............................................. 0.2 15-1 – 15-10 ........................................... 0.2 16-i – 16-xii............................................. 0.2 16-1 – 16-122 ......................................... 0.2 16A-i – 16A-iv ......................................... 0.2 16A-1 – 16A-24 ...................................... 0.2 17-i – 17-iv.............................................. 0.2 17-1 – 17-12 ........................................... 0.2 18-i – 18-ii............................................... 0.2 19-i – 19-iv.............................................. 0.2 19-1 – 19-20 ........................................... 0.2 20-i – 20-iv.............................................. 0.2 20-1 – 20-10 ........................................... 0.2 21-i – 21-iv.............................................. 0.2 21-1 – 21-16 ........................................... 0.2 22-i – 22-iv.............................................. 0.2 22-1 – 22-8 ............................................. 0.2 WA-1 – WA-26........................................ 0.2 APPA-i – APPA-ii .................................... 0.2 APPA-1 – APPA-6................................... 0.2 APPB-1 – APPB-10 ............................... 0.2 ANN-1 – ANN-4 ..................................... 0.2

*Zero in this column indicates an original page.

CONTENTS Chapter 1

AIRCRAFT GENERAL

Chapter 2

ELECTRICAL POWER SYSTEMS

Chapter 3

LIGHTING

Chapter 4

MASTER WARNING SYSTEM

Chapter 5

FUEL SYSTEM

Chapter 6

AUXILIARY POWER UNIT

Chapter 7

POWERPLANT

Chapter 8

FIRE PROTECTION

Chapter 9

PNEUMATICS

Chapter 10

ICE AND RAIN PROTECTION

Chapter 11

AIR CONDITIONING

Chapter 12

PRESSURIZATION

Chapter 13

HYDRAULIC POWER SYSTEMS

Chapter 14

LANDING GEAR AND BRAKES

Chapter 15

FLIGHT CONTROLS

Chapter 16

AVIONICS

Chapter 16A

WIDE AREA AUGMENTATION SYSTEM (WAAS)

Chapter 17

OXYGEN

Chapter 18

MISCELLANEOUS SYSTEMS

Chapter 19

MANEUVERS AND PROCEDURES

Chapter 20

WEIGHT AND BALANCE

Chapter 21

FLIGHT PLANNING AND PERFORMANCE

Chapter 22

CREW RESOURCE MANAGEMENT

WALKAROUND APPENDIX A APPENDIX B ANNUNCIATOR PANELS

1 AIRCRAFT GENERAL

KING AIR 350/350C PRO LINE 21 PILOT TRAINING MANUAL

CHAPTER 1 AIRCRAFT GENERAL CONTENTS Page INTRODUCTION ............................................................................................................... 1-1 GENERAL ........................................................................................................................... 1-2 Configuration................................................................................................................. 1-5 Specifications ................................................................................................................. 1-6 DOORS.................................................................................................................................. 1-8 Airstair Entrance........................................................................................................... 1-8 Emergency Exits ......................................................................................................... 1-10 Cargo Door.................................................................................................................. 1-11 350C Airstair Entrance .............................................................................................. 1-11 FLIGHT DECK ................................................................................................................ 1-12 Seats.............................................................................................................................. 1-12 Instruments/Controls.................................................................................................. 1-13 CABIN FEATURES ......................................................................................................... 1-20 Seats.............................................................................................................................. 1-20 Toilet............................................................................................................................. 1-20 AC Power..................................................................................................................... 1-21 BAGGAGE COMPARTMENT ...................................................................................... 1-21 CONTROL SURFACES................................................................................................... 1-22 GENERAL OPERATING INFORMATION............................................................... 1-23 Preflight Inspection..................................................................................................... 1-23 Tiedown and Securing ................................................................................................ 1-23

FOR TRAINING PURPOSES ONLY

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Taxiing .......................................................................................................................... 1-24 Servicing Data ............................................................................................................. 1-26 LIMITATIONS................................................................................................................... 1-26 Airspeed Limitations.................................................................................................. 1-26 Weight Limits .............................................................................................................. 1-26 Maximum Operating Limits ...................................................................................... 1-28 Maximum Outside Air Temperature Limits............................................................ 1-28 General Limitations.................................................................................................... 1-28 Cracked or Shattered Windshield............................................................................. 1-28 Crack in Side Window (Cockpit or Cabin) ............................................................. 1-29 Miscellaneous Airspeeds............................................................................................ 1-29 QUESTIONS ...................................................................................................................... 1-31

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ILLUSTRATIONS Figure

Title

Page

1-1

King Air 350........................................................................................................... 1-2

1-2

Dual Aft Strakes.................................................................................................... 1-3

1-3

King Air 350 General Arrangement ................................................................... 1-4

1-4

King Air 350 Cabin Seating Arrangement......................................................... 1-5

1-5

King Air 350 Dimensions ..................................................................................... 1-7

1-6

Airstair Door ......................................................................................................... 1-8

1-7

Door Lock .............................................................................................................. 1-8

1-8

Plunger-Out/Plunger-In ........................................................................................ 1-9

1-9

Visual Inspection Ports ......................................................................................... 1-9

1-10

Emergency Exit .................................................................................................. 1-10

1-11

Emergency Exit Placards ................................................................................... 1-10

1-12

Overhead Light Control Panel .......................................................................... 1-14

1-13

Glareshield ........................................................................................................... 1-14

1-14

Left Instrument Panel......................................................................................... 1-15

1-15

Right Instrument Panel ...................................................................................... 1-15

1-16

Center Instrument Panel .................................................................................... 1-16

1-17

Pilot Subpanels .................................................................................................... 1-17

1-18

Copilot Subpanels ............................................................................................... 1-17

1-19

Center Pedestal.................................................................................................... 1-18

1-20

Circuit Breaker Panel—Right Console ............................................................ 1-19

1-21

Fuel Control Panel—Left Console.................................................................... 1-19

1-22

Passenger Seats.................................................................................................... 1-20

1-23

Toilet Seat............................................................................................................. 1-20


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1 AIRCRAFT GENERAL


1-24

Flight Control Locks........................................................................................... 1-22

1-25

Preflight Inspection............................................................................................. 1-23

1-26

Tiedowns............................................................................................................... 1-24

1-27

Turn Radius and Danger Areas......................................................................... 1-25

1-28

Service Data......................................................................................................... 1-27

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CHAPTER 1 AIRCRAFT GENERAL

INTRODUCTION This training manual provides a description of the major airframe and engine systems in the King Air 350 Pro Line 21 aircraft. Information on the cargo (350C) and extended range (350ER) models is also included. This manual is an instructional aid. Its material does not supersede, nor is it meant to substitute for, any of the manufacturer operating manuals. Changes in aircraft appearance or system operation are covered during academic training and subsequent revisions to this manual. This introductory chapter presents an overall view of the aircraft for familiarization. Information includes general specifications and limitations, cabin features, and general cockpit layout.


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1 AIRCRAFT GENERAL


GENERAL The King Air 350 is a high performance pressurized twin-engine turboprop (Figure1-1). The aircraft is equipped for day or night IFR conditions and flight into known icing conditions in and out of small airports within operating limits stated in the Pilot Operating Handbook. FL 381, 383 and subsequent aircraft have the Pro Line 21 avionics package. In late 2 0 0 7, t h e 3 5 0 E R w a s c e r t i f i e d . I t h a s additional nacelle fuel tanks, heavy-weight l a n d i n g g e a r, a n d a m a x i m u m t a ke o ff w e i g h t i n c r e a s e. Th e 3 5 0 E R h a s a n extended range of 2,30 0 nm (4,260 km) and eight hour endurance. The structure is an all aluminum low-wing monoplane with fully cantilevered wings

and a T-tail empennage. The wings are an efficient, high-aspect ratio design. The airfoil provides an excellent combination of low drag for cruise conditions and easy handling for low speed terminal or small airport operations. The NASA-designed winglets further improve performance. All Pro Line 21 aircraft also include dual aft strakes (Figure1-2). The wing/body vortices normally disrupt airflow under the aft fuselage. This creates drag. The strakes eliminate this separation by channeling the vortices and accelerating the air. They are, in effect, pushing the aircraft through the air. The dual strakes eliminate or raise yaw damper limits to increase dispatch reliability. They permit flight with the yaw damper off until 19,0 0 0 ft.

Figure 1-1. King Air 350 (Sheet 1 of 2)

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The faired-oval nacelle on each side of the wing center section houses the engine and l a n d i n g g e a r. Th e n a c e l l e s m a x i m i z e propeller-to-ground clearance, minimize cabin noise, and provide a low drag installation of the powerplants on the wing. The pitot-type intakes and smaller frontal area of the exhaust stacks reduce drag to also boost performance. The distinctive T-tail provides improved aerodynamics, lighter control forces, and a w i d e r c e n t e r- o f - g ra v i t y ra n g e. M o d e l 350ER has an increased rudder area. The fuselage is a conventional monocoque structure with high strength aluminum alloys. The basic cross-sectional cabin is a favorable compromise between passenger comfort and efficient cruise performance. The squared-oval cabin allows passengers to sit comfortably. The floors are flat from side to side for passenger ease in entering and leaving the cabin (Figure 1-3).

Figure 1-2. Dual Aft Strakes

Figure 1-1. King Air 350 (Sheet 2 of 2)


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12

11 10 7 3

9 8

13

6

5 4

2 14 15

1

3

5

16

6

19 23 21 20

24

18 22 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12.

17

Weather Radar Antenna Communications, Navigation and Radar Equipment Outboard Flap Section Ground Escape Hatch Inboard Flap Section Liquid Storage Cabinet Lavatory Privacy Curtain Belted Lavatory Pressurization Safety and Outflow Valves Oxygen Bottle Emergency Locator Transmitter Elevator Trim Tabs

13. 14. 15. 16. 17. 18. 19. 20. 21. 22. 23. 24.

Rudder Trim Tab Baggage Area Airstair Door Aileron Trim Tab Box Section Fuel Tanks Leading Edge Fuel Tanks Auxiliary Fuel Tank Wing Ice Check Light Nacelle Fuel Tank PT6A Turboprop Engine Heated Pitot Mast Landing and Taxi Lights

Figure 1-3. King Air 350 General Arrangement

1-4


CONFIGURATION The King Air 350 is certificated for up to 17 people (15 passengers and 2 crew), but normal corporate configuration is 9 to 11 (Figure 1-4).

In addition to the standard configurations, Beechcraft offers optional items that are available at additional cost and weight. Basic specifications are detailed below. Refer to the appropriate aircraft POH for detailed, up-to-date information.

Figure 1-4. King Air 350 Cabin Seating Arrangement


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SPECIFICATIONS Crew—FAA Certificated ........................ 1 • Except where otherwise prescribed by the appropriate operating re g u l a t i o n s, o n e p i l o t w i t h FA A approved passenger seating configurations of nine or less; or one pilot and one copilot for all other approved configurations. Occupancy—Max. FAA Cert. (with crew) ................................................ 17

Airstair Entrance Door Height (Min) ............................ 51.5 inches Cargo Door Width ...................... 49 inches Cargo Door Height .................... 52 inches Pressure Vessel Volume ................................ 443 cubic feet Potential Cargo area volume ................................ 303 cubic feet

Specific Loadings

Passengers— Normal Configuration .................... 9 to 11

Wing Loading: 48.4 pounds per square foot

Engines—P & W Turboprop, 1050 SHP ................................ 2 PT6A-60A

Power Loading: 7.14 pounds per shaft horsepower.

Propellers—4 Blade, Reversible.................................... 2 Hartzell

Figure 1-5 illustrates the King Air 350 dimensions.

Landing Gear—Retractable, Tricycle, Dual Main Wheels .... Hydraulic Wing Area .................................... 310 sq. ft.

Cabin and Entry Dimensions Cabin Width (Max) .................... 54 inches Cabin Length (Max between pressure bulkheads) .................... 24 feet, 10 inches Cabin Height (Max) .................. 57 inches Airstair Entrance Door Width (Min) .......................... 26.75 inches

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Figure 1-5. King Air 350 Dimensions


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DOORS AIRSTAIR ENTRANCE The cabin entry airstair door is on the left side of the fuselage, just aft of the wing (Figure 1-6). The swing-down door, hinged a t t h e b o tt o m , p ro v i d e s a c o nv e n i e n t stairway for entry and exit. Two of the four steps are movable and automatically fold flat against the door in the closed position. A self-storing platform that automatically folds down over the doorsill when the door opens provides a stepping platform for door seal protection. A plastic encased cable supports the door in the open position. It also provides a handhold and a m e a n s t o c l o s e t h e d o o r f ro m i n s i d e. Additional handhold cable is available as an option.

A hydraulic damper permits the door to lower gradually. Because excessive weight could damage the door attach fitting, no more than one person should be on the airstair door at a time. The door can be locked with a key for security on the ground.

Airstair Locking Mechanism Either one of two vertically staggered handles, one inside and one outside, lock the d o o r . Th e h a n d l e s a r e m e c h a n i c a l l y interconnected. When either is rotated per placard instructions, two bayonet pins on each side of the door and two hooks at the top engage the door frame to secure the door.

Opening the Door A button next to the door handle must be depressed before the handle can be rotated to open the door. As an additional safety measure, a differential pressures e n s i t i v e d i a p h ra g m i s i n t h e re l e a s e button mechanism.

Securing the Door To secure the airstair door inside, rotate the handle clockwise as far as it will go. The release button should pop out. The handle should be pointing down (Figure 1-7).

Figure 1-7. Door Lock Figure 1-6. Airstair Door

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Attempt to rotate the handle counterclockwise without depressing the release button to check security. The handle should not move. Next lift the folded airstep just below the door handle. Ensure that the safety lock is in position around the diaphragm shaft when the handle is in the locked position. To observe this area, depress a red switch near the window. This illuminates a lamp inside the door (Figure 1-8). If the arm is properly positioned around the shaft, proceed to check indication in each of the visual inspection ports near each corner of the door. Ensure the green stripe on the latch bolt is aligned with the black pointer in the visual inspection port (Figure 1-9).

Figure 1-9. Visual Inspection Ports

To check the upper door hook engagement, view the hooks through two inspection openings in the headliner just above the fore and aft upper corners. To illuminate the h o o k e n g a g e m e n t a r e a s, d e p r e s s t h e CABIN DOOR HOOK, OBSV LT SW button between the two inspection openings in the headliner.

PLUNGER-OUT

WARNING Never attempt to unlock or check the security of the door in flight. If the CABIN DOOR annunciator illuminates in flight, or if the pilot has any reason to suspect the door may not be securely locked, instruct all occupants to remain seated with seatbelts fastened. Re d u c e c a b i n p re s s u re t o t h e lowest practical value (considering altitude first). After the aircraft has made a full-stop landing, a crewmember should check the security of the airstair door. Perform the “Cabin Door Annunciator Circuitry Check” in the POH Normal Procedures section prior to the first flight of the day. If any condition specified in this procedure is not met, DO NOT TAKE OFF.

PLUNGER-IN

Figure 1-8. Plunger-Out/Plunger-In


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1 AIRCRAFT GENERAL


EMERGENCY EXITS The emergency exits are on the left and right side of the fuselage at the forward ends of the passenger compartment (Figure 1-10). From inside, release the hatches with the EXIT-PULL pull-down handle. From the outside, a flush-mounted, pullout handle releases the hatches. The nonhinged, plugtype hatches can be removed completely from the frame into the cabin when the latches are released (Figure 1-11). The hatch can be locked so that it cannot be removed or opened from the outside. The hatch is locked when the lock lever inside is in the down or locked position.

When the aircraft is parked, lock the hatch for security. Prior to flight, the lock lever should be in the up or unlocked position to allow removal of the hatch from the outside in an emergency. Removal of the hatch from inside is possible at all times with the EXIT-PULL handle because it is not locked by the lock lever. An exit lock placard on the lock lever can be read when the lever is in the locked position.

EXIT-PULL

INSIDE EMERGENCY EXIT PUSH

1. PULL HANDLE 2. PUSH IN AFTER RELEASE

OUTSIDE

Figure 1-10. Emergency Exit

1-10

Figure 1-11. Emergency Exit Placards


CARGO DOOR A large, swing-up cargo door is hinged at the top to provides access for loading large items. Two handles operate the door lock system. One is in the upper aft area of the door, and the other is in the lower forward area of the door. Two separate access covers must be opened to operate the two handles. There are no lock handles on the outside of the cargo door. It can be opened and closed only from inside the aircraft. To move the upper aft handle out of the locked position, depress the black release button in the handle. Then rotate the yellow handle upward as far as it will go. This movement transmits via cables to two hollow, crescent latches on the forward side and two on the aft side. The latches rotate to release latch posts in the cargo door frame. To move the lower lock handle out of the locked position (forward), lift the orange lock hook from the stud and rotate the handle aft as far as it will go. This movement transmits via linkage to four latch pins on the bottom of the cargo door. The pins move aft to disengage latch lugs at the bottom of the cargo door frame.

CAUTION After unlocking the bottom latch pins, close the forward lock handle access cover. If this cover is left open, it rotates on its hinge until a portion of it extends below the bottom of the cargo door when the cargo door is opened. When the cargo door is subsequently closed, the access cover breaks. To open the cargo door after it is unlocked, push out on the bottom of the door. After the cargo door is manually opened a few feet, gas springs raise the door to the fully open position.

To close the cargo door, pull it down and inboard. The gas springs resists the closing effort until the door is only open a few feet. Then, as the springs move over center, they begin applying a closing force to the door. An inflatable rubber seal around the perimeter of the cargo door seats against the door frame when closed. When the cabin is pressurized, air seeps into the rubber seal through small holes in the outboard side of the seal. The higher the cabin differential pressure, the more the seal inflates. This is a passive seal system and has no mechanical connection to a bleed air source.

WARNING Never attempt to unlock or check the security of the door in flight. If the door annunciator illuminates in flight, or if the pilot has any reason to suspect the door may not be securely locked, instruct all occupants to remain seated with seatbelts fastened. Re d u c e c a b i n p re s s u re t o t h e lowest practical value (considering altitude first). After the aircraft has made a full-stop landing, a crewmember should check the security of the airstair d o o r. Pe r f o r m t h e “ C i r c u i t r y Check” in the POH Normal Procedures section prior to the f i r s t f l i g h t o f t h e d a y. I f a n y condition specified in this procedure is not met, DO NOT TAKE OFF.

350C AIRSTAIR ENTRANCE The airstair door is built into the cargo door. It is hinged at the bottom and swings downward when opened. The stairway is built onto the inboard side. Two of the stairsteps fold flat against the door when it is closed. When the door is opened, a self-storing platform automatically folds down over the door sill to protect the rubber door seal.


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1 AIRCRAFT GENERAL


A hydraulic damper ensures the door swings down slowly when it opens. While the door is open, a plastic-encased cable that serves as a handrail supports the door. Additionally, this cable is used when closing the door from inside. A n i n f l a t a b l e r u b b e r s e a l a ro u n d t h e perimeter seats against the door frame as t h e d o o r i s c l o s e d . Wh e n t h e c a b i n i s pressurized, air seeps into the rubber seal through small holes in the outboard side of the seal. The higher the cabin differential pressure, the more the seal inflates. This is a passive-seal system with no mechanical connection to a bleed air source. The outside door handle can be locked with a k e y, f o r s e c u r i t y o f t h e a i r c r a f t o n the ground.

CAUTION Only one person should be on the airstair door stairway at any one time.

Locking Mechanism Rotating either outside or inside door handle locks the door. The handles move simultaneously. Three hollow, crescent latches on each side of the door rotate to capture or release latch posts in the cargo door to secure the airstair door. When l o c ke d , t h e a i r s t a i r d o o r b e c o m e s a n integral part of the cargo door. Whether unlocking the door from outside or inside, depress and hold the release button adjacent to the door handle before rotating the handle. Inside, rotate the handle counterc l o c k w i s e ; o u t s i d e, ro t a t e t h e h a n d l e clockwise. Unlocking the door is a two-hand operation requiring deliberate action.

1-12

The release button acts as a safety device to prevent accidental opening. As an additional safety measure, a differential-pressuresensitive diaphragm is in the release-button m e c h a n i s m . Th e o u t b o a rd s i d e o f t h e diaphragm is open to atmospheric pressure; the inboard side opens to cabin air pressure. As the cabin-to-atmospheric pressure differential increases, it becomes increasingly difficult to depress the release button because the diaphragm moves inboard when either the outboard or inside release button is depressed.

FLIGHT DECK SEATS The pilot and copilot sit side by side in individual chairs, separated by the control pedestal. The seats are adjustable fore, aft, and vertically with release levers beneath the seats. Depressing the release lever on the side of the seat adjusts the angle of the seat. A button on the lower inboard side of the seat back controls the firmness of the lower seat back for lumbar control. After adjusting the seat back to a comfortable position, move forward on the seat to remove all the weight from the seat back. Hold the button in until the support fully inflates. Release the button and lean back in the seat. If the support is too firm, hold the button in until the desired degree of firmness is obtained. Each seat has seat belts and inertia-type shoulder harnesses. The shoulder harness consists of a Ystrap mounted to an inertia reel in the lower seatback. One strap is w o r n o v e r e a c h s h o u l d e r . Th e s t r a p terminates with a fitting that inserts into a rotary buckle.


Release the shoulder harness straps and inboard lap belt simultaneously by rotating the buckle release 1/8 of a turn in a clockwise direction. The armrests have angular adjustment and vertical stowing. To stow the armrest, release the lever on its forward end and rotate the armrest aft to the vertical position.

Sun Visors Each crewmember has a sun visor. If the visor is stowed, push straight back and allow the visor to rotate down. Move it along the track to desired place. Pivot it out near the windshield or window. Rotate knob clockwise to lock. To change positions, rotate knob counterclockwise to unlock. Then move to desired location and position; relock.

INSTRUMENTS/CONTROLS Due to conventional dual controls, the aircraft can be flown by either pilot. The controls and instruments are arranged for convenient single-pilot operation, or pilot and copilot crew. The instrument panel poster that accompanies this manual illustrates a typical cockpit a r r a n g e m e n t . Th e a n n u n c i a t o r p a n e l chapter at the end of this manual locates specific annunciators and control panels. Each system chapter describes in detail the controls and instruments appropriate to that system. Fi g u r e s 1- 1 2 t h r o u g h 1- 2 1 i l l u s t r a t e each section.

To stow the visor, rotate knob counterclockwise and then move it along the track to recessed area of headliner. Pivot the visor up and press forward until the catch retains the assembly.


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Figure 1-12. Overhead Light Control Panel

Figure 1-13. Glareshield

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Figure 1-14 . Left Instrument Panel

Figure 1-15. Right Instrument Panel


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Figure 1-16. Center Instrument Panel

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Figure 1-17. Pilot Subpanels

Figure 1-18. Copilot Subpanels


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Figure 1-19. Center Pedestal

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Figure 1-20. Circuit Breaker Panel—Right Console

Figure 1-21. Fuel Control Panel—Left Console


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CABIN FEATURES SEATS Passenger seats are installed on continuous tracks mounted on the floor. A placard on the horizontal leg cross brace denotes each seat as FRONT or AFT FACING. All passenger seats have adjustable headrests and shoulder harnesses. The seats are adjustable fore and aft (7 inches [17.8 cm]) and laterally (2 1/2 inches [6.35 cm]). Seat backs may be adjusted for maximum comfort. Some seats may swivel through approximately 45º (Figure 1-22).

Figure 1-22. Passenger Seats

A two-position lever on the forward face of the inboard armrest and a button on the inboard side of the armrest adjust the seats. Moving the lever upward releases the seat for fore and aft and/or lateral movement. Release the lever to lock seat in desired position. Depressing the button adjusts the seatback. Release the button when the back is in the desired position. If no weight is applied to the seatback when the button is depressed, the seatback returns to the upright position. Before takeoff and landing, lateral tracking seats should be in the outboard position, all seatbacks positioned upright, and all headrests fully extended.

1-20

The inboard armrest on each seat can be stowed if desired. Lift the armrest to the full-up position to unlatch the mechanism. Lower the armrest to stow. To use the armrest, simply raise it to the fullup position. Then allow it to settle to the locked position. If the armrest does not lock in the up position, cycle it fully down and back to the up position to reset the locking mechanism.

Foyer Seat A hinged seat cushion on the top of the toilet forms an extra passenger seat when the toilet is not is use (Figure 1-23).

Figure 1-23. Toilet Seat

TOILET On B350 models, the side facing toilet in the foyer faces the airstair door. On B350C models, the forward facing toilet is in the baggage compartment. Raise the hinged lid to access the toilet. If a Monogram electrically flushing toilet is installed, the sliding knife valve should be open at all times except during servicing. Open the cabinet below the toilet to access the knife valve actuator handle.


Relief Tubes A relief tube is in the seat shroud of the side facing toilet (B350) or on the baggage compartment wall forward of the toilet (B350C). An optional relief tube may also be installed in the cockpit and stowed under the pilot seat. A valve lever is on the side of the relief tube horn. When the tube is in use, the lever must be depressed at all times. The relief tubes are for use during flight only.

The inverter also shuts down for an output short circuit. Following a short circuit shutdown, the inverter can be manually reset by cycling the furnishing switch off and on.

Furnishing Switch A switch on the cockpit overhead panel controls the inverter. The two-position switch FURN ON/OFF is standard. An optional switch has the following positions: FURN COFFEE ON/FURN ON/OFF. The inverter operates when the switch is in FURN ON or FURN COFFEE ON position.

AC POWER The aircraft has four AC power outlets to provide 115 VAC for laptop computers. The outlets are on each side of the cabin beneath the cabin tables. Access by lifting the cover placarded 115 VAC. One 115-volt, 60-Hz inverter powers the outlets. The inverter is in the right center section wing just outboard of the nacelle. The left generator bus supplies 28 VDC for the inverter through the INVERTER circuit breaker in the DC Power distribution panel under the center aisle floor. A 115 VAC-5 AMP circuit breaker adjacent to the inverter protects its output. For normal operation, input current to the inverter can vary from approximately 0.5 amperes to approximately 20 amperes depending on the load. The inverter is capable of providing a continuous output of 4 amperes. The total electrical load connected to the four outlets must not exceed 4 amperes. Excess load may cause the inverter input circuit breaker to open.

BAGGAGE COMPARTMENT On Model 350, the entire aft-cabin area aft of the foyer may be used as a baggage compartment. A nylon web restrains loose items. O n M o d e l 3 5 0 C, a s e p a r a t e b a g g a g e compartment is aft of the passenger compartment. A partition separates it from the passenger area. The toilet is on the aft wall of the baggage compartment. A nylon web restrains items. Unless authorized by applicable Departmen of Transportation Regulations, do not carry hazardous material anywhere in the aircraft. D o n o t c a r r y c h i l d re n i n t h e b a g g a g e compartment unless secured in a seat. Secure baggage and other objects with webs to prevent shifting in turbulent air.

The inverter shuts down for input over voltage, under voltage and high internal temperature conditions. It automatically resets when the conditions are corrected.


1-21

1 AIRCRAFT GENERAL


1 AIRCRAFT GENERAL


CONTROL SURFACES The King Air 350 has conventional ailerons and rudder. A T-tail horizontal stabilizer and elevator mounted at the extreme top of the vertical stabilizer. Conventional dual controls in the flight deck operate the cablecontrol surfaces.

Any time the aircraft is parked overnight or in windy conditions, install the rudder g u s t p i n a n d c o n t ro l l o c k s t o p re v e n t damage to the control surfaces and hinges or the controls (Figure 1-24). Dual push rod actuators are installed on all pilot controlled trim tabs.

Figure 1-24. Flight Control Locks

1-22


GENERAL OPERATING INFORMATION PREFLIGHT INSPECTION The preflight inspection procedure has been divided into five areas as shown in Figure 1-25. The inspection begins in the flight compartment, proceeds aft, then moves clockwise around the aircraft.

TIEDOWN AND SECURING When the aircraft is parked overnight o r d u r i n g h i g h w i n d s, i t s h o u l d b e securely moored with protective covers (Figure 1-26). Place wheel chocks fore and aft of the main gear wheels and nose wheel. Using the mooring points, tie the aircraft down with suitable chain or rope. Install the control surface lock. Ensure flaps are up.

Figure 1-25. Preflight Inspection


1-23

1 AIRCRAFT GENERAL


1 AIRCRAFT GENERAL


Secure the propellers to prevent windmilling. This aircraft has free-spinning propellers that could be hazardous if not restrained. Allowing engine gears and bearings to windmill without lubrication is not a good practice. Install the engine inlet cover if there is blowing dust or rain. Before towing the aircraft, release the parking brake (brake handle pushed in) just under the left corner of the subpanel. Remove the rudder gust lockpin from the pinhole in the pilot floorboard. Serious damage to the tires, brakes, and steering l i n k a g e c a n re s u l t i f t h e s e i t e m s a r e not released.

TAXIING The ground turning radii are predicated on the use of partial braking action and differential power. Locking the inside brake can cause tire or strut damage. If the wingtip clears obstacles when turning the aircraft, the tail also clears. Because of the propeller windstream, an area directly to the rear of the engines can be hazardous to persons or parked aircraft when taxiing, turning, and starting t h e e n g i n e s. Wh i l e t h e v e l o c i t i e s a n d temperatures cannot be accurately measured, exercise reasonable care to prevent incidents within these danger areas (Figure 1-27).

PARKING BRAKE

PROPELLER TIEDOWNS

Figure 1-26. Tiedowns

1-24


1 AIRCRAFT GENERAL


Figure 1-27. Turn Radius and Danger Areas


1-25

1 AIRCRAFT GENERAL


SERVICING DATA

Air Min Control Speed (V MCA )

The Handling, Servicing, and Maintenance section of the POH outlines requirements for maintaining the King Air 350 in a condition equal to its original manufacture. This information sets time intervals at which the aircraft should be taken to a Hawker/ Beechcraft Service Center for periodic servicing or preventive mainten a n c e. A l l l i m i t s, p r o c e d u r e s, s a f e t y practices, time limits, servicing, and maintenance requirements contained in the POH are mandatory. This section of the POH also includes a consumable materials chart that lists approved and recommended materials for servicing the aircraft. The servicing data diagram (Figure 1-28) lists and illustrates servicing points and materials required. This chart is for reference only and is always superseded by the POH information.

LIMITATIONS Model 350ER limitations are in parenthesis where applicable.

AIRSPEED LIMITATIONS Maneuvering Speed (V A) 184 (182) KIAS Max Flap Extension/Extended Speeds (V FE ):

Propeller Feathered/ Flaps Up ..................... 94 (10 1) KIAS Propeller Feathered/ Flaps Approach............ 93 (98) KIAS Maximum Operating Speed V MO ............................. 263 (245) KIAS M MO ....................................... 0.58 Mach

Airspeed Indicator Display Red line ............................................ V MCA Solid red bar ........ Impending stallspeed low speed cue DN (white) ...................... Maximum speed permissible with flaps extended beyond approach APP (white) .................. Maximum speed permissible with flaps in approach position Blue line .................................. One-engine inoperative best rate-of-climb speed Solid red bar at top .............. V MO marker

WEIGHT LIMITS Max Ramp Weight .. 15,10 0 (16,600) lbs Max Takeoff Weight 15,000 (16,500) lbs

Approach ............................... 202 KIAS

Max Landing Weight 15,000 (15,675) lbs

Full Down .............................. 158 KIAS

Max Zero Fuel Weight 12,500 (13,000) lbs

Maximum Landing Gear Extended Speed (V LE )................................. 184 (182) KIAS M a x La n d i n g G e a r O p e ra t i n g S p e e d s (V LO )

Max Weight in Baggage Compartment: .......... 550 lbs Max Weight in Wing Lockers ........ 300 lbs

Extension ................... 184 (182) KIAS Retraction .................. 166 (164) KIAS

1-26


1 AIRCRAFT GENERAL


Figure 1-28. Service Data


1-27

1 AIRCRAFT GENERAL


MAXIMUM OPERATING LIMITS Normal Operation ...................... 35,0 0 0 ft Yaw Damp System .... 5,000 or 19,0 0 0 ft (strakes) With Aviation Gasoline: Both Standby Pumps Operative 35,000 ft Either Standby Pump Inoperative ................................ Prohibited Climb without crossfeed capability........................................ 20,0 0 0 ft

MAXIMUM OUTSIDE AIR TEMPERATURE LIMITS Sea Level to 25,000 ft Pressure Altitude ...................... ISA +37 C Above 25,000 ft Pressure Altitude ...................... ISA +31 C

GENERAL LIMITATIONS Acrobatic maneuvers, including spins, are prohibited.

Th e f o l l o w i n g l i m i t a t i o n s a p p l y when continued flight is required with a cracked outer or inner ply of the windshield. • Flight limited to 25 flight hours • Crack(s) must not impair visibility • Crack(s) must not interfere with use of windshield wipers for flights requiring use of wipers • Wi n d s h i e l d a n t i - i c e m u s t b e operational for flights into icing conditions • Following placard must be installed in view of the pilot: MAXIMUM AIRPLANE ALTITUDE IS LIMITED TO 25,0 0 0 FEET. CABIN ∆P MUST BE MAINTAINED BETWEEN 2.0 AND 4.6 PSI DURING FLIGHT Windshields that have cracks in both the inner and outer plies must be replaced prior to the next flight unless a special flight permit is obtained from the local FAA Flight Standards District Office.

Seat back of each occupied aft-facing seat must be in the upright position and headrest fully extended for takeoff and landing. All cargo must be properly secured by an FAA-approved cargo restraint system. Cargo must be arranged to permit free access to all exits and emergency exits.

CRACKED OR SHATTERED WINDSHIELD Windshields with a shattered inner ply have numerous cracks that obstruct forward vision. Small particles or flakes of glass can break free of the windshield and interfere with the crew's vision. These windshields must be replaced prior to the next flight unless a special flight permit is obtained f ro m t h e l o c a l FA A F l i g h t S t a n d a rd s District Office.

1-28


CRACK IN SIDE WINDOW (COCKPIT OR CABIN)

Airspeeds for Safe Operation 15,000 (16,500) Lbs

The following limitations apply when continued flight is required with a cracked outer or inner ply in any side window. These limitations do not apply to minor compression-type chips (clamshell) that may occur on the milled edge of cockpit side windows. Refer to the maintenance manual for the disposition of such chips.

Max Demonstrated Crosswind Component .................................... 20 KIAS

• Limited to 25 flight hours. • Flights must be conducted with cabin depressurized. • Following placard must be installed in clear view of the pilot:

Two-Engine Best Angle-of-Climb (V X ).................................... 125 (135) KIAS Two-Engine Best Rate-of-Climb (V Y ) .................................. 140 (135) KIAS Cruise Climb: • Sea Level to 10,0 0 0 feet 170 KIAS • 10,0 0 0 to 15,0 0 0 feet ...... 160 KIAS • 15,0 0 0 to 20,0 0 0 feet ...... 150 KIAS • 20,0 0 0 to 25,0 0 0 feet ...... 140 KIAS • 25,0 0 0 to 30,0 0 0 feet ...... 130 KIAS

PRESSURIZED FLIGHT IS PROHIBITED DUE TO A CRACKED SlDE WINDOW. CONDUCT FLIGHT WITH THE CABIN PRESSURE SWITCH IN THE DUMP POSITION

• 30,0 0 0 to 35,0 0 0 feet ...... 120 KIAS Turbulent Air Penetration ...... 170 KIAS Intentional One-Engine Inoperative Speed (V SSE ) .............................. 110 (135) KIAS

Overspeed Warning MISCELLANEOUS AIRSPEEDS Emergency Airspeeds 15,000 (16,500) lbs Model 350ER airspeeds are in parenthesis where applicable.

An overspeed warning horn sounds when the airspeed exceeds the barber pole by no more than 6 knots or .0 1 Mach, whichever is less. A test switch on the copilot left subpanel allows the pilot to test the overspeed warning prior to flight.

One-Engine-Inoperative Best Angle-ofClimb (V XSE ) .............................. 125 KIAS One-Engine-Inoperative Best Rate-ofClimb (V YSE ) .............................. 125 KIAS Air Minimum Control Speeds (V MCA ): Flaps Up .................................... 94 KIAS Flaps Approach ....................... 93 KIAS One-Engine-Inoperative Enroute Climb ............................................ 125 KIAS Emergency Descent .................. 184 KIAS Maximum Range Glide ............ 135 KIAS


1-29

1 AIRCRAFT GENERAL


1 AIRCRAFT GENERAL


INTENTIONALLY LEFT BLANK

1-30


QUESTIONS 1. Aircraft equipped with dual strakes require yaw damper operation above _________ feet: A. 13,000 B. 15,000 C. 19,000 D. 20,000

6. Single pilot operations require: A. The pilot to use a headset with a boom microphone. B. A flight attendant. C. Operations not to be conducted under 14 CFR Part 135. D. Operations only during Day VFR.

2. Lateral-tracking seats must be in the full _______ position for _______ A. Outboard; takeoff only. B. Inboard; landing. C. Inboard; takeoff and landing. D. Outboard; takeoff and landing.

7. With appropriate equipment, the kinds of operations allowed: A. Permit flight at night. B. Prohibit flight at night. C. Pe r m i t f l i g h t i n i c e d u r i n g d a y operations only. D. Prohibit flight in ICE during night operations.

3. Illumination of the red master warning annunciator [DOOR UNLOCKED] indicates: A. The emergency escape hatch is open or not secure. B. The airstair door is open or not secure. C. The emergency or airstair door is open or not secure. D. Both the emergency and airstair doors are open or not secure. 4. Th e m a x i m u m a l l o w e d o p e r a t i n g altitude limit is ________ feet. A. 30,000 B. 35,000 C. 37,000 D. 41,000 5. Th e m a x i m u m a l l o w e d o p e r a t i n g temperature limit above 25,000 feet is ISA + ______°C. A. 25 B. 27 C. 31 D. 37

8. Passenger briefing cards are required at one per seat for: A. All operations. B. 14 CFR Part 135 operations. C. 14 CFR Part 135 operations with out a flight attendant. D. Single pilot operations only. 9. V XSE is _______ KIAS. A. 84 B. 125 C. 135 D. 140 10. V MCA for Flaps Approach is ______ KIAS. A. 85 B. 93 C. 94 D. 140


1-31

1 AIRCRAFT GENERAL



CHAPTER 2 ELECTRICAL POWER SYSTEMS

Page INTRODUCTION ............................................................................................................... 2-1 GENERAL ........................................................................................................................... 2-1 COMPONENTS ................................................................................................................... 2-2 Battery ............................................................................................................................ 2-2 Starter/Generators ........................................................................................................ 2-3 Ammeters....................................................................................................................... 2-5 CIRCUIT BREAKERS ...................................................................................................... 2-5 Buses ............................................................................................................................... 2-7 OPERATION ....................................................................................................................... 2-9 Protection ....................................................................................................................... 2-9 Starting ......................................................................................................................... 2-11 Normal Operation....................................................................................................... 2-13 EXTERNAL POWER ...................................................................................................... 2-16 EMERGENCY AND ABNORMAL INDICATIONS ................................................ 2-17 Battery .......................................................................................................................... 2-18 Circuit Breaker Tripped ............................................................................................. 2-19 Generators ................................................................................................................... 2-19 System Distribution Schematics................................................................................ 2-22 CIRCUIT BREAKER LISTING .................................................................................... 2-27 QUESTIONS ...................................................................................................................... 2-31


2-i

2 ELECTRICAL POWER SYSTEMS

CONTENTS


ILLUSTRATIONS Title

Page

2-1

Basic Electrical Symbols....................................................................................... 2-2

2-2

Battery Installation ............................................................................................... 2-2

2-3

Starter/Generator Installation ............................................................................. 2-3

2-4

Pilot Subpanel........................................................................................................ 2-4

2-5

Overhead Light Control and Meter Panel......................................................... 2-5

2-6

Left Circuit Breaker Panel................................................................................... 2-5

2-7

Copilot Sidewall Circuit Breaker Panel ............................................................. 2-6

2-8

King Air 350 Electrical System Component Location ..................................... 2-7

2-9

Electrical System ................................................................................................... 2-8

2-10

BAT Switch ON................................................................................................... 2-11

2-11

Right Engine Start .............................................................................................. 2-12

2-12

Cross Generator Start ........................................................................................ 2-13

2-13

Both Generators On........................................................................................... 2-14

2-14

Both Generators On — Generator Ties Open................................................ 2-15

2-15

External Power.................................................................................................... 2-16

2-16

BAT TIE OPEN.................................................................................................. 2-18

2-17

L/R GEN TIE OPEN......................................................................................... 2-19

2-18

L/R DC GEN Annunciators.............................................................................. 2-19

2-19

Dual Generator Failure...................................................................................... 2-21

2-20

Battery Off........................................................................................................... 2-22

2-21

Right Generator On ........................................................................................... 2-23

2-22

Bus Sense Test with Both Generator On......................................................... 2-24

2-23

Left Generator Bus Isolated ............................................................................. 2-25


2-iii


Figure


2-24

Center Bus Isolated ........................................................................................... 2-26

2-25

Triple-Fed Bus Isolated ..................................................................................... 2-27

TABLES 2 ELECTRICAL POWER SYSTEMS

Table

Title

Page

2-1

King Air 350 Load Management...................................................................... 2-20

2-2

Circuit Breakers.................................................................................................. 2-28

2-iv




CHAPTER 2 ELECTRICAL POWER SYSTEMS

INTRODUCTION A thorough understanding of the aircraft electrical system eases pilot workload in normal operations and prepares him for any electrical malfunctions that may occur. This chapter describes the electrical system components and operations so the pilot can quickly locate switches and circuit breakers for appropriate corrective actions in abnormal and emergency situations.

GENERAL The electrical system is a 28-volt DC system with the negative lead of each power source grounded to the main aircraft structure. Tw o s t a r t e r- g e n e r a t o r s c o n n e c t e d i n parallel and a battery provide the direct current.

An external power receptacle is available for an external power unit to provide electricity while the aircraft is on the ground. Power from these sources is distributed to the individual electrical loads with a multibus system. Each power source electrically


2-1



connects to the distribution system through relays and line contactors. Bus tie relays and individual bus relays interconnect the buses.

COMPONENTS

The electrical system provides maximum protection against loss of electrical power if a ground fault (or short) occurs.

The battery for the King Air 350 is a 42ampere-hour sealed lead acid battery. It is in the right wing center section in an aircooled box (Figure 2-2).

BATTERY

The schematics in this chapter use basic electrical symbols to illustrate the system. (Figure 2-1) provides a key to those symbols. BATTERY

FUSE

CURRENT LIMITER (OR ISOLATION LIMITER) THIS ACTS AS A LARGE, SLOW-BLOW FUSE

DIODE THE DIODE ACTS AS A ONE-WAY "CHECK VALVE" FOR ELECTRICITY. (Triangle points in direction of power flow. Power cannot flow in opposite direction.)

CIRCUIT BREAKER

SWITCH - TYPE CIRCUIT BREAKER

The battery is used for engine starting and as a final redundant power source if both generators fail. To meet specified battery duration times, the battery charge current must be 10 amps or less prior to takeoff. Takeoff with a battery charge current above 10 amps is permitted at the discretion of the pilot.

RELAY OPEN

The BAT switch and BAT BUS switch on the pilot left subpanel control the battery. With both switches in OFF, the battery disconnects from all electrical loads.

RELAY CLOSED

BUS TIE & SENSOR

Figure 2-1. Basic Electrical Symbols

2-2

Figure 2-2. Battery Installation

BAT BUS Switch The BAT BUS switch controls a remote c o n t ro l c i rc u i t b re a ke r i n t h e battery compartment that functions as a battery bus contactor.



When the switch is in the EMER OFF position, the remote control circuit breaker o p e n s t o i s o l a t e t h e b a tt e r y f ro m t h e battery bus.

BAT Switch When the BAT switch is in ON, it closes the battery relay to apply power to the triplefed bus. The battery bus tie closes to apply power to the center bus. In the OFF position, both the battery relay and the battery bus tie relay open to disconnect the battery from all buses except the battery bus.

STARTER/GENERATORS

units (Figure 2-3). The unit is used as a starter to drive the engine during engine start and as a engine-driven generator to provide electrical power. A series starter winding is used during starter operation; a shunt field winding is used during generator operation. The regulated output voltage of the generator is 28.25 (±0.25) volts with a maximum continuous load rating of 30 0 amperes. In addition to the starter/generators, the generator system consists of control switches, generator control units (GCU), line contactors and loadmeters.

Starter Function The center bus provides starter power through a starter relay. A three-position IGNITION AND ENGINE START switch for each engine on the pilot left subpanel controls the operations. Switch positions are ON-OFF-STARTER ONLY.

Th e t w o 2 8 - v o l t , 3 0 0 - a m p e re s t a r t e r / generators are dual-purpose, engine-driven

Figure 2-3. Starter/Generator Installation


2-3


When the switch is in the NORM position, battery power is applied to the battery bus. Because the battery bus powers such items a s e n t r y l i g h t s a n d c l o c k s, t h i s i s t h e normal position.



Actuating a switch to either the STARTER ONLY or ON position supplies a signal to the start relay and generator field sense relay. The start relay energizes the starter. The generator field sense relay disables the shunt field to prevent generator operation during the start cycle. The starter drives the compressor section of the engine through accessory gearing.

Generator Function The generating function is self-exciting and does not require battery power for o p e ra t i o n . I t u s e s g e n e ra t o r re s i d u a l voltage for initial generator buildup.

the GEN RESET position, the generator voltage builds up to 28 volts and the line contactor is open. When the generator switch is released to ON, the line contactor closes.

Generator Control Unit Tw o g e n e ra t o r c o n t ro l u n i t s ( G C U s ) control generator operation. The GCU below the center aisle floor makes constant v o l t a g e a va i l a b l e t o t h e b u s e s d u r i n g variations in engine speed and electrical load requirements. The GCUs provide the following functions: • Voltage regulation/line contactor control

GEN Switch The L GEN and R GEN switches in the pilot left subpanel are under the MASTER SWITCH gang bar (Figure 2-4). Switch positions are GEN RESET–ON–OFF.

• Overvoltage/overexcitation protection • Paralleling/load sharing • Reverse-current protection

Placing the switch momentarily in GEN RESET and then releasing to the ON position brings the generators on-line. In

• Cross-generator-start current limiting

Figure 2-4. Pilot Subpanel

2-4



AMMETERS Left and right loadmeters on the overhead meter panel display the load on each generator (Figure 2-5).

Voltage on each bus may also be monitored on the voltmeter with the VOLTMETER BUS SELECT switch adjacent to the voltmeter. Selector positions include EXT PWR, CTR , L GEN, R GEN, TPL FED, BAT. Move the selector switch to appropriate position and then read the voltage on the adjacent loadmeter.

CIRCUIT BREAKERS DC power is distributed to the various systems via circuit breakers that protect most of the components in the aircraft. Two of these circuit breaker panels are in the cockpit. Each of the circuit breakers has its amperage rating printed on it. The smaller breaker panel is to the left of the pilot below the fuel management panel (Figure 2-6). The larger circuit breaker panel is on the copilot sidewall (Figure 2-7).

Figure 2-5. Overhead Light Control and Meter Panel

Figure 2-6. Left Circuit Breaker Panel


2-5


Each of these protection features is discussed in detail in the Operation portion of this chapter.


A color-coded ring around each circuit breaker indicates the bus to which the circuit breaker connects. The triple-fed bus and battery bus circuit breakers are colorcoded yellow; left generator bus circuit breakers are blue; right generator bus circuit breakers are green; and the standby bus circuit breakers are red. 2 ELECTRICAL POWER SYSTEMS

Circuit breaker switches on the pilot right subpanel protect components such as exterior lighting and ice protection e q u i p m e n t . Th e s e s w i t c h e s h a v e t h e amperage rating stamped on the end of the switch. A typical listings of all buses and circuit breakers is at the end of this chapter.

Procedures for handling tripped circuit breakers and other related electrical system warnings are in the Emergency and Abnormal Procedures section of the Pilot’s Operating Handbook. As a general rule if a nonessential circuit breaker trips in flight, do not reset it. Resetting a tripped breaker could cause further damage to the component or system. If an essential system circuit breaker such as an avionics breaker trips, let it cool and then reset it. If it fails to reset, do not attempt to reset it again. Take corrective action according to the procedures in the appropriate section of the POH.

Figure 2-7. Copilot Sidewall Circuit Breaker Panel

2-6



• Ba tt e r y b u s — Ba tt e r y t h ro u g h a remote control circuit breaker

BUSES Electrical loads are divided among the buses. Equipment on the buses is arranged so that all items with duplicate functions (such as right and left landing lights) connect to different buses (Figure 2-8).

• Left and right generator bus—Left and right generators

• Center bus—Both generator buses and battery

In normal operation, all buses are automatically tied into a single-loop system where all sources supply power through individual protective devices.

The generator buses connect to the center bus with the left and right bus tie relays. The battery connects through the battery bus tie, which closes when the BAT switch is in

Buses and main power sources are the following:

LEGEND ABBREVIATIONS USED L = LEFT R = RIGHT B = BATTERY BT = BUS TIE LC = LINE CONTACTOR SB = SUB BUS SR = STARTER RELAY BB = BATTERY BUS DFB = DUAL FED BUS

EPR = EXTERNAL POWER RELAY STR/GEN = STARTER GENERATOR GEN CONT = GENERATOR CONTROL EXT PWR = EXTERNAL POWER CTR BUS = CENTER BUS RG = RIGHT GENERATOR LG = LEFT GENERATOR RCCB = REMOTE CONTROL CIRCUIT BREAKER

STR/ GEN

STR/ GEN

L L C

L G B U S

R L C

L S R

L B T

DUAL FED BUS

L S B

CTR BUS

BATT BUS

R S R

TRIPLE FED BUS

R B T

R S B

R G B U S

EXT PWR

RCCB EPR B R

B B T

BBS BATTERY BS GEN GEN CONT CONT

Figure 2-8. King Air 350 Electrical System Component Location


2-7


• Triple-fed bus—Battery and both generators buses


comes on line. If the battery is the only power source on line, both generator bus ties open to isolate the left and right generator buses from the battery. Equipment that remains operational during battery only operations has a white ring around the control switch.

ON. The battery is then available for center bus loads or recharging (Figure 2-9).

GEN TIES Switch In the OPEN position, both the left and right bus tie relays open to isolate both generator buses from the center bus. The NORM position allows automatic closure of the left and right bus tie relays when either generator or external power LEFT STARTER RELAY

TO GENERATOR FIELD

BUS SENSE GEN TIES RESET MAN CLOSE SPRING LOADED TO CENTER

STARTER/ GENERATOR

TEST L DC GEN

LOAD METER

L GEN TIE OPEN

LEFT STARTER RELAY

SPRING LOADED FROM MAN CLOSE TO CENTER LEVER LOCK OUT OF CENTER

OPEN R GEN TIE OPEN

LEFT GENERATOR SWITCH

LEFT LINE CONTACTOR

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN

LOAD METER

BAT TIE OPEN

RIGHT GENERATOR SWITCH

RIGHT LINE CONTACTOR

MAN TIES CLOSE

GENERATOR CONTROL

GENERATOR CONTROL 275

275 250

H E D

LEFT GEN BUS

ESIS BATT 5

ESIS BATT BUS BAT BUS SWITCH NORMAL

60

250

H E D

CENTER BUS

RIGHT GEN BUS RIGHT GENERATOR BUS TIE

EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

EXT PWR RELAY

BATTERY BUS TIE

BAT BUS

RCCB FROM BAT BUS

DUAL-FED BUS 275


Momentarily placing the MAN TIE switch in the MAN CLOSE position during battery operation closes both generator bus ties. The battery then powers the generator buses.

HED

BATTERY AMMETER BATT SWITCH

BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-9. Electrical System

2-8


60


BUS SENSE Switch Bus current sensors sense current to each generator bus from the center bus and current to the center bus from the battery. If either generator bus sensor senses a high current condition, it opens the corresponding generator bus tie to isolate the bus. If the battery bus sensor senses a high battery discharge current, it opens the battery bus tie to isolate the battery. The battery bus sensor does not work during engine starts and landing gear operation. Th e BU S S E N S E s w i t c h o n t h e p i l o t subpanel resets and tests the sensors. The RESET position resets the bus current sensors if they have tripped because of a test or an actual high current condition. The momentary TEST position opens the bus current sensors for the generator bus ties and battery ties. The yellow caution L and R GEN TIE OPEN and BAT TIE OPEN annunciators illuminate.

The AVIONICS MASTER circuit breaker in the right circuit breaker panel provides the power to control the avionics relays. If the avionics buses become disconnected as a result of a control circuit fault, the AVIONICS MASTER circuit breaker can be pulled to restore power.

OPERATION Th e D C p o w e r d i s t r i b u t i o n s y s t e m i s commonly called a triple-fed system because most buses receive power from three sources. The triple-fed bus powers many systems. Th r e e s o u r c e s ( g e n e r a t o r b u s e s a n d battery) power the triple-fed bus. It only receives power; it does not transfer electricity from one part of a system to another. That is a function of the the center bus. Because of this arrangement, a backup power source is available to most of the aircraft electrical systems. In normal operation, all buses are automatically tied together so that the battery and two generators collectively supply power through individual protective devices.

PROTECTION

AVIONICS MASTER POWER Switch Th re e a v i o n i c s b u s e s a re e l e c t r i c a l l y connected to the main distribution system through avionics relays. The AVIONICS MASTER POWER switch on the pilot subpanel controls these relays. The ON position opens the control circuit so the relays are in their normally closed p o s i t i o n s. Th i s s u p p l i e s p o w e r t o t h e avionics buses. The OFF position applies control power to the relays to disconnect the avionics buses.

The bus tie system protects the electrical system from excessively high current flow. The abilities to isolate a bus and load shed are equally important protective features. The system automatically removes excess loads (generator buses) when the power source is reduced to battery only. When both generators fail, the generator bus ties open to shed generator bus loads. The battery continues to power the center, triple-fed, and battery buses. If necessary, use the GEN TIE switch to manually close the generator ties. This restores power to the generator buses.


2-9


The green advisory MAN TIES CLOSE annunciator illuminates to indicate the generator bus ties have been manually closed during battery operation.


When load shedding occurs in flight, land as soon as practical unless the situation can be remedied and at least one generator brought back online. Refer to the Abnormal Indications discussion in this section and emergency procedures section of the POH for more details. 2 ELECTRICAL POWER SYSTEMS

GCU Protection Voltage Regulation/Line Contactor Control

detect which generator is producing excessive voltage output and attempting to a b s o r b a l l e l e c t r i c a l l o a d s. Th e G C U overexcitation circuit disconnects that generator from the electrical system. The overexcitation portion of the GCU activates if generator voltage increases without control, but does not reach an overvoltage condition. If the generator field reaches the limitation value, this circuitry removes the affected generator from the bus.

The generators are normally regulated to 28.25 (±.25) VDC. When the GEN switch i s h e l d i n R E S ET, g e n e ra t o r re s i d u a l voltage is applied through the GCU to the generator shunt field. This causes generator output voltage to rise.

Paralleling/Load Sharing

When the switch is released to ON, the 28volt regulator circuit takes over. It controls the generator shunt field to maintain a c o n s t a n t o u t p u t v o l t a g e. Th e v o l t a g e regulator circuit varies shunt field excitation to maintain a constant 28-volt output from the generator for all rated conditions of generator speed, load, and temperature.

The paralleling circuits sense the interpole winding voltages of both generators to provide an indication of the load. The voltage regulator circuits are then biased up or down as required to increase or decrease generator loads until both generators share the load equally. The GCUs balance loads to within 10%.

When the GEN switch is released to ON, the GCU enables the line contactor control circuit. The GCU compares generator output voltage with aircraft bus voltage. If output voltage is within 0.5 volts of bus voltage, the GCU closes the line contactor to connect the generator to the aircraft bus. It also closes both generator ties to connect the center bus and generator buses. The generator can now recharge the aircraft battery and power all aircraft electrical loads.

Reverse-Current Protection

When a generator fails or is turned off, the GCU opens the line contactor to isolate the inoperative generator from its bus.

Overvoltage/Overexcitation The GCU provides overvoltage protection to prevent excessive generator voltage to electrical equipment. If a generator output reaches the maximum allowable 32-volts, the overexcitation circuits of the GCU 2-10

The paralleling circuit averages the output of both generators to equalize load levels. This feature is operative when both generators are online.

When a generator becomes underexcited or cannot maintain bus voltage for some reason (i.e., low generator speed during engine shutdown), it begins to draw current (rev e rse curre nt) from the e lectrical system. The GCU senses reverse current by comparing generator output voltage to generator bus voltage. When bus voltage exceeds output voltage, the GCU opens the line contactor to protect the generator.

Cross-Generator Start Current Limiting Wh e n t h e I G N I T I O N A N D E N G I N E START switch on the second engine is activated to ON during a cross-generator start, a signal from the switch is applied to the GCU of the operating generator.



STARTING

Th i s a c t i va t e s t h e c ro s s - s t a r t c u r re n t limiting circuit to limit output of the operating generator to no more than 400 amps. This protects the 250-amp current limiter on the operating generator side. When a starter is selected, the bus tie sensors are disabled to prevent them from opening their respective bus tie relays. When using STARTER ONLY to motor the engine, the same functions occur.

• Through the battery relay to the triple-fed bus • Through the battery bus tie relay to the center bus • To b o t h s t a r t e r re l a y s t o p e r m i t starting either engine

LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR

TEST L DC GEN

LOAD METER

LEVER LOCK OUT OF CENTER

OPEN

L GEN TIE OPEN

LEFT STARTER RELAY

SPRING LOADED FROM MAN CLOSE TO CENTER

R GEN TIE OPEN

STARTER/ GENERATOR

R DC GEN

LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR

TO GENERATOR FIELD



MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

5


60


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY

BATTERY BUS TIE

60

RCCB FROM BAT BUS

DUAL-FED BUS 275

ESIS BATT

250

H E D

CENTER BUS

HED


BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-10. BAT Switch ON FOR TRAINING PURPOSES ONLY

2-11


When BAT switch is turned to ON (Figure 2-10), the battery relay and battery bus tie re l a y s c l o s e. Ba tt e r y p o w e r i s ro u t e d as follows:


Normally one engine is started on battery power alone; the second engine uses a cross-generator start (Figure 2-12).

Without generator or external power, neither generator bus is powered because the generator bus ties are normally open when only battery power is available.

LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR

TEST L DC GEN

LOAD METER

L GEN TIE OPEN

LEFT STARTER RELAY


OPEN R GEN TIE OPEN


LEFT LINE CONTACTOR

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN

LOAD METER

BAT TIE OPEN



MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

ESIS BATT 5


60

250

H E D

CENTER BUS


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY

BATTERY BUS TIE RCCB

FROM BAT BUS DUAL-FED BUS 275


The operating GCU limits its generator output to no more than 40 0 amps during a cross-generator start. This ensures that the 250-amp current limiter on the operating generator side does not open due to transient surges.

The starter relay connects the battery to the starter/generator during engine starts. With one engine running and its generator online (Figure 2-11), the opposite engine can be started with power from the battery and operating generator channeled through the starter relay. This is called a crossgenerator start.

HED


BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-11. Right Engine Start

2-12


60


In addition, while a starter is selected, the bus tie sensors are disabled to prevent them from opening their respective bus tie relays.

NOTE Th e a b o v e l i m i t a t i o n i s o n l y applicable when the starter is driving the engine, not when the engine is driving the starter.

CAUTION

LEFT STARTER RELAY

TO GENERATOR FIELD

After either engine has been started and its generator switch has been moved to RESET, the GCU brings the generator up


STARTER/ GENERATOR

TEST L DC GEN

LOAD METER


OPEN

L GEN TIE OPEN

LEFT STARTER RELAY


R GEN TIE OPEN

STARTER/ GENERATOR

R DC GEN

LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR

TO GENERATOR FIELD



MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

5


60


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY

BATTERY BUS TIE

60

RCCB FROM BAT BUS

DUAL-FED BUS 275

ESIS BATT

250

H E D

CENTER BUS

BATTERY AMMETER

HED

BATT SWITCH

BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-12. Cross Generator Start


2-13


NORMAL OPERATION

Do not exceed the starter motor operating time limits of 30 seconds ON, five minutes off, 30 seconds ON, five minutes off, 30 seconds ON, then 30 minutes off.


center bus, electricity flows to the battery through the battery bus tie and to the left generator bus through the left generator bus tie and 250-amp current limiter. Power is also fed to the triple-fed bus from the right generator bus.

to normal system voltage. Releasing the spring-loaded GEN switch to the center ON position closes the generator line contactor. This powers the the generator bus and closes both generator ties automatically (the green MAN TIES CLOSED annunciator extinguishes if the generator ties have been manually closed). This action distributes power through the right 250-amp current limiter and generator bus tie relay to the center bus. From the LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR

TEST L DC GEN

LOAD METER

L GEN TIE OPEN

LEFT STARTER RELAY


OPEN R GEN TIE OPEN


LEFT LINE CONTACTOR

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN

LOAD METER

BAT TIE OPEN



MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

ESIS BATT 5


60

250

H E D

CENTER BUS


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY




When both generators are operating, each generator directly feeds its own generator bus which, in turn, feeds the center bus, triple-fed bus, battery bus, and battery, if it is discharged (Figure 2-13).

HED


BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-13. Both Generators On

2-14


60


LEFT STARTER RELAY

TO GENERATOR FIELD

BUS SENSE GEN TIES RESET MAN CLOSE

TEST

STARTER/ GENERATOR L DC GEN

LOAD METER

STARTER/ GENERATOR

R GEN TIE OPEN

R DC GEN

LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR

TO GENERATOR FIELD


OPEN

L GEN TIE OPEN

LEFT STARTER RELAY


SPRING LOADED TO CENTER



MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

ESIS BATT 5


(Figure 2-14) depicts the system with the generator ties open.

The center bus ties the generator bus and battery together. The triple-fed bus is powered (or fed) from the battery and each generator bus through 60-amp limiters and through diodes that provide fault isolation protection between the power sources.


60

250

H E D

CENTER BUS


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY BATTERY BUS TIE

60

RCCB

FROM BAT BUS 275

DUAL-FED BUS HED


BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-14. Both Generators On - Generator Ties Open


2-15


2-15). It is recommended that the battery be online (BAT switch in ON) whenever the external power is in use.

EXTERNAL POWER The external power receptacle under the right wing outboard of the nacelle facilitates connecting a 28 VDC external power unit to the aircraft electrical system. An EXT PWR control switch in the pilot left subpanel controls the external power relay. The external power relay closes when the switch is in the ON position (Figure LEFT STARTER RELAY

TO GENERATOR FIELD


TEST

STARTER/ GENERATOR L DC GEN

LOAD METER

L GEN TIE OPEN

R GEN TIE OPEN

LEFT LINE CONTACTOR

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN

BAT TIE OPEN


LEFT STARTER RELAY


OPEN

LOAD METER RIGHT GENERATOR SWITCH


EXT PWR

MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

ESIS BATT 5


60

250

H E D

CENTER BUS


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

EXT PWR RELAY

BATTERY BUS TIE

BAT BUS

RCCB FROM BAT BUS

DUAL-FED BUS 275


Before selecting the ON position of the EXT PWR switch, verify the external power voltage is within acceptable limits (28.0 – 28 4 volts). Turn the VOLTMETER BUS SELECT switch in the overhead panel to the EXT PWR position and read the voltage.

HED


BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-15. External Power

2-16


60


The overvoltage protection circuit opens the external power relay to electrically disconnect the external power from the aircraft if an over voltage occurs. After an overvoltage disconnection occurs, turn the EXT PWR switch to off to reset the overvoltage circuit. The yellow caution EXT PWR annunciator illuminates to indicate the state of the external power. Steady illumination indicates that the external power is electrically connected and supplying power to the aircraft electrical system. A flashing EXT PWR annunciator indicates that an external power plug is connected to the aircraft, but the external power output voltage is low or the external power is electrically disconnected from the aircraft electrical system. Correct either condition to prevent depleting the battery. Observe the following precautions when using an external power source.

CAUTION Th e r e c o m m e n d e d m i n i m u m indicated battery voltage prior to connecting external power is 23 volts; however, never connect an e x t e r n a l p o w e r s o u rc e t o t h e aircraft unless a battery indicating a charge of at least 20 volts is in the aircraft. If the battery voltage is less than 2 0 v o l t s, t h e b a t t e r y m u s t b e re c h a rg e d o r re p l a c e d w i t h a battery indicating at least 20 volts before connecting external power. Use only an external power source fitted with an AN-type plug.

Voltage is required to energize the avionics master power relays to remove power from the avionics equipment. Therefore, never apply external power to the aircraft without first applying battery voltage. The battery may be damaged if exposed to voltages higher than 30 volts for extended periods of time. Refer to the Normal Procedures section of the POH for using external power.

EMERGENCY AND ABNORMAL INDICATIONS Electrical fires are covered in the Emergency Procedures section of the P i l o t ’s O p e ra t i n g H a n d b o o k ( P O H ) . Emergency and abnormal in-flight situations are described in the Emergency and Abnormal Procedures section of the POH. Generator and battery irregularities are described there under Electrical System Failures. G e n e ra t o r b u s a n d c e n t e r b u s f a u l t s / malfunctions are normally associated with illumination of the corresponding bus tie a n n u n c i a t o r s. A t r i p l e - f e d b u s f a u l t / malfunction does not have an associated b u s t i e i n d i c a t i o n . I n a l l c a s e s, a b u s problem should be investigated by referencing other bus indications, such as loss of related equipment, and by checking bus voltages with the VOLTMETER BUS SELECT and/or loadmeter indications. Complete bus loss is a highly unlikely situation. The multi-bus electrical system has protection devices that normally isolate a fault with minimum equipment loss.


2-17


Reverse polarity protection and overvoltage protection are provided. The reverse polarity protection circuit prevents the external power relay from closing if the external power polarity is different than the aircraft electrical system.


Battery malfunctions are extremely rare. There have been a few cases in aircraft with similar installations where malfunctions have occurred; however, the battery monitoring system has provided sufficient warning to the pilot for timely corrective action to be completed before the situation could deteriorate to a more serious condition. 2 ELECTRICAL POWER SYSTEMS

Th e re a re t w o re a s o n s f o r t h i s s a f e t y margin. First, battery malfunctions are historically slow to develop and can be identified early with the charge monitoring system. Second, sufficient cooling is provided to the battery in flight to diminish the likelihood of serious damage to the aircraft or its electrical system. Therefore, when identified early, a battery malfunction will not deteriorate into a serious condition as long as th e p i l o t complies with the proper procedures as outlined in the POH.

BATTERY BAT TIE OPEN Annunciator Illuminated If the BAT TIE OPEN annunciator is illuminated, the battery bus tie relay is open (Figure 2-16). This indicates a possible center bus fault/malfunction. Check the center bus for proper voltage indication.

If it is within normal limits (24 to 28 volts), attempt to reset the bus tie by momentarily actuating the BUS SENSE switch to RESET. If this is successful, a transient spike in the electrical system tripped the sensor and opened the battery tie relay. If this procedure was unsuccessful, there is a probable malfunction within the battery bus tie circuitry that cannot be reset at this time. If center bus voltage is 0, there is a possible center bus fault. Open the generator bus tie relays by moving the GEN TIES switch to OPEN. Check to ensure that both GEN TIE OPEN annunciators illuminate. Pull the LANDING GEAR RELAY circuit breaker on the pilot right subpanel.

WARNING If electrical power is applied to the landing gear hydraulic pump motor relay with the center bus shorted, damage may occur to the electrical system.

NOTE It will not be possible to charge the battery and the landing gear will have to be manually extended. With the center bus unpowered, turn the air conditioning system off prior to landing. Refer to the POH for procedural details.

CIRCUIT BREAKER TRIPPED Figure 2-16. BAT TIE OPEN

If one of the circuit breakers in the cockpit trip and it is a nonessential circuit, do not reset in flight. If it is an essential circuit, push to reset. If the circuit breaker trips again, do not reset. In this situation, a connected item may be inoperative.

2-18



L or R GEN TIE OPEN Annunciator Illuminated

L or R DC GEN Annunciator Illuminated (Generator Inoperative)

Illumination of a L or R GEN TIE OPEN annunciator indicates a left or right generator tie is open (Figure 2-17). This signals a possible generator bus fault/malfunction.

If the L or R DC GEN yellow caution annunciator (Figure 2-18) illuminates during flight, verify with the associated loadmeter. Then push the corresponding generator switch to RESET and after one second to ON.

Figure 2-17. L/R GEN TIE OPEN

Figure 2-18. L/R DC GEN Annunciators

Monitor the corresponding loadmeter. If it is less than 100 percent and a normal indication, move the BUS SENSE switch to RESET. If it is greater than 100 percent or an abnormal indication, turn the appropriate generator OFF and monitor the opposite loadmeter not to exceed 10 0 percent. If the generator bus tie relay does not reset, monitor the loadmeters.

If the generator does not reset, turn it off and rely on the other generator. Monitor the loadmeter so the load on the remaining generator does not exceed 100 percent. Tu r n o f f a l l n o n e s s e n t i a l e l e c t r i c a l equipment as necessary.

In this situation, the generator paralleling circuit is open. The generators, therefore, are not sharing the aircraft electrical loads equally and the loadmeters need not be within 10% of each other. Monitor the loadmeters to ensure that neither exceeds 100% total load. Refer to the POH for procedural details.


2-19


GENERATORS


Load Management for Dual Generator Failure


The equipment listed in Table 2-1 remains operable after a dual generator failure (Figure 2-19). With only the equipment o p e ra t i n g l i s t e d a s c o n t i n u o u s i n t h e OPERATING TIME column, the battery duration is approximately 30 minutes (based upon a 50-amp load and a 75% battery capacity). Use of the equipment with prescribed operating times reduces battery duration by the approximate times listed. Multiple usage of this equipment is additive. Table 2-1. KING AIR 350 LOAD MANAGEMENT

Table 2-1. KING AIR 350 LOAD MANAGEMENT (Cont)

EQUIPMENT

OPERATING TIME (MIN)

REDUCTION IN MAIN BATTERY DURATION (MIN)

Annunciator Panel

As Required

-----

Instrument Indirect/ Emergency Lights

Continuous

-----

Cabin Lights

5

2

Ice Light

5

0.5

Continuous

----

Taxi Lights

1

0.5

Digital OAT

Continuous

----

Fuel Quantity Indicator

Continuous

----

5

1

Beacon Lights

OPERATING TIME (MIN)

REDUCTION IN MAIN BATTERY DURATION (MIN)

Air-driven Attitude Gyro

Continuous

----

Standby Attitude Gyro

Continuous

None*

Left Bleed Air Valve

Continuous

----

Inverter 1

Continuous

-----

Pressurization Control

Continuous

----

3

0.5 ----

Continuous

-----

Nav 1

Continuous

----

Cabin Temperature Control

Continuous

Pilot Audio

RMI 2

Continuous

----

Engine Ignition

0.5

0.1

Pilot Altimeter

Continuous

----

Surface Deice

6 cycles

0.1

Pilot ADI (Electromechanical)

Continuous

----

Left and Right Main Engine Anti-ice

Single Operation

0.1

Pilot EADI (EFIS)

Not Operational

----

Manual Prop Deice

3

3

Pilot HSI (Electromechanical)

Continuous

----

Windshield Wiper

1

0.1

Pilot EHSI (EFIS)

Not Operational

----

Left Pitot Heat

Continuous

----

Landing Gear

0.5

Continuous

----

Single Operation

EQUIPMENT

Comm 1 Xmit

Turn & Slip Indicator

Single Standby Fuel Pump

* Optional equipment. Powered by Auxiliary battery.

2-20



NOTE WARNING

Equipment that remains operable is designated with a WHITE C I RC L E a r o u n d t h e c o n t r o l switch. Attitude reference will depend upon the specific instrument panel equipment. Refer to t h e LOA D M A NAG E M E N T table to determine which attitude instruments will remain operable with a dual generator failure. BUS SENSE GEN TIES RESET MAN CLOSE SPRING LOADED TO CENTER

STARTER/ GENERATOR

TEST L DC GEN

LOAD METER


OPEN

L GEN TIE OPEN

LEFT STARTER RELAY


R GEN TIE OPEN

LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN



MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

5


60

250

H E D

CENTER BUS


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY

BATTERY BUS TIE

60

RCCB FROM BAT BUS

DUAL-FED BUS 275

ESIS BATT


LEFT STARTER RELAY

TO GENERATOR FIELD

Do not place the GEN TIES s w i t c h i n t h e M A N C LO S E position. This action reconnects the left and right generator bus l o a d s a n d s e v e re l y l i m i t s t h e battery duration.

HED


BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-19. Dual Generator Failure


2-21


• Figure 2-22: Bus Sense Test with Both Generators On

SYSTEM DISTRIBUTION SCHEMATICS

• Fi g u re 2 - 2 3 : L e f t G e n e ra t o r B u s Isolated

The following pages present a variety of electrical scenarios. These include the following:

• Figure 2-24: Center Bus Isolated • Figure 2-25: Triple-Fed Bus Isolated

• Figure 2-20: Battery Off

LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR

TEST L DC GEN

LOAD METER

LEFT STARTER RELAY


OPEN

L GEN TIE OPEN

R GEN TIE OPEN

LEFT LINE CONTACTOR

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN

LOAD METER

BAT TIE OPEN




MAN TIES CLOSE

GENERATOR CONTROL


275 250

H E D

LEFT GEN BUS

ESIS BATT 5


60

250

H E D

CENTER BUS


EXT PWR RECEPTACLE

BAT BUS CONTROL .5A

EXT PWR RELAY

BATTERY BUS TIE

BAT BUS

RCCB FROM BAT BUS

DUAL-FED BUS 275


• Figure 2-21: Right Generator On

HED


BATTERY

BATTERY RELAY

60

20A

TRIPLE FED BUS

Figure 2-20. Battery Off

2-22


60

LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR


TEST L DC GEN

LOAD METER


OPEN R GEN TIE OPEN

L GEN TIE OPEN

LEFT STARTER RELAY


LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR

250

H E D

5


60

250

H E D

CENTER BUS


BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY

BATTERY BUS TIE

EXT PWR RECEPTACLE 60

RCCB FROM BAT BUS

DUAL-FED BUS 275

ESIS BATT

GENERATOR CONTROL

275

275

LEFT GEN BUS



MAN TIES CLOSE

GENERATOR CONTROL

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN



BATTERY AMMETER

HED

BATT SWITCH BATTERY

BATTERY RELAY

60

20A TRIPLE FED BUS

Figure 2-21. Right Generator On


2-23



LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR


TEST L DC GEN

LOAD METER


OPEN R GEN TIE OPEN

L GEN TIE OPEN

LEFT STARTER RELAY


STARTER/ GENERATOR

R DC GEN

LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR


ESIS BATT

5

BAT BUS SWITCH NORMAL

60

GENERATOR CONTROL

275 H E D

ESIS BATT BUS



MAN TIES CLOSE

275

LEFT GEN BUS

TO GENERATOR FIELD

250

H E D

CENTER BUS


BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY


FROM BAT BUS 275

DUAL-FED BUS BATTERY AMMETER

HED

BATT SWITCH BATTERY

BATTERY RELAY

60

20A TRIPLE FED BUS

Figure 2-22. Bus Sense Test with Both Generator On

2-24



LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR


TEST L DC GEN

LOAD METER


OPEN R GEN TIE OPEN

L GEN TIE OPEN

LEFT STARTER RELAY


LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR


5

MAN TIES CLOSE

H E D


60

GENERATOR CONTROL

275

250

H E D

CENTER BUS


BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY

BATTERY BUS TIE


RCCB FROM BAT BUS

DUAL-FED BUS 275

ESIS BATT



275

LEFT GEN BUS

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN



BATTERY AMMETER

HED

BATT SWITCH BATTERY

BATTERY RELAY

60

20A TRIPLE FED BUS

Figure 2-23. Left Generator Bus Isolated


2-25



STARTER/ GENERATOR


TEST L DC GEN

LOAD METER

LEFT STARTER RELAY


OPEN R GEN TIE OPEN

L GEN TIE OPEN

LEFT LINE CONTACTOR

LOAD METER

250

ESIS BATT

5

H E D


60

GENERATOR CONTROL

275

275

LEFT GEN BUS



MAN TIES CLOSE

GENERATOR CONTROL

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN

BAT TIE OPEN


250

H E D

CENTER BUS


BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY




LEFT STARTER RELAY

TO GENERATOR FIELD

BATTERY AMMETER

HED

BATT SWITCH BATTERY

BATTERY RELAY

60

20A TRIPLE FED BUS

Figure 2-24. Center Bus Isolated

2-26




CIRCUIT BREAKER LISTING

LEFT STARTER RELAY

TO GENERATOR FIELD


STARTER/ GENERATOR


TEST L DC GEN

LOAD METER


OPEN R GEN TIE OPEN

L GEN TIE OPEN

LEFT STARTER RELAY


LOAD METER

BAT TIE OPEN


LEFT LINE CONTACTOR

250

H E D

5


60

250

H E D

CENTER BUS


BAT BUS CONTROL .5A

BAT BUS

EXT PWR RELAY

BATTERY BUS TIE


RCCB FROM BAT BUS

DUAL-FED BUS 275

ESIS BATT

GENERATOR CONTROL

275

275

LEFT GEN BUS



MAN TIES CLOSE

GENERATOR CONTROL

TO GENERATOR FIELD

STARTER/ GENERATOR

R DC GEN


The Table 2-2 provides a handy reference of the buses and their circuit breakers.

BATTERY AMMETER

HED

BATT SWITCH BATTERY

BATTERY RELAY

60

20A TRIPLE FED BUS

Figure 2-25. Triple-Fed Bus Isolated


2-27


Table 2-2. CIRCUIT BREAKERS LEFT GENERATOR BUS AVIONICS

ENVIRONMENTAL

FURNISHINGS

Pilot Cntl Instr Lights

Pilot PFD Heater

Bleed Air Control, R

Cigar Lighter

Pilot PFD & DCP Lights

DBU

Blower,Cabin Fwd

Furnishings Master Control

Plt Flt Instr & Side Pnl Lights

Radiant Heat (Cargo Door)

Inverter (Cabin Outlets)

Tail Flood Lights


ELECTRICAL Bus Tie Power

Refreshment Bar ESIS

LIGHTS

WEATHER

ESIS Battery Charge

FGP Lights

Engine Anti-Ice, Standby, L

ENGINES

Landing, L

Fuel Vent, L

Chip Detector, L

FLIGHT CONTROLS

MFD & RTU Lights

Pilot Wdshld Anti-Ice Contro

DCU 1

Flap Ind & Control

Nav Lights

Pilot Wdshld Anti-Ice Pwr

EDC 1

Flap Motor

No Smk, FSB, & Baggage

Prop Deice, Auto

ESIS BATTERY BUS ESIS Bus Control

ESIS Disp

ESIS Lights

HDG Snsr

CENTER BUS ELECTRICAL

Test Jack

LANDING GEAR

Bus Tie Control

ENVIRONMENTAL

Landing Gear Motor

Taxi Lights

Bus Tie Indicator

Condenser Blower

LIGHTS

WEATHER

Bus Tie Power

Elec Heat

Beacon Lights

Man Prop Deice, L & R

Ice Lights

RIGHT GENERATOR BUS AVIONICS

ENVIRONMENTAL

LIGHTS

WEATHER

EGPWS

Air Cond Clutch

CDU 1/CDU 2

Brake Deice (opt.)

MFD Heater

Blower, Cabin Aft

Copilot PFD & DCP Lts

Copilot Windshield Anti-Ice

ELECTRICAL

Blower,Cockpit

Copilot PFD & DCP Lts

Engine Anti-Ice, Stby, R

Bus Tie Power, R Gen

Copilot Flight Instr Lts

Fuel Vent, R

ENGINES

Landing, R

Pitot Heat, R

Chip Detector, R

FLIGHT CONTROLS

Pedestal Light Control

Stall Warn Heat

DCU 2

Pitch Trim

Reading Lights

Window Defog

EDC 2

Recognition Lights

Prop Gov Test

FURNISHINGS

Strobe Lights

Prop Sync

Toilet

Subpnl, Ovhd, Cons Lts

BATTERY BUS ELECTRICAL

Bat Relay

Gnd Com

Avionics

Bat Bus Cont

Gnd Heat

DUAL-FED BUS ENGINES Eng Fire Ext, L

2-28

Eng Fire Ext, R

LIGHTS Cabin Entry Lts



Table 2-2. CIRCUIT BREAKERS (Cont) TRIPLE-FED BUS Gen Reset

FLIGHT INSTRUMENTS

WARNINGS/ANNUNCIATORS

ENGINES

Outside Air Temp

Annunciator Ind

Aural Warn

Autofeather

FUEL

Annunciator Power

Avionics Master

DCU 1 and 2 Secondary Aux Fuel XFR & Warn, L & R

Avionics Annunciator

Cabin Audio

Fire Detect, L & R

Crossfeed

Bleed Air Warning, L & R

* CCP

Ignitor Power, L & R

Firewall Valve, L & R

Ldg Gear Ind

DC Converter 2

Oil Press, L & R

Fuel Press Warn, L & R

Ldg Gear Warn

FGC 1 Servo

Start Control, L & R

Fuel Qty, L & R

Oil Press Warn, L & R

FGC 2 Servo

Torque Meter, L & R

Fuel Qty Warn, L & R

Stall Warn

IAPS, L & R

ENVIRONMENTAL

Stby Pump, L & R

MFD

Bleed Air Control, L

LANDING GEAR

WEATHER

Pilot Audio

Cabin Alt High

LDG Gear Control Control

Eng Anti-Ice, Main, L & R

Pilot Audio Control

Cabin Diff Press


AVIONICS AHC 2 Secondary

Manual Prop Deice Cont

Voice Rcdr

Oxygen Control

LIGHTS

Pitot Heat, L

ELECTRICAL

Press Control

Cabin Lights

Surface Deice

Bus Tie Power

Temp Control

Instrument Indirect Lights

Wshd Wiper

LEFT GENERATOR AVIONICS BUS AVIONICS

* FSU

HF COM (opt.)

CDU 1

* FSU FAN

Radar

TEL

DIALER

GPS 1

SELCAL (opt.)

ENVIRONMENTAL

DME 1

HF ANT (opt.)

TCAS

Nose Equipment Cooling

TRIPLE-FED AVIONICS BUS AVIONICS

AHC 1 Secondary

ADC 1 AHC 1

DC Converter 1

Pilot PFD

ATC 1

NAV 1

RTU

COM 1

Pilot DCP

RIGHT GENERATOR AVIONICS BUS AVIONICS

Copilot Audio

GPS 2 (opt.)

ADC 2

Copilot Audio Control

NAV 2

AHC 2

Copilot DCP

Radio Altimeter

ATC 2

Copilot PFD

* XM WX (opt.)

CDU 2 (opt.)

Copilot PFD Heater

ENVIRONMENTAL

COM 2

* CMU (opt.)

Flight Instr Pnl Cooling

* COM 3 (opt.)

DME 2 (opt.)

IEC

* If Installed


2-29




2-30



1. During a battery start, prior to selecting ON with the IGNITION AND ENGINE START switch and before starting the second engine, the DC percent loadmeter should read approximately _______ percent or less. A. 50 B. 55 C. 65 D. 75 2. The minimum battery voltage required for an external power start is _______ volts. A. 17 B. 18 C. 20 D. 23 3. Control switches which are operable during a dual generator failure are indicated by ______________ the switch. A. A white circle around B. Th e a b s e n c e o f a w h i t e c i r c l e around C. A number engraved on the tip of D. The absence of a number engraved on the tip of

5. The external power cart will be set to _______ volts and be capable of generating a minimum of 10 0 0 amps momentarily and 30 0 amps continuously. A. 20.0 – 20.4 B. 24.0 – 20.4 C. 28.0 – 28.4 D. 29.0 – 29.4 6. The maximum sustained generator load at 30,0 0 0 feet is _______ percent. A. 65 B. 70 C. 95 D. 10 0 7. The first immediate action item for a DUAL GENERATOR FAILURE is: A. Generators ....RESET, THEN ON B. ECS Mode .................................OFF C. Instrument Emergency Lights (if requied) ..................................ON D. Non-essential Equipment .....OFF

4. A generator bus tie will open automatically to protect the electrical system from a malfunction when excessive current is sensed on _____________________ bus. A. The center B. The off-side generator C. The same-side generator D. On either generator


2-31


QUESTIONS


CHAPTER 3 LIGHTING CONTENTS Page INTRODUCTION ............................................................................................................... 3-1 INTERNAL LIGHTING.................................................................................................... 3-1 Cockpit ........................................................................................................................... 3-1 Cabin Lighting ............................................................................................................... 3-2 EXTERIOR LIGHTING.................................................................................................... 3-4

Wing Ice Lights.............................................................................................................. 3-5 Anti-collision/Strobe Lights......................................................................................... 3-6 Navigation Lights .......................................................................................................... 3-6 Recognition Lights ........................................................................................................ 3-6 Floodlights...................................................................................................................... 3-6 QUESTIONS ........................................................................................................................ 3-7


3-i

3 LIGHTING

Landing/Taxi Lights ...................................................................................................... 3-5



Title

Page

Cabin Light Control Panel ................................................................................... 3-2

3-2

Threshold Light Switch......................................................................................... 3-3

3-3

Baggage Compartment Light Switch.................................................................. 3-3

3-4

Exterior Lights Control........................................................................................ 3-4

3-5

Landing/Taxi Lights .............................................................................................. 3-5

3-6

Wing Ice Light ....................................................................................................... 3-5

3-7

Anti-Collision/Strobe Light ................................................................................. 3-6

3 LIGHTING

3-1


3-iii


3 LIGHTING

CHAPTER 3 LIGHTING

INTRODUCTION The King Air 350 lighting system consists of cockpit-controlled interior and exterior lights. Interior lights are in the cockpit and passenger cabin. They also include entry and exit threshold lights and baggage area lights. Exterior lights consist of lights for navigation and identification. The aircraft is also equipped with emergency lights.

INTERNAL LIGHTING COCKPIT The overhead panel contains a functional arrangement of all lighting controls for the cockpit (Figure 3-1). The controls are easily accessible to the pilot and copilot. A master

switch turns all the lights on. Each light group then has its own BRT–OFF rheostat for individual adjustment.


3-1


3 LIGHTING

Figure 3-1. Cabin Light Control Panel

The top row of rheostats are the PILOT and C O P I LOT F LO O D l i g h t s. Th e c e n t e r rheostat INSTR INDIRECT is for the center instrument panel.

Annunciator Adjustment

The next row of rheostats, left to right, include the following:

Emergency Lighting

• PILOT PNL • PILOT DISPLAYS • OVHD PED & SUBPANEL, INSTR

To the right of the top row is an adjustment pushbutton for the cockpit annunciators.

An INSTRUMENT EMERG LTS switch is to the right of the electrical gages at the base of the overhead panel. If the normal cockpit lighting is not working, this switch powers the lights from the dual-fed bus.

• SIDE PANEL • COPILOT DISPLAYS

CABIN LIGHTING

• COPILOT INSTR PANL

A three-position CABIN LIGHTS switch controls the indirect fluorescent cabin lights. The switch has three positions: BRIGHT-DIM-OFF.

The MASTER PANEL LIGHTS switch at the extreme left of the overhead panel controls all the cockpit lights.

3-2



A t w o - s e c t i o n ( + i n c re a s e / – d e c re a s e ) switch in the center of the cabin headliner midway between the exit signs can control light intensity.

Wh e n t h e a i r s t a i r d o o r i s c l o s e d a n d latched, all the lights controlled by this switch are extinguished. The dual-fed bus powers these lights.

If the CABIN LIGHTS switch is placed in either the BRIGHT or OFF position, the dim switch is overridden. When the CABIN LIGHTS master switch is on, passengers may turn the individual reading lights along the top of the cabin on or off with a switch in the sidewall tables.

Baggage Compartment Two reading lights in the headliner illuminate the aft compartment when the three-position BAGGAGE switch is placed in BAGGAGE. The switch is just inside the airstair door aft of the door frame (Figure 3-3). The INDIRECT position of this switch is operable only when the triple-fed bus is powered. The BAGGAGE position connects directly to the dual-fed bus.

3 LIGHTING

When the cockpit CABIN LIGHTS switch is moved from the OFF position to the DIM position, the cabin indirect lights illuminate in the full bright mode. Dim control is enabled. The intensity of the cabin indirect lights may then be changed by momentarily touching the appropriate section of the headliner switch.

Threshold Light A threshold light is forward of the airstair door at floor level. In addition, two aisle lights at floor level are on both sides of the spar cover. A switch adjacent to the threshold light turns these lights on and off (Figure 3-2). This switch also activates the exterior entry light under the left wing center section and the lights under each step on the airstair door.

Figure 3-2. Threshold Light Switch

Figure 3-3. Baggage Compartment Light Switch


3-3


Seat Belt-No Smoking Signs A switch to the right of the CABIN LIGHTS activates the no-smoking/fastenseat-belt signs in the cabin. Accompanying chimes also sound. The switch has three positions: NO SMK FSB–OFF–FSB. No smoking configurations have the same switch even though the NO SMK position is inoperative.

Exit Lights Two exit lights are in the center of the cabin headliner. One is in the forward cabin between the emergency exit; the second is in the aft cabin at the airstair door.

position, the light illuminates. It extinguishes when the switch is momentarily placed in the OFF-RESET position. An internal switch automatically activates the internal light source if rapid deceleration is sensed.

EXTERIOR LIGHTING The pilot right subpanel contains switches for the exterior lights (Figure 3-4). These include the following: • Left and right landing lights

3 LIGHTING

Each light has two light sources. During normal operation, the aircraft’s electrical systems power one light source. During abnormal conditions, internal batteries power the other source.

• Taxi light • Wing ice lights • Anti-collision/strobe lights • Navigation lights

A three-position rocker switch spring-loaded to the center OFF position controls the internal light source. When the switch is m o m e n t a r i l y p l a c e d i n t h e O N - T E ST

• Recognition lights • Flood lights

Figure 3-4. Exterior Lights Control

3-4



LANDING/TAXI LIGHTS

WING ICE LIGHTS

Two landing lights and one taxi light are on the nose landing gear (Figure 3-5).

Wing ice lights are on the outboard side of each nacelle (Figure 3-6). They illuminate the wing leading edges so the flight crew can determine ice buildup.

Th e LA N D I N G s w i t c h e s L E F T a n d RIGHT control these through 10-amp circuit breakers. The TAXI switch controls the 15-amp circuit breaker for the taxi light.

The ICE circuit-breaker switch on the pilot right sub-panel controls these lights.

3 LIGHTING

Use the wing ice lights as required during night flight to check for wing ice accumulation. Because these lights operate at a high temperature, do not use for prolonged periods while the aircraft is on the ground.

Figure 3-6. Wing Ice Light

Figure 3-5. Landing/Taxi Lights


3-5


ANTI-COLLISION/STROBE LIGHTS The standard anti-collision lights consist of an upper light in the upper empennage and a lower light on the bottom of the fuselage. The BEACON switch controls the 10-amp circuit breaker. An optional highintensity anti-collision light may be in each wing tip along with a flashtube and socket assembly in the tail. The pulsating strobe lights provide a means of recognition for the aircraft during night fights. The STROBE switch controls a 5amp circuit breaker.

3 LIGHTING

Each strobe has a separate power supply. The wing strobe light power source is just outboard of each nacelle near the wing leading edge. The wing tips incorporate the navigation, recognition, and strobe light systems (Figure 3-7).

NAVIGATION LIGHTS The navigation light consists of two red lights in the left wing tip, two green lights in the right wing tip, and a light in the aft end of the empennage. The NAV switch controls the 7.5 amp circuit breaker for these lights. If optional high intensity anti-collision lights are installed, the tail navigation light is a halogen lamp in the same assembly as the flash tube.

RECOGNITION LIGHTS The optional recognition lights are in each wing tip just forward and inboard of the strobe lights. They are focused to the front and outboard of the aircraft. The RECOG switch controls a 7.5 amp circuit breaker.

FLOODLIGHTS Tail floodlights are part of the horizontal stabilizers. They illuminate both sides of the vertical stabilizer. A flush-mounted floodlight forward of the flaps in the bottom of the left wing illuminates the area around the airstair door. This is for the convenience of the passengers at night.

Figure 3-7. Anti-Collision/Strobe Light

The separate power supply for the tail strobe light is within the tail fairing just forward of the tail strobe light. The power supply units are wired to operate the strobe lights wing tips and tail synchronously.

3-6

It is controlled by the threshold light switch just inside the door on the forward doorframe. It extinguishes automatically whenever the cabin door is closed and latched.



QUESTIONS 1. Selecting the landing light switches on will illuminate both landing: A. Lights if the gear is extended. B. A n d t a x i l i g h t s i f t h e g e a r i s extended. C. Lights regardless of gear position. D. And taxi lights regardless of gear position.

3 LIGHTING

2. Both wing ice lights are required to be operable during flight during _______ operations. A. VFR night B. IFR day C. IFR night D. Icing conditions 3. The EXIT signs automatically illuminate during normal flight operations when: A. Battery power is lost. B. Generated power is lost. C. Rapid acceleration is sensed. D. Rapid deceleration is sensed.


3-7


CHAPTER 4 MASTER WARNING SYSTEM CONTENTS Page INTRODUCTION ............................................................................................................... 4-1 GENERAL ........................................................................................................................... 4-1 ANNUNCIATOR SYSTEM............................................................................................... 4-2 Master Warning and Warning Annunciators ............................................................. 4-3 Master Caution and Caution Annunciators .............................................................. 4-4 Advisory and Status Annunciators ..................................................................................... 4-4 Dimming......................................................................................................................... 4-4 Testing............................................................................................................................. 4-5 ANNUNCIATOR DESCRIPTIONS................................................................................ 4-6

4 MASTER WARNING SYSTEM

QUESTIONS ...................................................................................................................... 4-11


4-i



Title

Page

4-1

Master Warning System........................................................................................ 4-2

4-2

Master Warning and Master Caution Flashers.................................................. 4-3

4-3

Warning Annunciators.......................................................................................... 4-3

4-4

Caution/Advisory/Status Annunciator Panel..................................................... 4-4

4-5

Annunciator Lamp Replacement........................................................................ 4-5

TABLES Table

Title

Page

King Air 350 Warning Annunciators .................................................................. 4-6

4-2

King Air 350 Caution Annunciators ................................................................... 4-7

4-3

King Air 350 Advisory Annunciators ................................................................. 4-9

4-4

King Air 350 Status Annunciators .................................................................... 4-10


4-1


4-iii


INTRODUCTION This chapter presents a description of the warning system on the King Air 350. The warning system includes flashing annunciators to alert the crew of a problem and a series of warning, caution, advisory, and status annunciators. The description of the annunciator panels includes an explanation for the illumination of each annunciator.

GENERAL Warning and caution annunciators are the first indication of trouble or malfunction in a system or component of the aircraft. Crewmembers should be completely f a m i l i a r w i t h t h e s e a n n u n c i a t o r s. Fo r

warning and caution annunciators, the crew should also know the action required to c o r re c t t h e p ro b l e m o r c o p e w i t h t h e situation until the problem can be corrected or a safe landing can be made.


4-1


CHAPTER 4 MASTER WARNING SYSTEM


ANNUNCIATOR SYSTEM Th e a n n u n c i a t o r s y s t e m ( Fi g u re 4 - 1 ) consists of flashers, a red annunciator warning panel centrally located in the glareshield, and a caution/advisory/status annunciator panel on the center subpanel.

A red MASTER WARNING flasher is in the glareshield in front of the pilot. A second red flasher is in front of the copilot. A yellow MASTER CAUTION flasher is just inboard of the MASTER WARNING flasher on ea ch side of the cockpit.

MASTER WARNING AND CAUTION FLASHERS

WARNING ANNUNCIATORS

PRESS TO TEST

4 MASTER WARNING SYSTEM CAUTION/ADVISORY/STATUS ANNUNCIATORS

Figure 4-1. Master Warning System

4-2



Whenever a fault condition covered by the annunciator system occurs, a signal illuminates the appropriate annunciator.

An illuminated lens in the warning annunciator panel remains on until the fault is corrected.

A PRESS TO TEST switch is immediately to the right of the warning annunciator panel.

Th e M AS T E R WA R N I N G f l a s h e r s, however, can be extinguished even if the fault is not corrected. Depress the face of either flasher to extinguish the light and reset the circuit. If an additional warning annunciator illuminates, the MASTER WARNING flashers re-activate.

MASTER WARNING AND WARNING ANNUNCIATORS If a fault requires the immediate attention and reaction of the pilot, both MASTER WARNING annunciators begin flashing (Figure 4-2). The appropriate annunciator in the warning annunciator panel also illuminates (Figure 4-3).

When a warning fault is corrected, the affected annunciator extinguishes. The flashers continue flashing until one of them is depressed.


Figure 4-2. Master Warning and Master Caution Flashers

KING AIR 350

Figure 4-3. Warning Annunciators


4-3


MASTER CAUTION AND CAUTION ANNUNCIATORS

These annunciators extinguish when the condition indicated by the illuminated lens changes.

If a fault requires the pilot’s attention but not his immediate reaction, the appropriate yellow caution annunciator illuminates (Figure 4-4). Both MASTER CAUTION annunciators also begin flashing.

The green advisory annunciators confirm that a pilot-initiated operation occurred (e.g., the tail deice has been turned on).

Depress the face of either flashing MASTER CAUTION to extinguish the light and reset the circuit. Subsequently, if any other caution annunciator illuminates, the MASTER CAUTION flashers are activated again. An illuminated caution annunciator remains on until the fault condition is corrected. One of the MASTER CAUTION annunciators must be depressed to extinguish the flashers.

ADVISORY AND STATUS ANNUNCIATORS The annunciator panel also contains green advisory and white status annunciators. There are no master flashers associated w i t h t h e g re e n o r w h i t e a n n u n c i a t o r s because they indicate functional situations that do not demand immediate attention or reaction.

Two green AUTOFEATHER annunciators adjacent to the torquemeters on the instrument panel function the same even though they are not on the caution/ advisory/status annunciator panel. The white status annunciators indicate a condition that can be normal or abnormal and may or may not have been initiated by the pilot (e.g., notice that the N 1 flow to the air conditioning system is too low).

DIMMING The red and yellow flashers, annunciator panels,firewall fuel valve pushbuttons, landing gear handle lights, and gear position lights have a BRIGHT and a DIM mode of variable intensity as sensed by a photoelectric cell.


Figure 4-4. Caution/Advisory/Status Annunciator Panel

4-4



The DIM mode is selected automatically whenever all of the following conditions are met: • Generator on line • Pilot or copilot overhead floodlights OFF • MASTER PANEL LIGHTS switch ON • Pilot flight lights ON • Ambient light level in the cockpit (as sensed by a photoelectric cell in the overhead light control panel) is below a preset value

TESTING Test the lamps in the annunciator system before every flight and any time the integrity of a lamp is in question. Depress the PRESS TO TEST button to the right of the warning annunciator panel i n t h e g l a r e s h i e l d . A l l t h e M AS T E R WARNING and MASTER CAUTION lamps should flash. In addition, all the annunciators on the warning annunciator panel and the caution/advisory/ status a n n u n c i a t o r p a n e l s h o u l d i l l u m i n a t e. Replace any lamp that fails to illuminate (Figure 4-5).

Unless all these conditions are met, the BRIGHT mode is selected automatically.


The annunciators in the fire extinguisher pushbuttons do not have a DIM mode. A rheostat on the instrument panel dims the AUTOFEATHER annunciators.

Figure 4-5. Annunciator Lamp Replacement


4-5


Lamp Replacement To replace a lamp, depress the center of the annunciator with your finger. Release your finger. The lens pops out slightly. P u l l t h e a n n u n c i a t o r f ro m t h e p a n e l . Remove the lamp from the rear of the annunciator. Replace the failed bulb with a spare lamp from an unused annunciator.

ANNUNCIATOR DESCRIPTIONS Tables 4-1 through 4-4 list all the warning, caution, advisory, and status annunciators. The cause for illumination is included beside each annunciator.

Replace the annunciator. Depress until it locks in place.

Table 4-1. KING AIR 350 WARNING ANNUNCIATORS ANNUNCIATOR DOOR UNLOCKED

L FUEL PRES LO


4-6

CAUSE FOR ILLUMINATION AIRSTAIR DOOR OR CARGO DOOR IS OPEN OR NOT SECURE

FUEL PRESSURE FAILURE ON THE LEFT SIDE

CABIN ALT HI

CABIN PRESSURE ALTITUDE EXCEEDS 1,200 FEET

CABIN DIFF HI

CABIN DIFFERENTIAL PRSSURE EXCEEDS 6.9 PSI

R FUEL PRES LO

FUEL PRESSURE FAILURE ON THE RIGHT SIDE

L OIL PRES LO

LOW OIL PRESSURE LEFT ENGINE

R OIL PRES LO

LOW OIL PRESSURE RIGHT ENGINE

L BLEED FAIL

MELTED OR FAILED PLASTIC LEFT BLEED-AIR FAILURE WARNING LINE

R BLEED FAIL

MELTED OR FAILED PLASTIC RIGHT BLEED-AIR FAILURE WARNING LINE



Table 4-2. KING AIR 350 CAUTION ANNUNCIATORS

L DC GEN

CAUSE FOR ILLUMINATION LEFT GENERATOR IS OFF LINE

L GEN TIE OPEN

LEFT GENERATOR BUS IS ISOLATED FROM THE CENTER BUS

HYD FLUID LOW

HYDRAULIC FLUID IN THE POWER PACK IS LOW

RVS NOT READY

PROPELLER LEVERS ARE NOT IN THE HIGH-RPM, LOW-PITCH POSITION WITH THE LANDING GEAR EXTENDED

R GEN TIE OPEN

RIGHT GENERATOR BUS IS ISOLATED FROM THE CENTER BUS

R DC GEN

L CHIP DETECT L NO FUEL XFR

BAT TIE OPEN

DUCT OVERTEMP

RIGHT GENERATOR IS OFF LINE

METAL CONTAMINATION IN LEFT ENGINE OIL IS DETECTED

NO LEFT AUXILIARY FUEL TRANSFER

BATTERY IS ISOLATED FROM THE GENERATOR BUSES

DUCT AIR IS TOO HOT

R NO FUEL XFR

NO RIGHT AUXILIARY FUEL TRANSFER

R CHIP DETECT

METAL CONTAMINATION IN RIGHT ENGINE OIL IS DETECTED

L ENG ICE FAIL

LEFT ENGINE SELECTED ANTI-ICE SYSTEM IS INOPERATIVE

L FUEL QTY

ELEC HEAT ON

EXT PWR

R FUEL QTY


ANNUNCIATOR

LEFT FUEL QUANTITY—LESS THAN 30 MINUTES REMAINING AT MAXIMUM CONTINUOUS POWER

UNCOMMANDED OPERATION OF ELECTRIC HEAT (FL-544 AND SUBSEQUENT)

EXTERNAL POWER CONNECTOR IS PLUGGED IN

RIGHT FUEL QUANTITY—LESS THAN 30 MINUTES REMAINING AT MAXIMUM CONTINUOUS POWER


4-7


Table 4-2. KING AIR 350 CAUTION ANNUNCIATORS (Cont) ANNUNCIATOR

CAUSE FOR ILLUMINATION

R ENG ICE FAIL

RIGHT ENGINE SELECTED ANTI-ICE SYSTEM IS INOPERATIVE

L BL AIR OFF

LEFT BLEED-AIR VALVE SWITCH IS NOT OPEN

AUTOFTHER OFF

AUTOFEATHER SWITCH IS NOT ARMED, AND LANDING GEAR IS EXTENDED

OXY NOT ARMED

OXYGEN ARMING HANDLE HAS NOT BEEN PULLED, OR SYSTEM FAILED TO CHARGE

RUD BOOST OFF

RUDDER BOOST SWITCH IS OFF

R BL AIR OFF

RIGHT BLEED-AIR VALVE SWITCH IS NOT OPEN

L PITOT HEAT

LEFT PITOT HEAT IS INOPERATVE OR SWITCH IS IN THE OFF POSITION

PROP GND SOL

R PITOT HEAT

ONE OR BOTH GROUND IDLE LOW-PITCH-STOP SOLENOIDS ARE POWERED BY 28 VOLTS RIGHT PITOT HEAT IS INOPERATVE OR SWITCH IS IN THE OFF POSITION


4-8



Table 4-3. KING AIR 350 ADVISORY ANNUNCIATORS ANNUNCIATOR L IGNITION ON

L ENG ANTI-ICE

FUEL CROSSFEED

CAUSE FOR ILLUMINATION LEFT IGNITION AND ENGINE START SWITCH IS ON, OR LEFT AUTOIGNITION SYSTEM IS ARMED WITH LEFT ENGINE TORQUE BELOW 17% LEFT ENGINE ANTI-ICE VANES ARE IN POSITION FOR ICING CONDITIONS

FUEL CROSSFEED IS SELECTED

R ENG ANTI-ICE

RIGHT ENGINE ANTI-ICE VANES ARE IN POSITION FOR ICING CONDITIONS

R IGNITION ON

RIGHT IGNITION AND ENGINE START SWITCH IS ON, OR LEFT AUTOIGNITION SYSTEM IS ARMED WITH LEFT ENGINE TORQUE BELOW 17%

L BK DEICE ON MAN TIES CLOSE

R BK DEICE ON

TAIL DEICE

WING SURFACE DEICE SYSTEM IS IN OPERATION

LEFT BRAKE DEICE SYSTEM IS IN OPERATION

MANUALLY CLOSED GENERATOR BUS TIES

RIGHT BRAKE DEICE SYSTEM IS IN OPERATION

HORIZONTAL STABILIZER SURFACE DEICE SYSTEM IS IN OPERATION


WING DEICE


4-9


Table 4-4. KING AIR 350 STATUS ANNUNCIATORS ANNUNCIATOR L PROP PITCH

CAUSE FOR ILLUMINATION LEFT PROPELLER IS BELOW THE FLIGHT IDLE STOP

CABIN ALTITUDE

CABIN ALTITUDE EXCEEDS 10,000 FEET

LDG/TAXI LIGHT

LANDING LIGHTS OR THE TAXI LIGHT IS ON WITH THE LANDING GEAR UP

PASS OXYGEN ON

PASSENGER OXYGEN SYSTEM IS CHARGED

AIR COND N1 LOW

RIGHT ENGINE N1 IS TOO LOW FOR THE AIR-CONDITIONING LOAD

R PROP PITCH

RIGHT PROPELLER IS BELOW THE FLIGHT IDLE STOP


4-10



QUESTIONS 1. The MASTER WARNING FLASHERS illuminate when ___________ annunciator illuminate(s). A. A red warning B. An amber caution C. A red warning or amber caution D. A red warning and amber caution 2. A r e d w a r n i n g a n n u n c i a t o r w i l l extinguish when: A. Th e M a s t e r Wa r n i n g f l a s h e r i s canceled. B. The fault is no longer sensed. C. A n e w f a u l t i s s e n s e d , c a u s i n g illumination of a new red warning annunciator. D. The appropriate checklist proce dure is accomplished.



3. Faults that illuminate the ______________ annunciators require immediate attention and reaction of the pilot. A. Red warning B. Amber caution C. Green advisory D. White status

4-11


CHAPTER 5 FUEL SYSTEM CONTENTS INTRODUCTION ............................................................................................................... 5-1 GENERAL ........................................................................................................................... 5-1 FUEL STORAGE AND CAPACITY............................................................................... 5-2 Main Tank System ......................................................................................................... 5-2 Auxiliary Tank System .................................................................................................. 5-3 King Air 350ER Saddle Tank ...................................................................................... 5-4 Fuel Capacity ................................................................................................................. 5-4 Fuel Tank Vents ............................................................................................................. 5-4 FUEL COMPONENTS....................................................................................................... 5-6 Pumps ............................................................................................................................. 5-6 Firewall Fuel Valves ..................................................................................................... 5-8 CONTROLS AND INDICATIONS.................................................................................. 5-9 Fuel Quantity Indications .......................................................................................... 5-10 Fuel Pressure Indication............................................................................................. 5-12 Fuel System Operation ............................................................................................... 5-12 Normal Operation....................................................................................................... 5-12 Transfer......................................................................................................................... 5-13 Crossfeed...................................................................................................................... 5-16

PREFLIGHT AND SERVICING ................................................................................... 5-19 Drain System ............................................................................................................... 5-19 Fuel Handling Practices ............................................................................................. 5-20


5-i

5 FUEL SYSTEM

Fuel Manifold Purge System...................................................................................... 5-18


Fuel Types and Additives ........................................................................................... 5-21 Filling the Tanks .......................................................................................................... 5-22 Defueling the Aircraft ................................................................................................ 5-22 LIMITATIONS................................................................................................................... 5-23 QUESTIONS ...................................................................................................................... 5-25

5 FUEL SYSTEM

5-ii




Title

Page

5-1

Main Fuel Tank System......................................................................................... 5-2

5-2

Auxiliary Fuel Tank System.................................................................................. 5-3

5-3

350ER Saddle Tank ............................................................................................... 5-4

5-4

Fuel Vents ............................................................................................................... 5-5

5-5

Fuel System Schematic Diagram ......................................................................... 5-7

5-6

Firewall Fuel Valves............................................................................................... 5-8

5-7

Fuel Control Panels ............................................................................................... 5-9

5-8

Fuel Quantity Indication System....................................................................... 5-10

5-9

Auxiliary Fuel Transfer System—Operating .................................................... 5-13

5-10

Auxiliary Fuel Transfer System—Override ...................................................... 5-14

5-11

Auxiliary Fuel Transfer System—Empty .......................................................... 5-15

5-12

Crossfeed Schematic ........................................................................................... 5-17

5-13

Fuel Manifold Purge System Schematic ........................................................... 5-18

5-14

Fuel Drain Locations .......................................................................................... 5-19

5-15

Main and Auxiliary Filler Caps ......................................................................... 5-23

5-16

Saddle Tank Filler Cap ....................................................................................... 5-23

TABLE 5-1

Title

Page

Fuel Drain Locations .......................................................................................... 5-19 5 FUEL SYSTEM

Table


5-iii


CHAPTER 5 FUEL SYSTEM

INTRODUCTION A complete understanding of the fuel system is essential to competent and confident operation of the aircraft. Management of fuel and fuel system components is a major everyday concern of the pilot. This section presents a description of the fuel system components and operation including physical layout of fuel cells, vents, and drains. Specific procedures such as taking fuel samples are also presented. The chapter discussion also includes information on the King Air 350ER model with extended fuel capabilities.

The King Air 350 fuel system simplifies cockpit flight procedures and provides easy access for ground servicing. A crossfeed system connects the two wing main fuel systems. Each wing also has an auxiliary fuel tank.

The King Air 350ER has a supplemental fuel system that includes two extended range fuel tanks that increase fuel supply.


5-1

5 FUEL SYSTEM

GENERAL


FUEL STORAGE AND CAPACITY

Gravity feed lines connect all the cells to allow fuel to flow into the nacelle tank that pumps fuel directly to each engine.

MAIN TANK SYSTEM

The filler cap is near the wingtip by the l e a d i n g e d g e. A n a n t i s i p h o n v a l v e i s installed at each filler port to prevent the loss of fuel or collapse of fuel tank bladder if the filler cap is improperly installed.

The main fuel system in each wing consists of two wing leading edge bladder-type cells, two box-section bladder-type cells, one wet wing integral-type cell, and the nacelle tank (Figure 5-1).

A crossfeed line connects each nacelle tank to the engine on the opposite side. Fuel in either wing system is available to either engine during single-engine operation.

5 FUEL SYSTEM

Figure 5-1. Main Fuel Tank System

5-2



AUXILIARY TANK SYSTEM The auxiliary fuel system consists of a fuel tank on each side of the aircraft in the wing center section (Figure 5-2). Because the cells are lower than the nacelle tank, they cannot gravity feed into the nacelle tanks. A jet pump adjacent to the outlet strainer and drain transfers fuel to the nacelle tank. The auxiliary transfer system is automatic. If the auxiliary tank contains any usable fuel, the system transfers it. Auxiliary tank fuel is used first during normal operation for the King Air 350.

On the 350ER model, the auxiliary system begins operation after the fuel in the extended range tanks is depleted. If the tank has fuel but does not transfer because of some system discrepancy, a yellow caution NO FUEL XFR annunciator illuminates. An override switch backs up the automatic system. Each auxiliary tank has its own filler opening with an antisiphon valve.

5 FUEL SYSTEM

Figure 5-2. Auxiliary Fuel Tank System


5-3


KING AIR 350ER SADDLE TANK

Vent Float Valve

A supplemental fuel tank is added to each side of the aircraft (Figure 5-3). The saddle tanks add 236 gallons to the aircraft for extended range capability.

The vent float valve in the integral fuel cell and the one in the top of the nacelle tank allow air to flow in either direction. However, when the fuel level rises up to the level of the vent float valve, the float rises with the fuel to seal the vent system at that point. This prevents fuel from flowing into the vent system.

Float Check Valve

Figure 5-3. 350ER Saddle Tank

FUEL CAPACITY The main system has a capacity of 380 gallons or 190 gallons on each side of usable fuel. The auxiliary system has 159 gallons or 79.5 gallons on each side. Total capacity for each side is 539 gallons. Approximately 2,546 pounds are available in the main system with about 1,273 pounds on each side. An estimated 1,065 pounds are available in the auxiliary system with about 533 pounds on each side. Total usable fuel is 539 gallons or 3,611 pounds.

King Air 350ER The ER model has an additional 236 gallons with the extended tanks for a total of 775 gallons or 5,192 pounds.

FUEL TANK VENTS 5 FUEL SYSTEM

The main and auxiliary fuel systems are vented through a recessed vent coupled to a heated ram vent on the underside of the wing adjacent to the nacelle (Figure 5-4). One vent is recessed to prevent icing. The other vent is heated to prevent icing.

5-4

The float check valve accomplishes the same job as the vent float valve. The float check valve looks like an ordinary check valve, but it is not installed in any tank or cell. It is in a short vent line between the recessed and heated ram vents and the auxiliary tank. When fuel is not present at the float check valve, the float is down to allow air to pass in either direction. When fuel is present, the float rises with the fuel and seals the check valve to prevent fuel flow through it.

Vent Lines Air is vented into or out of the auxiliary fuel cell through a line that extends from the recessed and heated ram vents through the float check valve to the auxiliary fuel cell. The wing cells are cross-vented with one another through a float-operated vent valve on the integral fuel cell. Air enters the wing cells through four passages. The first two are primary; the last two are applicable only in flight if the first two passages are plugged. • Line extending from heated ram vents through the leading edge of the wing to the vent float valve on the integral fuel cell • Line extending from the heated ram vents through the float check valve and then through the center of the wing to the vent float valve on the integral fuel cell



VENT FLOAT VALVE PRESSURE RELIEF VALVE

VENT FLOAT VALVE

RECESSED VENT AIR INLET

INTERGRAL FUEL CELL HEATED RAM VENT

FLOAT CHECK VALVE

FLAME ARRESTOR

Figure 5-4. Fuel Vents

• Line that bypasses the vent float valve altogether to extend from the air inlet on the underside of the wing near the tip through a tee and check valve to the integral fuel cell

Vent Operation Air vents into the nacelle tank through a vent float valve in the tip of the tank and/or through a tube next to the vent float valve.

Both the vent float valve and the tube next to it have a check valve downstream to prevent air or fuel from expanding out of the nacelle tank through these passages. Air flows to these passages and into the nacelle tank from the ram vents through the float check valve to a tee that is just prior to the auxiliary tank and then through a vent line that leads to the top of the nacelle tank. Another tee on top of the nacelle tank divides this line into the passages that lead to the vent float valve or to the tube next to the vent float valve. Air can escape from the nacelle tank through the vent float valve and then through the fuel return line leading to the auxiliary tank. From there air is vented overboard.


5-5

5 FUEL SYSTEM

• Line extending from the air inlet on the underside of the wing near the tip through a check valve to the vent float valve on the integral fuel cell; this passage is primarily a siphon break that prevents siphoning of fuel from the auxiliary tank through the wing tip to the heated ram vents when the aircraft is shut down on the ground


When the aircraft is shut down on the ground and all tanks are full of fuel that is colder than the ambient air temperature, the fuel must expand overboard. Fuel expands from the wing cells through the gravity feed line to the nacelle tank. It then expands out the top of the nacelle tank through the fuel return tube, pressure relief valve, and fuel return line to the auxiliary tank. When fuel expands out of the auxiliary fuel tank to the float check valve, the float closes the check valve to prevent some of the excess fuel from being discharged through the vents. When the check valve closes, the auxiliary fuel expands through a line routed outboard from the check valve through the center of the wing to the wing tip. It then continues down the wing leading edge vent line to the recessed and heated ram vents and onto the ground. When the fuel has expanded fully, the siphon break prevents continued siphoning of fuel.

FUEL COMPONENTS Components to operate the fuel system include three pumps, a firewall shutoff valve, various switches and gauges (Figure 5-5).

PUMPS Engine-Drive Pumps The engine-driven high-pressure fuel pump mounts on the accessory case of each engine in conjunction with the fuel control unit. An internal 200-mesh strainer protects the pump against fuel contamination.

Primary Boost Pumps The primary fuel boost pump is also enginedriven. It mounts on a drive pad on the aft accessory section of each engine. The boost pump has an operating capacity of 1,250 pounds per hour at a pressure of 30 psi. Any time the gas generator (N 1 ) is turning, this pump operates to provide sufficient fuel to the engine-driven fuel pump for all conditions except operation with crossfeed or operation with aviation gasoline above 20,0 0 0 feet.

Standby Boost Pumps An electrically driven standby pump in the bottom of each nacelle tank backs up the boost pump. It also provides additional pressure required for fuel crossfeed from one side of the aircraft to the other. The boost pump or the standby boost pump is capable of supplying fuel to the enginedriven fuel pump at the minimum pressure requirements. L e v e l l o c k STA N D BY P U M P t o g g l e s switches on the fuel control panel to control electrical power to the standby pumps. The respective generator bus supplies the fuel subpanel circuit breakers. Two 10ampere circuit breakers below the fuel control panel protect the circuit. The triplefed bus is the other source of power to the standby pumps. A diode network prevents interaction between the two power sources. The engine can operate with the failure of one or both boost pumps; failure, however, of the engine-driven high-pressure fuel pump causes the engine to flame out.

5 FUEL SYSTEM

This pump along with the FCU regulates fuel flow to the fuel nozzles in the engine. The pump has an output pressure of up to a maximum of 1050 psi that varies with N 1 rpm and FCU operation.

5-6


/

P3 AIR LINE FOR FUEL PURGE

ENGINE FUEL MANIFOLD FUEL CONTROL UNIT

FUEL FLOW TRANSMITTER AND INDICATOR


P3 BLEED-AIR LINE ENGINE-DRIVEN HIGH PRESS FUEL PUMP

FUEL PURGE TANK FIREWALL FUEL FILTER

FUEL HEATER AIR FILTER FUEL CONTROL PURGE LINE LEFT FUEL PRESSURE ANNUNCIATOR PRESSURE SWITCH

DRAIN VALVE

ENGINE DRIVEN BOOST PUMP CHECK VALVE

FIREWALL SHUTOFF VALVE STANDBY BOOST PUMP NACELLE TANK 54 GALLONS

GRAVITY FLOW CHECK VALVE MOTIVE FLOW VALVE PRESSURE SWITCH FOR LEFT NO FUEL TRANSFER LIGHT ON CAUTION PANEL FUEL LOW LEVEL SENSOR WS 290.92

VENT FLOAT VALVE

DRAIN VALVE

L FUEL QTY

FUEL QUANTITY PROBE

PRESSURE RELIEF VALVE

FUEL QUANTITY PROBE

CROSSFEED VALVE WING LEADING EDGE 13 GALLONS

WING LEADING EDGE 40 GALLONS

25 GALLONS BOX SECTION

INTEGRAL (WET CELL) 35 GALLONS

25 GALLONS BOX SECTION

AIR INLET VENT FLOAT VALVE DRAIN

FUEL QUANTITY PROBE

RECESSED VENT HEATED RAM VENT

AUXILIARY

79.5 GALLONS FLOAT CHECK VALVE DRAIN VALVE FLAME ARRESTOR

5-7

Figure 5-5. Fuel System Schematic Diagram

5 FUEL SYSTEM

TRANSFER JET PUMP

STRAINER, DRAIN AND FUEL SWITCH


TO FLOW DIVIDER


Jet Transfer Pumps

FIREWALL FUEL VALVES

A jet pump transfers fuel from the auxiliary cell to the nacelle tank.

The supply line from the nacelle tank is routed from the inboard side of the nacelle tank forward to the engine-driven boost pump through a normally open firewall shutoff valve in the fuel line immediately behind the engine firewall.

The pump in the sump of the auxiliary cell is mounted adjacent to the outlet strainer and drain. The fuel line that supplies the motive flow to the jet transfer pump is routed along the outboard side of the nacelle through the jet pump motive control valve just aft of the firewall.

King Air 350ER Additional Jet Pumps Each of the saddle supplementary tanks has a jet pump that transfers fuel from the tank to the nacelle tank. Fuel pressure from the engine-driven boost pump provides the motive flow to drive the jet pump.

Th e F / W VA LV E P U S H a n n u n c i a t o r switch on the instrument panel glareshield closes its respective firewall shutoff valve to shut off the flow of fuel to the engine ( Fi g u r e 5 - 6 ) . Th e l e g e n d i l l u m i n a t e s CLOSED to indicate the firewall valve is closed. When the valve is in the open position, the legend is extinguished. A flashing annunciator indicates the vale is not in the selected position. When either annunciator switch is d e p re s s e d , t h e re d E X T I N G U I S H E R PUSH annunciator in the corresponding fire extinguisher switch illuminates to indicate the fire extinguisher is armed.

Figure 5-6. Firewall Fuel Valves 5 FUEL SYSTEM

5-8



CONTROLS AND INDICATIONS The fuel panel on the pilot side panel contains a fuel quantity gage for each engine, a placard stating the usable fuel, and the following switches (Figure 5-7): • Left and right STANDBY PUMP with ON–OFF postions to control the standby boost pumps

• AUX TRANSFER with OVERRIDEAU TO p o s i t i o n s t o c o n t ro l f u e l t ra n s f e r. Fo r t h e 3 5 0 E R , X F R OV E R R I D E w i t h AU X – AU TO – ER positions • C R O S S F E E D F LOW w i t h O N OFF positions • F U E L Q UA N T I T Y s w i t c h w i t h T E S T – M A I N – AU X I L I A RY positions Each of these switches is discussed in detail in the appropriate operation section.

5 FUEL SYSTEM

KING AIR 350

KING AIR 350ER

Figure 5-7. Fuel Control Panels


5-9


FUEL QUANTITY INDICATIONS

Fuel Quantity Probes

The fuel quantity system is a capacitance gaging system with one quantity indicator per wing (Figure 5-8). A spring-loaded selector allows the pilot to individually check tank quantity.

Each side of the aircraft has an independent gaging system consisting of the following fuel quantity (capacitance) probes:

The system compensates for specific gravity and reads in pounds on a linear scale. An electronic circuit in the system processes the signals from the fuel quantity (capacitance) probes in the various fuel cells for an accurate readout by the fuel quantity indicators. A Density Variation of Aviation Fuel g r a p h i s i n t h e We i g h t a n d B a l a n c e Equipment List section of the POH to allow more accurate calculations of weights for all approved fuels.

• One in nacelle fuel cell • One in aft inboard fuel cell • Two in integral (wet wing) fuel cell • Two in inboard leading edge fuel cell • Two in center section fuel tank The fuel quantity probe is a variable capacitor composed of two concentric tubes. The tubes serve as fixed electrodes. The fuel in the space between the tubes acts as the dielectric of the fuel quantity probe.

NACELLE TANK INSTALLATION AUX FUEL

GRAVITY FEED FROM OUTBOARD MAIN TANKS

FUEL PROBE

FUEL LOW LEVEL SENSOR

STANDBY BOOST PUMP

OUTLET STRAINER CAPACITANCE PROBES FUEL LOW LEVEL SENSOR

5 FUEL SYSTEM

PROBES

Figure 5-8. Fuel Quantity Indication System

5-10



Fu e l d e n s i t y a n d e l e c t r i c a l d i e l e c t r i c constant vary with respect to temperature, fuel type, and fuel batch. The capacitance gaging system senses and compensates for these variables. The capacitance of the probe varies with respect to the change in the dielectric that results from the ratio of fuel to air in the fuel cell. As the fuel level between the inner and outer tubes rises, air with a dielectric constant of one is replaced by fuel with a dielectric constant of approximately two, thus increasing the capacitance of the probe. This variation in the volume of fuel in the fuel cell produces a capacitance variation that is a linear function of that volume. This is converted to linear current that actuates the fuel quantity indicator.

King Air 350 ER Fuel Quantity Th e F U E L QUA N T I TY s w i t c h i n t h e 350ER is essentially the same except that the switch upper portion is for reading the quanlity of the ER tanks. A separate TEST switch is to the lower left of the panel. While transmitting on the HF system, deviations of ER fuel quanity may be observed throughout the usable HF frequency band (2 MHz to 30 MHz). The frequency at which the deviations are observed may vary, due to issues such as the actual fuel quantity in the ER fuel tank, the tuned frequency on which the HF system is transmitting, etc. The displayed ER fuel quantity valve returns to the correct value after the HF transmission ceases.

Fuel Quantity Gauges

The FUEL QUANITY switch is springloaded to the center position. The TEST position provides a test function of the L and R FUEL QTY fiber optic sensing circuitry and caution annunciators. When the FUEL QUANTITY switch is in the MAIN position, the 5-ampere QTY IND circuit breaker supplies power through the fuel quantity gage to the capacitance probes in the main fuel system tanks. When the switch is in the AUXILIARY position, the circuit breaker supplies power to ground through the coil of the gage switching relay. Power is then supplied through the fuel quantity gage to the c a p a c i t a n c e p ro b e s i n e a c h a u x i l i a r y fuel tank.

Low Fuel Quantity Indication Fiber optic sensors in both nacelle fuel tanks alert the pilot to a low fuel situation. When fuel quantity remaining in the main system is below approximately 30 0 pounds (45 gallons) or 30 minutes of fuel at maximum continuous power, the corresponding yellow caution L or R FUEL QTY annunciator illuminates. This also triggers the MASTER CAUTION flashes. A five- to seven-second delay is built into the annunciator circuit to reduce the likelihood of fuel sloshing that might cause transient indications. A holding circuit keeps the annunciator illuminated for three to five seconds once it does illuminate.

Testing the System Test the system by holding the FUEL QUANTITY switch in TEST. A simulated low quantity signal is sent to the fiber optic sensor in the nacelle tank; the L and R FUEL QTY annunciators illuminate after a five second delay.


5-11

5 FUEL SYSTEM

When the FUEL QUANTITY selector switch is in the MAIN position, the fuel quantity gages indicate the amount of fuel remaining in the main tank. When the switch is in the AUXILIARY position, the gages indicate the fuel remaining in the auxiliary tank


When the switch is released, it springs back to center. The annunciators extinguish after approximately five seconds.

FUEL PRESSURE INDICATION The fuel pressure switch is directly above the firewall-mounted fuel filter and indicates fuel boost pressure. At 9 to 11 psig of decreasing pressure, the switch closes and actuates the red warning L or R FUEL PRESS annunciator in the warning annunciator panel. With low boost pressure indicated, switch on the electric standby boost pump unless a fuel leak is indicated.

CAUTION Operation with the FUEL PRESS light on is limited to 10 hours between overhaul or replacement of the engine-driven high pressure fuel pump. Should both boost pumps fail, suction lift operation may be employed; however, suction lift operation is restricted to 10 hours total time between h i g h p re s s u re p u m p o v e r h a u l periods. If the pump is operated on suction lift beyond the 10-hour limit, overhaul or replacement of the high-pressure pump is necessary. Windmilling time is not equivalent to operation of the engine at high power with respect to the effects of cavitation on fuel pump components. Consequently, windmilling time is not to be included in the 10-hour limit on engine operation without a boost pump. 5 FUEL SYSTEM

Th e re d F U E L P R E S S a n n u n c i a t o r extinguishes with 9 to 11 psig of increasing fuel pressure.

5-12

FUEL SYSTEM OPERATION Fuel flow from each wing outer main cell and auxiliary tank system is automatic without pilot action (see Figure 5-1). Fuel in the auxiliary tank is used first followed by the fuel in the main tanks. On the King Air 350ER, the fuel in the saddle tanks is used first followed by the fuel in the auxiliary tanks, and by fuel in the main tanks. The outer wing cells gravity-feed into the nacelle tank. The line extends from aft i n b o a rd w i n g c e l l , f o r w a rd a l o n g t h e outboard side of the nacelle tank, and aft of the firewall immediately under the motive flow valve. A gravity flow check valve in the end of the gravity feed line prevents any backflow of fuel into the outer wing cells.

NORMAL OPERATION The supply line from the nacelle tank is routed from the inboard side of the nacelle tank through a motorized firewall fuel valve immediately behind the engine firewall. From the firewall fuel valve, fuel is routed to the engine-driven boost pump and then to the main fuel filter on the lower center of the engine firewall. The filter has a bypass valve that permits fuel flow in case of plugging and a drain valve to drain the filter prior to each flight. A pressure switch mounted directly above the filter senses boost pump fuel pressure at the filter. From the main filter, fuel is routed through the fuel heater that uses heat from the engine oil to warm the fuel. The fuel is then routed to the high pressure pump and fuel control unit (FCU) that regulates fuel flow to the fuel nozzles. A fuel flow transmitter located adjacent to the FCU sends a signal to an electric DC powered fuel flow gage in the cockpit.



TRANSFER

Transfer Operation

The jet pump transfers fuel from the sump of the auxiliary tank to the nacelle tanks. Fuel pressure from the engine-driven boost pump or the electrical standby boost pump provides the motive flow for the jet transfer pump.

The AUX TRANSFER lever-lock toggle switch on the fuel control panel actuates the jet transfer pumps. In the AUTO position, the automatic fuel tra nsfer module a pplie s power to the normally closed motive flow valve to open it (Figure 5-9). The module applies the power when the boost pump pressure switch senses fuel pressure and the float switch senses fuel in the auxiliary tank.

The fuel line that supplies the motive flow is routed along the outboard side of the nacelle through the motive flow valve just aft of the firewall to the jet pump. A check valve in the motive flow line immediately aft of the motive flow control valve prevents the engine from taking in air when the boost pump is not operating. AUX TRANSFER

AUX TRANSFER SWITCH OVERRIDE

AUTO

FLOAT SWITCH

MOTIVE FLOW PRESSURE SWITCH

L NO FUEL XFER LIGHT

NOT EMPTY 11± 2 SEC DELAY EMPTY

CROSSFEED ON

AUTOMATIC FUEL TRANSFER MODULE

IGNITION ON

(ON ONLY)

6.5 SEC DELAY JET TRANSFER PUMP

NC AUTO IGNITION

MOTIVE FLOW VALVE

BOOST PUMP PRESSURE SWITCH

TO NACELLE TANK

FROM AUX TANK SUMP PRESSURE WARNING

FROM BOOST PUMP

TO ENGINE

Figure 5-9. Auxiliary Fuel Transfer System—Operating


5-13

5 FUEL SYSTEM

L FUEL PRESS LOW LIGHT


gizes the motive flow valve; the valve closes. The time delay prevents cycling of the motive flow valve because of sloshing fuel (Figure 5-10).

Once the motive flow valve opens, the jet transfer pump pumps fuel from the sump of the auxiliary fuel tank into the nacelle fuel cell for as long as there is fuel in the auxiliary tank and the engine-driven boost pump or electrical standby boost pump operate.

The OVERRIDE position of the AUX TRANSFER position bypasses the fuel transfer module to apply power directly to the motive flow valve.

A motive flow pressure actuates at 6 (±1) psi to confirm motive flow fuel pressure. The switch is in the fuel line between the motive flow valve and check valve.

Overflow Line Th e f u e l t ra n s f e r ra t e i s g re a t e r t h a n normal engine fuel consumption. As a result, an overflow return line is required. The overflow line is plumbed from the nacelle tank back to the auxiliary tank to provide a return for excess fuel.

When the auxiliary fuel is depleted, the float switch sends a signal after a six- to seven-second time delay to the automatic fuel transfer module. The module deenerAUX TRANSFER SWITCH

AUX TRANSFER

OVERRIDE

AUTO

FLOAT SWITCH




CROSSFEED ON

X


IGNITION ON

AUTO IGNITION

(ON ONLY)


NC MOTIVE FLOW VALVE

BOOST PUMP PRESSURE SWITCH

TO NACELLE TANK



5 FUEL SYSTEM

FROM BOOST PUMP

TO ENGINE

Figure 5-10. Auxiliary Fuel Transfer System—Override

5-14



The overflow of fuel from the nacelle tank comes out of an overflow tube at the top of the nacelle tank. It then continues past a 11/2 psi pressure relief valve and into a fuel return line to the auxiliary tank.

The appropriate NO FUEL XFR annunciator illuminates when there is less than 6 (±1) psi of pressure and the float switch in the auxiliary tank does not sense an empty tank.

Abnormal Conditions

With fuel in the auxiliary tank, should this pressure switch not be actuated, the L or R NO FUEL XFR illuminates or remains illuminated to indicate that the motive flow v a l v e i s s t i l l c l o s e d . P l a c e t h e AU X TRANSFER switch in the OVERRIDE position (Figure 5-11).

NO FUEL XFR Annunciator Illumination of the yellow caution L or R NO FUEL XFR can signal several different conditions in the auxiliary fuel system transfer. The annunciator has an 11 second delay to prevent transients from t r i g g e r i n g b o t h i t a n d t h e M AS T E R CAUTION flashers.

AUX TRANSFER

AUX TRANSFER SWITCH OVERRIDE

AUTO

FLOAT SWITCH




CROSSFEED ON


IGNITION ON

(ON ONLY)


NC AUTO IGNITION

MOTIVE FLOW VALVE BOOST PUMP PRESSURE SWITCH

TO NACELLE TANK


FROM BOOST PUMP

TO ENGINE

Figure 5-11. Auxiliary Fuel Transfer System—Empty


5-15

5 FUEL SYSTEM



The auxiliary fuel system does not feed into the main fuel system if there is a failure of both boost pumps or a failure of the motive flow valve. This condition is visible on the auxiliary tank FUEL QUANTITY gage and with the illumination of the NO FUEL XFER annunciator. Any time the engine ignition circuit is powered through t h e AU T O I G N I T I O N o r S TA RT & IGNITION switch, the automatic fuel transfer module removes power from the motive flow valve. If the system is transferring fuel, the valve closes; the appropriate NO FUEL XFR annunciator illuminates. Selecting crossfeed also causes the fuel transfer module to interrupt electricity and close the motive flow valve. The appropriate NO FUEL XFER light also illuminates if there is fuel in the auxiliary tank.

power to a 30-second time-delay relay. This relay closes the extended range motive flow valve and opens the valves associated with the auxiliary fuel tank. Upon exhaustion of the extended range fuel tank and auxiliary fuel tank, a float switch in the auxiliary fuel tank sends a signal to close all valves associated with fuel transfer. Normal gravity transfer of the main wing fuel into the nacelle tanks begins. When the XFR OVERRIDE switch is in the AUTO position and the extended range fuel tank is empty, the automatic fuel transfer module along with additional relay logic simultaneously remove power and close the extended range motive flow valve. This prevents continued operation of the jet pump.

ER Switch Positions Low Boost Pressure If fuel boost pressure drops below 10 psi (FUEL PRESS annunciator illuminated), the automatic fuel transfer module removes power to close the motive flow valve. This prevents continued operation of the jet transfer pump. The jet transfer pump is not damaged by operating after the tank is dry, but extended operation with an empty auxiliary tank tends to draw unnecessary moist air into the main fuel system from the empty, vented auxiliary tanks.

King Air 350ER Transfer Operation

5 FUEL SYSTEM

During transfer of extended range fuel, the auxiliary tanks and nacelle tanks are maintained full. A check valve in the gravity feed line from the outboard wing prevents reverse fuel flow from the nacelle tank. When all usable fuel in the extended range tank is transferred, a float switch toward the aft end of the tank actuates and supplies

5-16

When the XFR OVERRIDE switch is in the ER position and the extended range fuel tank is empty, the XFR OVERRIDE switch must be manually positioned to the AU T O o r AU X p o s i t i o n . Th e AU T O position returns control to the automatic fuel transfer module; the AUX position commands fuel to be supplied from the auxiliary fuel tank. The extended range fuel system does not feed into the main fuel system if there is a failure of both boost pumps (engine-driven and electrical) or a failure of the extended range motive flow valve. The NO FUEL XFR annunciator illuminates for the same conditions as the auxiliary transfer system.

CROSSFEED A crossfeed line connects each nacelle tank to the engine on the opposite wing. The line is routed from the inboard side of the nacelle aft to the center wing section a n d a c ro s s t o t h e i n b o a rd s i d e o f t h e opposite nacelle.



When the CROSSFEED switch on the fuel control panel is actuated, a 5-ampere circuit breaker on the fuel control panel supplies power to the solenoid that opens the crossfeed valve.

A valve connected into the line at the aft inboard corner of the left nacelle controls the crossfeed line (Figure 5-12). Crossfeed requires standby boost pump operation on the side from which crossfeed is desired. Its operation ensures an adequate flow of fuel to the receiving engine. It also maintains motive flow for the jet transfer pump on the supply side.

TO FLOW DEVIDER

The automatic fuel transfer module simultaneously energizes the standby pump on the side from which crossfeed is desired and deenergizes (closes) the motive flow valve on the side being crossfed.

LOW PRESSURE ENGINE-DRIVEN FUEL PUMP

TO FLOW DIVIDER

FIREWALL SHUTOFF VALVE MOTIVE FLOW VALVE

MOTIVE FLOW VALVE

STANDBY BOOST PUMP

FUEL CROSSFEED

5 FUEL SYSTEM

CROSSFEED VALVE

LOW FUEL QUANTITY PROBE

Figure 5-12. Crossfeed Schematic


5-17


Crossfeed does not transfer fuel from one cell to another; its primary function is to supply fuel from one side to the opposite engine during an engine-out condition. If the standby boost pumps on both sides are operating and the crossfeed valve is open, fuel is supplied to the engines in the normal manner because pressure on each side of the crossfeed valve is equal. When crossfeed is selected, the green advisory FUEL CROSSFEED annunciator illuminates indicating that the crossfeed valve has opened.

Precautions When performing crossfeed, be aware of the following precautions: • AUX TRANSFER switch must be in AUTO for the side receiving fuel. If the switch is in OVERRIDE, the motive flow valve remains open. In addition, incoming fuel would start filling the tanks through the auxiliary transfer line and could result in fuel being dumped overboard.

FUEL MANIFOLD PURGE SYSTEM Th i s a i r c r a f t i s e q u i p p e d w i t h a f u e l m a n i f o l d p u rg e s y s t e m t o e n s u re a n y residual fuel in the fuel manifold is consumed during engine shutdown (Figure 5-13). During engine operation, compressor discharge (P3 air) is routed through a filter and check valve to pressurize a small air tank on the engine truss mount. On engine shutdown, the pressure differential between the air tank and the fuel manifold causes air to be discharged from the air tank through a check valve and into the fuel manifold system. The air forces all residual fuel remaining in the fuel manifold o u t t h ro u g h t h e n o z z l e s a n d i n t o t h e combustion chamber. Th e f u e l f o r c e d i n t o t h e c o m b u s t i o n chamber is consumed, which in turn causes a momentary rise in engine speed.

• Both STANDBY PUMP switches should be in the OFF position. The crossfeed system automatically turns on the pump it needs to establish crossfeed. • If the firewall fuel valve was closed on t h e i n o p e ra t i v e e n g i n e d u r i n g shutdown, the FUEL PRESSURE annunciator remains illuminated, and any auxiliary fuel on that side is unusable due to lack of motive flow pressure.

5 FUEL SYSTEM

Figure 5-13. Fuel Manifold Purge System Schematic

5-18



PREFLIGHT AND SERVICING

Table 5-1. FUEL DRAIN LOCATIONS DRAINS

LOCATION

DRAIN SYSTEM

Flush fuel drain

During each preflight, the fuel drains on the tanks, lines, and filters should be drained to check for fuel contamination.

Gravity line drain

Outboard of nacelle underside of wing

Fuel drain

Outboard of nacelle underside of wing

The main and auxiliary fuel systems have five sump drains, a standby pump drain manifold, and a firewall filter drain in each wing. The drain valve for the firewall fuel filter is to the right of the filter at the firewall on the underside of the nacelle.

Strainer drain

Bottom of nacelle

Filter drain

Forward of wheel well

Inboard of fuel tank drain

Underside of wing by wing root

5 FUEL SYSTEM

Each wing has four tank drains, one line drain, and one filter drain (Figure 5-14). See Table 5-1.

Underside of wing forward of aileron

Figure 5-14. Fuel Drain Locations


5-19


The nacelle tank has two drains on the bottom of the nacelle forward of the wheel well. The inboard drain is for the standby boost pump and the outboard drain is for the nacelle fuel sump and strainer. Do not drain the standby pump drain on preflight.

FUEL HANDLING PRACTICES

The leading edge tank has a drain on the underside of the wing just outboard of the nacelle. The integral (wet wing) fuel tank has a sump drain approximately midway on the underside of the wing aft of the main spar. The drain for the auxiliary tank is at the wing root midway between the main and aft spars.

Kerosene, with its higher specific gravity, tends to absorb and suspend more water than aviation gasoline. Along with the water, it suspends rust, lint, and other foreign materials longer. Given sufficient time, these suspended contaminants settle to the bottom of the tank. The settling time for kerosene is five times that of aviation gasoline; therefore, jet fuels require good fuel handling practices to ensure the aircraft is serviced with clean fuel.

The gravity feed line from the wing tanks to the nacelle tank also has a drain line that extends aft along the outboard side of the main gear wheel well to a drain valve just aft of the wheel well. Because jet fuel and water are of similar densities, water does not settle out of jet fuel as easily as from aviation gasoline. For maximum water and fuel separation, the aircraft should sit perfectly still with no fuel being added for approximately four hours prior to draining the sumps. If there is a substantial amount of water in the fuel, however, water and fuel separation does occur soon after fueling or moving the aircraft. Although turbine engines are not as critical as reciprocating engines regarding water ingestion, remove water periodically to prevent formations of fungus and contamination-induced inaccuracies in the fuel gaging system. When draining the flush-mounted drains, do not turn the draining tool. Turning or twisting unseats the O-ring seal causing a leak. 5 FUEL SYSTEM

King Air 350ER Drains The extended range fuel tanks have one drain valve on the lower aft end of each fuel tank. Flapper valves inside each tank prevent the fuel from surging forward.

5-20

All hydrocarbon fuels contain some dissolved and some suspended water. The quantity of water in the fuel depends on temperature and type of fuel.

If recommended ground procedures are carefully followed, solid contaminants settle and free water can be reduced to 30 parts per million (ppm). This value is currently accepted by the major airlines. Since most suspended matter can be removed from the fuel by sufficient settling time and proper filtration, it is not a major problem. Dissolved water has been found to be the major fuel contamination problem. Its effects are multiplied in aircraft operating primarily in humid regions and warm climates. Dissolved water cannot be filtered from the fuel by micronic-type filters. It can be released by lowering the fuel temperature; this occurs in flight. For example, kerosene fuel may contain 65 ppm (8 ounces per 1,0 0 0 gallons) of dissolved water at 80°F. When the fuel temperature is lowered to 15°F, only about 25 ppm remain in solution. The difference of 40 ppm has been released as supercooled water droplets that need only a piece of solid contaminant or an impact shock to convert them to ice crystals. Tests indicate that these water droplets do not settle during flight; they pump freely



through the system. If they become ice crystals in the tank, they do not settle because the specific gravity of ice is approximately equal to that of kerosene. Forty ppm of suspended water seems like a very small quantity, but when added to suspended water in the fuel at the time of delivery, it is sufficient to ice a filter. While the critical fuel temperature range is from 0°F to –2°F, which produces severe system icing, water droplets can freeze at any temperature below 32°F.

3. Pe r f o r m f i l t e r i n s p e c t i o n s determine if sludge is present.

to

4. M a i n t a i n g o o d h o u s e k e e p i n g b y periodically flushing the fuel tank system. The frequency of flushing is determined by the climate and presence of sludge. 5. Aviation gas is an emergency fuel. O b s e r v e t h e 15 0 h o u r s m a x i m u m operation on aviation gasoline. 6. Use only clean fuel servicing equipment.

Although this aircraft uses bladder-type fuel cells in addition to an integral (wet wing) fuel cell in each wing and all metal parts (except the standby boost pumps and jet transfer pumps) are mounted above the settlement areas, the possibility of filter clogging and corrosive attacks on fuel pumps exists if contaminated fuels are consistently used. The primary means of fuel contamination control by the owner/operator is good housekeeping. This applies not only to fuel supply, but to keeping the aircraft system clean. The following is a list of steps to recognize and prevent contamination problems: 1. Know your supplier. It is impractical to assume fuel free from contaminants is always available. But it is feasible to exercise caution and be watchful for signs of fuel contamination. 2. Ensure as much as possible that the fuel obtained has been properly stored, that it is filtered as it is pumped to the truck, and again as it is pumped from the truck to the aircraft.

7. After refueling, allow a settle period of at least four hours whenever p o s s i b l e, a n d t h e n d r a i n a s m a l l amount of fuel from each drain.

CAUTION Re m o v e s p i l l e d f u e l f ro m t h e ramp area immediately to prevent the contaminated surface from causing tire damage. Even if the fuel does not contain water, there is still a possibility of fuel icing at very low temperatures.

FUEL TYPES AND ADDITIVES Jet A, Jet A-1, Jet B, JP-4,-5, and -8 fuels may be mixed in any ratio. Aviation Gasoline Grades 80/87, 91/96, 10 0 LL, 10 0/130 and 115/145 are emergency fuels and may be mixed in any ratio with the normal fuels when necessary. Use of the lowest octane rating available is suggested because of its lower lead content. The use of aviation gasoline shall be limited to 150 hours operation during each time between overhaul (TBO) period.


5-21

5 FUEL SYSTEM

Water in jet fuel also creates an environment favorable to the growth of a microbiological sludge in the settlement areas of the fuel cells. This sludge, in addition to other contaminants in the fuel, can cause corrosion of metal parts in the fuel system as well as clog the fuel filters.


The Fuel Brands and Type Designations c h a r t i n t h e H a n d l i n g , S e r v i c e, a n d Maintenance section of the POH gives fuel refiner’s brand name along with the corresponding designations established by the American Petroleum Institute (API) a n d t h e A m e r i c a n S o c i e t y o f Te s t i n g Material (ASTM). The brand names are listed for ready reference and are not specifically recommended by the aircraft manufactuter. Any product conforming to the recommended specification may be used.

Anti-icing Fuel Additive Engine oil heats the fuel before it enters the FCU. Because no temperature measurement is available for the fuel at the nacelle tank, it must be assumed to be the same as the outside air temperature (OAT). If the OAT is below -45°C, ice formation could occur during takeoff or in flight. An anti-icing additive per MIL-I-27686 should be mixed with the fuel at refueling to ensure safe operation. Refer to the POH and manufacturer’s maintenance manual for procedures to follow when blending antiicing additive with the aircraft fuel.

When filling the aircraft fuel tanks, always observe the following: 1. E n s u r e t h e a i r c r a f t i s s t a t i c a l l y grounded to the servicing unit and to the ramp. 2. Service the main tanks on each side first. The main filler caps are in the outboard fuel cell on the leading edge of each wing near the wingtip. The auxiliary filler caps are on top of the center section, inboard of each nacelle (Figure 5-15); filler caps for the King Air 350 ER are on top of the saddle tank (Figure 5-16). 3. Allow a four-hour settling period w h e n e v e r p o s s i b l e. Th e n d ra i n a sufficient amount of fuel from each d ra i n p o i n t t o re m o v e w a t e r a n d contaminants.

DEFUELING THE AIRCRAFT As an integral part of the nacelle fuel tank, a defueling adapter aft of the standby pump contains a check valve to prevent fuel drainage when the plug is removed. Drain each wing fuel system as follows:

Fuel Biocide Additive Fuel biocide-fungicide BIOBOR JF in concentrations of 135 ppm or 270 ppm may be used in the fuel. BIOBOR JF may be used as the only fuel additive, or it may be used with the anti-icing additive conforming to MIL-I-27686 specification. Used together, the additives have no detrimental effect on the fuel system components.

5 FUEL SYSTEM

S e e t h e m a n u f a c t u r e r ’s m a i n t e n a n c e manual for concentrations to use and for procedures for adding BIOBOR JF to the aircraft fuel.

5-22

FILLING THE TANKS

1. Remove cover on the bottom of the nacelle to access the adapter plug. 2. Remove plug and screw the long end of an AN832-12 union into the adapter. The fuel begins draining as the union unseats the check valve. 3. The fuel may be gravity-drained or, to facilitate defueling, pumped out with the aid of a fuel truck. Refer to the manufacturer’s maintenance manual for more details.



LIMITATIONS The limitations that pertain to the fuel system are briefly summarized below. Refer t o t h e P i l o t ’s O p e r a t i n g H a n d b o o k , Maintenance Manual, and other specific topics in this section for more details. 1. Operation with a fuel pressure light illuminated is limited to ten hours before overhaul or replacement of the engine-driven high-pressure fuel pump. 2. The King Air 30 0 maximum zero fuel weight is 11,50 0. The King Air 350 maximum zero fuel weight is 12,500.

Figure 5-15. Main and Auxiliary Filler Caps

3. Maximum operation with aviation g a s o l i n e i s l i m i t e d t o 15 0 h o u r s between engine overhauls. Use of aviation gasoline is limited to 150 hours due to lead deposits which form in the turbine section during aviation gas consumption and cause power degradation. Since the aviation gas will probably be mixed with jet fuel already in the tanks, it is important to record the number of gallons of aviation gas taken aboard. As a rough approximation, it is expected that the PT6A-60A will have an average fuel consumption of 55 gallons per hour per engine, therefore each time 55 gallons of aviation gasoline are added, one hour of the 150-hour limitation is being used for that engine. Consult t h e m a n u f a c t u r e r ’s m a i n t e n a n c e manual for more details.

5 FUEL SYSTEM

Figure 5-16. Saddle Tank Filler Cap


5-23


4. If the tanks have been serviced with aviation gasoline, operation is prohibited if either standby boost pump is inoperative. The chart found in the Weight and Balance section of the Pilot’s Operating Handbook shows that the density of aviation gasoline is considerably less than that of jet fuel. B e c a u s e i t i s l e s s d e n s e, a v i a t i o n gasoline delivery is much more critical than jet fuel delivery. Aviation gasoline feeds well under pressure feed but does not feed well on suction feed, particularly at high altitudes. For this reason, two alternate means of pressure feed must be available for aviation gasoline at high altitude. These two means are the standby boost pump and crossfeed f ro m t h e o p p o s i t e s i d e. Th u s, a crossfeed capability is required for climbs above 20,0 0 0 feet pressure altitude.

9. M i n i m u m k n o w n o r f o r e c a s t a i r temperature for operation without fuel anti-icing additive is –45°C.

WARNING One operative standby fuel pump is required for takeoff when using recommended engine fuels, but in such a case, crossfeed of fuel will not be available from the side of the inoperative standby fuel pump.

5. When fueling the Super King Air 30 0 or 350, the main fuel tanks should be f u l l b e f o re a n y f u e l i s p u t i n t h e auxiliary tanks to reduce the structural bending moment in flight. 6. The Super King Air 300 and 350 have a maximum fuel imbalance of 30 0 pounds between wing fuel systems. 7. Takeoff is prohibited when the fuel quantity indicator needles are in the yellow arc or when there is less than 265 pounds of fuel in each main system. 8. Crossfeeding of fuel is permitted only when one engine is inoperative.

5 FUEL SYSTEM

5-24



QUESTIONS 1. I f a u x i l i a r y f u e l i s r e q u i r e d , t h e auxiliary tank _______ be filled _______ filling the main fuel tanks. A. May; after B. May; before C. Must; before D. Must; after 2. Illumination of the amber [L/R FUEL QTY] annunciator indicates less than 30 minutes of fuel remaining: A. In the appropriate auxiliary fuel tank. B. In the appropriate main fuel tank. C. At maximum continuous power. D. At maximum range power.

5. The approved military grade fuels are: A. JP-4, JP-5, and JP-8. B. 100LL and 115/145. C. Jet A and Jet A-1. D. Jet A and Jet B. 6. The maximum allowed lateral fuel imbalance is _______ lbs. A. 100 B. 300 C. 500 D. 700

3. Illumination of the red [L/R FUEL P R E S S LO ] w a r n i n g a n n u n c i a t o r during normal flight operations indicates: A. Insufficient pressure at the fuel pressure switch. B. D u r i n g a l l o p e r a t i o n s w i t h emergency fuel. C. C r o s s f e e d o p e r a t i o n i s n o t available. D. Powerplant failure is imminent.


5 FUEL SYSTEM

4. According to the checklist, crossfeed is selected: A. When fuel transfer is required. B. O n l y d u r i n g s i n g l e e n g i n e operations. C. A n y t i m e a f u e l i m b a l a n c e i s exceeds a limitation during normal operations. D. When the auxiliary fuel is empty after being used.

5-25

6 AUXILIARY POWER SYSTEM


The material normally covered in this chapter is not applicable to this aircraft.


6-i


CHAPTER 7 POWERPLANT CONTENTS INTRODUCTION ............................................................................................................... 7-1 GENERAL ........................................................................................................................... 7-1 Engine Ratings .............................................................................................................. 7-1 Engine Stations.............................................................................................................. 7-2 Engine Terms ................................................................................................................. 7-4 POWERPLANT ................................................................................................................... 7-4 General Principles......................................................................................................... 7-5 General Operation ........................................................................................................ 7-6 Engine Airflow .............................................................................................................. 7-8 Ignition System .............................................................................................................. 7-9 Accessory Section ....................................................................................................... 7-10 Lubrication System ..................................................................................................... 7-13 Engine Fuel System .................................................................................................... 7-16 Engine Power Control................................................................................................ 7-21 Engine Instruments .................................................................................................... 7-23 Engine Limitations ..................................................................................................... 7-25 PROPELLER .................................................................................................................... 7-28 Blade Angle ................................................................................................................. 7-31 Primary Governor....................................................................................................... 7-31 Overspeed Governor.................................................................................................. 7-44 Fuel Topping Governor .............................................................................................. 7-45


7-i

7 POWERPLANT

Page


Power Levers ............................................................................................................... 7-45 Propeller Control Levers ........................................................................................... 7-46 Propeller Feathering ................................................................................................... 7-46 Synchrophaser ............................................................................................................. 7-51 QUESTIONS ...................................................................................................................... 7-53 7 POWERPLANT

7-ii



ILLUSTRATIONS Title

Page

7-1

PT6A-60A Specifications...................................................................................... 7-2

7-2

Engine Cutaway..................................................................................................... 7-3

7-3

PT6A-60A Powerplant Installation..................................................................... 7-4

7-4

Engine Modular Concept ..................................................................................... 7-5

7-5

Engine Gas Flow and Stations............................................................................. 7-7

7-6

Jet-Flap, Compressor Bleed Valve, and Swing Check Valve ............................ 7-9

7-7

Engine Start and Ignition Switches................................................................... 7-10

7-8

Typical PT6A Engine.......................................................................................... 7-11

7-9

Front and Rear Accessory Drive....................................................................... 7-12

7-10

Accessory Gearbox Geartrain ........................................................................... 7-12

7-11

Engine Lubrication Diagram............................................................................. 7-14

7-12

Magnetic Chip Detector..................................................................................... 7-13

7-13

Engine Oil Dipstick ............................................................................................ 7-15

7-14

Simplified Fuel System Diagram ....................................................................... 7-16

7-15

Simplified Fuel Control System......................................................................... 7-18

7-16

Fuel Pressure Annunciator................................................................................. 7-20

7-17

Fuel Flow Indicator............................................................................................. 7-20

7-18

Control Pedestal (Typical).................................................................................. 7-21

7-19

Control Levers..................................................................................................... 7-22

7-20

Engine Display .................................................................................................... 7-23

7-21

ITT Reading......................................................................................................... 7-24

7-22

Torquemeter......................................................................................................... 7-24

7-23

Gas Generator Tachometer................................................................................ 7-24

7-24

Engine Limits Chart ........................................................................................... 7-25


7-iii

7 POWERPLANT

Figure


7 POWERPLANT

7-25

Overtorque Limits............................................................................................... 7-26

7-26

Overtemperature Limits..................................................................................... 7-27

7-27

In-Flight Engine Data Log................................................................................. 7-27

7-28

Hartzell Propeller................................................................................................ 7-29

7-29

Propeller System Complete................................................................................ 7-30

7-30

Propeller Blade Angle Diagram........................................................................ 7-31

7-31

Propeller Pitch Diagram..................................................................................... 7-32

7-32

Primary Governor ............................................................................................... 7-33

7-33

Complete Propeller System................................................................................ 7-33

7-34

Propeller Onspeed Diagram .............................................................................. 7-34

7-35

Propeller Overspeed Diagram........................................................................... 7-35

7-36

Propeller Underspeed Diagram ........................................................................ 7-35

7-37

Low Pitch Stop Diagram .................................................................................... 7-37

7-38

GROUND FINE Range and REVERSE Diagram ....................................... 7-38

7-39

Propeller Positioning—Flight Idle to Ground Low Pitch Stop ..................... 7-40

7-40

King Air 350 Ground Idle Stop Electrical Circuit .......................................... 7-43

7-41

Overspeed Governor Diagram.......................................................................... 7-44

7-42

Power Levers ...................................................................................................... 7-45

7-43

Propeller Control Levers.................................................................................... 7-46

7-44

Autofeather Diagram—Armed ......................................................................... 7-47

7-45

Autofeather Diagram—Test .............................................................................. 7-48

7-46

Autofeather Diagram—Left Engine Failure Armed ...................................... 7-49

7-47

Autofeather Test Diagram (Right Engine)—Low Power and Feathering... 7-49

7-48

Propeller Synchrophaser System ....................................................................... 7-51

7-iv



7 POWERPLANT

CHAPTER 7 POWERPLANT

INTRODUCTION In-depth knowledge of the powerplant and propeller systems is essential to good power management. Operating within the design parameters extends engine life and ensures safety. To better equip the pilot for effective power management, this chapter describes the basic components of the engines and propellers along with their limits. It also discusses details of engine operation so the pilot can familiarize himself with normal and abnormal conditions.

GENERAL ENGINE RATINGS In turboprop engines, power is measured in shaft horsepower (SHP) and equivalent shaft horsepower (ESHP). SHP is determined by propeller rpm and torque applied to turn the propeller shaft.

Power transmitted through the propeller shaft, however, is only a portion of the total thrust created by the engine. Hot exhaust gases exiting the engine also develop some kinetic energy similar to a turbojet engine. This additional thrust created by the exhaust amounts to about 10% of the total engine


7-1


horsepower. ESHP is the term applied to the total horsepower delivered—including the exhaust thrust.

7 POWERPLANT

Turboprop engine specifications usually show both ESHP and SHP along with limiting ambient temperatures. (Figure 71) lists the engine rating and temperatures and (Figure 7-2) illustrates the various engine sections.

ENGINE STATIONS

at a specific point, the appropriate station number is used. For example, temperature of the airflow measured between the compressor and first stage power turbine at engine station number 5 is called T 5 , which is read in the cockpit as ITT. Engine bleed air after t h e c e n t r i f u g a l c o m p re s s o r s t a g e a n d prior to entering the combustion chamber i s r e f e r r e d t o a s P 3 a i r . Th i s a i r i s for cabin heat, pressurization, and the pneumatic system.

To identify points in the engine, station n u m b e r s a re e s t a b l i s h e d . To r e f e r t o pressure or temperature in the airflow path

Figure 7-1. PT6A-60A Specifications

7-2


SINGLE-STAGE COMPRESSOR TURBINE

EXHAUST OUTLET COMPRESSOR SECTION POWER SECTION FOR TRAINING PURPOSES ONLY

INTAKE AIR

COMBUSTION SECTION

ENGINE AIR INLET

7-3

Figure 7-2. Engine Cutaway

7 POWERPLANT


TWO-STAGE POWER TURBINE


POWERPLANT

ENGINE TERMS Several basic terms aid in the general understanding of the PT6A series engines: • N 1 or N g —Gas generator rpm in percent of turbine speed

7 POWERPLANT

• N p —Propeller rpm • N 2 or N f —Power turbine rpm (not indicated on engine instruments) • P 2.5 —Air pressure between engine stations 2 and 3. Also referred to as axial stage air or compressor interstage air • P 3 —Air pressure at engine station 3; the source of bleed air used for some aircraft systems • ITT or T5—Interstage turbine temperature in degrees centigrade at engine station 5.

The powerplant for the King Air 350 is the Pratt and Whitney Canada PT6A-60A freeturbine-turboprop engine that drives a fourbladed propeller (Figures 7-3). The engine is flat-rated to 1,050 shaft horsepower. Th e p o w e r p l a n t s a r e e q u i p p e d w i t h conventional, four-blade, full-feathering, reversing, constant-speed propellers. The propellers mount on the output shaft of the engine reduction gearbox. Engine oil pressure controls the propeller pitch and speed through single-action, engine-driven propeller governors. The propellers feather automatically when the engines are shut down. They unfeather when the engines are started.

Figure 7-3. PT6A-60A Powerplant Installation

7-4



The PT6A-60A engine consists basically of a free-turbine, reverse-flow engine that d r i v e s a p ro p e l l e r t h ro u g h p l a n e t a r y gearing. The term free-turbine refers to the turbine sections of the single-stage engine. There are two turbine sections: the compressor turbine drives the engine compressor and accessories and the dualpower turbine drives the power section and propeller. The power turbine section has no physical connection to the compressor turbine. The compressor and power turbines, mounted on separate shafts, are driven in opposite directions by the gas flow across them.

The term reverse flow refers to airflow through the engine. Inlet air enters the compressor at the aft end of the engine. It then moves forward through the combustion section and the turbines. Finally, it is exhausted at the front of the engine.

Engine Modular Concept An important feature of the PT6A-60A engine is its modular construction. The engine is basically divided into two modules: a gas generator section and a power section (Figure 7-4). The gas generator section includes the compressor and the combustion section. Its function is to draw air into the engine and add energy to it in the form of burning fuel to produce the gases necessary to drive the compressor and power turbines.

Figure 7-4. Engine Modular Concept


7-5

7 POWERPLANT

GENERAL PRINCIPLES


7 POWERPLANT

The function of the power section is to convert the gas flow from the gas generator section into mechanical action to drive the propeller. An integral planetary gearbox converts the high speed and low torque of the power turbine to the low speed and high torque required at the propeller. The reduction ratio from power turbine shaft rpm (N f ) to propeller rpm (N p ) is approximately 17.6:1. The engine requires a minimum of maintenance. A hot section inspection (HSI) is u s u a l l y c a r r i e d o u t a t m i d - T B O. Th i s involves splitting the engine between the compressor and power turbines. Since it is not necessary to remove the engine from the aircraft to carry out the HSI, the inspection is both simple and fast. Th e m o d u l a r d e s i g n a l l o w s c o m p l e t e replacement of either the gas generator section or the combustion section independently of the other section. This permits easy maintenance, modular overhaul, and onwing HSI.

GENERAL OPERATION Another important feature of the PT6A60A engine is the reverse flow. Inlet air enters the rear of the engine through an annular plenum chamber formed by the compressor inlet case. The air is directed forward to the compressor. (Figure 7-5). The compressor consists of three axial stages combined with a single centrifugal stage. They are assembled as an integral unit on a common shaft. A row of stator vanes between each stage of compression diffuses the air, raises its static pressure, and then directs it to the next stage of compression. The compressed air passes through diffuser tubes that turn the air through 90° in direction and convert velocity to static pressure. The diffused air then passes through straightening vanes to the annulus surrounding the combustion chamber liner.

7-6

The flow of air changes direction 180° as it enters and mixes with fuel in the combustion chamber. The combustion chamber liner has varying size perforations that allow entry of compressor delivery air. Approximately 25% of the air mixes with fuel to support combustion. The remaining 75% enters the flame in the combustion chamber can and internally cools the engine. The fuel/air mixture is ignited. The resultant expanding gases are directed to the t u r b i n e s. Th e u n i q u e l o c a t i o n o f t h e combustion chamber liner using flow reversal eliminates the need for a long shaft between the compressor and the compressor turbine.This reduces the overall length and weight of the engine. For ease of starting, fuel is injected into the combustion chamber liner through 14 simplex nozzles arranged in two sets. A dual fuel manifold of primary and secondary transfer tubes and adapters supplies the fuel. Fo r s t a r t i n g o n l y, t w o s p a r k i g n i t e r s that protrude into the liner ignite the f u e l / a i r m i x t u r e. A f t e r s t a r t i n g , t h e igniters are turned off bcause combustion is self-sustaining. The resultant gases expand from the liner, reverse direction in the exit duct zone, and pass through the compressor turbine inlet guide vanes to the single-stage compressor drive turbine. The guide vanes ensure that the expanding gases impinge on the turbine blades at the correct angle with minimum loss of energy. The expanding gases are then directed forward to drive the power turbine section.


7 POWERPLANT


Figure 7-5. Engine Gas Flow and Stations


7-7


The compressor turbine extracts approximately 60% of the energy from the combustion gases. The power turbines extract the remaining energy. The dual-stage power turbine consists of inlet guide vane and turbines that drive the propeller shaft through a reduction gearbox. 7 POWERPLANT

The compressor and power turbines are in the approximate center of the engine with their respective shafts extending in opposite directions. This feature simplifies the installation and inspection procedures.

If the compressor bleed valve remains closed at low N 1 speeds, compressor stalls would result as the engine attempts to accelerate to takeoff power. If the valve remains open at high N 1 speeds, ITT would be higher than normal and torque considerably lower than normal. This would reduce power output as the engine becomes temperature-limited at reduced torque. Therefore, at both low speeds and high speeds, proper compressor bleed valve operation is critical to normal engine operation.

The exhaust gas from the power turbines is directed through an annular exhaust plenum to the atmosphere through twin opposed exhaust ports provided in the exhaust duct.

Jet-Flap Intake System

ENGINE AIRFLOW

In this engine, it is incorporated in a jet flap system. A jet flap, or slot, is machined into one side of each hollow strut that secures the accessory section to the compressor section of the engine.

Compressor Bleed Valve The compressor bleed valve is a pneumatic piston that references the pressure differential between the axial and centrifugal stages. Looking forward, the valve is at the 3 o’clock position. The function of this valve is to prevent compressor stalls and surges in the N 1 rpm range. At low N 1 rpm, the compressor axial stages produce more compressed air than the centrifugal stage can use. To compensate for this excess airflow at low rpm, the open compressor bleed valve overboards or bleeds axial-stage air (P 2.5 ). This reduces back pressure on the axial stages (Figures 7-5 and 7-6). The pressure relief helps prevent compressor stall. At low N 1 rpm, the valve is open. At takeoff and cruise above approximately 90% N 1 rpm, the bleed valve is closed.

7-8

A unique feature of the PT6A-60A engine is its efficient utilization of P 2.5 air. In most other PT6A engines, it is ported overboard.

The jet flap intake system (Figure 7-6) functions as a variable inlet guide vane without variable geometry. Each hollow core provides a passageway for compressor interstage air (P 2.5 ) to exit through the narrow slot parallel to the engine centerline. The interstage air passing into the intake zone provides a swirl effect on inlet air entering the compressor. This pre-swirl effect improves low speed compressor characteristics. It also eliminates one of the compressor bleed valves found in most of the PT6A series engines.

Swing Check Valve A swing check valve is on the right side at the 3 o’clock position on the compressor bleed valve cover. It is a plate valve hinged at the upper edge and capable of pivoting through approximately 90°.



IGNITION SYSTEM The combustion chamber has two sparktype igniters to provide positive ignition during engine start.

Wh e n t h e I G N I T I O N A N D E N G I N E START switch is moved to the ON position, these igniters activate. Although the engine is equipped with two igniters, it needs only one to start. The system is designed so that if one igniter is open or shorted, the remaining igniter continues to function. Because combustion is self-sustaining, the igniters may be turned off once the engines start. Move the IGNITION AND ENGINE START switch to the OFF position.

Figure 7-6. Jet-Flap, Compressor Bleed Valve, and Swing Check Valve


7-9

7 POWERPLANT

The valve relieves excess P 2.5 pressure that can not be used by the jet flap system when the compressor bleed valve is open.


The spark ignition provides the engine with an ignition system capable of quick lightups over a wide temperature range.

Components and Controls

7 POWERPLANT

The system consists of an airframe-mounted ignition exciter, two individual high tension cable assemblies, and two spark igniters. It is energized from the aircraft nominal 28-volt DC supply. The system operates in the 9- to 30-volt range. The igniter control box produces up to 3,500 volts. Switches for start and ignition are are on the pilot left subpanel (Figure 7-7). The IGNITION AND ENGINE START switches have three positions: ON, OFF, and STARTER ONLY. The ON position activates both the starter and igniters. The STARTER ONLY position is a holddown position, spring- loaded to center (OFF). It only provides motoring to clear the engine of unburned fuel. With the switch in this position, there is no ignition.

The ignition system features an automatic b a c k u p f o r e m e r g e n c i e s. Th e AU TO IGNITION switches should be moved to the ARM position in turbulence, precipitation, and icing conditions. If engine torque falls below approximately 17% and auto-ignition is armed, the igniters automatically energize to attempt a start if an engine flames out. Th e g r e e n a d v i s o r y I G N I T I O N O N annunciators illuminate when the system is armed.

ACCESSORY SECTION All of the engine-driven accessories e x c e p t t h e p ro p e l l e r t a c h o m e t e r a n d propeller governors are mounted on the accessory gearbox. The accessory gearbox is at the rear of the engine (Figures 7-8). The compressor shaft (N 1 ) drives the accessories through a coupling shaft.

Figure 7-7. Engine Start and Ignition Switches

7-10


7 POWERPLANT


Figure 7-8. Typical PT6A Engine


7-11


The lubricating and scavenge oil pumps are mounted inside the accessory gearbox. Two scavenge pumps are externally mounted.

The starter/generator, high-pressure fuel pump, Ng tachometer generator, and other optional accessories are mounted on pads on the rear of the accessory drive case. There are several such mounting pads, each with its own different gear ratio (Figures 7-9 and 7-10).

7 POWERPLANT

Figure 7-9. Front and Rear Accessory Drive

Figure 7-10. Accessory Gearbox Geartrain

7-12



The oil cooler is thermostatically controlled to maintain the desired oil temperature. Another externally mounted unit, the oil-to-fuel heat exchanger, uses hot engine oil to heat fuel before it enters the engine fuel system. This prevents icing at the pump filter.

LUBRICATION SYSTEM The PT6A engine lubrication system has a dual function (Figure 7-11). Its primary function is to cool and lubricate the engine bearings and bushings. Its second function is to provide oil to the propeller governor and propeller reversing control system.

A magnetic chip detector is in the bottom of each engine nose gearbox (Figure 7-12). This detector activates a yellow caution L C H I P D ET E C T o r R C H I P D ET E C T annunciator on the annunciator panel to alert the pilot of possible oil contamination.

The main oil tank houses a gear- t y p e e n g i n e - d r i v e n p r e s s u r e p u m p, a n o i l pressure regulator, and an oil filter. The engine oil tank is an integral part of the compressor inlet case and is in front of the accessory gearbox. As oil is pumped from the tank, it passes through the pressure and temperature sensing bulbs mounted on or near the rear accessory case. The oil then proceeds to the various bearing compartments and nose case through an external oil transfer line below the engine.

Illumination of a CHIP DETECT annunciator indicates possible metal contamination in the engine oil supply. Although the annunciator indicates a possible or pending engine failure, illumination of a CHIP DETECT annunciator is not in itself cause for an engine to be shut down. Monitor engine parameters for abnormal indications.

Scavenge oil returns from the nose case and bearing compartments through the g e a r- t y p e o i l s c a v e n g e p u m p s i n t h e accessory case, through external oil transfer lines, and through the external oil cooler below the engine.

If parameters are abnormal, a precautionary shut down may be made at the pilot’s discretion. After illumination of a CHIP DETECT annunciator, determine cause of malfunction and correct prior to the next flight.

VALVE

MAGNETIC POLES

PREFORMED PACKING VALVE SEAT PREFORMED PACKING

ADAPTER ASSEMBLY

PREFORMED PACKINGS

INSULATION

ADAPTER RETAINING NUT

VALVE HOUSING

DETECTOR HOUSING

ELECTRICAL CONNECTOR

Figure 7-12. Magnetic Chip Detector


7-13

7 POWERPLANT

Magnetic Chip Detector

Components and Operation

7-14



Figure 7-11. Engine Lubrication Diagram

7 POWERPLANT


The oil tank has a filler neck and integral quantity dipstick housing. The cap and dipstick are secured to the filler neck that passes through the gearbox housing and accessory diaphragm and into the tank. The markings on the dipstick indicate the number of U.S. quarts of oil less than full (Figure 7-13). The engine oil system has a total capacity of four U.S. gallons including the 2.5 gallon oil tank. Maximum oil consumption is one quart every10 hours of operation. Normal oil consumption may be as little as one quart per 50 hours of operation. Most PT6A engines normally seek an oil level of one to two quarts down on the dipstick with hot oil, and approximately one quart lower than that when oil is cold. Do not overfill.

When adding oil between oil changes, do not mix types or brands of oil due to the possibility of chemical incompatibility and loss of lubricating qualities. A placard inside the engine cover shows the brand and type of oil needed. Although the preflight checklist requires checking the oil level, the best time to check oil quantity is shortly after shutdown because oil levels are most accurately indicated at that time. Oil level checks during preflight may require motoring the engine for a brief period for an accurate level reading. When gas generator speeds are above 72% N 1 , normal oil pressure is between 90 and 135 psi.

Figure 7-13. Engine Oil Dipstick


7-15

7 POWERPLANT

Servicing


ENGINE FUEL SYSTEM

• High-pressure fuel pump

The fuel control system for PT6A-60A engines is a fuel governor that increases or decreases fuel flow to the engine to maintain selected engine operating speeds. The engine fuel control system consists of the main components shown in the block diagram (Figure 7-14).

• Fuel control unit

7 POWERPLANT

Components These components include the following: • Engine-driven boost pump

• Fuel flow transmitter • Minimum pressure flow valve • Flow divider • Dual fuel manifolds with 14 simplex nozzles The engine-driven boost pump operates when the gas generator shaft (N 1) is turning to provide sufficient fuel head pressure (approximately 30 psi) to the high-pressure pump. This prevents cavitations.

• Firewall fuel filter and pressure switch • Oil-to-fuel heat exchanger

Figure 7-14. Simplified Fuel System Diagram

7-16



After fuel passes through the oil-to-fuel heat exchanger, it flows into the highpressure, engine-driven fuel pump and on into the fuel control unit (FCU). The high-pressure fuel pump is an enginedriven, gear-type pump with an inlet and outlet filter. Flow rates and pressures vary with gas generator (N 1 ) rpm and FCU operation. The high-pressure pump supplies fuel up to a maximum pressure of 1,050 psi to the fuel-receiving side of the FCU. Its primary purpose is to provide sufficient pressure at the fuel nozzles for a good spray pattern in all modes of engine operation. A fuel-purge line positioned at the output side of the high-pressure fuel pump constantly directs a small amount of fuel back to the gravity-feed line between the wing and nacelle tanks. This ensure the FCU stays clear of vapors and bubbles. Also located in the FCU is the pump u n l o a d i n g va l v e. Th e c o n d i t i o n l e v e r controls this valve. It is either open to unload the pressure or closed. There is no intermediate position. The minimum pressurizing valve is located in conjunction with the flow divider. It blocks fuel flow during starts until fuel pressure builds sufficiently to maintain a proper spray pattern in the combustion chamber. About 100 psi is required to open t h e m i n i m u m p re s s u r i z i n g va l v e. Th e engine-driven high-pressure fuel pump maintains this required pressure. If the pump fails, the valve closes and the engine flames out.

For starting, fuel flows initially through the flow divider to the primary fuel spray nozzles in the combustion chamber. As the engine accelerates through approximately 35 to 40% N1, fuel pressure increases sufficiently to also supply the secondary fuel nozzles. At this time, all 14 nozzles are delivering atomized fuel to the combustion chamber. This progressive sequence of primary and secondary fuel nozzle operation provides cooler starts. During engine starts, there may be an increased acceleration in N 1 speed when the secondary fuel nozzles start delivering fuel.

Fuel Manifold Purge System A fuel manifold purge system disposes of residual fuel in the flow divider and fuel manifold after engine shutdown. It consists of a P 3 pressure tank with connections for P 3 air input at one end and a discharge to the flow divider at the other end. During normal engine operation, P 3 air enters the tank through a check valve to pressurize the tank. Fuel pressure against the discharge check valve prevents the air from escaping as long as the engine is running. As fuel pressure drops to zero during shutdown, P3 air escapes through the flow divider into the fuel manifold and nozzles. The airfow pushes any residual fuel into the combustion chamber where it is burned. As a result, the pilot may notice a one- to two-second delay in initial engine spool down after shutdown.


7-17

7 POWERPLANT

The oil-to-fuel heat exchanger uses warm e n g i n e o i l t o m a i n t a i n a d e s i re d f u e l temperature at the fuel pump inlet.This prevents icing at the pump filter. It occurs automatically rquiring no pilot action.


• Power lever that selects gas generator speed (N 1 ) between idle and maximum through the 3D cam, cam follower lever, and fuel valve

Fuel Control Unit The fuel control unit (FCU) has multiple functions. Its main purpose is to meter the proper fuel amount to the nozzles in all modes of engine operation.

• Flyweight governor that controls fuel flow to maintain selected speed

The fuel control consists of the following major components (Figure 7-15): 7 POWERPLANT

• Pneumatic bellows that control the acceleration schedule and act to reduce gas generator speed if a propeller overspeed occurs

• Condition lever that selects start, low idle, and high idle functions

FUEL CONDITION LEVER

ULTIMATE RELIEF VALVE

HIGH IDLE CAM PO PUMP UNLOADING VALVE

FUEL INLET (FROM OIL-TO-FUEL HEATER)

BYPASS REG VALVE

P2

74-MICRON FILTER

P1 FUEL VALVE FOLLOWER

BYPASS VALVE

PT

PO

3D CAM FOLLOWER

POWER LEVER

P1 PZ FILTER BYPASS VALVE

P3

P2

NG GOVERNOR

FUEL PUMP

MIN FLOW ADJ (P3) SENSOR BELLOWS ASSEMBLY

P3 FILTER

PY MINIMUM PRESSURIZING AND SHUTDOWN VALVE

P3 AIR

NF GOVERNOR

FUEL FLOW DIVIDER AND DUMP VALVE

LEGEND P1 UNMETERED PUMP DELIVERY FUEL P2 METERED FUEL P3 COMPRESSOR DISCHARGE AIR PO BYPASS FUEL PT FUEL SERVO PRESSURE PY GOVERNING AIR PRESSURE PZ INTERMEDIATE FUEL PRESSURE

Figure 7-15. Simplified Fuel Control System

7-18

10-MICRON FILTER



A minimum flow adjustment set to approximately 90 pounds/hour guarantees s u ff i c i e n t f u e l f l o w t o s u s t a i n e n g i n e operation at minimum power. The FCU is mounted on the rear flange of the fuel pump. A splined coupling between the pump and the FCU transmits a speed signal, proportional to gas generator shaft speed (N1), to the governing section in the FCU. The FCU determines the fuel schedule for the engine to provide power required by controlling gas generator speed. Engine power output is directly dependent upon gas generator speed (N 1 ), which is controlled by regulating the amount of fuel to the combustion section of the engine. Compressor discharge pressure (P 3 ) is sensed by the FCU is used to establish acceleration fuel flow limits. This fuel limiting function prevents overtemperature conditions in the engine during starting and acceleration.

FCU Operation A b r i e f, s i m p l i e d d i s c u s s i o n o f F C U operation follows. For detailed description and operation, refer to the Pratt & Whitney Maintenance Manual for this engine. The condition lever selects the LOW IDLE to HIGH IDLE N 1 speeds. The power lever selects speeds between idle and maximum, 104% N 1 and positions a 3D cam in the FCU. The cam, through a cam follower and lever, determines fuel flow corresponding to the selected N 1 speed. The gas generator (N 1 ) governor that c o n t ro l s e n g i n e s p e e d c o n t a i n s t w o flyweights mounted on an engine-driven

ballhead. The flyweight governor is the feedback element of the speed select system. It controls the on-speed condition by positioning the 3D cam in response to speed variations in the gas generator. As N 1 speed increases or decreases, the resulting flyweight action changes the 3D cam setting. This, in turn, changes the fuel flow valve setting to maintain the selected N 1 speed. The N 1 governor maintains these forces in balance continually so the axial position of the 3D cam always represents engine speed. The cam follower and arm transmit motion of the 3D cam to the fuel valve. As the 3D cam moves upward, fuel flow to the engine is increased and N 1 speed increases. Downward movement of the 3D cam decreases fuel flow and N 1 speed. The N 1 governor, in response to variations in power lever position, maintains N 1 speed. The governor adjusts fuel flow as required. Compressor discharge pressure (P 3 air) is a second input affecting the fuel flow valve position during acceleration or deceleration to maintain the selected speed condition of the gas generator. An increase in P 3 causes the fuel flow valve to increase f u e l f l o w i n re s p o n s e t o i n c re a s e d P 3 pressure until N 1 speed is stabilized. A decrease in P 3 causes the fuel flow valve to decrease fuel flow until N1 speed is stabilized at the lower selected value.

Overspeed Condition In an overspeed condition, increasing pressure by the governor flyweights moves the 3D cam downward. This results in decreased fuel flow from the fuel flow valve to the engine. A balance point is reached when the N1 speed is reduced to the selected speed, and the cam is stationary at the new speed position. Underspeed Condition In an underspeed condition, decreasing pressure by the governor flyweights moves the


7-19

7 POWERPLANT

The FCU is calibrated for starting flow rates, acceleration, and maximum power. It compares gas generator speed (N 1) with the power lever setting and regulates fuel to the engine fuel nozzles. The FCU also senses compressor section discharge pressure, compares it to rpm, and establishes acceleration and deceleration fuel flow limits.


3D cam upward. This results in increased fuel flow from the fuel flow valve to the engine until the system is in equilibrium again.

In the event of an engine-driven fuel pump (high pressure) failure, the engine flames because this high-pressure fuel is required to open the minimum pressurizing valve.

Fuel Pressure Indicators

7 POWERPLANT

If a primary engine-driven boost pump fails, the approrpiate red warning FUEL PRESS annunciator in the warning annunciator p a n e l i l l u m i n a t e s ( Fi g u re 7- 16 ) . Th e MASTER WARNING lights also flash.

Fuel Flow Indicator A transmitter in the engine fuel supply line between the FCU and the flow divider senses fuel flow information. This information displays on the multi-function display in the center of the instrume nt panel (Figure 7-17). The displays indicate fuel flow in pounds-per-hour units.

Figure 7-17. Fuel Flow Indicator

Fuel Additives

Figure 7-16. Fuel Pressure Annunciator

The FUEL PRESS annunciator illuminates when outlet pressure at the engine-driven boost pump decreases below 10 psi. Switching on the standby fuel boost pump should increase fuel pressure above 11 (±2) psi and extinguish the warning.

CAUTION Engine operation with the FUEL PRESS annunciator on is limited to 10 hours between overhaul or replacement of the engine-driven high-pressure fuel pump. Boost pump fuel pressure is needed to l u b r i c a t e, c o o l a n d p r e v e n t cavitations of the high pressure fuel pump.

7-20

Two fuel additives are approved for the King Air 350. An anti-icing additive conforming to specification MIL-I-27686 is required when flying into known forecast c o n d i t i o n s b e l o w – 4 5 ° C. I t s h o u l d b e blended in accordance with the procedures outlined in the POH. The fuel biocide Biobor JF is also approved for prevention of microorganism growth within fuel tanks and lines. It should be blended as outlined in the King Air Maintenance Manual.

ENGINE POWER CONTROL The power lever acting on the gas generator governor (N 1 ) controls torque. When the position of the power lever calls for more torque, the governor settings prevent the bleed-off of internal pressure and some of the P 3 air in the FCU. This moves the metering nozzle to allow the necessary fuel



flow into the spray nozzles to meet the power condition called for.

UP FLAPS

20 TAKEOFF AND APPROACH

60

The propeller lever adjusts the propeller governor to the desired propeller speed. The propeller maintains the set speed by varying the blade angle as more or less torque is applied.

80

1

2 CABIN CLIMB

4

THDS FT PER MIN

0 .5

6 1

2

4

40 35 30 25

100

0

T AL

0F

T

5

1 7

2

6

5

20

4

3

10

15

7 POWERPLANT

DOWN

.5

Power Management Power management is relatively simple with two primary operating limitations: temperature and torque. Engine torque and ITT operating parameters are affected by ambient temperature and altitude. During operation requiring maximum engine performance at cold temperature or low altitude, torque limits power. At hot temperature or high altitude, ITT limits power. Whichever reaches its limit first determines the power available.

Control Pedestal The control pedestal extends between pilot and copilot (Figure 7-18). The three sets of control levers are, left to right, the power levers, propeller rpm and feather levers, and the condition levers.

Power Levers The power levers control engine power from idle to takeoff power through the operation of the gas generator (N1) governor in the FCU. Increasing N 1 rpm results in increased engine power. Th e p o w e r l e v e r s h a v e t h re e c o n t ro l regions: forward thrust, ground fine, and reverse. When the levers are lifted over the IDLE gate and pulled back into the GROUND FINE range, they hold engine power at the selected idle speed and control propeller blade angle.

Figure 7-18. Control Pedestal (Typical)


7-21


7 POWERPLANT

The GROUND FINE range is normally used for taxi. When the levers are lifted over the GROUND FINE gate into the REVERSE range, they control engine power and propeller blade angle.

Condition Levers

Propeller Levers

At LOW IDLE, engine gas generator speed (N 1 ) is a minimum of 62%; at HI IDLE it is 70%. The levers can be set between these two values for any proportional speed between 62% and 70% N 1 speed.

The propeller levers are conventional in setting the required rpms for takeoff and cruise positions (Figure 7-19). The normal governing range is 1,450 to 1,700 rpm. POWER LEVERS

PROPELLER LEVERS CONDITION LEVERS

The condition levers have three positions: F U E L C U TO F F, LOW I D L E , a n d H I IDLE (Figure 7-19). At FUEL CUTOFF position, fuel flow to the engines is cut off.

Control Lever Operation The propeller, power, and condition levers control the engines from the cockpit. Both the power and condition levers are connected to the N 1 governing section of the FCU. Either lever resets the FCU to maintain a new N 1 rpm. For starting, the power levers are at IDLE position. Once the condition levers are moved to LOW IDLE, the fuel cutoff valve allows fuel to flow to the nozzles. The N 1 g o v e r n o r i s s e t a t LOW I D L E . Th e condition levers are continuously variable from LOW IDLE at 62% to HI IDLE at 70% N 1 . This variable idle speed with power levers at IDLE enhances engine cooling by maintaining a steady airflow through the engines.

Figure 7-19. Control Levers

This aircraft is equipped with both manual and automatic propeller feathering systems. To feather a propeller manually, pull the propeller lever back past the friction detent into the red and white striped section of the quadrant. To unfeather, push the lever back into the governing range. The propellers go to the feathered position when the engines shut down because of the loss of oil pressure in the propeller hub.

7-22

With the condition levers at LOW IDLE, the power levers select N 1 rpm from 62% t o 1 0 4 % , t h e m a x i m u m f o r t a k e o f f. However, if the condition levers are at HI IDLE, the power levers can select N 1 rpm only from 70% to 104%. Moving either the power levers or the condition levers changes N 1 rpm. As the power or condition levers are advanced, ITT, torque and fuel flow increases. These indicators are by-products of the N 1 speed maintained by the FCU. With power levers in a fixed position, N 1 remains constant e v e n i n a c li mb o r d e s c e n t. Ho wever, I T T, t o r q u e a n d f u e l f l o w v a r y w i t h altitude, ambient air temperature, and propeller setting.



ENGINE INSTRUMENTS Figure 7-20 presents the engine display along with their operating limits.

TORQUE METER

400-820°C NORMAL OPERATING RANGE

0 TO 100% NORMAL OPERATING RANGE

820°C MAXIMUM CONTINUOUS LIMIT

100% MAXIMUM LIMIT

7 POWERPLANT

INTERSTAGE TURBINE TEMPERATURE GAGE

1000°C MAXIMUM STARTING ONLY LIMIT

PROPELLER TACHOMETER (NP SPEED) NO LIMITATIONS MARKINGS

1450-1700 RPM NORMAL OPERATING RANGE 1700 RPM MAXIMUM LIMIT

GAS GENERATOR TACHOMETER (N1 SPEED) 62-104% NORMAL OPERATING RANGE 104% MAXIMUM LIMIT

OIL TEMPERATURE SCALE 0-99°C NORMAL OPERATING RANGE 99-110°C CAUTION RANGE 110° MAXIMUM OIL TEMPERATURE LIMIT

OIL PRESURE SCALE 60 PSI MINIMUM LIMIT 60-90 PSI CAUTION RANGE 90-135 PSI NORMAL OPERATING RANGE 200 PSI MAXIMUM LIMIT

Figure 7-20. Engine Display


7-23


ITT Gage The ITT gage monitors the interstage turbine temperature at Station 5 (Figure 721). ITT is a prime limiting indicator of the amount of power available from the engine under varying ambient temperature and altitude conditions.

Torque is measured by a hydromechanical torquemeter in the first stage of the reduction gearcase. Rotational force on the first-stage ring gear compresses oil in the torquemeter chamber. The difference between the torquemeter chamber pressure a n d re d u c t i o n g e a r i n t e r n a l p re s s u re accurately indicates the torque being produced at the propeller shaft.

7 POWERPLANT

The torquemeter transmitter measures this torque and displays it on the center multifunction display. Figure 7-21. ITT Reading

Th e ITT readout is in the center of the engine display on the multifunction display. The normal operating range is 40 0°C to 820°C. These limits also apply to maximum continuous power.

Gas Generator Tachometer Th e g a s g e n e r a t o r ( N 1 ) t a c h o m e t e r measures rotational speed of the compressor shaft in percent rpm,based on 37,50 0 rpm at 10 0% (Figure 7-23). The N 1 indicator is self-generating.

The maximum starting-only temperature is 1,0 0 0°C .This starting limit is limited to five seconds. The engines can be damaged if limiting temperatures are exceeded. Figure 7-23. Gas Generator Tachometer

Torquemeter The torquemeter constantly measures rotational force applied to the propeller shaft (Figure 7-22). The maximum permissible sustained torque is 10 0%. Cruise torques vary with altitude and temperature.

The tachometer generator sensing unit in the engine accessory section supplies the information to the N 1 display to indicate the percent of N 1 revolutions. Maximum continuous gas generator speed is limited to 39,0 0 0 rpm, which is 104% on the N 1 indicator.

Figure 7-22. Torquemeter

7-24



Aircraft and engine limits are described in the Limitations section of the POH (Figure 7-24). These limitations have been approved by the Federal Aviation Administration and must be observed in the operation of the King Air 350. OPERATING TORQUE MAXIMUM CONDITION % ITT (1) °C

The ENGINE OPERATING LIMITS chart g i v e s t h e m a j o r o p e ra t i n g l i m i t s. Th e P OW E R P LA N T I N ST RU M E N T MARKINGS chart lists the minimum, normal, and maximum limits. Figure 7-20 for a l i s t i n g o f t h e o p e ra t i n g ra n g e s f o r torque, ITT, propeller, gas generator, oil temperature and pressure. GAS GENERATOR RPM % N1

PROP RPM N2

OIL OIL PRESS TEMP PSI (2) °C (3) (4) 0 to 200 -40 (min)

STARTING

---

1000 (5)

---

---

IDLE

---

750 (6)

62 (min)

1050 (min)

60 (min)

TAKEOFF

100 (10)

820

104

1700 (9)

90 to 135 0 to 110

MAX CONT

100 (10)

820

104

1700 (9)

90 to 135 0 to 110

CRUISE CLIMB

(7) (10)

785

104

1700 (9)

90 to 135 0 to 110

MAX CRUISE

(7) (10)

820

104

1700 (9)

90 to 135 0 to 110

---

760

---

1650

90 to 135 0 to 99

156 (8)

850 (8)

104

1870 (8)

MAX REVERSE TRANSIENT

200

7 POWERPLANT

ENGINE LIMITATIONS

-40 to +110

0 to 110

FOOTNOTES: (1) Torque limit applies within range of 1000 - 1700 propeller rpm (N2). Below 1000 propeller rpm, torque is limited to 62%. (2) Normal oil pressure is 90 to 135 psi at gas generator speeds above 72%. With engine torque below 62%, minimum oil pressure is 60 psi at normal oil temperature (60 degrees to 70 degrees C). Oil pressures under 90 psi are undesirable. Under emergency conditions, to complete a flight, a lower oil pressure limit of 60 psi is permissible at a reduced power, not to exceed 62% torque. Oil pressures below 60 psi are unsafe and require that either the engine be shut down or a landing be made at the nearest suitable airport, using the minimum power required to sustain flight. Fluctuations of plus or minus 10 psi are acceptable. During extremely cold starts, oil pressure may reach 200 psi. In flight, oil pressures above 135 psi but not exceeding 200 psi are permitted only for the duration of the flight. (3) A minimum oil temperature of 55°C is recommended for fuel heater operation at take-off power. (4) Oil temperature limits are -40°C and +110°C. However, temperatures between 99°C and 110°C are limited to a maximum of 10 minutes. (5) This value is time limited to 5 seconds. (6) High ITT at ground idle may be corrected by reducing accessory load and/or increasing N1 rpm. (7) Cruise torque values vary with altitude and temperature. (8) These values are time limited to 20 seconds. (9) To account for power setting accuracy and steady state fluctuations, inadvertent propeller RPM excursions up to 1735 RPM are time limited to 7 minutes. (10) To account for power setting accuracy and steady state fluctuations, inadvertent torque excursions up to 102% is time limited to 7 minutes.

Figure 7-24. Engine Limits Chart FOR TRAINING PURPOSES ONLY

7-25


During engine start, temperature is the most critical limit. The ITT starting limit of 1,000°C is limited to five seconds. For this reason, it is helpful during starts to keep the condition lever out of the LOW IDLE detent so that the lever can be quickly pulled back to FUEL CUTOFF. 7 POWERPLANT

Monitor oil pressure and oil temperature. During the start, oil pressure should come up to the minimum 60 psi quickly, but should not exceed the maximum of 200 psi. During normal operation the oil temperature and pressure should be green, from 90 to 135 psi. Fluctuations of ± 10 psi are acceptable. Oil pressure between 60 and 90 psi is undesirable; it should be tolerated only for completion of the flight, and then only at a reduced power setting. Oil pressure below 60 psi is unsafe; it requires that either the engine be shut down or that a landing be made as soon as possible using minimum power required to sustain flight. A minimum oil temperature of 55°C is recommended for oil-to-fuel heater operation at takeoff power. Oil temperature limits are –40°C and +110°C during IDLE, and 0°C to +110°C during normal o p e r a t i o n s. H o w e v e r , t e m p e r a t u r e s between +99°C and +110°C are limited to a maximum of 10 minutes.

the IDLE limit of 750°C, the N 1 loads may be restored as desired as long as ITT stays below 750°C. During normal flight o p e r a t i o n s, t h e I T T s h o u l d n e v e r b e allowed to exceed the maximum continuous limit of 820°C . During the climb, torque decreases; ITT may increase slightly. The cruise climb ITT limit is not placarded. Torque, ITT, N 1 , and propeller limits are the same in maximum cruise as they are for takeoff; however, cruise torque values vary with altitude and temperature. Transient limits provide buffers for surges during engine acceleration. Torque and ITT have an allowable excursion duration of 20 seconds. A momentary peak of 156% and 850°C is allowed for torque and ITT respectively during acceleration. Th e OV E RT O R Q U E L I M I T S c h a r t (Figure 7-25) shows actions required if t o r q u e l i m i t s a re e x c e e d e d u n d e r a l l conditions. If the torque limits are exceeded for more than a few minutes, the gearbox can be damaged. The chart shows the specific limits and action required if they are exceeded.

During ground operations, ITT temperatures are critical. They can be controlled by the N 1 rpm, air conditioning, P3 use, and the generator load. With the condition levers at LO IDLE, high ITT can be corrected by reducing the generator and other N 1 loads, then increasing the N 1 rpm by advancing the condition levers. The air conditioner, for example, draws a heavy load on the r i g h t e n g i n e N 1. I t m a y h a v e t o b e temporarily turned off. At approximately 70% N 1 rpm, the HI IDLE condition lever position normally reduces the ITT. Once ITT is reduced below

7-26

Figure 7-25. Overtorque Limits



The Generator Limits table in the POH s h o w s l i m i t s f o r g ro u n d o p e ra t i o n a t various N 1 rpms. Any time these limits are exceeded, the accessory load must be reduced or N 1 increased to the limits shown in the table. A generator load of 75% is maximum for ground operation at an N 1 of between 62% and 70%.

The overtemperature chart (Figure 7-26) shows the specific actions required if ITT limits are exceeded during starting conditions. For Area A, determine and correct the cause of overtemperature. If it is during a start, have the engine visually inspected through the exhaust duct, then record the action in the engine logbook (Figure 7-27).

7 POWERPLANT

The Inflight limits are 100% generator load from sea level to 34,000 feet. Above 34,000 feet, maximum sustained generator load limit is 95%. Maximum continuous and maximum cruise share the same generator limits, but due to N 1 loading, certain limits must not be exceeded as indicated in the Before Takeoff (Final Items) checklist in the POH.

Figure 7-26. Overtemperature Limits

Overtemperature in Area B requires that a hot section inspection be performed. During a hot section inspection, the combustion chamber and turbine areas and components are examined and replaced. Parts should be repaired or replaced as necessary.

Figure 7-27. In-Flight Engine Data Log


7-27


In Area C, overtemperatures may require that the engine be returned for overhaul. Exceeding ITT limits in this area for more than a few seconds may cause extensive engine damage.

Starter Operating Temperature Limits 7 POWERPLANT

The engine starters are time-limited during the starting cycle if for any reason multiple starts are required in quick sequence. The starter is limited to 30 seconds ON then five minutes OFF for cooling before the next sequence of 30 seconds ON, five minutes OFF. After the third cycle of 30 seconds ON, the starter must stay OFF for 30 minutes. If these limits are not observed, overheating may damage the starter. The second starter cycle is used for clearing the engine of residual fuel.

Trend Monitoring During normal operations, gas turbine engines are capable of producing rated power for extended periods of time. Engine o p e ra t i n g p a ra m e t e r s, s u c h a s o u t p u t torque, interstage turbine temperature, compressor speed, and fuel flow for individual engines, are predictable under specific ambient conditions. O n P T 6 A e n g i n e s, t h e s e p r e d i c t a b l e characteristics may be taken advantage of by establishing and recording individual engine performance parameters. These parameters can then be compared periodically to predicted values to provide day-to-day visual confirmation of engine efficiency. The engine condition trend monitoring system recommended by Pratt & Whitney is a process of periodically recording engine instrument readings and then comparing them to a set of typical engine character-

7-28

istics. Such comparisons produce a set of deviations in ITT, compressor speed, and fuel flow. Readings that should be collected include the torque, ITT, compressor speed, and fuel flow. C o r r e c t t h e r e a d i n g s f o r a l t i t u d e, outside air temperature, and airspeed, if applicable.

Data Collection The trend monitoring procedure used specifies that flight data be recorded on each flying day, every five flight hours, or other flight period. Select a flight with long-established cruise, preferably at a representative altitude and airspeed. Wi t h e n g i n e p o w e r e s t a b l i s h e d a n d stabilized for a minimum of five minutes, record the following data on a form similar to this inflight engine data log (see Figure 7-27): • Indicated airspeed (IAS)—Knots • Outside air temperature (OAT)— Degrees Centigrade • Pressure altitude (ALT)—Feet • Propeller speed (N p )—RPM • Torque (T q )—Percentage • Gas generator speed (N g or N 1 )— Percent of gas generator speed • Interturbine temperature (ITT)— Degrees Centigrade • Fuel flow (WF)—Pounds per hour

PROPELLER This section on the description, operation, and testing of the propeller system should increase the pilot’s understanding of the propeller and system checks in the POH.



Oil pressure controls the propeller pitch t h ro u g h a re d u c t i o n g e a r b o x- d r i v e n propeller governor. An oil pump that is part of the propeller governor boosts engine oil pressure to move the propeller toward the low pitch (high rpm) position and into reverse (Figure 7-29). Without oil pressure to counteract the counterweights and feathering springs, the propeller blades would move into feather. Counter-weights and feathering springs move the propeller blades toward high pitch (low rpm) and into the feathered position. Because there are no high pitch stop locks, the propeller feathers after engine shutdown.

A mechanically actuated hydraulic stop determines low pitch propeller position. Power levers on the pedestal adjust the lower pitch stop position to control the GROUND FINE and REVERSE blade angles in the GROUND FINE and REVERSE range. Three governors that may control propeller rpm include the primary, overspeed, and fuel topping governors. The primary and overspeed governors use oil pressure to change propeller blade angle so that the propeller rpm is adjusted or limited. The fuel topping governor limits fuel to limit propeller rpm. The propeller control lever adjusts the p r i m a r y g o v e r n o r t h ro u g h i t s n o r m a l governing range of 1,450 to 1,70 0 rpm. If the primary governor malfunctions, the

Figure 7-28. Hartzell Propeller


7-29

7 POWERPLANT

Each engine is equipped with a conventional four-blade, constant speed, fullf e a t h e r i n g , r e v e r s i n g p r o p e l l e r . Th e propeller is mounted on the output shaft of the reduction gearbox (Figure 7-28).

7 POWERPLANT

7-30 PUSH-PULL CONTROL REVERSING CAM

SPEEDER SPRING PILOT VALVE


Py BETA VALVE

AIR BLEED LINK

MAXIMUM STOP

BETA ROD

MINIMUM GOVERNOR ADJ.

TO SUMP

FCU ARM TEST SOLENOID

HYDRAULIC LOW PITCH ADJ.

FEATHERING VALVE

OVERSPEED GOVERNOR

COUNTERWEIGHT TO SUMP

Figure 7-29. Propeller System Complete


RESET POST

overspeed governor prevents the propeller s p e e d f ro m e x c e e d i n g a p p ro x i m a t e l y 1,768 rpm.

in-corporated in a blade to increase its efficiency, blade angle is different near the hub than it is near the tip.

If the propeller blade angle cannot be changed by either the primary or overspeed governor, the fuel top p in g governor intervenes. The fuel topping governor attempts to limit the propeller rpm to 106% of selected propeller rpm when the power lever is in the forward thrust range. In the GROUND FINE and R E V E R S E r a n g e s, t h e f u e l t o p p i n g governor resets to approximately 95% of selected propeller rpm. This ensures that the primary governor remains in an underspeed condition while in the REVERSE range on the ground.

In the propellers, the cord 42 inches out f ro m t h e p ro p e l l e r ’s c e n t e r h a s b e e n selected as the position at which blade angle is measured. This position is referred to as the 42-inch station. All blade angles in this chapter are approximations based on the 42-inch station (Figure 7-30).

PRIMARY GOVERNOR

BLADE ANGLE

The primary governor mounted on top of the engine reduction gearbox converts a variable pitch propeller into a constant speed propeller. It does this by changing blade angle to maintain the propeller speed the operator has selected.

Blade angle is the angle between the chord of the propeller and the propeller’s plane of rotation. Because of the normal twist

The primary governor can maintain any selected propeller speed from 1,450 rpm to 1,700 rpm. BLADE ANGLES

79.5° (300) 79.3° (350) FEATHER

CRUISE 30°–45° –14° MAXIMUM REVERSE

+13° (300) FLIGHT LOW PITCH STOP +12° (350)

+1° (300) GROUND LOW PITCH STOP +2° (350)

–3° GROUND FINE (ZERO THRUST) 0°

Figure 7-30. Propeller Blade Angle Diagram


7-31

7 POWERPLANT



If an aircraft is in normal cruising flight with the propeller turning at 1,50 0 rpm and the pilot trims the aircraft down into a descent without changing power, the airspeed increases. This decreases the angle of attack of the propeller blades and causes less drag on the propeller. Its rpm begins to increase. 7 POWERPLANT

If the propeller has variable pitch capabilities and is equipped with a governor set at 1,50 0 rpm, the governor senses this overspeed condition and increases blade angle to a higher pitch. The higher pitch i n c re a s e s t h e b l a d e’s a n g l e o f a tt a c k , slowing it back to 1,500 rpm or onspeed. This propeller control process occurs many times per second. Likewise, if the aircraft moves from cruise to climb airspeeds without a power change, the propeller rpm tends to decrease. The governor responds to this underspeed condition by decreasing blade angle to a lower pitch (Figure 7-31). Rpm returns to its original value. Thus, the governor gives constant speed c h a ra c t e r i s t i c s t o t h e va r i a b l e p i t c h propeller. Power changes, as well as airspeed changes, cause the propeller to momentarily experience overspeed or underspeed conditions. The governor reacts to maintain the onspeed condition. Because the actions of the governor are smooth, the pilot notices few, if any, of these minor adjustments.

Primary Governor Operation Th e p ro p e l l e r l e v e r s a d j u s t p r i m a r y propeller governor settings between 1,450 rpm and 1,700 rpm. The propeller governor can select any constant propeller rpm within the range of 1,450 to 1,700. The propeller governor adjusts propeller rpm by controlling the oil supply to the propeller hub mechanism (Figure 7-32). An integral part of the primary propeller governor is the governor pump. The N p shaft drives this pump that can raise engine

7-32

Figure 7-31. Propeller Pitch Diagram

oil pressure from its normal range to a maximum of 375 psi. The greater the oil pressure sent to the propeller dome, the lower the propeller pitch and the higher the propeller rpm. Oil pressure is always trying to maintain a low pitch. The feathering springs and counterweights, however, are trying to send the propeller to feather.



A transfer gland surrounds the propeller shaft. This transfer gland allows the oil to enter and exit the propeller dome as the propeller pitch is adjusted (Figure 7-33). The primary propeller governor uses a set of rotating flyweights geared to the propeller shaft. The flyweights provide a

comparison between the desired reference speed (as requested by the propeller levers in the cockpit) and the actual speed the propeller is turning. These flyweights connect to a free-floating pilot valve. The slower the flyweights are turning in relation to the desired reference speed, the lower the position of the pilot valve. If the propeller and flyweights turn faster, the additional centrifugal force

Figure 7-32. Primary Governor

Figure 7-33. Complete Propeller System


7-33

7 POWERPLANT

Propeller control is a balancing act of opposing forces.


m a ke s t h e p i l o t va l v e r i s e i n s i d e t h e g o v e r n o r . Th e p i l o t v a l v e p o s i t i o n determines how much oil pressure is sent to the propeller dome (Figure 7-33).

Propeller Onspeed

7 POWERPLANT

If a propeller rpm of 1,500 is selected and the propeller is actually turning at 1,50 0, the flyweights are in their center or onspeed condition. The pilot valve is in the middle position (Figure 7-34). This maintains a constant oil pressure to the propeller dome to create a constant pitch and constant rpm.

Propeller Overspeed If the aircraft enters a descent or if engine power is increased without any change to the propeller levers, there is tendency for airspeed to increase and the propeller to turn faster. The flyweights, in turn, rotate faster. The additional centrifugal force makes the pilot valve rise. Notice that oil can now escape via the pilot valve (Figure 7-35).

Lower oil pressure results in a higher pitch and a reduction of propeller rpm. The propeller returns to its original rpm setting. The flyweights then slow down; the pilot valve returns to the middle position to maintain selected propeller rpm.

Propeller Underspeed If the aircraft enters a climb or if engine power is decreased without any change in the propeller controls, airspeed decreases. The propeller tends to slow down. The flyweights in the propeller governor also slow down due to a loss in centrifugal force. The pilot valve moves lower (Figure 7-36). This allows more oil pressure to the propeller dome. Higher oil pressure results in a lower pitch. This, in turn, causes an increase in propeller rpm. As the propeller increases to its original rpm setting, the flyweights speed up. The pilot valve returns to its middle or onspeed position.

Figure 7-34. Propeller Onspeed Diagram

7-34


7 POWERPLANT


Figure 7-35. Propeller Overspeed Diagram

Figure 7-36. Propeller Underspeed Diagram


7-35


The flyweights and pilot valve are always making small adjustments so that the propeller rpm is held constant by changing the propeller blade angles.

7 POWERPLANT

The cockpit propeller lever adjusts when the onspeed condition occurs. The pilot can select any constant propeller rpm from 1,450 to 1,700 rpm, that is used for takeoff. Maximum range power and recommended cruise power use 1,500 rpm. If a failure in the governor control linkage occurs, an external spring on top of the governor moves the governor adjustment to 1,700 rpm propeller speed. If the blade angle could decrease all the way to 0° or reverse, the propeller would c re a t e s o m u c h d ra g, a i rc ra f t c o n t ro l w o u l d b e d r a m a t i c a l l y r e d u c e d . Th e propeller, acting as a large disc, would create excessive drag and blank the airflow around the wing and tail surfaces. To prevent undesirable flight characteristics, a mechanism stops the governor from selecting blade angles that are too low for safety. This mechanism provides for an adjustable low pitch stop.

oil pressure to the dome. Propeller pitch then decreases as power and airspeed are reduced.

NOTE Momentary periods of underspeed are not being considered. In this situation, propeller rpm is stabilized below selected governor rpm. Assuming the propeller is not feathered or in the process of being feathered when propeller rpm is below the selected governor rpm, the propeller blade angle is at the low pitch stop. If the aircraft is on the ground, it is called the ground low pitch stop. If the aircraft is in flight, it is called the flight low pitch stop. On many aircraft, the low pitch stop is simply the low pitch limit of travel determined by propeller construction. But with a reversing propeller, the extreme travel in the low pitch direction is past 0º into reverse or negative blade angles. The low pitch stop then can be moved or repositioned when reversing is desired.

Low Pitch Stop

A mechanical linkage senses blade angle and creates the low pitch stop. The linkage closes a valve that stops the flow of oil into the propeller dome. Because oil flow causes low pitch and reversing, a low pitch stop is created when it is blocked. The low pitch stop valve, commonly referred to as the beta valve, is quite positive in its mechanical operation. Furthermore, the valve is spring-loaded to cut off the flow of oil and dump it out of the propeller dome if valve control is lost because of linkage failure.

In situations such as final approach where power and airspeed are being reduced, the primary governor cannot maintain the selected propeller rpm. With the progressive reduction of power and airspeed on final, the propeller and rotating flyweights tend to go to the underspeed condition where the pilot valve drops and increases

The propeller dome is connected by four spring-loaded polished rods to the feedback ring behind the propeller (Figure 7-37). A carbon block riding in the feedback ring transfers the movement of the latter through the propeller reversing lever to the beta valve on the governor. The initial forward motion of the beta valve blocks off

A s t h e g o v e r n o r d e c re a s e s t h e b l a d e angle, the flight low pitch stop is eventually reached where blade angle becomes fixed and cannot continue to a lower pitch. The governor is, therefore, incapable of restoring the onspeed condition. Propeller rpm decreases below the selected primary governor rpm.

7-36



The power lever controls the position of the low pitch stop. When the power lever is at IDLE or above, the flight low pitch stop is set at 12°. The ground low pitch stop is set at 2°. Bringing the power lever aft of IDLE progressively repositions the low pitch stop to smaller blade angles. The geometry of the power lever linkage through the cam box

means that power lever increments from IDLE to full forward thrust have no effect on the position of the beta valve. When the power lever is moved from IDLE into the GROUND FINE and REVERSE ranges, it pulls the reverse lever and the beta valve aft. The blade angle decreases because this action opens the beta valve to increase oil pressure to the propeller. As the blade angle decreases, the distance traveled by the propeller dome is fed back to the beta valve through the rods, ring, and reverse lever, which pulls the beta valve forward. This closes the valve and stops oil flow into t h e p r o p e l l e r d o m e. Th e b l a d e a n g l e stabilizes at the selected position. The opposite occurs when the power lever is moved forward to IDLE. The power lever pushes the reverse lever and beta valve fully forward to relieve oil pressure from the propeller dome. This increases blade

IDLE IDLE GATE ON GROUND

COUNTERWEIGHT FEATHER RETURN SPRINGS

CARBON BLOCK

+1° OR 2°

RING, ROD END

GROUND FINE

GROUND FINE GATE

GROUND FINE GATE

-3°

FEEDBACK RING

POLISHED ROD

REVERSE RETURN SPRING

LOW-PITCH STOP NUT

MAXIMUM REVERSE REVERSE

-14°

Figure 7-37. Low Pitch Stop Diagram


7-37

7 POWERPLANT

the flow of oil to the propeller. Further motion forward dumps the oil from the propeller into the reduction gearbox sump. A mechanical stop limits the forward m o t i o n o f t h e b e t a v a l v e. Re a r w a r d movement of the beta valve does not affect normal propeller control. When the propeller is rotating at a speed slower than that selected, the governor pump provides oil pressure to the propeller dome. It also decreases the pitch of the propeller blades until forward motion on the feedback ring pulls the beta valve into a position that blocks supply of oil to the propeller. This prevents further pitch reduction.


angle. As the blade angle increases, the distance traveled by the propeller dome is fed back to the beta valve through the rods, ring, and reverse lever. This pushes the beta valve aft until the port relieving oil pressure closes. The blade angle is now stabilized at the selected position.

GROUND FINE Range The region between the ground low pitch stop and –3° blade angle is the GROUND FINE range (Figure 7-38). In this range, t h e e n g i n e’ s c o m p r e s s o r s p e e d ( N 1 ) remains at the value it had when the power lever was at IDLE (62% to 70% based on condition lever position).

7 POWERPLANT

Figure 7-38. GROUND FINE Range and REVERSE Diagram

7-38


To enter the GROUND FINE range, the power lever must be lifted beyond the I D L E g a t e a n d m o v e d a f t . Wi t h a f t movement of the power levers in GROUND FINE, blade angle moves progressively from the ground low pitch stop to –3° (GROUND FINE gate). When a power lever is lifted up and over the IDLE gate into the GROUND FINE range, it is pulling aft on the top of the reverse lever. As the reverse lever moves aft, the beta valve is pushed aft to reestablish oil flow to the propeller dome. This moves the propeller blade angle below the ground low pitch stop. As the propeller blade angle continues below the ground low pitch stop, the propeller dome and feedback ring continue forward. They eventually pull the beta valve forward to the oil cutoff position.

REVERSE Range The region between –3° and –14° blade angle is the REVERSE range. In this range, N 1 progressively increases to a maximum value of 87 ±1% while blade angle decreases. To enter the REVERSE range, the power lever must be lifted beyond the GROUND FINE gate and moved aft. With aft movement of the power levers to REVERSE, blade angle progressively decreases from –3° (GROUND FINE gate) to –14° (maximum reverse).

When a power lever is moved forward, away from or out of the GROUND FINE or REVERSE ranges toward IDLE, it pushes the reverse lever forward. This, in turn, pulls the beta valve fully forward. This opens a port so oil is dumped from the p ro p e l l e r d o m e t o t h e n o s e c a s e. Th e propeller blade angle then increases until the rods and ring moving aft with the propeller dome have pushed the reverse lever and the beta valve far enough aft to cut off the oil.

PROP PITCH Annunciators The white status L PROP PITCH and R PROP PITCH annunciators indicate propeller blade angle has decreased below the flight low pitch stop. The FAA requires these ammunciators to alert the pilot any time the propeller pitch changes more than 8° below the flight low pitch stop without any direct pilot action. The system uses a magnetic proximity sensor to sense the position of the feedback ring and, thereby, the position of the propeller dome and blade angle.


7-39

7 POWERPLANT



Low Pitch Stop Operation (Figure 7-39) illustrates the sequence of flight idle to ground stop low pitch stop.

7 POWERPLANT

As power levers are advanced above 68 to 70% N 1 , a microswitch on each power lever breaks the circuit to the respective ground low pitch stop solenoid. The blade angle changes from the ground low pitch stop of

IDLE

+2° to the flight low pitch stop of +12º. W h e n t h e b l a d e a n g l e c h a n g e s, t h e propeller rpm decreases momentarily because of the increased rotational drag. This helps prevent surging as power is added.

IDLE

OIL

IDLE

OIL

OIL

12°

12° 2°

FLIGHT IDLE

IN TRANSIT

AS AIRCRAFT DECREASES AIRSPEED, THE PROPELLER BLADE ANGLE DECREASES TO MAINTAIN PROPELLER GOVERNOR RPM AS THE BLADE ANGLE APPROACHES 13°, THE BETA VALVE IS PULLED INTO THE CUTOFF POSITION, CREATING A LOW PITCH STOP.

AS THE AIRCRAFT TOUCHES DOWN, A SQUAT SWITCH ACTUATED ELECTRICAL SOLENOID REPOSITIONS THE BETA LEVER AFT, MOVING THE BETA VALVE INTO THE OPEN POSITION. THIS ALLOWS OIL TO FLOW TO THE PROPELLER DOME.

2°

GROUND LOW PITCH STOP AS THE PROPELLER DOME MOVES FORWARD, THE BETA VALVE REPOSITIONS THE LOW PITCH STOP. OIL IS AGAIN TRAPPED IN THE PROPELLER DOME, EFFECTIVELY CREATING THE GROUND LOW PITCH STOP.

AS THE PROPELLER DOME FILLS WITH OIL, IT MOVES FORWARD, CARRYING THE BETA ROD AND LEVER ASSEMBLY, AND CONSEQUENTIALLY ROTATING THE PROPELLER BLADES TO A LOWER ANGLE.

Figure 7-39. Propeller Positioning—Flight Idle to Ground Low Pitch Stop (Sheet 1 of 3)

7-40


WARNING

CAUTION

Do not lift the power levers at the IDLE gate in flight. Doing so will energize the ground low pitch stop solenoids and cause the blade angle, if the primary governor is in an under-speed condition (the indicated propeller rpm is less than that selected with the propeller control levers), to decrease from the flight low pitch stop to the ground low pitch stop. This will cause excessive drag and the aircraft will develop a high sink rate.

Attempting to pull the power levers into the GROUND FINE and REVERSE ranges with the propellers in feather will cause damage to the reversing linkage of the power lever. Also, pulling the power levers into GROUND F I N E a n d R E V E RS E o n t h e ground with the engines shut down will damage the reversing system.

GROUND IDLE

7 POWERPLANT


GROUND FINE OIL

OIL

2°

OIL

–3°

2°

–3°

GROUND LOW PITCH STOP

IN TRANSIT

GROUND FINE GATE

THE GROUND FINE IS A RANGE IN WHICH OPTIMUM AIRCRAFT CONTROL AND ENGINE PERFORMANCE ARE MAINTAINED DURING TAXI. ONCE THE AIRCRAFT IS ON THE GROUND, THE PILOT MOVES THE POWER LEVER AFT.

THIS MOVES THE BETA LEVER AFT AND REPOSITIONS THE BETA VALVE TO THE OPEN POSITION ALLOWING OIL TO FLOW TO THE PROPELLER DOME. AS THE PROPELLER DOME FILLS WITH OIL, IT MOVES FORWARD, CARRYING THE BETA ROD AND LEVER ASSEMBLY, FURTHER ROTATING THE PROPELLER BLADES TO A LOWER ANGLE.

THIS ACTION MOVES THE BETA VALVE TO THE CLOSED POSITION, TRAPPING OIL IN THE PROPELLER DOME, EFFECTIVELY CREATING A STOP AT ZERO THRUST OR GROUND FINE.



7-41


7 POWERPLANT

In flight, do not lift the power levers at the IDLE gate. Doing so energizes the ground low pitch stop solenoids (Figure 7-40). If the primary governor is in an underspeed condition, the blade angle decreases from the flight low pitch stop to the ground low pitch stop. This causes excessive drag, and the aircraft develops a high sink rate. Attempting to pull the power levers into t h e G R O U N D F I N E a n d R E V E RS E ranges with the propellers in feather causes damage to the reversing linkage of the power lever. Also, pulling the power levers into GROUND FINE and REVERSE on the ground with the engines shut down damages the reversing system.

When the propeller is not in feather during normal operation on the ground, maintain propeller rpm above 1,050 rpm. With the engine idling, operation with the propeller feathered is permissible because the speed is below the resonance rpm range. Avoid sustained operation in feather on the ground, however, because excessive heat may build up in the nacelle, nose avionics area, and fuselage. While on the ground, the minimum blade angle is approximately 2° at IDLE.

GROUND FINE

FULL REVERSE OIL

OIL

–3°

OIL

–14°

–14° –3°

GROUND FINE GATE

IN TRANSIT

FULL REVERSE

TO ENTER THE REVERSE RANGE FROM GROUND FINE, THE PILOT MUST LIFT UP ON THE POWER LEVER AND MOVE THE POWER LEVER AFT.

AGAIN, THIS MOVES THE BETA LEVER AFT, MOVING THE BETA VALVE TO THE OPEN POSITION, ALLOWING OIL TO FLOW TO THE PROPELLER DOME, MOVING IT FORWARD. AS THE PROPELLER DOME MOVES FORWARD, IT CARRIES THE BETA ROD AND LEVER ASSEMBLY, FURTHER ROTATING THE PROPELLER BLADES TO A NEGATIVE ANGLE.

THIS ACTION MOVES THE BETA VALVE TO THE CLOSED POSITION, TRAPPING OIL IN THE PROPELLER DOME, EFFECTIVELY CREATING A STOP FOR FULL REVERSE.


7-42


7 POWERPLANT


Figure 7-40. King Air 350 Ground Idle Stop Electrical Circuit

Th e r i g h t l a n d i n g g e a r s a f e t y s w i t c h n o r m a l l y c o n t ro l s t h e l o w p i t c h s t o p. During the landing flare at idle power, the propeller blade angle is at the flight low pitch stop. Upon touchdown the squat switch causes the propeller blade angle to immediately decrease to the ground low pitch stop. Decreasing to the ground low pitch stop occurs because the ground low pitch stop solenoid is energized after the primary governor is in an underspeed c o n d i t i o n . Th e s o l e n o i d p u l l s a s m a l l distance aft on the REVERSE lever. This,

in turn, pushes the beta valve aft to open. Oil pressure increases to the propeller dome. The propeller dome moves forward; the blade angle decreases. As the blade angle decreases, the propeller dome pulls the polished rods, feedback ring, REVERSE lever, and beta valve forward. When the blade angle reaches approximately 2°, the beta valve has been pulled far enough forward to cut off oil pressure to the dome to stabilize the blade angle at the ground low pitch stop.


7-43


7 POWERPLANT

As a backup for the right squat, a switch in the power quadrant ensures the ground low pitch stop solenoids are activated when either or both power levers are lifted at the IDLE gate. The propeller blade angle remains at the ground low pitch stop until the power lever is moved aft of the IDLE gate. The PROP GOV TEST circuit breaker on the right circuit breaker panel protects the electric circuits of the ground low pitch stop solenoids. After the solenoids are energized, resistors reduce voltage to the g ro u n d l o w p i t c h s t o p s o l e n o i d f ro m approximately 28 volts to approximately 14 volts. This ensures that excess heat does not build up on the solenoids.

Propeller Resonance To avoid propeller resonance, maintain the propeller rpm above 1,050 or below 40 0 during ground operations. The most severe resonance is in the range of 850 to 900 rpm. It is especially severe with a quartering tail wind. Sustained operation in this rpm zone exposes the propeller to increased resonance and stress.

OVERSPEED GOVERNOR The overspeed governor (Figure 7-41) provides protection against excessive propeller speed if a primary governor malfunctions. Because the PT6 propeller is driven by a free turbine independent of the engine compressor, overspeed can occur rapidly, if a primary governor fails. The overspeed governor is on the left side of the propeller reduction gearbox.

Overspeed Governor Operation The overspeed governor is set at approximately 1,768 rpm. Its operation is very similar to that of the primary governor with two major differences: • Pilot cannot select a particular speed except when the overspeed governor is tested • Overspeed governor only reduces oil pressure to the propeller dome. Only the primary governor can increase oil pressure.

Figure 7-41. Overspeed Governor Diagram

7-44



From a pilot’s point of view, a propeller tachometer stabilized at approximately 1,768 indicates failure of the primary governor and proper operation of the overspeed governor.

When the power lever is in the GROUND FINE or REVERSE range, the value FTG attempts to limit propeller rpm to approximately 95% of selected propeller rpm. This ensures that the primary governor remains in an underspeed condition while in the REVERSE range on the ground.

POWER LEVERS The power levers (Figure 7-42) are on the power lever quadrant (first two levers on the left side) on the center pedestal. They are mechanically interconnected through a cam box to the fuel control unit, reverse lever, beta valve, and the fuel topping governor.

Pre-Takeoff Check For pre-takeoff check purposes, the set point of this governor can be rescheduled down to approximately 1,565 rpm with the GOV test switch on the pilot left subpanel.

FUEL TOPPING GOVERNOR The primary propeller governor contains a fuel topping governor (FTG) that helps protect against propeller overspeeding if the primary or overspeed governors have no effect on blade angle (e.g., a frozen propeller hub assembly). The FTG attempts to limit propeller speed by limiting the amount of fuel flowing to the fuel nozzles. It reduces the P 3 air p re s s u re i n t h e F C U. Th i s u l t i m a t e l y reduces the power applied to the propeller shaft and, hopefully, limits propeller rpm. When the power lever is in the forward thrust range, the value at which the FTG attempts to limits propeller rpm is approximately 106% of selected propeller rpm.

Figure 7-42. Power Levers

Th e p o w e r l e v e r q u a d r a n t p e r m i t s movement of the power lever in the forward thrust range from IDLE to maximum thrust and in the GROUND FINE or REVERSE range from IDLE to maximum reverse. Gates in the power lever quadrant at the IDLE and GROUND FINE positions prevent inadvertent movement of the lever into the GROUND FINE or REVERSE range. The pilot must lift the power levers up and over these gates to select the GROUND FINE or REVERSE range.


7-45

7 POWERPLANT

If a propeller overspeeds up to the value of the overspeed governor, it is safe to assume the blade angle is too low for the amount of power applied to the propeller. It is also assumed that the reason for the low blade angle is too much oil pressure in the propeller dome. If a propeller speed reaches approximately 1,768 rpm, the overspeed governor’s flyweights rotate fast enough to overcome the preset speeder spring tension. The flyweights move out and, in turn, pull the overspeed governor pilot valve up. This allows oil pressure to be dumped from the propeller dome back to the case to increase propeller blade angle and slow the propeller down.


The function of the power levers in the forward thrust range is to select a gas generator rpm through the FCU. The FCU thens set a fuel flow that produces and maintains the selected N 1 rpm.

7 POWERPLANT

In the GROUND FINE range, the power lever is used to: (1) select a propeller blade angle proportionate to the aft travel of the lever, thus reducing residual propeller t h r u s t , a n d ( 2 ) re s e t t h e f u e l t o p p i n g governor from its normal 106 percent to approximately 95 % of selected propeller rpm. N1 rpm is not affected in the GROUND FINE range. In the REVERSE range, the power lever functions to (1) select a propeller blade angle proportionate to the aft travel of the lever, (2) select an N 1 that sustains the selected reverse power, and (3) reset the fuel topping governor from its normal 106 percent to approximately 95 percent of selected propeller rpm. Therefore, propeller rpm in the GROUND FINE or REVERSE range is a function of N 1 and power lever position. It may be limited by the FTG acting through the FCU t o l i m i t f u e l f l o w a n d , c o n s e q u e n t l y, propeller rpm in relation to power lever position.

PROPELLER CONTROL LEVERS Propeller rpm within the primary governor range of 1,450 to 1,700 rpm is set by the position of the propeller control levers (Figure 7-43). These levers, one for each engine, are between the power levers and the fuel cutoff levers on the center pedestal. The full forward position sets the primary governor at 1,700 rpm. In the full aft position forward of the feathering detent, the primary governor is set at 1,450 rpm. Intermediate propeller rpm positions can be selected by

7-46

Figure 7-43. Propeller Control Levers

moving the propeller levers to select the desired rpm as indicated on the propeller t a c h o m e t e r . Th e s e t a c h o m e t e r s r e a d directly in revolutions per minute. A detent at the low rpm position has red and white stripes across the lever slot to prevent inadvertent movement of the propeller lever into the FEATHER detent.

PROPELLER FEATHERING Move the propeller lever full aft into the feather detent to feather the propeller. This action opens the feathering dump valve. All oil pressure quickly drains from the propeller dome; the propeller feathers. In this type of turbine engine, the propeller shaft and N 1 shaft are not connected. Thus, the propeller can be feathered with the engine at idle power without damaging the engine or gearbox.



At the feather position, the propeller lever positions the feathering dump valve to dump oil pressure from the propeller dome. This allows the counterweights and springs to position the propeller blades at the feather position.

Autofeather System

A n AU T O F E AT H E R s w i t c h o n t h e subpanel controls the system. When the switch is in the ARM position (Figure 7-44), the completion of the arming phase occurs when both power levers are advanced above 88% N 1 and both torque indications are above 17%. At this point, both the right and left green AUTOFEATHER annunciators indicate a fully armed system. When the switch is in the ARM position, the system is inoperative as long as either power lever is retarded below 88% N 1 position.

Autofeather Operation

The automatic feathering system provides a means of immediately dumping oil from t h e p ro p e l l e r d o m e. Th i s e n a b l e s t h e feathering spring and counterweights to feather the propeller blades if an engine fails.

Autofeather is required to be operable for all flights and armed for takeoff, climb, and a p p ro a c h . Wi t h t h e AU TO F E AT HER switch in the ARM position and starting with both power levers above 88% N 1 as well as both torque indications above 17%, the feathering springs and counterweights

Figure 7-44. Autofeather Diagram—Armed


7-47

7 POWERPLANT

If the engine is shut down in cruise flight without the autofeather system armed, the propeller stays onspeed and in sync unless it is manually feathered or all oil pressure is lost.


move the propeller blades for a particular engine toward feather. This occurs after the following occur:

7 POWERPLANT

• Torque manifold oil pressure for that engine dropped below the t o r q u e m e t e r v a l u e o f 17 % . Th i s disarms the opposite engine autofeather electrical circuit as evidenced by its AUTOFEATHER annunciator extinguishing. • Torque continues to drop below 10%. Th i s f e a t h e r s i t s p ro p e l l e r a n d i s evidenced by the extinguishing of its AUTOFEATHER annunciator. If the torque for the last operating engine drops below 17% or either power lever is retarded below 88% N 1 , the autofeather system is completely disarmed. When the AUTOFEATHER switch is in the TEST position (Figure 7-45), it works the same as above except:

• Power levers do not have to be above 88% N 1. However, both torque indications must start from above 17%, preferably 22%, because the TEST position on the switch bypasses both power lever switches. • When an engine failure is simulated with a power lever (Figure 7-46), the associated annunciator flashes off and on when the power lever is pulled to IDLE because there is still idle power.

Autofeather SystemTest Th e s y s t e m m a y b e t e s t e d a s f o l l o w s (Figure 7-47): 1. Power Levers—Approximately 22% torque; 22% torque is a power setting sufficient to simulate normal operation of both engines. 2. AUTOFEATHER Switch—Hold to TEST. This action completes a circuit bypassing the 88% autofeather arming switches inside the pedestal to allow

Figure 7-45. Autofeather Diagram—Test

7-48


7 POWERPLANT


Figure 7-46. Autofeather Diagram—Left Engine Failure Armed

Figure 7-47. Autofeather Test Diagram (Right Engine)—Low Power and Feathering


7-49


the test at power settings below the normal operating range. Both yellow c a u t i o n AU T O F E AT H E R O F F annunciators illuminate when switch moved to TEST.

7 POWERPLANT

3. Power Levers—Simulate engine-out situations by retarding each engine power lever individually. At approximately 17 % t o r q u e, the AUTOFEATHER OFF annunciator extinguishes for the engine not being re t a rd e d . Th i s c h e c k v e r i f i e s t h e autofeather system disarms itself on the operative engine. Continuing to retard power lever results in initiation of the feathering action at approximately 10% torque. At this time the propeller on the simulated inoperative engine tries to feather indicated by a blinking annunciator light and fluctuating torque. The AUTOFEATHER OFF annunciators cycle on and off with each fluctuation of torque as the propeller moves in and out of feather.

Then, with the condition levers in LOW IDLE, propeller feathering (manual, as compared to AUTOFEATHER) is checked. In this free-turbine engine, the propeller may be allowed to completely feather with the compressor operating at LOW IDLE with no engine damage sustained. Operation on the ground and in feather for extended periods of time may overheat the fuselage and possibly damage nosemounted avionics because hot exhaust gases are not being blown aft by the propeller’s air blast. Green AFX annunciators display in the ITT torque indicator for each engine on the MFD. Illumination of this annunciation indicates the respective system is armed and that the power lever is advanced above 90% N 1 .

4. Power Levers—Retard both to IDLE. Upon completion of Step 3 for each power lever, check that with both levers retarded while holding the AUTO-FEATHER switch to TEST neither propeller feathers. The loss of both AUTOFEATHER OFF annunciators verifies this disarming of both systems due to the intentional reduction of power. Upon completion of this test, arm the system for takeoff.

7-50



The propeller synchrophaser (Figure 7-48) automatically matches the rpm of the two propellers. It also maintains the blades of one propeller at a predetermined relative position to the blades of the other propeller. It is not a designated masterslave system because it functions to match the rpm of the slower propeller to the faster propeller and establish a blade phase relationship between them. Its actions reduce propeller beat from unsynchronized propellers to minimize cabin noise.

The synchrophaser system is an electronic system certified for use during all flight operations including takeoff and landing. The synchrophaser has a limited range of authority in relation to the primary governor setting. The maximum increase possible is approximately 20 rpm. In no case does the rpm fall below that selected by the propeller control lever. Normal governor operation is unchanged; the synchrophaser simply monitors propeller rpm continuously and resets either governor as required.

RH PROP LH PROP

LH PRIMARY GOVERNOR

RPM AND PHASE

RH PRIMARY GOVERNOR

RPM AND PHASE

CONTROL BOX

ON

PROP SYNC

5A OFF

Figure 7-48. Propeller Synchrophaser System


7-51

7 POWERPLANT

SYNCHROPHASER


Components A magnetic pickup adjacent to each propeller spinner bulkhead senses propeller rpm and position. This magnetic pickup transmits electrical pulses once per revolution to a control box forward of the pedestal. 7 POWERPLANT

The control box converts any pulse rate differences into correction commands that w h e n t r a n s m i t t e d t o c o i l s, c l o s e t h e flyweights of each primary governor. By varying coil voltage, the governor speed settings are biased until the propeller rpms exactly match.

If the synchrophaser is on but does not adjust the propeller rpms to match, the system has reached the end of its range. Increase the setting of the slow propeller or reduce the setting of the fast propeller to bring the speeds within the limited synchrophaser range. If preferred, turn the synchrophaser switch off, resynchronize manually, and turn the synchrophaser on. In the synchrophaser off position, the governors operate at the manual speed settings selected by the pilot.

A PROP SYNC pushbutton above and to the left of the LDG GEAR CONTROL turns the system on. When depressed, the green ON legend illuminates.

Synchrophaser Operation To operate the synchrophaser system, synchronize the propellers manually or establish a maximum of 20 rpm difference between the propellers. First turn the synchrophaser on. The system may be on at all times unless a malfunction is indicated. To change rpm with the system on, adjust both propeller controls at the same time.

7-52



1. The minimum N 1 required to select LOW IDLE on the condition lever during engine start is: A. 10%. B. 12% C. 14%. D. 16%. 2. Overfilling the oil may cause: A. Discharge until a satisfactory level is reached. B. Discharge until an unsatisfactory level is reached. C. Inconsistent propeller operation. D. Inconsistent propeller operation in reverse and ground fine operation. 3. If the compressor bleed valve fails to close as static take-off power is set, t o r q u e w i l l i n d i c a t e ________ t h a n normal and ITT will indicate _______ than normal. A. Lower; lower B. Higher; higher C. Higher; lower D. Lower; higher 4. The shaft horse power rating of 1,050 is a direct function of: A. Torque only. B. Propeller RPM only. C. Torque and Propeller RPM. D. Torque, RPM and exhaust thrust.

5. I g n i t i o n o p e ra t i o n o c c u r s d u r i n g engine start and during operations of _____________ or less when engine auto ignition is ____________. A. 17% torque; armed or off B. 17% torque; armed C. 70% N 1 ; armed or off D. 70% N 1 ; armed 6. The minimum oil temperature limit allowed for engine start is _____°C. A. –40 B. –30 C. –27 D. 0 7. The maximum allowed continuous ITT for takeoff is _______°C. A. 750 B. 820 C. 850 D. 1000 8. The minimum allowed oil pressure for idle is _______ PSI. A. 60 B. 70 C. 80 D. 100 9. Oil temperatures between 99°C and 110°C are limited to _______ minutes. A. Two B. Four C. Eight D. Ten


7-53

7 POWERPLANT

QUESTIONS


7 POWERPLANT

10. The maximum gas generator N 1 RPM limit for takeoff is: A. 100. B. 104. C. 106. D. 168.

14. The propeller governor is scheduled to control RPM between _______ RPM. A. 1050–1450 B. 1250–1450 C. 1450–1700 D. 1050–1700

11. The first immediate action item for an ENGINE FIRE OR FAILURE IN FLIGHT is affected engine: A. Prop Lever ...................FEATHER. B. Condition Lever ....................FUEL CUTOFF. C. Generator ..................................OFF. D. Starter Switch ..STARTER ONLY.

15. The autofeather system will feather the inoperative engine’s propeller when the opposite engine torque drops below: A. 89% N 1 . B. 69% N 1 . C. 17% torque. D. 10% torque.

12. The immediate action items for and E N G I N E FA I L U R E D U R I N G TAKEOFF (AT OR BELOW V 1 ) – TAKEOFF ABORTED are: A. Power Levers ....GROUND FINE. B. Brakes .........................MAXIMUM. C. Power Levers ....GROUND FINE, Brakes .........................MAXIMUM, ATC ....................................NOTIFY. D. Power Levers ....GROUND FINE, Brakes .........................MAXIMUM.

16. Th e f u e l t o p p i n g g o v e r n o r l i m i t s propeller RPM in flight to _______ percent of selected RPM. A. 96 B. 100 C. 104 D. 106

13. In order to select ground fine after landing, the pilot: A. Lifts the power lever and moves them aft to the first gate. B. Leaves the power lever at flight idle position and ground fine is automatically engaged. C. Lifts the propeller levers over the low RPM gate. D. Engages the ground fine switch on control yoke.

7-54

17. Th e o v e r s p e e d g o v e r n o r l i m i t s propeller RPM to a maximum of: A. 1700. B. 1768. C. 1800. D. 1876. 18. The maximum allowed continuous RPM for takeoff is _______ RPM. A. 1500 B. 1600 C. 1700 D. 1768



CHAPTER 8 FIRE PROTECTION CONTENTS Page INTRODUCTION ............................................................................................................... 8-1 GENERAL ........................................................................................................................... 8-1 FIRE DETECTION ........................................................................................................... 8-2 Components ................................................................................................................... 8-2 Operation ....................................................................................................................... 8-3

Components ................................................................................................................... 8-5 Operation ....................................................................................................................... 8-6 Portable Extinguishers.................................................................................................. 8-7 QUESTIONS ........................................................................................................................ 8-9


8-i

8 FIRE PROTECTION

FIRE EXTINGUISHING................................................................................................... 8-5



Title

Page

Engine Fire Detection System ............................................................................ 8-2

8-2

ENGINE FIRE Annunciators ........................................................................... 8-3

8-3

Engine Fire Detection System Simplified Schematic ....................................... 8-4

8-4

Fire-Extinguishing System ................................................................................... 8-5

8-5

EXTINGUISHER Annunciators ...................................................................... 8-6

8-6

Portable Fire-Extinguishers ................................................................................. 8-7

8 FIRE PROTECTION

8-1


8-iii


8 FIRE PROTECTION

CHAPTER 8 FIRE PROTECTION

INTRODUCTION Fire detection and fire-extinguishing systems provide fire protection in both engine compartments. Detection is automatic, but the crew must manually activate the extinguishing system.

GENERAL Th e K i n g A i r 3 5 0 h a s a n e n g i n e f i r e detection system that automatically alerts the crew if an engine fire or overtemperature situation occurs.

The extinguishing system consists of a cylinder with extingushant for the engine exhaust area and the engine accessory area. For a fire in the cabin or cockpit, two portable fire extinguishers are available.


8-1


FIRE DETECTION An engine fire detection system provides an immediate visual warning if a fire occurs in either engine compartment.

COMPONENTS Sensing Cable The main element of the system is a temperature sensor cable that loops continuously around the engine and terminates in the responder unit (Figure 8-1).

The cable consists of a hermetically sealed, corrosion-resistant, stainless-steel outer tube filled with an inert gas. The inner hydride core is filled with an active gas.

NOTE The fire sensor cable section in the plenum chamber area is two feet long and is not a fire or overheat zone. Activation temperature in this zone is approximately 90 0°F.

8 FIRE PROTECTION

Figure 8-1. Engine Fire Detection System

8-2



Th e r e s p o n d e r u n i t i s i n t h e e n g i n e accessory area approximately at the 10 o’clock position. The responder contains two sets of contacts. The integrity switch contacts are for the continuity test functions of the fire d e t e c t i o n c i rc u i t r y. Th e a l a r m s w i t c h contacts complete the circuit to activate the fire warning system.

Detection Pushbuttons The fire lights are on the glareshield just below the warning annunciator panel. These guarded pushbuttons have split lenses (Figure 8-2).

L or R ENGINE FIRE) annunciator triggers the MASTER WARNING flasher. The lower lens (L or R FW VALVE PUSH) indicates position of the firewall fuel valve. When the battery switch is on, this annunciator can indicate three situations. • Annunciator extinguished—Firewall fuel valve open • Annunciator illuminated—Firewall fuel valve closed • Annunciator flashing—Firewall valve position does not agree with firewall fuel valve switch position

OPERATION For fire detection/protection purposes, critical areas around the engine have been divided into three zones; • Zone 1—The accessory compartment • Zone 2—The plenum chamber area • Zone 3—The engine exhaust area The fire detector is preset at the factory to activate the alarm when any of the following conditions occur: • Any one-foot section of tube heated to 80 0°F

Figure 8-2. ENGINE FIRE Annunciators

They serve the dual function of monitoring the firewall fuel valve position and providing a visual warning of an overtemperature condition in either engine area.

• Average temperature of entire tube reaches 360°F • Temperature in accessory compartment reaches 545°F • Temperature in hot section compartment reaches 540°F

The red top lens (L or R ENGINE FIRE illuminates when the fire detector tube senses an overtemperature condition. The


8-3

8 FIRE PROTECTION

Responder Unit


If a fire or overtemperature occurs, the temperature around the sensor tube increases. The gases within the tube begin to expand. When pressure from the expanding gases reach a factory preset point, the contacts of the responder alarm switch close (Figure 8-3). S i g n a l s f ro m e a c h re s p o n d e r u n i t a re transmitted to a printed circuit board forward of the main spar underneath the center aisle floor. From the printed circuit board, the signal is routed to illuminate the appropriate red L or R ENGINE FIRE annunciator (Figure 8-3). The MASTER WARNING annunciators flash. A red annunciation of FIRE appears on the MFD in the appropriate engine’s TORQUE indicator. 8 FIRE PROTECTION

Fire Detection System Test

of the fire detection system. The switches have three positions: DET–OFF–EXT. When either switch is placed in the DET position, electrical current flows from a 5amp RIGHT or LEFT FIRE DET circuit breaker on the right CB panel. The current then flows through the engine fire detector circuitry to activate the red L or R ENG FIRE light. The MASTER WARNING annunciators also flash. In addition, a red annunciation of FIRE is visible on the MFD in the ITT/TORQUE indicator for the appropriate engine. If either annunciator fails to illuminate during the test for that side, a malfunction w i t h t h e s e n s o r t u b e, re s p o n d e r u n i t , annunciator, or electrical portion of the system is indicated. The malfunction must be corrected prior to flight.

Two ENG FIRE TEST toggle switches on the copilot left subpanel test the integrity

Figure 8-3. Engine Fire Detection System Simplified Schematic

8-4



The gases within the tubes form a pressure barrier to keep the contacts of the responder integrity switch closed for the test.

FIRE EXTINGUISHING The aircraft contains a fire extinguishing system and two portable fire extinguishers.

A fire extinguisher supply container is mounted on brackets aft of the main spar in each wheel well of each main landing gear (Figure 8-4). Each cylinder is charged with 2.50 pounds of bromotrifluoromethane (CBrF3) pressurized to 450 to 475 psi at 72°F. The line from the container runs along the side of the nacelle and branches into eight spray tubes strategically located about the engine to diffuse the extinguishing agent.

8 FIRE PROTECTION

COMPONENTS

Extinguishing Cylinder

Figure 8-4. Fire-Extinguishing System


8-5


Four of the nozzles discharge into the e n g i n e e x h a u s t a r e a . Th e o t h e r f o u r discharge into the accessory area. Once activated, the entire supply of extinguishing agent is discharged for that side. Th e c y l i n d e r a l s o h a s a p y r o t e c h n i c cartridge that discharges the extinguishant.

Extinguisher Pushbuttons The EXTINGUISHER pushbuttons on the glareshield activate the system (Figure 8-5). This guarded pushbutton is safetied.

OPERATION Illumination of either of ENG FIRE annunciators and a red annunciation of F I R E i n t h e I T T / TO R QU E i n d i c a t o r indicate a fire or overheat condition in that particular engine. Lifting the plastic guard of the ENG FIRE switch and depressing the lens illuminates t h e F W VA LV E P U S H l e n s a n d t h e EXTIN-GUISHER PUSH annunciator. Illumination of the CLOSED annunciator indicates the firewall valve for that side is now closed. The extinguisher for that side is armed.

8 FIRE PROTECTION

To discharge the extinguisher, raise the guard of the EXTINGUISHER PUSH annunciator and depress the switch. This completely discharges the appropriate fire e x t i n g u i s h e r c y l i n d e r . Th e a m b e r DISCHARGED annunciator illuminates. The DISCHARGED annunciator remains illuminated regardless of battery switch p o s i t i o n u n t i l t h e e x p e n d e d cy l i n d e r is replaced. Figure 8-5. EXTINGUISHER Annunciators

A 5-amp RIGHT or LEFT ENG FIRE EXT circuit breaker on the FUEL SYSTEM CB panel powers the appropriate switch. The push-to-activate pushbuttons have two indicator lights: red EXTINGUISHER PUSH and amber DISCHARGED. The red EXTINGUISHER PUSH portion indicates an electrical circuit from the FW VA LV E C LO S E D p u s h b u t t o n i n t h e detection system to the extinguishing system is complete.

Fire Extinguisher System Test The ENG FIRE TEST toggle switches on the copilot left subpanel also test the functions of the fire extinguishing system. When either switch is placed in the EXT position, a circuit completes from the 5amp circuit breaker on the battery bus through the following: • EXT test switch • EXTINGUISHER PUSH annunciator • Cartridge on the container • Ground

The DISCHARGED portion indicates the pyrotechnic cartridge in the container has been discharged.

8-6



Failure of either annunciator to illuminate on a particular side indicates a malfunction of the engine fire extinguisher system. The malfunction must be corrected prior to flight.

PORTABLE EXTINGUISHERS Two portable fire extinguishers are in the aircraft. One extinguisher is in the aft cabin on the lower side of the door frame. The other extinguisher is in the cockpit on the bottom of the co-pilot seat (Figure 8-6).

The bottles have hand-operated actuating valves. The portable extinguishers may be recharged at locally approved fire equipment service shops. Halon 1211 or bromochlorodifluoromethane is in a liquefied gas state while contained under pressure in the fire extinguisher. Upon release, the liquid quickly turns to a vapor that dissipates into the air and leaves no residue to clean up. As it changes from liquid to vapor, a rapid temperature drop to below freezing occurs. Do not direct the discharge at exposed skin or eyes. When used on fires of intense heat, decomposition of the Halon 1211 vapor may be accompanied by a sharp acrid odor. This odor warns the operator of excessive exposure to the products of combustion and to take evasive action.

Both extinguishers are mounted on red quick-release brackets.

WARNING

The portable extinguishers contain two pounds of 1211 extinguishing agent and 10 0 psi of nitrogen for pressurization. Halon 1211 is a chemical agent effective against combustible fires (Class A), f l a m m a b l e l i q u i d f i re s ( C l a s s B ) , a n d electrical fires (Class C). The smaller size extinguishers do not contain enough agent to qualify for a Class A rating.

Liquefied Halon 1211 can cause f r o s t b i t e. Av o i d c o n t a c t w i t h exposed skin or eyes. High concent ra t i o n s c a n p ro d u c e t o x i c b y products when applied to fire. Av o i d i n h a l a t i o n o f t h e b y products by evacuating and ventilating the area. Do not use in confined space with less than 311 cubic feet per extinguisher.

Figure 8-6. Portable Fire-Extinguishers


8-7

8 FIRE PROTECTION

Illumination of the EXTINGUISHER PUSH annunciator indicates the cartridge circuit is in good order. The test also checks the cartridge sensor for that side. The DISCHARGED annunciator should also illuminate to indicate the sensor is in good order.


8 FIRE PROTECTION


8-8



QUESTIONS 1, Engine fire detection and extinguishing is available when the battery bus switch is selected to _________ and the battery switch to __________. A. EMERG OFF; OFF B. EMERG OFF; ON C. NORM; OFF D. NORM; ON

8 FIRE PROTECTION

2. Engine fire extinguishing is available for the engine: A. Compartment. B. Compartment and wheel well. C. Compartment and wing locker (if installed). D. Compartment, wheel well, and wing locker (if installed). 3. With the hot battery bus powered, an engine fire extinguisher may be discharged: A. Anytime by depressing the fire extinguisher discharge switch. B. After arming the switch by depressing either firewall fuel valve shutoff switch. C. After arming the on-side switch and closing the firewall fuel valve. D. A f t e r d e p r e s s i n g t h e o n - s i d e firewall fuel valve switch. 4. Th e f i r s t i m m e d i a t e a c t i o n f o r E N V I R O N M E N TA L SYS T E M SMOKE OR FUMES is: A. Oxygen Mask(s) ......................DON B. Land ................................NEAREST SUITABLE AIRPORT C. Passenger Manual Drop Out .........................PULL ON D. Descend ...............AS REQUIRED


8-9


CHAPTER 9 PNEUMATICS CONTENTS Page INTRODUCTION ............................................................................................................... 9-1 GENERAL ........................................................................................................................... 9-1 Controls and Indications .............................................................................................. 9-2 OPERATION ....................................................................................................................... 9-3 Pneumatic....................................................................................................................... 9-3 Vacuum ........................................................................................................................... 9-5 SYSTEM USERS.................................................................................................................. 9-5 Engine Bleed-Air Warning System............................................................................. 9-5 Cabin Windows/Cockpit Side Window Defogging ................................................... 9-7 Hydraulic Fill Can Pressure ......................................................................................... 9-7

9 PNEUMATICS

QUESTIONS ........................................................................................................................ 9-9


9-i



Title

Page

BLEED AIR VALVE Switches .......................................................................... 9-2

9-2

Pneumatic Pressure Gages ................................................................................... 9-3

9-3

Pneumatic System Diagram ................................................................................. 9-4

9-4

Bleed-Air Warning System Diagram .................................................................. 9-5

9-5

Bleed-Air Warning Plastic Tubing ...................................................................... 9-6

9-6

L/R BLEED FAIL Annunciator......................................................................... 9-6

9-7

Cabin Windows/Cockpit Side Windows Defogging.......................................... 9-7

9 PNEUMATICS

9-1


9-iii


CHAPTER 9 PNEUMATICS

The aircraft pneumatic and vacuum systems accomplish many small but important tasks. This chapter presents a description of those tasks along with bleed air sources, indications, and normal and abnormal operations.

GENERAL High-pressure P 3 bleed air from each engine compressor is routed through the normally-open, firewall-mounted shutoff valves into the fuselage. The bleed air then is regulated to 18 psi to supply pressure for the pneumatic system and provide a vacuum source. Vacuum is derived from a bleed air ejector.

The pneumatic system supports the following: • Flight hour meter • Brake deice (see Chapter 10) • Bleed air warning • Window defogging • Hydraulic fill can pressure


9-1

9 PNEUMATICS

INTRODUCTION


In addition, pneumatic pressure creates a vacuum source for the air-driven gyros, pressurization control, and deflation of the deice boots.

pressurization and environmental functions (refer to Chapters 11 and 12 for details).

One engine normally can supply sufficient bleed air for all pneumatic and vacuum systems. During single-engine operation, a check valve in the bleed air line from each engine prevents pressure loss back through the supply line on the inoperative engine.

When the switches are in the OPEN position, both environmental flow control valves and pneumatic instrument air valves are open. When the switches are in the ENVIR OFF position, the environmental flow control valves close while the pneumatic instrument air valves remain open. In the PNEU & ENVIR OFF position, all valves are closed.

CONTROLS AND INDICATIONS

Gages

Switches

A pneumatic pressure gage on the copilot r i g h t s u b p a n e l i n d i c a t e s a i r p re s s u re available to the pneumatic manifold in pounds per square inch (Figure 9-2).

Two BLEED AIR VALVE switches (Figure 9-1) control bleed air entering the cabin for

9 PNEUMATICS

Figure 9-1. BLEED AIR VALVE Switches

9-2



To the left of the pressure gage is a suction gage calibrated in inches of mercury that indicates instrument vacuum. Below these two is the hour meter.

temperature of approximately 70 0°F. By the time it reaches the tee in the fuselage, heat transfer in the pneumatic plumbing cools the airflow to approximately 70° above ambient temperature. This regulated pneumatic air (Figure 9-3):

OPERATION

• Supplies pressure to inflate the surface deicers

PNEUMATIC P 3 bleed air at 90 to 120 psi pressure from both engines flows through pneumatic lines to a tee junction in the fuselage. Check valves prevent reverse flow during singleengine operation.

• Operates the flight hour meter and bleed air failure warning system

Downstream from the tee, the P 3 air passes through an 18 psi regulator. The regulator has a relief valve set to operate at 21 psi in case of failure.

Ordinarily, the pneumatic system pressure regulator under the right seat deck immediately forward of the main spar provides 18 ± 1 psi with the engine running at 70% to 80% N 1 . The pneumatic pressure gage allows the crew to monitor system pressure.

• Provides flow for vacuum ejector

9 PNEUMATICS

Bleed air extracted from the fourth stage of the engine compressor is at a maximum

• Provides pressure to the hydraulic system

Figure 9-2. Pneumatic Pressure Gages


9-3


LEGEND

TO PNEUMATIC PRESSURE GAGE (IN COCKPIT)

HIGH PRESSURE BLEED AIR REGULATED BLEED AIR VACUUM

FLIGHT HOURS GAGE

RIGHT SQUAT SWITCH

DEICE DISTRIBUTOR VALVE

PRESSURE SWITCH

LANDING GEAR HYDRAULIC FILL CAN

EXHAUST OVERBOARD EJECTOR LEFT SQUAT SWITCH

AIRSTAIR DOOR SEAL LINE

TO DEICE BOOTS

CLOSED ON GROUND (NO) LEFT BLEED-AIR WARNING SYSTEM

VACUUM REGULATOR GYRO INSTRUMENTS (PRIOR TO PROLINE 21) PRESSURATION CONTROLLER, TO GYRO OUTFLOW AND SUCTION SAFETY VALVES (IN COCKPIT) RIGHT BLEED-AIR WARNING SYSTEM 18 PSI PRESSURE REGULATOR

9 PNEUMATICS

CABIN AND COCKPIT SIDE WINDOWS LEFT P3 AIR

CHECK VALVE

WINDOW DEFOGGING REGULATOR RIGHT P3 AIR

CHECK VALVE

PNEUMATIC AIR VALVE (NO)

PNEUMATIC AIR VALVE (NO)

RIGHT BRAKE DEICE VALVE (NC)

LEFT BRAKE DEICE VALVE (NC)

Figure 9-3. Pneumatic System Diagram

9-4



VACUUM Vacuum is obtained from pneumatic air passing through the vacuum ejector. The ejector is a pressure line that constricts to form a venturi. At the apex or point of lowest pressure, the vacuum line is attached. As P 3 air passes through the ejector, it draws air from the attached vacuum line to create suction for the vacuum system. The ejector is capable of supplying from 15 inches of mercury (in. Hg) vacuum at sea level to 6 in. Hg vacuum at 31,000 feet. Th e v a c u u m s y s t e m s u p p l i e s v a c u u m regulated 4.3 to 5.9 in. Hg for the deice boots, air-driven gyro instruments, and pressurization control system. Even with one engine running at 70% to 80% N 1, the vacuum gage normally reads approximately 5.9(+0,-0.2) in. Hg at sea level.

The vacuum line for the instruments is routed through a regulator that maintains sufficient vacuum for proper operation of the instruments. Th e v a c u u m r e g u l a t o r i s i n t h e n o s e compartment on the left side of the pressure b u l k h e a d . A f o a m f i l t e r f o r t h e va l v e provides a filtered and sheltered air source for the air-driven gyros.

SYSTEM USERS ENGINE BLEED-AIR WARNING SYSTEM Th e e n g i n e b l e e d - a i r w a r n i n g s y s t e m visually warns the crew of a rupture in the bleed air lines. This allows the crew to shut off the affected bleed air valves before heat from the escaping air damages the skin and structure adjacent to the break (Figure 9-4).

PLUGS

PRESSURE SWITCHES

AMBIENT AIR

ENVIRONMENTAL BLEED-AIR SHUTOFF VALVE

PNEUMATIC BLEED-AIR SHUTOFF VALVE

ENVIRONMENTAL BLEED-AIR SHUTOFF VALVE

ENGINE FIREWALL PLUGS

ENGINE P3 BLEED-AIR CONNECTOR AMBIENT AIR

ENGINE FIREWALL PLUGS

PNEUMATIC BLEED-AIR SHUTOFF VALVE

WHEEL WELL

WHEEL WELL

BLEED-AIR WARNING LINES ENVIRONMENTAL BLEED-AIR LINES

18 PSI PRESSURE REGULATOR

PNEUMATIC BLEED-AIR LINES

Figure 9-4. Bleed-Air Warning System Diagram


9-5

9 PNEUMATICS

ENGINE P3 BLEED-AIR CONNECTOR


EVA Tubes Ethylene vinyl acetate (EVA) tubes are in c l o s e p ro x i m i t y t o t h e b l e e d a i r l i n e s (environmental and pneumatic) leading from the engines to the cabin. Pneumatic air at 18 psi tapped off from the pneumatic manifold pressurizes these tubes.

Pressure Switches The system has two pressure switches, one for each side, mounted under the cockpit f l o o r b o a rd . A r u p t u re d b l e e d a i r l i n e produces excessive heat on the tubing (Figure 9-5).

Figure 9-6. L/R BLEED FAIL Annunciator

Corrective Action When the indication of a bleed air failure becomes evident, turn off all bleed air for that side by placing the respective BLEED AIR VALVE switch in the PNEU & ENVIR O F F p o s i t i o n . Wi t h t h e s w i t c h i n t h i s position, both environmental and pneumatic shutoff valves close. This stops bleed air flow at the engine firewall. Next place the ECS switch in MAN HEAT position. Hold the TEMP switch in decrease position for 30 seconds. Then increase to maintain cabin/cockpit temperature. This action does not extinguish the BL AIR FAIL annunciator.

Figure 9-5. Bleed-Air Warning Plastic Tubing 9 PNEUMATICS

When the tubing melts, the air escapes and pressure inside the tubing decreases. When it drops below 2 psi, the normally-closed switch in the line closes to complete a circuit to the respective L BLEED FAIL and R BLEED FAIL annunciator in the warning panel (Figure 9-6).

9-6



CABIN WINDOWS/COCKPIT SIDE WINDOW DEFOGGING Engine bleed air provides defogging for the cabin windows and cockpit side windows when one or both engines are running (Figure 9-7). Bleed air tubes direct bleed air directly onto the windows. The WINDOW DEFOG switch on the copilot subpanel controls this function. Refer to Chapter 10.

CAUTION Caution must be used when removing or installing the emergency exits to avoid damaging the bleed air lines. Make certain the bleed air lines are properly connected when installing the hatches. Air for the cockpit side windows is routed forward from the pressure regulator to the cockpit below the floorboard. At this point the line tees, running to the sides of the aircraft and up the sidewall to the windows. The warm dry air is then directed to the inside surface of the windows through a manifold in the window frame.

HYDRAULIC FILL CAN PRESSURE

A pressure regulator mounted under the floorboard regulates the bleed air pressure. From the regulator, the air flows to the right of center and forward where the air lines divide. At this point, part of the air flows to the sides of the aircraft and up the sidewall to just below the windows where it tees and runs fore and aft under the cabin windows.

Pneumatic air pressure is injected into the landing gear hydraulic fill can. It provides positive pressure to the reservoir to minimize vaporization of hydraulic fluid. It also provides positive feed to the hydraulic pump. Wh e n t h e e n g i n e s a re s h u t d o w n , t h e pneumatic pressure in the hydraulic fill bleeds off through a small orifice.

Beneath each cabin window there is another tee in the system that delivers the air to the inside surface of the exterior window through two tubes.


9-7

9 PNEUMATICS

Figure 9-7. Cabin Windows/Cockpit Side Windows Defogging



9 PNEUMATICS

9-8



QUESTIONS 1. Regulated pneumatic air pressure is used to: A. Deice the brakes. B. Inflate the deice boots. C. Operate the air conditioner. D. Heat the aft cabin. 2. After selecting the bleed air valve to pneumatic and environmental off after illumination of a single [L or R BLEED FAIL] red master warning annunciator, the annunciator will: A. Extinguish. B. Remain illuminated. C. Extinguish provided the opposite system is functioning properly. D. Remain illuminated if the system continues to detect an overheat condition.

9 PNEUMATICS

3. Vacuum air is provided for: A. Door seal inflation. B. Flight hour meter operation. C. Cross bleed engine starts. D. Wing deice boot hold-down.


9-9


CHAPTER 10 ICE AND RAIN PROTECTION CONTENTS Page INTRODUCTION............................................................................................................. 10-1 GENERAL ......................................................................................................................... 10-1 Controls........................................................................................................................ 10-2 ENGINE PROTECTION................................................................................................. 10-4 Engine Air Inlet .......................................................................................................... 10-4 Inertial Separators ...................................................................................................... 10-5 Auto-Ignition System.................................................................................................. 10-7 SURFACE DEICE ............................................................................................................ 10-8 Components................................................................................................................. 10-8 Operation................................................................................................................... 10-10 BRAKE DEICE SYSTEM............................................................................................. 10-10 Components .............................................................................................................. 10-10 Operation................................................................................................................... 10-11 PROPELLER DEICE ................................................................................................... 10-12 Operation................................................................................................................... 10-13 WINDSHIELD ANTI-ICE ............................................................................................ 10-14 Components .............................................................................................................. 10-14

Windshield Wipers ................................................................................................... 10-18 WINDOW DEFOGGING ............................................................................................. 10-18 Cabin Windows ......................................................................................................... 10-18


10-i

10 ICE AND RAIN PROTECTION

Operation................................................................................................................... 10-15


Cockpit Side Windows ............................................................................................. 10-19 FUEL SYSTEM ANTI-ICE ........................................................................................... 10-19 Heated Vents ............................................................................................................. 10-20 Heat Exchanger ........................................................................................................ 10-20 ELECTRICAL HEATING ........................................................................................... 10-20 Pitot Heat .................................................................................................................. 10-20 Stall Warning Vane ................................................................................................... 10-21 PRECAUTIONS DURING ICING CONDITIONS ................................................. 10-22 Exterior Inspection................................................................................................... 10-22 Before Taxi................................................................................................................. 10-23 QUESTIONS.................................................................................................................... 10-25


10-ii




Title

Page

King Air 350 Anti-Icing and Deicing Components ........................................ 10-2

10-2

Ice and Rain Protection Controls ..................................................................... 10-3

10-3

Engine Inlet Lip Heat......................................................................................... 10-4

10-4

Inertial Separator ................................................................................................ 10-5

10-5

Ice Vane Controls ................................................................................................ 10-6

10-6

Caution Annunciators......................................................................................... 10-6

10-7

Auto-Ignition Switches ....................................................................................... 10-7

10-8

Surface Deice Boot Installation ........................................................................ 10-8

10-9

Surface Deice System Diagram ......................................................................... 10-9

10-10

Brake Deice ....................................................................................................... 10-10

10-11

Brake Deice Controls ....................................................................................... 10-11

10-12

Brake Deice Schematic (System On) ............................................................. 10-11

10-13

Propeller Deice Boots ...................................................................................... 10-12

10-14

Propeller Deice System .................................................................................... 10-13

10-15

Windshield Anti-Ice Switches.......................................................................... 10-15

10-16

Windshield Anti-Ice Diagram—Normal Heat .............................................. 10-16

10-17

Windshield Anti-Ice Diagram—High Heat ................................................... 10-17

10-18

Windshield Wiper.............................................................................................. 10-18

10-19

Fuel System Anti-Ice ........................................................................................ 10-19

10-20

Pitot Mast and Heat Controls ......................................................................... 10-20

10-21

Stall Warning Vane and Heat Control............................................................ 10-21 10 ICE AND RAIN PROTECTION

10-1


10-iii


CHAPTER 10 ICE AND RAIN PROTECTION

INTRODUCTION The King Air 350 is FAA approved for flight in known icing conditions when the required equipment is installed and operational. The Kinds of Operations Equipment List in the Limitations section of the Pilot’s Operating Handbook lists the necessary equipment. Flight in known icing conditions requires knowledge of conditions conducive to icing as well as knowledge of the aircraft anti-ice and deice systems that prevent excessive ice from forming. This chapter identifies these systems and controls.

GENERAL • Auto-Ignition


The aircraft has a variety of ice and rain protection systems for operation in inclement weather conditions. These include (Figure 10-1):

• Windshield Anti-Ice • Windshield Wipers

• Engine Inlet Lip Heat

• Side Window Defog

• Inertial Separators (Ice Vanes)

• Propeller Deice


10-1


An oil-to-fuel heat exchanger warms the fuel prior to its entrance to the fuel control unit.

• Fuel System Anti-Ice • Pitot Heat • Stall Warning Heat • Surface Deice (Leading-Edge Boots) • Brake Deice (Optional) Engine bleed is available for the leading edge lip of the engine air inlet, wings, horizontal stabilizer, window defogging, and brake deice. Electrical heating elements protect the pitot masts, windshield, stall warning vane, and fuel vent. An inertial separation system with electric motors and actuators provide engine ice protection. Electrothermal boots on each blade protect the propellers against icing.

An auto-ignition system ensures positive engine ignition during turbulence or penetration into icing or percipitations conditions. And, finally, heavy-duty windshield wipers for both the pilot and copilot increase visibility during rainy flight conditions.

CONTROLS Crew controls for the ice protection features are on the pilot subpanel. These control panels include the following (Figure 10-2): • ENGINE ANTI-ICE • ICE PROTECTION


Figure 10-1. King Air 350 Anti-Icing and Deicing Components

10-2




Figure 10-2. Ice and Rain Protection Controls


10-3


• ENG AUTO IGNITION The control knob for the windshield wipers is on the overhead panel.

ENGINE PROTECTION ENGINE AIR INLET

A scoop in the left engine exhaust stack deflects a small portion of the hot exhaust gases downward into the hollow lip tube that encircles the air inlet. The gases are expelled into the right exhaust stack where they move out with the engine exhaust gases. No pilot action is necessary. Engine exhaust air flows through the inlet heat duct whenever the engine is running.

Hot exhaust gases prevent ice formation around the lips of both engine cowling air inlets (Figure 10-3).


Figure 10-3. Engine Inlet Lip Heat

10-4



INERTIAL SEPARATORS An inertial separation system in each engine air inlet prevents moisture from entering the inlet plenum during freezing conditions. The system consists of two electrically actuated movable vanes.

BYPASS DOOR FORWARD ICE VANE

OIL COOLER INLET

OIL COOLER BYPASS DUCT

Figure 10-4. Inertial Separator


INLET LIP ANTI-ICE

INDUCTION AIR

In normal operation, the vanes are positioned with the forward vane retracted (up) and the aft vane extended (down) to direct all incoming air into the engine air plenum (Figure 10-4).


10-5


Components An electrically operated linear actuator simultaneously positions the vanes for each engine through a linkage system. The actuator t ra v e l i s m a x i m u m i n t h e d i re c t i o n commanded; it has no intermediate positions.

primary motor actuators (Figure 10-5). The AC T UAT O R S s w i t c h e s b e l o w t h e m energize the primary (MAIN position) or secondary system (STANDBY position).

The actuator is a primary and secondary motor with a single actuator rod assembly. The primary motor normally drives the system. If a malfunction occurs in the p r i m a r y m o t o r, t h e s e c o n d a r y m o t o r p r o v i d e s p o w e r . Th e m o t o r s a r e interchangeable. The only difference is their power source. The triple-fed bus powers the primary motor while the corresponding generator bus powers the secondary motor.

CAUTION Should the actuator primary motor malfunction, the cause must b e d e t e r m i n e d a n d c o r re c t e d before the next flight.

Controls and Indicators The L and R ENG ANTI-ICE switches on the left outboard subpanel energize the

Figure 10-5. Ice Vane Controls

A position sense switch on each vane linkage illuminates the green advisory L or R ENG ANTI-ICE annunciator in the caution/ advisory annunciator panel when the ENGINE ANTI-ICE switches are in the ON position (Figure 10-6).


Figure 10-6. Caution Annunciators

10-6



When the ENGINE ANTI-ICE switches are activated, a second position sense s w i t c h o n e a c h f o r w a rd va n e l i n k a g e energizes a 30 to 40 second time delay circuit. If full vane extension is not attained in this time, the yellow caution L and/or R ENG ICE FAIL annunciator illuminates. This indicates a fault in the primary motor of the designated actuator. To complete vane extension and extinguish the yellow annunciator, place the appropriate ACTUATORS switch to STANDBY. When the ENGINE ANTI-ICE switches are in the OFF position, the actuators move in the opposite direction to return the vanes to non-icing position. The green annunciators extinguish.

AUTO-IGNITION SYSTEM The engine auto-ignition system provides electrical power to the igniters in the e v e n t o f a n e n g i n e p o w e r l o s s. Th e system must be armed during periods of turbulence and penetration of icing or precipitation conditions. The ENG AUTO IGN switches above the ENGINE ANTI-ICE switches control the system. In the ARM position, auto-ignition automatically attempts to reignite the engine if an engine flameout occurs (Figure 10-7).

Operation Extend the inertial separators (ice vanes) whenever there is visible moisture at ambient temperatures of +5°C or below. When the ice vanes are extended, the green ENGINE ANTI-ICE annunciators illuminate. Because the airflow into the engine is now restricted, there may be a decrease in torque and a slight increase in ITT. Expect a decrease in overall cruise performance with vanes extended. When the vane doors retract, the annunciators extinguish; ITT and torque is restored.

Limitation

Operation With the system armed, electrical power energizes the engine igniters if engine torque falls below approximately 17% for any reason. I f t h i s o c c u r s, t h e g r e e n a d v i s o r y IGNITION ON annunciator on the caution/advisory panel illuminates. During ground operations, ensure the switches are in the OFF position to prolong igniter life.


ENGINE ANTI-ICE shall be OFF for takeoff operations in ambient temperatures of and above +10°C.

Figure 10-7. Auto-Ignition Switches


10-7


SURFACE DEICE

Control Switch

Pneumatic inflatable boots break off ice that collects on the leading edge of the wings and horizontal stabilizer (Figure 108). Alternately inflating and deflating the deice boots accomplishes the ice removal.

Figure 10-8. Surface Deice Boot Installation

Pneumatic system pressure inflates the boots; system vacuum deflates the boots and holds them down while not in use. A deice distributor valve controls the cycle.

COMPONENTS Deice Boots Each wing has an inboard and an outboard boot. The horizontal stabilizer has one boot on each of the left and right segments. The vertical stabilizer is not deiced because no adverse effects from icing were found during certification.

A three-position SURFACE DEICE toggle switch on the ICE PROTECTION panel actuates the deicer system (Figure 10-9). The switch is spring-loaded to return to the OFF position from either the MANUAL or SINGLE position. When the switch is pushed to the SINGLE position, one complete cycle of deicer operation automatically follows as the distributor valve opens to inflate the deicer boots. After an inflation period of approximately six seconds for the wings and four seconds for the horizontal stabilizer, a timer relay switches the distributor valve off (vacuum) for deflation of the boots. Total cycle time is approximately 10 seconds. When the switch is pushed to the MANUAL position, all the boots inflate. They hold in the inflated position as long as the switch is held in position. Upon release of the switch, the distributor valve returns to the OFF position. The deicer boots remain deflated until the switch is actuated again.

Pressure Switches/ Annunciators Two pressure sensing switches, one for each wing, monitor deice boot pressure. When both switches sense 14–15 psi wing deice boot pressure, the green advisory WING DEICE annunciator illuminates to signal boots are inflated. A third pressure switch senses tail deice boot pressure. It illuminates the green TAIL DEICE annunciator when it senses 14–15 psi tail deice boot pressure.


10-8




Figure 10-9. Surface Deice System Diagram


10-9


OPERATION

COMPONENTS

Wing ice lights aid the crew in detecting ice formation on the wing leading edge. The lights are on the outboard side of each engine nacelle. A 5-amp WING ICE circuit breaker switch on the pilot inboard subpanel controls the light.

A pneumatic line on the outboard side of each nacelle carries the engine P3 air to a shutoff valve.

Use the wing ice light to check accumulation. For most effective deicing operation, allow at least 1/2 inch, but no more than one inch, of ice to form before attempting ice removal. Very thin ice may crack and cling to the boots instead of shedding. Subsequent cyclings of the boots then have a tendency to build a shell of ice outside the contour of the inflated boot. This makes ice removal efforts ineffective. If ice is allowed to build to a depth greater than one inch, removal with the deice boots may be impossible.

The normally closed solenoid shutoff valve in each wheel well allows hot bleed air to enter the brake deice lines. An electrically powered module controls the shutoff valve. Th e m o d u l e i s u n d e r t h e c e n t e r a i s l e floorboard immediately aft of the partition between the cockpit and cabin. When the shutoff valves open, a signal illuminates the green advisory L and R BRAKE DEICE ON annunciators on the advisory panel. A distributor manifold attached to the brake piston and axle assembly directs the heated air through orifices around each r i n g o f t h e m a n i f o l d o n t o t h e b ra ke s (Figure 10-10).

Electrical Power The distributor valve requires electrical power to inflate the boots in either singlecycle or manual operation. If power is lost, vacuum holds them against the leading edge surfaces. Th e t r i p l e - f e d b u s p o w e r s t h e S U R F DEICE circuit breaker on the copilot CB sidepanel to supply electrical power.

BRAKE DEICE SYSTEM The brake deice system prevents ice and slush build-up between the wheels. This build-up freezes the brakes. The pneumatic system supplies bleed air from the compressor of each engine for brake deicing. 10 ICE AND RAIN PROTECTION

Figure 10-10. Brake Deice

10-10



When the switch is up, the right generator bus supplies power through a 5-ampere circuit breaker in the copilot sidepanel to the control module.

OPERATION The BRAKE DEICE lever switch on the pilot right subpanel controls the brake deice operation (Figure 10-11).

The module supplies current to open the shutoff valves so hot bleed air can enter the brake deice lines. The BRAKE DEICE ON annunciators illuminate. Fro m t h e s h u t o ff va l v e, t h e h o t a i r is plumbed through an insulated line down the back of the main gear strut to the distributor manifold. The hot P 3 air is directed on the brakes (Figure 10-12).

Figure 10-11. Brake Deice Controls

18-PSI PRESSURE REGULATOR “T” IN WHEEL WELL

PNEUMATIC LEFT P3 AIR

“T” IN WHEEL WELL

PNEUMATIC RIGHT P3 AIR

RIGHT BRAKE DEICE VALVE (N.C.)

LEFT BRAKE DEICE VALVE (N.C.)

TO BRAKE DEICE VALVES

BRAKE DEICE SWITCH ON

BRAKE DEICE CIRCUIT BREAKER 28 VDC

OFF

LEFT BRAKE DEICE MANIFOLD

10 MINUTE TIMER


RIGHT BRAKE DEICE MANIFOLD

UP NOT UP

LEFT MAIN GEAR UPLOCK SWITCH

Figure 10-12. Brake Deice Schematic (System On)


10-11


Wh e n t h e a i rc ra f t i s i n f l i g h t a n d t h e B R A K E D E I C E s w i t c h u p, a c i r c u i t completes through the left ma in g e a r uplock switch to a timing circuit in the control module. This timing circuit cycles the deice system off after 10 minutes of operation by closing the solenoid valves. This shuts off P 3 airflow to the brakes so adjacent components in the wheelwell are not damaged through overheating. The system cannot be activated again until the landing gear is cycled from the up and locked position.

Limitation Th e b ra ke d e i c e s y s t e m i s n o t t o b e c o n t i n u o u s l y o p e ra t e d a b o v e 15 ° C ambient temperature.

Figure 10-13. Propeller Deice Boots

PROPELLER DEICE

CAUTION

The propeller deice system includes the following components: an electrically heated boot for each propeller blade, slip rings, brush assemblies, timer, two on-off switches, and an ammeter.

Propeller deice must not be operated when the propellers are static or the slip rings around the propeller shaft will pit and burn, eventually becoming useless.

The heating elements in the deice boots reduce the adhesion of the ice (Figure 1013). The centrifugal force of the spinning propeller and blast of the airstream then removes the ice.

Th e P R O P s w i t c h e s a r e o n t h e I C E PROTECTION panel on the pilot right subpanel. The PROP ammeter is on the overhead panel. When the switch is on (up position), current flows through a timer and then through the brush assemblies to the slip ring where it is distributed to the individual propeller deice boots.


10-12



When the PROP-AUTO switch is turned on, the ammeter registers the amount of current (normally 26 to 32 amperes) passing through the system. If the current rises beyond normal limits, the circuit breaker switch shuts off power to the deicer timer (Figure 10-14).

Power to the deice boots is cycled in two 90-second phases. The first 90-second phase simultaneously heats all of the boots on one propeller. The second phase heats the boots on the other propeller. The deice t i m e r c o m p l e t e s o n e f u l l cy c l e e v e r y three minutes.


OPERATION

Figure 10-14. Propeller Deice System


10-13


As the deice timer moves from one phase to the next, a slight momentary deflection of the propeller ammeter needle may occasionally be noted. Once the system is turned on for automatic operation, it cycles continuously. The heating sequences for the propeller deice boots are the sequences in evidence during normal operation. If power is turned off, however, the timer moves ahead to the start of the cycle on the next propeller. It may restart on either propeller.

COMPONENTS Windshields The windshields are composed of three physical layers. The inner layer is a thick panel of glass that acts as the structural member. The middle layer is a polyvinyl sheet that carries fine wire heating grids. The outer layer is a protective layer of glass bonded to the first two layers. The outside of the windshield is treated with a static discharge film called a NESA coating.

Heating Elements

Manual System A manual propeller deicer system is a back up to the automatic system. When the PROP-MANUAL override switch is activated, power is applied to all heating elements on both props. This momentary switch must be held in place until the ice has been dislodged from the propeller surface. Although the propeller ammeter does not indicate prop boot load in manual mode, the loadmeters do indicate approximately a 10% increase in load when the manual deicer system operates.

WINDSHIELD ANTI-ICE Th e p i l o t a n d c o p i l o t w i n d s h i e l d s each have independent controls and electrothermal circuits.

Electrical heating elements in the lamination of the windshields protect them against icing. The elements consist of transparent material (stannic oxide) with high electrical resistance. Th e re s i s t i v e m a t e r i a l i s a r ra n g e d t o p ro v i d e p r i m a r y h e a t e d s u r f a c e s a n d secondary heated surfaces. Each is also fitted with electrical connections for temperature sensing elements. The heating elements connect at terminal blocks in the corners of the glass to the wiring of the WSHLD ANTI-ICE control switches.

Temperature-Sensing Elements A temperature-sensing element embedded in each windshield and a temperature controller in each windshield circuit automatically control the windshield temperature. The temperature controllers attempt to maintain the temperature of the windshield between 90 and 110°F.


10-14



WSHLD ANTI-ICE Switches The WSHLD ANTI-ICE switches on the ICE PROTECTION panel include one for the PILOT windshield and one for the COPILOT windshield (Figure 10-15).

The primary areas are smaller areas that heat faster. Each switch must be lifted over a detent before it can be moved into the HIGH position. This lever-lock feature prevents inadvertent selection of the HIGH position when moving the switches from NORMAL to the OFF (center) position.

OPERATION When the NORMAL position is selected (Figure 10-16), an automatic temperature controller senses the windshield temperature. It then attempts to maintain it at approximately 90 to 110°F by energizing the power relay as necessary. In this mode, almost the entire windshield is heated. When HIGH is selected (Figure 10-17), the same temperature controller senses the windshield temperature and attempts to maintain it at 90 to 110°F. In this mode, however, the controller also energize a high heat relay switch to apply the electrical heat to a more concentrated and more essential viewing area of the windshield. In HIGH, only the outboard two-thirds of the windshield is heated. A 50-ampere circuit breaker in the power distribution panel under the center aisle floor protects the power circuit of each system. A 5-ampere circuit breaker on the copilot CB panel protects each windshield heater control circuit.

When both switches are in the NORMAL position, the secondary areas of the windshields are heated. When the switches are in the HIGH position, the primary areas are heated.

In addition, windshield heat may be used to help in defrosting and defogging. The electrical field created by the heating elements may cause erratic operation of the magnetic compass.


10-15


Figure 10-15. Windshield Anti-Ice Switches

Windshield heat may be used at anytime and in any combination. It is best, however, to turn it on prior to entering icing conditions. This helps ensure the windshield is sufficiently warm to prevent ice formation.


Figure 10-16. Windshield Anti-Ice Diagram—Normal Heat 10 ICE AND RAIN PROTECTION

10-16




Figure 10-17. Windshield Anti-Ice Diagram—High Heat


10-17


Control Switch

CAUTION In the event of windshield icing during sustained icing conditions, it may be necessary to reduce the a i r s p e e d i n o rd e r t o ke e p t h e windshield ice-free; 226 KIAS is the maximum speed for effective windshield anti-icing.

WINDSHIELD WIPERS Separate windshield wipers are mounted on the pilot and copilot windshields. Use wipers as required on the ground or in flight, but do not operate them on a dry windshield because they may scratch a protective electrostatic (NESA) coating on the outer layer of glass. A single electric motor operates a mechanism that drives the dual wipers. All components are forward of the instrument panel.

The WINDSHIELD WIPER knob on the overhead panel (Figure 10-18) controls the wipers. Settings include OFF, SLOW, FAST and PARK. The windshield wiper circuit breaker is on the copilot circuit breaker panel in the WEATHER group.

WINDOW DEFOGGING Engine bleed air provides cabin window and cockpit side window defogging whenever one or both engines are running. A pressure regulator under the floorboard regulates the bleed air pressure for window defog.

CABIN WINDOWS From the regulator, the bleed air flows to the right of center and forward until the air lines divide. At this point part of the air is routed to the sides of the aircraft and up the sidewall to just below the windows. Here it tees and runs fore and aft under the cabin windows.


Figure 10-18. Windshield Wiper

10-18



Beneath each cabin window, another tee in the system delivers the bleed air through two tubes to the inside surface of the exterior windows.

CAUTION Caution must be used when removing or installing the emergency exits to avoid damaging the bleed air lines. Make certain the bleed air lines are properly connected when installing the hatches.

COCKPIT SIDE WINDOWS Air for the cockpit side windows is routed forward from the pressure regulator to the cockpit below the floorboard. At this point the line tees and runs to the side of the aircraft and up the sidewall to the windows. The warm dry air is now directed to the inside surface of the windows through a manifold in the window frame.

FUEL SYSTEM ANTI-ICE


Two anti-ice systems protect fuel flow through fuel lines to the engine (Figure 10-19). Without heat, moisture in the fuel could freeze. This would diminish or cut off fuel flow.

Figure 10-19. Fuel System Anti-Ice FOR TRAINING PURPOSES ONLY

10-19


HEATED VENTS Electrically heated vents in each wing prevent ice formation in the fuel vent system. The L and R FUEL VENT switches in the ICE PROTECTION group control the heat. The left and right generator buses power the system. Whenever ice is anticipated or encountered, turn the system on.

HEAT EXCHANGER An oil-to-fuel heat exchanger on the engine accessory section protects the fuel against ice. An engine oil line is next to the fuel line within the heat exchanger. Heat transfer occurs through conduction to melt ice particles that may have formed in the fuel. This operation is automatic whenever the engines are running.

Refer to the Fuel section of this manual for m o re d e t a i l s a b o u t f u e l h e a t a n d t h e Limitations section of the POH for temperature limitations concerning the oil-to-fuel heat exchanger.

ELECTRICAL HEATING PITOT HEAT Two pitot masts on the nose of the aircraft contain electric heating elements to ensure proper airspeed indication during icing conditions (Figure 10-20). Two PITOT LEFT-RIGHT circuit breaker switches on the pilot right subpanel control the heating elements. In the up position, the heating system is on; in the down position, the system is off. The triple-fed bus powers the left pitot heat; the right generator bus powers the right pitot heat.


Figure 10-20. Pitot Mast and Heat Controls

10-20



Th e p i t o t h e a t s y s t e m s h o u l d n o t b e operated on the ground except for testing or for short intervals to remove snow or ice from the mast. Turn pitot heat on for takeoff. It can be left on in flight during icing conditions or whenever icing conditions are expected. If a gradual reduction in airspeed indication occurs during flight at altitude, there may be pitot icing. If turning pitot heat on restores airspeed, leave the system on. Many pilot use pitot heat as standard practice during all flights at high altitude and/or cold temperatures to prevent pitot icing.

Illumination of the amber L or R PITOT HEAT annunciator indicates switch is off or the pitot heat system is inoperative. Icing of a pitot probe could impact validity of aircraft instruments.

STALL WARNING VANE Electric heat elements for the stall warning vane and plate protect them against freezeup during icing conditions (Figure 10-21). A two-position STALL WARN switch on the ICE PROECTION panel activates the system. Up position is on; down position is off.


Operation

Figure 10-21. Stall Warning Vane and Heat Control


10-21


The right generator bus supplies power. A safety switch on the left landing gear limits face plate and stall vane current flow to approximately 14 volts to prevent overheating while the aircraft is on the ground. In flight, after the left strut extends, the full 2 8 -v o lt c u r re n t i s a p p l i e d to th e s t a l l warning face plate and stall vane.

Precautions Even though the heating elements protect the lift transducer vane and face plate from ice, a buildup of ice on the wing may change or disrupt airflow. This prevents the system from accurately indicating an imminent stall. Remember stall speed increases whenever ice accumulates on any aircraft.

PRECAUTIONS DURING ICING CONDITIONS An aircraft needs special care and inspection before operating in cold or potential icing weather. When the aircraft is at rest, always cover pitot masts. Once covers are removed, ensure both masts and drains are free of ice or water. If they are clogged, faulty readings may result. When the engine is not operating, install tie-downs for propellers to ensure against damage to internal engine components not lubricated. Spinning propellers can also be a source of danger to crew, passengers, and ground support personnel. Propeller blades in their tie-down position channel moisture down the blades past the propeller hub and off the lower blade more effectively than in other positions or when left spinning. 10 ICE AND RAIN PROTECTION

During particularly icy ground conditions, inspect the propeller hubs for ice and snow accumulation. Turn the propellers by hand in their normal rotation direction to ensure they are free prior to engine start.

10-22

EXTERIOR INSPECTION During the normal exterior inspection, pay special attention to areas where frost and ice may accumulate. It is not the thickness of frost that matters, it is the texture. Any slightly irregular surface can substantially decrease proper airflow over the wings and stabilizers. Do not underestimate the damaging effects of frost. Remove all frost from leading edges of the wings, stabilizers, and propellers before the aircraft is moved. The windshield, control surfaces, hinges, pitot masts, fuel tank caps, and vents should also be free of frost. Use deicing fluid when needed.

Fuel Checks Test fuel drains for free flow. Water in the fuel system has a tendency to condense more readily during winter months. If left unchecked, large amounts of moisture may accumulate in the fuel tanks. Moisture does not always settle at the bottom of the tank. O c c a s i o n a l l y a t h i n l a y e r o f f u el gets trapped under a large mass of water that may deceive the tester. Make sure to take a good-sized sample of fuel. Add only the correct amount of anti-icing additive to the fuel. A higher concentration of anti-icer does not ensure lower fuel freezing temperatures. It may actually hinder the aircraft’s performance. Consult the Normal Procedures section of the Pilot’s Operating Handbook to determine correct blend.

Brakes and Tires C h e c k t h e b ra ke s a n d t i re - t o - g ro u n d contact for lockup. Do not use any anti-ice solution containing oil-based lubricant on the brakes.



I f t i re s a re f ro z e n t o t h e g ro u n d , u s e undiluted defrosting fluid or a ground heater to melt ice around tires. Move the aircraft as soon as the tires are free. Heat applied to tires should not exceed 160°F or 71°C.

BEFORE TAXI Brake deice may be turned on before taxi to help expel accumulated ice from the brake mechanisms. If brake deice is used, place the condition levers in HIGH IDLE. Keep flaps retracted to avoid throwing snow or slush into the flap mechanisms. This minimizes the possibility of damage to flap surfaces. When taxiing in extremely icy conditions, ensure the tires are rolling, not just sliding on the icy surface. Th e b r a k e d e i c e s y s t e m i s n o t t o b e c o n t i n u o u s l y o p e r a t e d a b o v e 15 º C ambient temperature.


Leave auto-ignition off until immediately before takeoff. This helps prolong the service life of the igniter units.


10-23




10-24



QUESTIONS 1. During flight in visible moisture, or at night when flight from visible moister cannot be assured, engine anti-ice must be on at temperatures below _____°C. A. 0 B. 5 C. 10 D. 15 2. In the event of windshield icing, reduce speed to ______ knots or below. A. 140 B. 170 C. 184 D. 226 3. Operating the propeller deice in the _______ mode provides _______ timer operation. A. MANUAL; automatic B. MANUAL; manual C. AUTO; manual D. AUTO; automatic 4. The surface deice system removes ice build up on the leading edge(s) of the: A. Horizontal stabilizer. B. Vertical stabilizer. C. Wing and horizontal stabilizer. D. Wing and vertical stabilizer.


5. The minimum airspeed for sustained flight in icing conditions is _____ knots. A. 140 B. 160 C. 226 D. 263


10-25

11 AIR CONDITIONING


CHAPTER 11 AIR CONDITIONING CONTENTS Page INTRODUCTION............................................................................................................. 11-1 GENERAL ......................................................................................................................... 11-1 FL-1, FL-4-492, FL-494-500............................................................................................... 11-4 Components................................................................................................................. 11-4 Operation ..................................................................................................................... 11-9 Heating....................................................................................................................... 11-10 Electric Heat ............................................................................................................. 11-11 Vent Blower Control ................................................................................................ 11-12 FL-493, FL-500, AND SUBSEQUENT........................................................................ 11-12 Components .............................................................................................................. 11-12 Compressor................................................................................................................ 11-13 Operation................................................................................................................... 11-20 LIMITATIONS................................................................................................................. 11-22 EMERGENCY/ABNORMAL ...................................................................................... 11-22 QUESTIONS.................................................................................................................... 11-23


11-i


Title

Page

11-1

Air Conditioning System (FL-1-492, FL-494-499)........................................... 11-2

11-2

Air Conditioning System (FL-493-500, and Subsequent) .............................. 11-3

11-3

Condenser and Receiver-Dryer Sight Gauge .................................................. 11-4

11-4

Air Conditioner Condenser Intake................................................................... 11-4

11-5

Floor and Ceiling Outlets................................................................................... 11-5

11-6

Cockpit Eyeball Outlets ..................................................................................... 11-5

11-7

Air Conditioning System Control Diagram ..................................................... 11-6

11-8

ENVIRONMENTAL Panel .............................................................................. 11-7

11-9

CABIN TEMP MODE Control Knob............................................................. 11-7

11-10

MANUAL TEMP INCR-DECR Switch ......................................................... 11-8

11-11

ELECT HEAT-OFF Switch............................................................................... 11-8

11-12

Annunciator Panel .............................................................................................. 11-9

11-13

PILOT AIR and COPILOT AIR Knobs....................................................... 11-11

11-14

Condenser and Receiver-Dryer Sight Gauge................................................ 11-13

11-15

Cockpit Eyeball Outlets................................................................................... 11-14

11-16

Supplemental Heat Vent.................................................................................. 11-14

11-17

Floor and Ceiling Outlets ................................................................................ 11-14

11-18

Air Conditioning System Control Diagram................................................... 11-15

11-19

ENVIRONMENTAL Panel ............................................................................ 11-16

11-20

Environmental System Control Knobs........................................................... 11-16

11-21

COCKPIT and CABIN BLOWER Control Knobs ..................................... 11-17

11-22

MAN TEMP INCR-DECR Switch ................................................................ 11-18

11-23

ENVIR BLEED AIR Switch .......................................................................... 11-18

11-24

Annunciator Panel ............................................................................................ 11-19


11-iii

11 AIR CONDITIONING


CHAPTER 11 AIR CONDITIONING

INTRODUCTION This chapter describes the air conditioning system on the King Air 350 that provides cooling, heating, and unpressurized ventilation. Electric heat is also available. The air conditioning system can be operated in the heating mode or cooling mode with either automatic or manual mode control.

GENERAL Th e a i r c o n d i t i o n i n g s y s t e m p ro v i d e s cooling, heating, and unpressurized ventilation inside the aircraft (Figures 11-1 and 1 1- 2 ) . A i rc ra f t F L - 4 - 4 9 2 , F L - 5 0 0 a n d subsequent includes the new Keith ECS system. The dual zone system allows the cabin temperature to be controlled independently of the cockpit temperature.

A refrigerant gas, vapor-cycle refrigeration system provides cabin cooling. Bleed air from the compressor of each engine flows into the cabin for heating and for pressurization. A supplemental electric heating system is available.


11-1

11 AIR CONDITIONING


11 AIR CONDITIONING

11-2

ENVIRONMENTAL BLEED–AIR SHUTOFF VALVE AMBIENT AIR PNEUMATIC MODULATING VALVE THERMOSTAT

REFRIGERANT COMPRESSOR

CABIN AIR CONTROL RETURN AIR VALVE FWD EVAPORATOR FWD EVAPORATOR AIR FILTER


VENT BLOWER FRESH AIR VALVE (CLOSED WHEN PRESSURIZED)

SIDE VIEW FWD

DETAIL A DOOR TO AFT FLOOR OUTLETS TO CEILING OUTLETS

FLOOR DUCT

AIR–CONDITIONED AIR FROM AFT EVAPORATOR

CEILING OUTLET

DOOR (COOLED AIR TO FLOOR OUTLETS) FLOOR OUTLET SAFETY/DUMP VALVE

NORMAL OUTFLOW VALVE

RAM–AIR SCOOP CONDENSER

CONDENSER BLOWER RECEIVER–DRYER

OUTLET AIR MIXING PLENUM WINDHSHIELD DEFROSTER (ON GLARESHIELD)

FWD PRESSURE BULKHEAD CREW HEAT DUCT ENVIRONMENTAL BLEED–AIR FLOW CONTROL UNIT INCLUDING MODULATING AND SHUTOFF VALVE

WINDSHIELD DEFROSTER CONTROL PILOT'S VENT AIR CONTROL INSTRUMENT PANEL

CEILING OUTLET

FLOOR CEILING OUTLET OUTLETS

FLOOR OUTLET

CEILING AFT PRESSURE BULKHEAD OUTLET

CABIN–HEAT CONTROL VALVE

ENVIRONMENTAL BLEED–AIR SHUTOFF VALVE

LEGEND HOT ENGINE BLEED AIR ENVIRONMENTAL BLEED AIR RECIRCULATED CABIN AIR (AIR CONDITIONED WHEN EVAPORATOR IS ON)

PNEUMATIC PNEUMATIC BLEED–AIR THERMOSTAT SHUTOFF AMBIENT AIR VALVE MODULATING VALVE

AIR INLET SCOOP

AIR–TO–AIR HEAT EXCHANGER

FIREWALL

Figure 11-1. Air Conditioning System (FL-1-492, FL-494-499)

AMBIENT AIR PRESSURE VESSEL


RETURN AIR FILTER

ENVIRONMENTAL BLEED– AIR FLOW CONTROL UNIT INCLUDING MODULATING AND SHUTOFF VALVE FIREWALL AIR–TO–AIR HEAT EXCHANGER PNEUMATIC BLEED–AIR SHUTOFF VALVE AFT HEATER CABIN–HEAT AIR INLET CONTROL REFRIGERANT LINES VALVE AFT CEILING EVAPORATOR DUCT/FLOOR AIR FILTER DUCT DIVIDER DUCT AFT FORWARD OVERTEMP COPILOT'S EVAPORATOR SENSOR HEATER VENT AIR FLAP– CONTROL FLOOR COPILOT'S PER CEILING OUTLET CEILING CABIN AIR OUTLET VALVE OUTLET CONTROL CEILING VALVE OUTLET



11-3

Figure 11-2. Air Conditioning System (FL-493-500 and Subsequent)

11 AIR CONDITIONING

11 AIR CONDITIONING


FL-1, FL-4-492, FL-494500

RECIEVERDRYER AND SIGHT GAUGE

COMPONENTS The environmental system has the following main components: • Belt-driven compressor (right engine) • Condenser blower • Evaporator • Aft evaporator • Forward vent blower

Figure 11-3. Condenser and ReceiverDryer Sight Gauge

• Mixing plenum • Floor outlet ducts • Ceiling eyeball outlets • Temperature-sensing device • Autotemperature controller • Flow control unit • Pilot/copilot outlets • Defroster • Air-to-air heat exchangers • Bleed air valves • Heating air outlets

Figure 11-4. Air Conditioner Condenser Intake

Compressor A belt-driven compressor on the right engine operates in either auto or manual cool modes. The compressor has built-in safety devices to prevent its operation in refrigerant over or underpressure conditions.

Condenser Blower The condenser sits slightly sideways in the nose crossover duct (Figure 11-3). Ram air passes through the condenser, then condenses, and cools the refrigerant gas passing through it into liquid form for use in cooling the cabin air (Figure 11-4). The condenser blower enhances airflow through the condenser for more efficient operation. 11-4

Forward Evaporator and Blower The cockpit blower motor recirculates cockpit air through the evaporator in the right side of the nose behind the crossover d u c t ( a l s o re f e r re d t o a s t h e f o r w a rd evaporator). Th e r e f r i g e r a n t f l o w s t h r o u g h t h e evaporator and absorbs heat from the recirculated cockpit air to cool the air passing through it.


11 AIR CONDITIONING


Aft Evaporators and Blowers The aft evaporator and blower are under the floor in the rear of the cabin. The blower draws in cabin air and blows it across the evaporator to the aft floor and ceiling outlets. It operates at high speed only.

Mixing Plenum The mixing plenum is in the right side of the nosecone under the copilot floorboard and aft of the forward evaporator. Within the plenum bleed air mixes with recirculated cabin air, before it is routed back into the cabin.

Cabin Floor Outlet Ducts Th e f l o o r o u t l e t d u c t s a r e b e t w e e n the passenger seats along the aircraft floorboards where they contact the interior sidewall of the aircraft cabin (Figure 11-5). Pressurization air heated as required by the environmental system enters the cabin through these vents. Figure 11-5. Floor and Ceiling Outlets

Ceiling Eyeball Outlets Eyeball outlets in the headliner provide cool air to the crew and passengers (Figures 11-5 and 11-6). Each outlet can be adjusted to direct the airstream as desired. Twisting the nozzle adjusts air volume from full open to closed. As the nozzle is twisted, a damper opens or closes to regulate airflow. The cockpit has two eyeball outlets; the cabin has seven such outlets.

Figure 11-6. Cockpit Eyeball Outlets


11-5

11 AIR CONDITIONING


Temperature-Sensing Device The cabin temperature sensor works with the CABIN TEMP MODE switch to achieve desired temperature (Figure 11-7). The sensor in the floor ducts monitors the bleed air temperature. If excessive temperature extreme is sensed, the sensor activates an annunciator in the cockpit.

Auto Temperature Controller

On the ground, these flow control units supply only bleed air to the environmental system.

Defrost System Two ducts provide warm air to the defroster below the windshields where they contact the top of the glareshield. The DEFROST AIR knob controls warm air flow through the ducts.

When the CABIN TEMP MODE switch is positioned to AUTO, the automatic temperature controller uses inputs from the cabin temperature sensor to adjust the system to maintain the desired temperature (Figure 11-7).

Air-to-Air Heat Exchangers

Flow Control Unit

Bleed Air Valves

In flight, flow control units on each engine firewall mix outside ambient air with bleed air to make bleed air temperature more manageable for the environmental system.

The bleed air valves are in the environmental flow control units on each engine firewall. These valves control bleed air flow into the aircraft and into the environmental, pressurization, and pneumatic systems (Figure 11-7).

An air-to-air heat exchanger is in the center section of each wing inboard of the engines. Bleed air passes through the air-to-air heat exchangers to reduce air temperature.

LH BYPASS VALVE MOTOR MANUAL TEMP INCR

MANUAL HEAT OR COOL

DECR

AIR TO AIR HEAT EXCHANGER

HEAT

COOL

HEAT LEFT ENGINE BLEED AIR

AUTO MANUAL COOL

TO CABIN

AUTO TEMP CONTROLLER

TO CABIN COOL AIR TO AIR HEAT EXCHANGER

TEMP SENSORS DUCT CABIN SELECTOR

RH BYPASS VALVE MOTOR

RIGHT ENGINE BLEED AIR AIR CONDITIONER LH BYPASS VALVE MOTOR SWITCH

Figure 11-7. Air Conditioning System Control Diagram

11-6


Cockpit Heating Air Outlets

CABIN TEMP MODE Knob

Two ducts under the instrument panel deliver warm air to the pilot and copilot. The PILOT AIR knob and CO-PILOT AIR knob control the warm air flow through these ducts.

The CABIN TEMP MODE knob has four positions (Figure 11-9):

Controls and Indications

• OFF—Air delivery system completely shut off; no bleed air input to cockpit or cabin

The ENVIRONMENTAL panel on the copilot left subpanel provides automatic or manual control of the air conditioning system (Figure 11-8).

• AU T O — A i r c o n d i t i o n i n g a n d heating systems operate automatically to establish pilot-requested temperature

BLEED AIR VALVES Switches Tw o B L E E D A I R VA LV E S s w i t c h e s control the inflow of pressurization air and are used for cockpit and cabin climate c o n t r o l . Th e s w i t c h e s a r e o n t h e ENVIRONMENTAL panel on the copilot left subpanel (Figure 11-8).

• MAN COOL—Air conditioning system operates in response to manual input; air conditioner operates as long as system pressures are acceptable and right engine N1 speed above 62% • M A N H E AT — H e a t i n g s y s t e m operates in response to manual input

Each switch has three positions: • OPEN—Allows bleed air into cabin for pressurization and climate control • ENVIR OFF—Restricts bleed air from the respective side environmental flow control unit from entering pressurization and air conditioning systems (for maximum cooling on ground, place switches in ENVIR OFF position) • INSTR & ENVIR OFF—Respective bleed air valve closes completely to deny bleed air to pressurization, air conditioning and pneumatic systems

Figure 11-9. CABIN TEMP MODE Control Knob

Figure 11-8. ENVIRONMENTAL Panel


11-7

11 AIR CONDITIONING


11 AIR CONDITIONING


MAN TEMP INCR-DECR Switch The MAN COOL or MAN HEAT position of the CABIN TEMP MODE switch allows manual adjustment of cockpit and cabin temperature. Momentarily positioning the MANUAL TEMP switch (Figure 11-10) to either INCR (increase) or DECR (decrease) repositions the bleed air valves to adjust cabin and cockpit temperature. When released, the switch returns to the OFF position.

Figure 11-11. ELECT HEAT-OFF Switch

AIR COND N1 LOW Annunciator Th e N 1 s p e e d s w i t c h ( e n g i n e s p e e d ) prevents compressor operation outside of established limitation parameters. The N 1 speed switch disengages the compressor clutch when engine speed is below 62% N 1 and air conditioning is requested. Figure 11-10. MANUAL TEMP INCR-DECR Switch

ELECT HEAT–OFF Switch The electric heat system is operated by a solenoid ELECT HEAT–OFF switch on the copilot left subpanel (Figure 11-11). The cabin can be warmed before engine start with the electric heat system running c o n c u r re n t l y w i t h a n a u x i l i a r y p o w e r unit.Use of the electric heat system is only permissible during ground operations. The system is squat switch protected from airborne operation.

If air conditioning is requested when the N 1 speed switch opens, the white AIR COND N 1 LOW annunciator illuminates (Figure 11-12).

DUCT OVERTEMP Annunciator If airflow in the ducts becomes too low, the amber DUCT OVERTEMP annunciator illuminates to indicate duct temperature has reached approximately 30 0°F (148°C) (Figure 11-12).

ELEC HEAT ON Annunciator Th e g re e n E L E C H E AT O N a d v i s o r y annunciator indicates the power relays are closed to apply power to the heating elements (Figure 11-12). Before blowers are selected OFF when electric heat is off, the ELEC HEAT ON annunciator must be extinguished to indicate power is removed from the heating elements.

11-8


11 AIR CONDITIONING


Figure 11-12. Annunciator Panel

L-R BL AIR OFF Annunciators Amber L-R BL AIR OFF annunciators illuminate whenever the respective BLEED AIR VALVES OPEN switch is in any position other than OPEN.

Airflow Control Knobs Four additional manual airflow push-pull knobs on the subpanels regulate cockpit and cabin comfort. When the cockpit door is closed and the cabin comfort level is satisfactory, each CABIN/COCKPIT AIR push-pull knob regulates airflow to the cockpit and cabin. When fully pulled out, each knob provides maximum airflow to the respective area. When fully pushed in, each knob provides minimum airflow.

OPERATION Automatic Mode Control The AUTO position on the CABIN TEMP MODE knob commands the automatic temperature control to modulate the bypass valves and activate the air conditioning compressor (see Figure 11-9).

For greater heating, bleed air bypasses the air-to-air heat exchangers in the wing center sections. For greater cooling, the bleed air passes through the air-to-air heat exchangers to reduce its temperature. In either case, the resulting bleed air mixes with recirculated cabin air that can be additionally cooled when the air conditioning compressor in the forward mixing plenum is in cooling mode.

Cooling The plumbing from the compressor on the right engine is routed through the right wing and then forward to the condenser coil, receiver-dryer, expansion valve, bypass valve, and forward evaporator, all of which are in the aircraft nose. The forward vent blower moves recirculated cabin air through the forward evaporator and into the mixing plenum, the floor-outlet ducts, and ceiling eyeball outlets. Approximately 75% of the recirculated air passes through the floor outlets while approximately 25% of the air bypasses the mixing plenum and flows through the ceiling outlets.


11-9

11 AIR CONDITIONING


With the system in AUTO, the forward vent blower normally runs at low speed. If the cooling mode is operating, refrigerant circulates through the forward evaporator to cool the output air. If either BLEED A I R VA LV E S s w i t c h i s p o s i t i o n e d t o OPEN, air entering the ceiling-outlet duct is cooler than air entering the floor outlets. The air discharges through the eyeball outlet in the cockpit and cabin (see Figures 11-5 and 11-6). Cool air also enters the floor-outlet duct. In order to provide cabin pressurization, however, warm bleed air also enters this duct any time either BLEED AIR VALVES switch is in the OPEN position. Therefore, pressurized air discharged from the floor outlets is always warmer than air discharged from the ceiling outlets no matter what temperature mode is used.

NOTE On aircraft with cargo doors, a lever on each floor outlet register (except forward facing register in baggage compartment) can be moved vertically to regulate airflow. A vane-axial blower in the nose section draws ambient air through the condenser to cool the refrigerant gas when the cooling mode is operating on the g ro u n d . Th i s b l o w e r s h u t s o ff automatically when gear is retracted. Th e r e c e i v e r- d r y e r a n d s i g h t g a u g e (glass) are in the upper portion of the condenser compartment. Remove the upper-compartment access panel on top of the nose section left of centerline to view these components. Th i s a c t i o n , h o w e v e r, i s n o t a n o r m a l preflight action.

11-10

If bubbles are seen through the sight glass (see Figure 11-3), then the refrigerant system is low on refrigerant gas. If, after adding more refrigerant gas, bubbles still appear in the sight glass, the system needs to be evacuated and recharged. A f t e va p o ra t o r s a n d b l o w e r s p ro v i d e additional cooling. The blowers recirculate cabin air across the evaporators and route it to the aft floor and ceiling outlets. Th e a f t e va p o ra t o r s i n c re a s e a i rc ra f t c o o l i n g c a p a c i t y f r o m 18 , 0 0 0 B T U (with the forward evaporator only) to 32,0 0 0 BTU. Refrigerant flows through the aft evaporator any time it flows through the forward evaporator. The additional cooling, however, is provided only when the aft blower is operating. During flights in warm air, such as short, low-altitude flights in the summer, all the cabin ceiling outlets must be fully open for maximum cooling.

HEATING Description Bleed air from the compressor of each engine flows into the cabin for heating and pressurization purposes. When the left landing gear safety switch is in the ground position, the ambient air valve in each flow control unit is closed. Therefore, only bleed air is delivered. When airborne, bleed air is mixed with outside ambient air from the ambient air valve in each flow control unit until a cold air temperature closes off the ambient flow. Then, only bleed air is delivered.


Operation In the cockpit, adjust either pilot damper to provide additional air. The PILOT AIR and COPILOT AIR knobs control the dampers (Figure 11-13). Movement of these k n o b s a ff e c t s c o c k p i t t e m p e ra t u re b y adjusting air volume.

The DEFROST AIR knob controls a valve on the pilot/copilot heat duct that admits air to two ducts delivering warm air to the defroster vents below the windshields. The rest of the air in the bleed-air duct mixes with recirculated cabin air and flows a f t t h ro u g h t h e f l o o r- o u t l e t d u c t t h a t handles 75% of the total airflow. During high-altitude flights, cool-night flights, and flights in cold weather, the ceiling outlets must be closed for maximum cabin heating.

ELECTRIC HEAT Operation

Figure 11-13. PILOT AIR and COPILOT AIR Knobs

The CABIN/COCKPIT AIR knob on the copilot left subpanel controls air volume to the cabin (see Figure 11-8). This knob controls the cabin air control valve. When pulled out of its stop, a minimum amount of air passes through the valve to the cabin to increase the volume of air available to the pilot and copilot outlets and defroster. When the knob is pushed all the way in, the valve opens to allow air in the duct to be directed into the cabin floor outlets.

Positioning the ELECT HEAT switch to ON energizes the heating elements in the forward duct and aft evaporator plenum (see Figure 11-11). The green ELEC HEAT ON annunciator illuminates to indicate power is being applied to the heating elements (see Figure 11-12). The electric heat system draws approximately 300 amps. During electric heat operation, the forward and the aft blowers must be operating. B e f o r e t h e E L E C T H E AT s w i t c h i s positioned to OFF and the BLOWER knob is positioned to OFF, the green ELEC H E AT O N a n n u n c i a t o r m u s t b e extinguished. This indicates the heating elements have been sufficiently deenergized for safety.


11-11

11 AIR CONDITIONING


11 AIR CONDITIONING


VENT BLOWER CONTROL Unpressurized Ventilation Fresh air is available during unpressurized flight with the CABIN PRESS switch in the DUMP position. This ambient (ram) air is obtained through the fresh air door and the ram-air scoop in the aircraft nose section (see Figure 11-4). This door is open only during unpressurized flight when the switch is in the DUMP position and there is 0 psi. This allows the forward blower to draw ram air into the cabin. This air is mixed with recirculated cabin air in the plenum chamber and then directed to both the floor registers and ceiling outlets. The CABIN AIR control knob regulates the air volume.

C a b i n t e m p e ra t u re va r i e s, t h e re f o re, according to the position of the cabin-heat control valves and whether or not the refrigerant system is working.

NOTE The air conditioner compressor does not operate unless the bypass valves are closed. To ensure that the valves are closed, select MAN COOL then hold the MANUAL TEMP switch in the DECR position for one minute.

FL-493, FL-500, AND SUBSEQUENT COMPONENTS

NOTE A flight conducted with the bleedair switches placed in any position other than OPEN also results in unpressurized flight, but the fresh air door is not open.

The environmental system has the following main components: • Belt-driven compressor (right engine) • Condenser blower • Evaporator

Manual Mode Control

• Aft evaporator

The MAN COOL or MAN HEAT position of the CABIN TEMP MODE control knob allows manual control of the cabin and cockpit temperature.

• Forward vent blower

Th e M A N UA L T E M P I N C R – D E C R switch returns to the center OFF position when released. When held in either position, it modulates the bypass valves in the bleed-air lines. Allow one minute (30 seconds per valve) for both valves to move fully open or fully closed.

• Ceiling eyeball outlets

Only one valve moves at a time to vary the amount of bleed air routed through the airt o - a i r h e a t e x c h a n g e r . Th i s c a u s e s a variance in bleed-air temperature. The bleed air mixes with recirculated cabin air in the mixing plenum and is then routed to the floor registers.

11-12

• Forward and aft mixing plenums • Floor outlet ducts • Temperature-sensing devices • Autotemperature controller • Flow control unit • Pilot/copilot outlets • Defroster • Air-to-air heat exchangers • Bleed air valves • Heating air outlets


COMPRESSOR A belt-driven compressor on the right engine operates in either auto or manual cool modes. The compressor has built-in safety devices that prevent its operation in cases of refrigerant over- or underpressure conditions.

Condenser Blower

Forward Evaporator and Blower The forward evaporator blower motor recirculates cockpit air through the forward evaporator in the right side of the nose behind the crossover duct. The refrigerant flowing through the evaporator absorbs heat from the recirculated cockpit air, cooling the air passing through it.

Two condensers joined together in a V are in the nose crossover duct. Ram air flowing through the condenser condenses and cools the refrigerant gas passing through it into liquid for use in cooling the cabin air (Figure 11-14).

Aft Evaporators and Blowers

The condenser blower enhances airflow through the condenser for more efficient operation and runs in the auto or manual cool modes when the air conditioner is operating.

Th e r e f r i g e r a n t f l o w i n g t h r o u g h t h e evaporator tubing absorbs heat from the recirculated air cooling it before it returns to the cabin. The cooled air reenters the cabin through the aft floor and ceiling outlets. When the air conditioning system is off, the blowers provide recirculated cabin air for ventilation.

High speed fans blow recirculated cabin air through two evaporators under the floorboards in the center aft cabin behind the main spar.

FWD and AFT Mixing Plenums

RECIEVERDRYER AND SIGHT GAUGE

Bleed air coming into the aircraft is routed i n t o f o r w a rd a n d a f t m i x i n g p l e n u m s beneath the cabin floorboards. The mixing plenums combine bleed air with recirculated cabin air to reduce bleed air temperature for passenger comfort. The conditioned air is then routed into the cabin.

Figure 11-14. Condenser and ReceiverDryer Sight Gauge


11-13

11 AIR CONDITIONING


11 AIR CONDITIONING


Cockpit Heating Air Outlets

Cabin Floor Outlet Ducts

Heating air outlets are under the instrument panel and outboard of the pilot and copilot seats on the floor (Figure 11-15). The C O C K P I T B LOW E R c o n t r o l k n o b controls air flow volume to these outlets.

The floor outlet ducts are between the p a s s e n g e r s e a t s a l o n g t h e a i rc ra f t floorboards where they contact the interior sidewall of the aircraft cabin (Figure 11-17). Pressurization air heated as required by the environmental system enters the cabin through these vents.

Ceiling Eyeball Outlets Eyeball outlets in the headliner provide cool air to the crew and passengers (Figure 11-17).

Figure 11-15. Cockpit Eyeball Outlets

Each outlet can be adjusted to direct the airstream as desired. Twisting the nozzle adjusts air volume from full open to closed. As the nozzle is twisted, a damper opens or closes to regulate airflow. The cockpit has two eyeball outlets; the cabin has seven such outlets.

The supplemental electric heat system discharges warm air directly aft of the cockpit center pedestal through a single floor outlet (Figure 11-16).

Figure 11-16. Supplemental Heat Vent

Figure 11-17. Floor and Ceiling Outlets

11-14


Temperature-Sensing Devices The cockpit and cabin temperature sensors work with the following to adjust the air to the desired temperatures (Figure 11-18):

Environmentally conditioned air flows constantly to the windshield defrost and glareshield outlets.

• Cockpit and cabin temperature control knobs

In AUTO mode, the air is regulated to a maximum temperature of 70°F (21°C). If more heat is required in colder environments, the temperature of the outlet air is allowed to increase to 105°F (41°C).

A sensor in the floor ducts monitors the bleed air temperature. If excessive temperature is sensed, the sensor activates an annunciator in the cockpit.

In MAN HEAT mode, the COCKIT TEMP knob controls glareshield and overhead temperatures. Airflow can be increased with the BLOWER knob.

Auto Temperature Controller

Flow Control Unit

With the ENVIRONMENT MODE switch p o s i t i o n e d t o AU T O, t h e a u t o m a t i c temperature controller uses inputs from cockpit and cabin temperature sensors to adjust the system and maintain the desired temperatures (Figure 11-18).

In flight, flow control units on each engine firewall mix outside ambient air with bleed air to make bleed air temperature more manageable for the environmental system.

• ENVIRONMENT MODE switch

On the ground, these flow control units supply only bleed air to the environmental system.

LH BYPASS VALVE MOTOR MANUAL TEMP INCR

MANUAL HEAT OR COOL

DECR

AIR TO AIR HEAT EXCHANGER

HEAT

COOL

HEAT LEFT BLEED AIR

AUTO MANUAL COOL

TO CABIN

AUTO TEMP CONTROLLER

TO CABIN COOL AIR TO AIR HEAT EXCHANGER

TEMP SENSORS DUCT CABIN SELECTOR

RH BYPASS VALVE MOTOR

RIGHT ENGINE BLEED AIR AIR CONDITIONER LH BYPASS VALVE MOTOR SWITCH

Figure 11-18. Air Conditioning System Control Diagram


11-15

11 AIR CONDITIONING


11 AIR CONDITIONING


Defrost System

BLEED AIR VALVES Switches

Two ducts provide warm air to the defroster below the windshields where they contact the top of the glareshield.

Tw o B L E E D A I R VA LV E S s w i t c h e s control the inflow of pressurization air for cockpit and cabin climate control (Figure 11-19). Each switch has three positions:

Air-to-Air Heat Exchangers An air-to-air heat exchanger is in the center section of each wing inboard of the engines. Bleed air passes through the air-to-air heat exchangers to reduce the air temperature.

Bleed Air Valves The bleed air valves are in the environmental flow control units on each engine firewall. The valves control bleed air flow into the aircraft and into the environmental, pressurization and pneumatic systems (Figure 11-19).

Controls And Indications

• OPEN—Allows bleed air into the cabin for pressurization and climate control • ENVIR OFF—Restricts bleed air from the respective side environmental flow control unit from entering the pressurization and air conditioning systems; for maximum cooling on the ground, place switches in ENVIR OFF position • PNEU & ENVIR OFF—Respective bleed air valve closes completely to deny bleed air to the pressurization, air conditioning and pneumatic systems.

The ENVIRONMENTAL panel on the copilot left subpanel provides automatic or manual control of the air conditioning system (Figure 11-19).

Figure 11-19. ENVIRONMENTAL Panel

11-16


11 AIR CONDITIONING


Environmental Control Knob The environmental control system knob has five positions (Figure 11-20): • OFF—Air completely shut off; no bleed air input to cockpit or cabin • AU T O — A i r c o n d i t i o n i n g a n d heating systems operate automatically to establish pilot-requested temperature • MAN COOL—Air conditioning system operates in response to manual input from pilot; air conditioner operates as long as system pressures acceptable and right engine N 1 speed above 62% • M A N H E AT — B o t h c a b i n a n d cockpit floor heat servos opened fully; accomplish cockpit and cabin temperatures through MAN TEMP switch to either INCR or DECR • E L E C H E AT — D i r e c t s a i r o v e r resistive heater elements into the cabin; operative on the ground only.

Figure 11-21. COCKPIT and CABIN BLOWER Control Knobs

Rotate the COCKPIT TEMP control knob as required to adjust cockpit temperature. A temperature sensor in the cockpit, in conjunction with the temperature setting, initiates a heat or cool command to the temperature controller. Rotate the CABIN TEMP control knob to adjust cabin temperature. A temperature sensor behind the first set of passenger oxygen masks, in conjunction with the temperature setting, initiates a heat or cool command to the temperature controller.

COCKPIT/CABIN BLOWER Knobs The COCKPIT and CABIN BLOWER knobs control the forward and aft vent blower (Figure 11-21). Each knob has two positions: Figure 11-20. Environmental System Control Knobs

COCKPIT/CABIN TEMP Knobs The COCKPIT and CABIN TEMP control knobs regulate the temperature in the AUTO and manual positions (Figure 11-21).

• AU TO — B l o w e r o p e ra t e s a t l o w speed if environmental control system knob is in any position except OFF and cabin/cockpit temperature has reached the set point as selected b y t h e c re w. I f t h e c a b i n / c o ckpit temperature is significantly different than the desired temperature, the blowers automatically come on high and then slow as the temperature approaches the desired set point.


11-17

11 AIR CONDITIONING


• Out of AUTO—Allows pilot to set desired blower speed. Wh e n t h e V E N T B LOW E R s w i t c h i s positioned to AUTO and the environmental control system knob is positioned to OFF, the blower ceases operation.

MAN TEMP INCR-DECR Switch The MAN COOL or MAN HEAT position of the environmental control system knob allows manual adjustment of the cockpit and cabin temperature (Figure 11-22). Momentarily positioning the MAN TEMP switch to the INCR or DECR regulates bleed air temperature as it enters the aircraft; it does not affect the flow rate.

• AU T O — A l l o w s e n v i r o n m e n t a l system controller to select flow setting automatically to maintain cockpit temperature or cabin pressure requirements; recommended position for most operations. • LOW—Default setting except when system demands additional heating I f t h e e n v i ro n m e n t a l c o n t ro l k n o b i s positioned to MAN HEAT, the bleed flow defaults to NORMAL. If the flow is positioned to MAN COOL, the bleed flow defaults to LOW.

CAUTION Always monitor cabin pressuri z a t i o n re q u i re m e n t s w h e n i n MAN COOL. Manual adjustments to the ENVIR BLEED AIR flow setting may be required. For maximum engine performance and/or high altitude takeoff requirements, position the ENVIR BLEED AIR switch to LOW.

Figure 11-22. MAN TEMP INCR-DECR Switch

ENVIR BLEED AIR Switch The ENVIR BLEED AIR switch on the copilot left subpanel controls bleed air flow volume (Figure 11-23). The switch has three positions:

Figure 11-23. ENVIR BLEED AIR Switch

• NORMAL—For increased heating or pressurization airflow.; use during climb to ensure optimum pressurization at higher altitude

11-18


AIR COND N1 LOW Annunciator

ELEC HEAT ON Annunciator

Th e N 1 s p e e d s w i t c h ( e n g i n e s p e e d ) prevents compressor operation outside of established limitation parameters. The white AIR COND N 1 LOW annunciator illuminates to indicate that the right engine speed is below 62% N 1 and air conditioning is requested (Figure 11-24).

DUCT OVERTEMP Annunciator If airflow in the ducts becomes too low, the amber DUCT OVERTEMP annunciator illuminates to indicate duct temperature has reached approximately 300°F (148°C) (Figure 11-24).

The amber ELEC HEAT ON annunciator indicates that the power relays are closed to apply power to the heating elements (Figure 11-24). Before blowers are selected OFF when electric heat is off, the ELEC HEAT ON annunciator must be extinguished to indicate power is removed from the heating elements.

BL AIR OFF L–R Annunciators Green BL AIR OFF L–R annunciators illuminate whenever the respective BLEED AIR VALVES OPEN switch is in any position other than OPEN.

ELEC HEAT Position The supplemental electric heat system is operated on the ground with the ELEC HEAT position on the ENVIRONMENTAL control system knob (see Figure 1120). The system is squat-switch protected from airborne operation.

Figure 11-24. Annunciator Panel


11-19

11 AIR CONDITIONING


11 AIR CONDITIONING


Protection controls prevent compressor operation if the following conditions occur:

OPERATION Automatic Mode Control

• Refrigerant pressure too high or low

The AUTO position of the ENVIRONMENTAL control knob allows the heating and air conditioning systems to operate automatically. The system adjusts bleed air temperature and blower speed and cycles the air conditioning compressor as necessary to maintain the selected temperature. The recommended setting on these knobs is the 12 o’clock (straight u p ) p o s i t i o n , w h i c h i s a p p ro x i m a t e l y 75°F (24°C).

• Right bleed air bypass valve reaches limit switch (indicates air conditioning not required because significant heat introduced into system)

If a different blower speed is desired, the respective COCKPIT or CABIN BLOWER k n o b c a n b e ro t a t e d f ro m t h e AU TO position to desired speed.

Cooling Plumbing from the compressor on the right engine is routed through the right wing and then forward to the condenser coil, receiver-dryer, expansion valve, bypass valve, and forward evaporator. All of these are in the nose of the aircraft. The forward vent blower moves recirculated cabin air through the forward evaporator, into the mixing boxes, into the cockpit distribution ducts, and then out the glareshield outlets and windshield defrost vents. The cabin blowers provide main cabin cooling by routing recirculated cabin air through two evaporators and into the cabin through the eyeball outlets in the cabin and cockpit headliner. When the system is commanded to provide little or no warmed air to the cabin, the majority of the warmer P3 air from the engines is routed to the aft of the cabin into the baggage compartment to avoid interference with the cooling process. The outflow valves quickly evacuate this warmer air overboard.

11-20

• Right engine speed below 62% N 1 ; w h i t e A I R C O N D N 1 LOW annunciator illuminates

Heating Bleed air from the compressor of each engine is deliv ered into the cabin for heating and pressurization purposes. When the left landing gear safety switch is in the ground position, the ambient air valve in each flow control unit is closed. Therefore, only bleed air is delivered. When airborne, bleed air is mixed with outside ambient air from the ambient air valve in each flow control unit until a cold air temperature closes off the ambient flow. Then, only bleed air is delivered. With the environmental control system k n o b i n AU T O, t h e t e m p e r a t u r e o f conditioned air is set to approximately 70°F (21°C). In colder temperature extremes where more heat is initially demanded, this is increased to approximately 105°F (41°C).

Electric Heat When the ELEC HEAT position is selected on the environmental control system knob, air is directed over several heater elements in a duct aft of the forward evaporator and into the cabin. (see Figure 11-20). Air is distributed through the electric heating duct by the cockpit blower that operates a u t o m a t i c a l l y w h e n t h e E L E C H E AT position is selected. The amber ELEC HEAT ON annunciator illuminates to advise the flight crew that power is being applied to the heating elements (see Figure 11-24). The electric heat system draws approximately 160 amps.


Heated air enters the cabin via a single flood outlet directly aft of the cockpit pedestal.

CAUTION The electric heat must not b e o p e ra t e d w i t h t h e c a b i n d o o r closed or the pedestal floor outlet blocked.

Defrost A constant flow of environmentally conditioned air is provided to the windshield defrost and glareshield outlets. In the AUTO mode, the temperature of this air is approximately 70°F (21°C). In extremely cold conditions, this air is allowed to reach 105°F (41°C).

Vent Blower Control During electrical heat operation, the blower operates at maximum speed regardless of the COCKPIT BLOWER knob setting. This electrical heat system is a supplemental heating system for ground operation only. B e f o r e d e s e l e c t i n g t h e E L E C H E AT position, the amber ELEC HEAT ON annunciator must be extinguished. This indicates the heating elements have been sufficiently deenergized for safety. If the a n n u n c i a t o r re m a i n s i l l u m i n a t e d , t h e system is not functioning correctly. To maintain adequate airflow across the h e a t i n g e l e m e n t s, t h e E L E C H E AT p o s i t i o n m u s t b e re s e l e c t e d u n t i l t h e engines have been shut down. Maintenance is required prior to flight. A f t e r d e s e l e c t i n g t h e E L E C H E AT position, safety devices in the heater assembly may continue to temporarily operate the blower at low speed. This allows proper cooling of the heater elements to avoid overheating the duct.

NOTE If after deselecting the ELEC HEAT position and initial blower shutdown, residual element heat causes the duct temperature to c o n t i n u e t o r i s e, t h e b l o w e r automatically cycles to cool the e l e m e n t s re g a rd l e s s o f BAT T switch position.

Unpressurized Ventilation Fresh air is available during unpressurized flight with the CABIN PRESS switch in the DUMP position. This ambient (ram) air is obtained through the fresh air door and the ram-air scoop in the aircraft nose section (see Figure 11-4). There is no fresh air scoop on the new Keith ECS system for unpressurized flight. This door is open only during unpressurized flight when the switch is in the DUMP position and there is 0 psi. This allows the forward blower to draw ram air into the cabin. This air is mixed with recirculated cabin air in the plenum chamber and then directed to both the floor registers and ceiling outlets. The CABIN AIR control knob regulates the air volume. There are no longer CABIN AIR CONTROL KNOBS on the new system.

NOTE A flight conducted with the bleedair switches placed in any position other than OPEN also results in unpressurized flight, but the fresh air door is not open. There is no fresh air door.

Manual Mode Control With the environmental knob in MAN HEAT, control of the cabin and cockpit temperatures is accomplished through the MAN TEMP switch. Moving the switch to either INCR or DECR regulates the bleed air temperature as it enters the cabin; flow rate is unchanged.


11-21

11 AIR CONDITIONING


11 AIR CONDITIONING


Sw i t c h i n p u t o f t w o t o t h re e s e c o n d s duration is recommended with a full 60 s e c o n d s i n b e t w e e n t o a v o i d o v e r- o r undershooting desired temperature. The time it takes the bleed air temperature to respond to switch input is proportional to the time the MAN TEMP switch is actuated (requiring approximately 30 seconds to go f ro m f u l l d e c re a s e t o f u l l i n c re a s e o r vice versa).

CAUTION Switch actuation longer than 2–3 seconds and less than 60 seconds in duration can result in duct overheating and illumination of the amber DUCT OVERTEMP annunciator. The CABIN TEMP or COCKPIT TEMP knobs fully control the temperature of the air to the glareshield and defrost vents when either MAN HEAT or MAN COOL is selected.

LIMITATIONS For specific information on limitations procedures, refer to the FAA-approved Airplane Flight Manual (AFM).

EMERGENCY/ ABNORMAL For specific information on emergency/ a b n o r m a l p r o c e d u r e s, r e f e r t o t h e appropriate abbreviated checklists or the FAA-approved AFM.

11-22


11 AIR CONDITIONING


QUESTIONS 1. The vapor-cycle refrigeration compressor is located: A. O n t h e r i g h t e n g i n e a c c e s s o r y section. B. O n t h e l e f t e n g i n e a c c e s s o r y section. C. In the baggage compartment. D. On the forward pressure bulkhead. 2. If the engine speed is too low for the air conditioning compressor to properly engage, the: A. W h i t e [ A I R C O N D N 1 LOW ] status annunciator illuminates. B. G r e e n [ A I R C O N D N 1 LOW ] advisory annunciator illuminates. C. Engine will automatically increase in speed to allow compressor operation. D. Compressor will engage and the white [AIR COND N1 LOW] status annunciator will illuminate. 3. Fo r m o re e ff i c i e n t c o o l i n g o n t h e ground, place the BLEED AIR VALVES switches to the __________ position. A. OPEN B. CLOSED C. ENVIR OFF D. PNEU & ENVIR OFF 4. In the MAN HEAT mode on the ECS, the pilot controls temperature with the: A. CABIN TEMP knob. B. COCKPIT TEMP knob. C. ENVIR BLEED AIR switch. D. MAN TEMP INCR DECR switch


11-23


CHAPTER 12 PRESSURIZATION CONTENTS INTRODUCTION ............................................................................................................. 12-1 GeNeRal.......................................................................................................................... 12-1 COMPONeNTS ................................................................................................................. 12-2 Flow Control Unit....................................................................................................... 12-2 CONTROlS aND INDICaTIONS ............................................................................... 12-5 Pressurization Controller ........................................................................................... 12-5 Switches ........................................................................................................................ 12-6 Gauges ......................................................................................................................... 12-7 annunciators ............................................................................................................... 12-7 OPeRaTION ..................................................................................................................... 12-8 Preflight Operation ..................................................................................................... 12-8 In-Flight Operation..................................................................................................... 12-8 Descent and landing Operation ............................................................................... 12-8 abnormal Operation .................................................................................................. 12-9 lIMITaTIONS ................................................................................................................... 12-9 Cabin Differential Pressure Gauge........................................................................... 12-9 PReSSURIZaTION PROFIleS ................................................................................. 12-10 MalFUNCTIONS aND TROUBleSHOOTING................................................... 12-15 QUeSTIONS.................................................................................................................... 12-17


12-i

12 PRESSURIZATION

Page


ILLUSTRATIONS Title

Page

12-1

Pressurization Controls ...................................................................................... 12-2

12-2

electronic Flow Control Unit............................................................................ 12-2

12-3

Outflow Valve ...................................................................................................... 12-3

12-4

Safety Valve ......................................................................................................... 12-4

12-5

Pressurization System Schematic ...................................................................... 12-5

12-6

BleeD aIR ValVeS Switches ...................................................................... 12-6

12-7

eNVIR BleeD aIR Switch ............................................................................ 12-6

12-8

CaBIN PReSS Switch ....................................................................................... 12-7

12-9

CaBIN alT Gauge............................................................................................ 12-7

12-10

annunciators ....................................................................................................... 12-7

12-11

Pressurization Controller Setting for landing ................................................ 12-9

12-12

Situation 1 .......................................................................................................... 12-10

12-13

Situation 2 .......................................................................................................... 12-11

12-14

Situation 3 .......................................................................................................... 12-12

12-15

Situation 4 .......................................................................................................... 12-13

12-16

Situation 5 .......................................................................................................... 12-14

TABLE Table 12-1

Title

Page

Descent and landing .......................................................................................... 12-8


12-iii

12 PRESSURIZATION

Figure


12 PRESSURIZATION

CHAPTER 12 PRESSURIZATION

INTRODUCTION This chapter describes the pressurization system on the King air 350 aircraft. The pressurization system provides a normal working pressure differential of 6.5 ±0.1 psi for cabin pressure altitudes of 2,800 feet at 20,000 feet, 8,600 feet at 31,000 feet, and 10,380 feet at 35,0 0 0 feet. The pressurization inflow system also provides fresh air ventilation.

GENERAL Bleed air from each engine pressurizes the cabin and cockpit areas. Pressurization is regulated through a pressurization controller, monitored by a cabin altimeter/psid indicator, and a rate-of-climb

indicator. Pressurization can be dumped with the CaBIN PReSS DUMP switch. The system includes a flow control unit as well as a vacuum line drain and outflow and safety valves (Figure 12-1).


12-1


TEST SQUAT SWITCH

PRESR NO. 1 DUAL-FED BUS

5A PRESS CONTROL

LEFT BLEED AIR VALVE

DUMP

12 PRESSURIZATION

PRESET SOLENOID (N.C.)

SAFETY VALVE OUTFLOW VALVE

PRESSURIZATION CONTROLLER

VAC/ PNEU MANIFOLD DUMP SOLENOID (N.C.)

OUTFLOW VALVE DRAIN

Figure 12-1. Pressurization Controls

COMPONENTS FLOW CONTROL UNIT an electronic flow control unit (FCU) in each engine nacelle controls the volume of the bleed air. It combines ambient air with the bleed air to provide a suitable air d e n s i t y f o r p r e s s u r i z a t i o n . Th e F C U controls the mass flow of both ambient and bleed air into the cabin (Figure 12-2). FIRESEAL

each unit consists of an ambient temperature sensor, an electronic controller, and an environmental air control valve assembly interconnected by a wire harness. The control valve assembly consists of the following: • Mass flow transducer • ambient flow motor and modulating valve • Check valve that prevents bleed air from escaping through ambient air intake

BLEED AIR FLOW TRANSDUCER

COCKPIT BLEED AIR VALVE SWITCH ELECTRONIC CONTROLLER

AMBIENT TEMPERATURE SENSOR

AMBIENT FLOW CONTROL MOTOR

SOLENOID (N.C.) ENVIRONMENTAL SHUTOFF VALVE (N.C.)

AMBIENT FLOW TRANSDUCER HP BLEED AIR (MASS FLOW SENSOR) CONTROL PRESSURE NO. 2 VENT AIR

AIR EJECTOR

CHECK VALVE ENGINE BLEED AIR

BLEED AIR (HIGH FLOW) BYPASS

Figure 12-2. Electronic Flow Control Unit

12-2

SQUAT SWITCH

BLEED AIR FLOW CONTROL MOTOR

AMBIENT AIR INLET

LEGEND

POWER


TO DUCT DISTRIBUTION SYSTEM

• Bleed air flow transducer • Bleed air flow motor and modulating valve (including bypass line) • air ejector • Flow control solenoid valve • environmental shutoff valve When the FCU is energized after engine start up, the bleed air modulating valve closes. When it is fully closed, it actuates the bleed air shaft switch. This signals the electronic controller to open the solenoid valve to enable P3 bleed air to pressurize the environmental shutoff valve open. The bleed air shaft continues to open until the desired bleed-air flow rate to the cabin is reached. The bleed-air flow transducer s e n s e s t h e f l o w r a t e. Th e e l e c t r o n i c controller controls the input of the ambient temperature sensor.

as the aircraft enters a cooler environment, ambient airflow is gradually reduced. Bleedair flow gradually increases to maintain a constant inflow and to provide sufficient heat for the cabin. at approximately 0°C (32°F) ambient temperature, ambient airflow is completely closed off. The bleed air valve bypass section opens as necessary to allow more bleed air flow past the fixed flow passage of the air ejector. The FCUs regulate the rate of airflow to the pressure vessel. The bleed air portion is variable from approximately 5 to 14 pounds per minute (ppm) depending upon ambient temperature. On the ground, since ambient air is not available, cabin inflow is variable and limited by ambient temperature. In flight, ambient air provides the balance of the constant airflow volume of 12 to 14 ppm.

Fro m h e re, t h e a i r f o r p re s s u r i z a tion, cooling, and heating flows into the pressure vessel to create differential, and then out through the outflow valve (Figure 12-3) on the aft pressure bulkhead.

SCHRADER VALVE

MAXIMUM DIFFERENTIAL DIAPHRAGM TO CONTROLLER CONNECTION

PLUG UPPER (CONTROL) DIAPHRAGM NEGATIVE RELIEF DIAPHRAGM

STATIC AIR

LEGEND

REAR PRESSURE BULKHEAD

CONTROL PRESSURE NO. 2 VENT AIR CONTROL PRESSURE NO. 3

Figure 12-3. Outflow Valve FOR TRAINING PURPOSES ONLY

12-3

12 PRESSURIZATION



12 PRESSURIZATION

To the left of the outflow valve (looking forward) is a safety valve (Figure 12-4). This valve provides pressure relief if the outflow va l v e f a i l s, d e p re s s u r i z e s t h e a i rc ra f t whenever the CaBIN PReSS DUMP switch is in DUMP, and keeps the aircraft unpressurized while it is on the ground with the left landing gear safety switch compressed.

When the BleeD aIR ValVeS switches are positioned to OPeN, the air mixture (bleed air and ambient air) from the FCU enters the aircraft. When the aircraft is on the ground, only bleed air enters the cabin because the safety switch causes the FCU to close a valve that allows ambient air to mix with the bleed air.

a negative pressure relief function that prevents outside atmospheric pressure from exceeding cabin pressure by more than 0.1 psi during rapid descents with or without bleed air flow is also incorporated into both valves.

at liftoff, the safety valve closes and, except for cold temperatures, ambient air begins to enter the FCU, and then the pressure vessel. as the left FCU ambient air valve opens, in approximately 6 to 8 seconds, the right FCU ambient air valve opens. By increasing airflow volume gradually (left first, then right), excessive pressure bumps are avoided during takeoff.

SCHRADER VALVE

MAXIMUM DIFFERENTIAL DIAPHRAGM SAFETY VALVE DUMP SOLENOID

CABIN AIR UPPER CONTROL DIAPHRAGM

NEGATIVE RELIEF DIAPHRAGM

STATIC AIR

LEGEND CONTROL PRESSURE NO. 2 CONTROL PRESSURE NO. 3

Figure 12-4. Safety Valve

12-4


REAR PRESSURE BULKHEAD


PRESSURIZATION CONTROLLER an adjustable cabin pressurization c o n t ro l l e r n e a r t h e t h ro tt l e q u a d ra n t modulates the outflow valve (Figure 12-5). a dual-scale dial in the center of the controller indicates the cabin pressure altitude on the outer scale (CaBIN alT) and maximum aircraft altitude on the inner

The RaTe control knob controls the rate at which cabin pressure altitude changes from the current value to the selected value. The selected rate of change can be from approximately 20 0 to 2,0 0 0 feet per minute (fpm).

STATIC

PLUG

LEGEND CABIN AIR VACUUM SOURCE STATIC AIR

OUTFLOW VALVE

CONTROL PRESSURE INTERNAL PRESSURE

350 ONLY ALTITUDE LIMIT CONTROLLER

FLOW CONTROL PRESSURE

MOISTURE ACCUMULATION DRAIN

ORIFICE

CABIN PRESET SOLENOID NO

FILTER

STATIC LG SAFETY SWITCH 350 ONLY

CONTROL SWITCH CABIN PRESSURE RESTRICTOR RATE

ALTITUDE

SAFETY VALVE

ORIFICE DUMP SOLENOID NC

ALTITUDE LIMIT CONTROLLER

VACUUM SOURCE FROM PNEUMATIC MANIFOLD

Figure 12-5. Pressurization System Schematic


12-5

12 PRESSURIZATION

scale (aCFT alT) at which the aircraft can fly without causing the cabin pressure t o e x c e e d m a x i m u m d i ff e re n t i a l . Th e engines maintain a 6.5 ± 0.1 psi differential that provides a nominal cabin pressure altitude of 10,380 feet at an aircraft altitude of 35,0 0 0 feet.

CONTROLS AND INDICATIONS


SWITCHES BLEED AIR VALVES Switches

12 PRESSURIZATION

The leFT–RIGHT BleeD aIR ValVeS switches are in the eNVIRONMeNTal group of the copilot subpanel (Figure 126). When either switch is positioned to eNVIR OFF or PNeU & eNVIR OFF, the environmental air valve is closed.

Figure 12-7. ENVIR BLEED AIR Switch

The aUTO position is the recommended position because it allows the environmental system controller to automatically s e l e c t t h e f l o w s e tt i n g b a s e d o n h e a t required to maintain temperature or cabin pressure requirements.

Figure 12-6. BLEED AIR VALVES Switches

When either switch is in OPeN, the air mixture flows through the environmental air valve toward the cabin.

ENVIR BLEED AIR Switch The eNVIR BleeD aIR switch on the copilot left subpanel controls the volume of bleed air (Figure 12-7). Th e lOW p o s i t i o n r e d u c e s b l e e d a i r extracted from the engines for environmental purposes to approximately half the normal amount. This position is normally used during operations in ambient temperatures above 10ºC to ensure takeoff power is available. For maximum engine performance and/or high altitude takeoff requirements, the switch should be in lOW. The NORMal position increases heating or increases pressurization airflow; it is usually selected during climb to ensure optimum performance of the system at higher altitudes.

12-6

In order for the aUTO position to function properly in response to the heating/cooling requirements, the environmental MODe c o n t r o l m u s t b e i n aU T O ( s e e a i r Conditioning chapter).

CAUTION always monitor cabin pressurization requirements if the environmental MODe switch is in MaN COOl because manual adjustm e n t s m a y b e re q u i re d t o the eNVIR BleeD aIR setting.

CABIN PRESS Switch The CaBIN PReSS switch left of the pressurization controller (Figure 12-8) has DUMP-PReSS-and TeST positions. The DUMP (forward lever locked) position opens the safety valve so the cabin can depressurize to approximately 13,500 feet. a b o v e t h a t a l t i t u d e, i t m a i n t a i n s t h e 13,500 feet.


Figure 12-9. CABIN ALT Gauge

ANNUNCIATORS annunciators for the pressurization system are on the warning panel and the caution/advisory panel (Figure 12-10): • W h i t e C a B I N a lT I T U D e — Illuminates to indicate cabin altitude exceeds 10,0 0 0 feet Figure 12-8. CABIN PRESS Switch

The PReSS (center) position pressurizes the cabin in flight depending on the controller setting. It closes the safety valve so the controller can take command of the outflow valve. The TeST (aft) position holds the safety valve closed, bypassing the landing gear safety switch, to allow cabin pressurization tests on the ground.

• amber l-R Bl aIR OFF— Illuminates to indicates flow control unit closed • Red CaBIN alT HI—Illuminates to indicates cabin pressure altitude exceeds 12,0 0 0 feet • Red CaBIN DIFF HI—Illuminates to indicate cabin differential pressure exceeds 6.9 psi

GAUGES CABIN ALT Gauge The CaBIN alT gauge is on the right side of the control panel with the annunciator panel. It continuously monitors actual cabin pressure altitude (outer scale) and cabin differential (inner scale) (Figure 12-9).

CABIN CLIMB Gauge The CaBIN ClIMB (cabin vertical speed) gauge is left of the CaBIN alT indicator. It continuously monitors the rate at which cabin pressure altitude is changing in feet per minute.

Figure 12-10. Annunciators


12-7

12 PRESSURIZATION



OPERATION

Table 12-1. DESCENT AND LANDING

PREFLIGHT OPERATION

12 PRESSURIZATION

Prior to takeoff, adjust the CaBIN alT selector knob until the aCFT alT (inner) scale on the dial reads an altitude of approximately 1,000 feet above the planned cruise pressure altitude or at least 500 feet above the takeoff field pressure altitude. adjust the RaTe control knob as desired. When the index mark is set at 12 o’clock position, the most comfortable rate of climb is maintained. Position the CaBIN PReSS DUMP switch to PReSS.

IN-FLIGHT OPERATION as the aircraft climbs, the cabin pressure altitude climbs at the selected rate of change until the cabin reaches the selected p r e s s u r e a l t i t u d e. Th e s y s t e m t h e n maintains cabin pressure altitude at the selected value. If the aircraft climbs to an altitude higher than the value indexed on the aCFT alT scale, the cabin-to-ambient pressure differential reaches the pressure relief settings of the outflow valve and the safety valve (6.5 psi cabin-to-ambient differential). The red CaBIN DIFF HI annunciator illuminates at 6.9 ±3 psi cabin-toambient differential pressure. If the flight plan requires an altitude change of 1,0 0 0 feet or more during cruise, the CaBIN alT dial must be readjusted.

DESCENT AND LANDING OPERATION During descent and in preparation for landing, set the CaBIN alT gauge to indicate a cabin altitude of approximately 50 0 feet above the landing field pressure altitude (Table 12-1). adjust the RaTe control knob as required to provide a comfortable cabin altitude rate of descent.

12-8

CLOSEST ADD TO ALTIMETER SETTING AIRPORT ELEVATION 28.00 28.10 28.20 28.30 28.40 28.50 28.60 28.70 28.80 28.90 29.00 29.10 29.20 29.30 29.40 29.50 29.60 29.70 29.80 29.90 30.00 30.10 30.20 30.30 30.40 30.50 30.60 30.70 30.80 30.90

+2,400 +2,300 +2,200 +2,100 +2,000 +1,900 +1,800 +1,700 +1,600 +1,500 +1,400 +1,300 +1,200 +1,100 +1,000 +900 +800 +700 +600 +500 +400 +300 +200 +100 –100 –200 –300 –400 –500

The aircraft rate of descent is controlled so the aircraft altitude does not catch up with the cabin pressure altitude until the cabin pressure altitude reaches the selected value and stabilizes. as the aircraft descends to and reaches the cabin pressure altitude, the outflow valve remains open. This keeps the vessel depressurized. as the aircraft continues to descend below the preselected cabin pressure altitude, the cabin remains depressurized and follows the aircraft rate of descent to touchdown.


0

ABNORMAL OPERATION If cabin pressure altitude reaches a value of 10,000 feet, the white CaBIN alTITUDe annunciator illuminates. an aural warning also sounds. Depress the CaBIN alT WaRN SIleNCe button on the copilot sub-panel to cancel the warning tone. If cabin pressure altitude reaches a value of 12,0 0 0 feet, the red CaBIN alT HI annunciator illuminates. In addition, the MaSTeR WaRNING flashes illuminate and an aural warning sounds. at 12,50 0 feet, the oxygen masks drop. The aural warning can be cancelled, but the CaBIN alTITUDe and CaBIN alT HI annunciator remain illuminated as long as the cabin pressure altitude remains above their respective actuation altitudes.

the landing field pressure altitude (Figure 12-11), and the rate control selector should be adjusted as required to provide a comfortable rate of descent for the cabin. The airplane rate of descent should be controlled so that the airplane altitude does not catch up with the cabin pressure altitude until the cabin pressure altitude reaches the selected value and stabilizes. as the airplane descends to and reaches the c a b i n p re s s u re a l t i t u d e, t h e n e g a tivepressure relief function modulates the outflow and safety valves toward the full open position, thereby equalizing the difference between ambient and cabin pressures. When the airplane continues to descend b e l o w t h e p re s e l e c t e d c a b i n p re s s u re altitude, the cabin will be unpressurized and will follow the airplane rate of descent to touchdown.

LIMITATIONS CABIN DIFFERENTIAL PRESSURE GAUGE The cabin differential pressure gauge has the following limitation markings: • Green arc (normal operating range— 0 to 6.6 psi • Re d a r c ( u n a p p r o v e d o p e r a t i n g range)—6.6 psi to end of scale Maximum cabin pressure differential is 6.6 psi.

Descent During enroute descents and in preparation for landing, the CaBIN alT selector should be set as appropriate for the lower altitude. For enroute descents, set the aCFT alT to 500–1000 feet above the level-off flight altitude unless it results in a CaBIN alT less than destination field pressure altitude. On normal descents, the CaBIN alT should be set to indicate a cabin altitude of approximately 500 feet above

Figure 12-11. Pressurization Controller Setting for Landing


12-9

12 PRESSURIZATION



PRESSURIZATION PROFILES The following pressurization situations are described in order to illustrate operation of the system using normal flight situations. In each case, the given conditions will be outlined on the profile diagram. 12 PRESSURIZATION

Situation 1 Climb from sea level to Fl310, then descend to a field pressure altitude of 1500 feet (Figure 12-12).

Conditions: • aircraft climbs at 20 0 0’/min to Fl20 0, then 10 0 0’/min to Fl310 • Cabin climbs at 50 0’/min • aircraft descends at 150 0’/min • Cabin descends at 50 0’/min Controller setup before takeoff—Prior to takeoff, set the inner dial (aCFT alT) on the pressurization controller to Fl315 (500 feet above cruise altitude) which will provide a 90 0 0-foot altitude on the outer dial (CaBIN alT). Set the rate knob between the twelve and one o’clock positions.

Figure 12-12. Situation 1

12-10



C o n t ro l l e r s e t u p f o r d e s c e n t — a s t h e aircraft starts to descend, set the outer dial ( Ca B I N a lT ) t o 2 0 0 0 f e e t p r e s s u r e altitude (50 0 feet above field pressure altitude). The rate knob should stay in the 12 to 1 o’clock position. Operation—as the aircraft descends to the field, which will take approximately 19 minutes, the cabin descends to 2000 feet, which takes approximately 14 minutes. as the aircraft passes 20 0 0 feet, the cabin descends unpressurized to 1500 feet with the aircraft.

Situation 2 Climb from sea level to Fl310, then descend to a field pressure altitude of 1500 feet (Figure 12-13).

Conditions: • aircraft climbs at 20 0 0’/min to Fl20 0, then 10 0 0’/min to Fl310 • Cabin climbs at 50 0’/min • aircraft descends at 150 0’/min • Cabin descends at 50 0’/min Controller setup—Same as Situation 1 except set aCFT alT dial to Fl310 (same as cruise altitude) which will put cabin at max. differential when aircraft gets to final a l t i t u d e. S e t i t a s i n S i t u a t i o n # 1 f o r the descent. Operation—everything is normal until the cabin gets to max. differential which then causes pressure bumps in the cabin. The condition normalizes on descent.

Remarks—all settings are normal and the system reacts properly.



12-11

12 PRESSURIZATION

O p e r a t i o n — a s t h e a i rc ra f t c l i m b s t o Fl310, which will take approximately 21 minutes, the cabin climbs to 9000 feet in approximately 18 minutes, thus the cabin always stays “ahead” of the aircraft during t h e c l i m b. ( S t a y i n g a h e a d m e a n s t h a t maximum differential is never achieved since the cabin climb rate is sufficient to make the final altitude on its own.)


Remarks—By not setting the pressurization controller properly, cabin pressure bumps are likely with resultant discomfort for passengers. The aCFT alT dial should be set to at least 50 0 feet above aircraft cruise altitude.

Situation 3 12 PRESSURIZATION

Climb from sea level to Fl310, then descend to a field pressure altitude of 1500 feet (Figure 12-14).

Conditions: • aircraft climbs at 20 0 0’/min to Fl200, then 1000’/min to Fl310 • Cabin climbs at 500’/min • aircraft descends at 1500’/min • Cabin descends at 500’/min

Controller setup—Same situation as 1 except set CaBIN alT dial to landing field p re s s u re a l t i t u d e p r i o r t o t a ke off. No readjustment required for descent. Operation—everything is normal until max. differential is reached as the aircraft passes approximately Fl190. as the aircraft continues to climb, the cabin climbs at a rate proportional to but less than the aircraft’s rate since it remains on max. differential. The cabin will be subject to pressure bumps while on max. differential. Once the aircraft starts to descend, the condition normalizes. Remarks—By setting the controller for landing prior to takeoff, a similar problem to that in Situation 2 occurred with resultant passenger discomfort.


12-12



The aircraft was held to 5000 feet for 15 minutes during climb, then cleared to Fl310. The aircraft was given a segmented descent to Fl250 before given final descent for landing (Figure 12-15).

Conditions: • aircraft climbs at 2000’/min to 5000’, l e v e l s o f f f o r 15 m i n u t e s t h e n continues to Fl310. • Cabin climbs at 50 0’/min to 450 0’, levels off until reset, then climbs at 500’/min to 9000’. • aircraft descends at 1500’/min with a 10 min level off at Fl250 before continuing down for landing. • Cabin descends at 50 0’/min with a brief level off at 590 0’ until being reset for landing, then descends at 500’/min.

C o n t ro l l e r s e t u p b e f o re t a ke o f f — S e t CaBIN alT dial to 500 feet below (4500’) the aircraft intermediate level off altitude (5000 feet). This prevents the cabin altitude from catching up to the aircraft altitude during climb. When finally cleared to FL310—Set the aCFT alT dial to 50 0 feet above the assigned flight level as in Situation 1. When cleared down to FL250—Set the aCFT alT dial to 500 feet above the newly assigned flight level. When cleared for final descent—Set the CaBIN alT dial to 50 0 feet above field pressure altitude. Operation—as the aircraft climbs to 5000 feet, the cabin climbs to 450 0 feet, thereby maintaining a slight pressurization differential. When the aircraft climbs to Fl310, the controller is reset as in Situation 1 and the cabin will climb accordingly. When the



12-13

12 PRESSURIZATION

Situation 4


aircraft starts to descend to Fl250, the cabin starts down and levels at approximately 590 0 feet. This allows the cabin pressurization to stay on the controller and not go to max. differential which avoids bumps as in Situation 2. Once commencing final descent, everything will proceed as normal in Situation 1. 12 PRESSURIZATION

Remarks—an alternate method in this situation is to set the rate knob to minimum until cleared to the final altitude or for final descent, at which time the rate knob is reset for normal situations as in Situation 1. One problem with this alternate method is that it disrupts the normal balance between the outflow and safety valves on takeoff and will result in a pressure bump shortly after takeoff. In addition, it is possible, although unlikely, that the cabin pressurization could still catch up to the aircraft. Using either method in this situation, forgetting to reset the controller will result in problems similar to those discussed in Situations 2 and 3.

Situation 5 Depart from an airport at a higher elevation (in this case 6000 feet), fly at an altitude of 16,0 0 0 feet, and land at a sea level airport 20 minutes later (Figure 12-16).

Conditions: • aircraft climbs at 2000’/min to 16,000’ • aircraft levels off for approx 5 min • Cabin climbs and descends at 50 0’/min Controller setting before takeoff—Set the CaBIN alT dial to 500 feet above takeoff field elevation (60 0 0 feet). Operation—Once the aircraft is stable in the climb and cabin altitude stabilizes at 650 0 feet, reset controller for landing, usually within a few minutes after takeoff. Controller setting in flight—Set the CaBIN alT dial to 50 0 feet above landing field pressure altitude.


12-14



Remarks—It is important to set the cabin p re s s u r i z a t i o n c o n t ro l l e r f o r a c a b i n a l t i t u d e a b o v e t a ke o ff f i e l d p re s s u re altitude. If it is set lower, a pressure bump will be experienced shortly after liftoff since the balance between the outflow and safety valves will be disrupted. Once cabin pressurization is stabilized after takeoff, the controller may be reset for landing, provided cruise altitude does not exceed the altitude in the aCFT alT window. If it d o e s, n o r m a l c o n t r o l l e r p r o c e d u r e s described earlier apply.

MALFUNCTIONS AND TROUBLESHOOTING The pressurization system in the Super King air 300 and 350 is derived from the proven and highly reliable system used in other King airs. It is well engineered for safety, comfort, reliability and ease of operation. Pilot controls are simple and straight-forward and workload is minimal. The pilot has sufficient controls readily available to either regain control or minimize the effect of most problems. With loss of pressurization in flight, follow the procedures outlined in the emergency Procedures section of the POH. Once the situation in the aircraft has stabilized, and if extreme care and good judgement are utilized, other corrective action may be taken using the techniques and procedures discussed in this section as a guide. For crew and passenger safety, pressurization troubleshooting should be accom plished below an aircraft altitude of 10,000 feet MSl whenever possible. In addition,

only minimal troubleshooting should be attempted with passengers onboard. It should also be noted that as the cabin climbs past 12,500 feet pressure altitude, the passenger oxygen masks should deploy. If they do not, the passenger manual dropout handle should be pulled. See the Oxygen section of this workbook for more details about the oxygen system. Most pressurization malfunctions will show up shortly after takeoff. They will show three general symptoms: rapid pressurization toward maximum differential, lack of pressurization (i.e., the cabin climbs at the same rate as the aircraft), and cabin leakdown (i.e., the cabin leaks pressurization slowly—50 0 feet/minute at low-pressure differentials and faster at high-pressure differentials.) The first two symptoms are generally caused by controller, control system, or outflow/safety valve malfunctions. The third is normally caused by air inflow problems. air inflow problems could be caused by a malfunction within the flow control units. Since it is unusual for both flow control units to fail simultaneously, one probably failed earlier and went undetected. a second cause of air inflow problems on takeoff can be the BleeD aIR ValVe switches. They could have been left in the eNVIR OFF or PNeU & eNVIR OFF. If they are OPeN, an electrical failure to the switches would cause these normally closed valves to close (Figure 12 7). In this case, the pilot should check the bleed air control circuit breakers on the copilot’s circuitbreaker panel. If cabin altitude descends rapidly shortly after takeoff, it is caused by closed outflow and safety valves (Figures 12 8 and 12 9). The safety valve normally closes shortly after takeoff, but the outflow valve should modulate open as directed by controller pressure. The problem lies somewhere in the plumbing and/or components in the outflow system.


12-15

12 PRESSURIZATION

Operation—The aircraft will continue to climb to altitude while the cabin starts to descend to 50 0 feet pressure altitude. aircraft levels at cruise, then descends for landing. By the time the aircraft is ready for landing, the cabin altitude is level at 50 0 feet (this assumes aircraft altitude vs. cabin altitude does not exceed 6.6 psid).


Possible sources of this problem include: a stuck preset solenoid, a cracked pressurization controller, a diaphragm failure in the pressurization controller, an open moisture drain, disconnected or leaking plumbing, a cracked outflow valve, or a failed diaphragm in the outflow valve. 12 PRESSURIZATION

In this situation, the pilot should turn off both bleed air valves, which will stop the P3 bleed air inflow and depressurize the cabin at its leak rate. Once the cabin is stabilized, cycling the CaBIN PReSS switch to TeST may free a stuck preset solenoid. The moisture drain can be checked through its access panel on the lower right sidewall of the baggage compartment. any additional troubleshooting should be accomplished on the ground. Failure of the cabin to pressurize shortly after takeoff is indicative of inflow and/or outflow malfunctions. If the cabin altitude climbs with the airplane, it indicates an outflow problem caused by an open outflow and/or safety valve. The outflow valve may have prematurely opened due to the preset solenoid opening on the ground. In this situation, the controller prematurely thought the aircraft took off. Controller pressure will hold the outflow valve wide open until the aircraft “catches up” to the controller pressure. This may not occur until the aircraft climbs through the altitude in the CaBIN alT window on the pressurization controller. If this condition is suspected, the pilot can turn the rate knob full counterclockwise to the minimum position, which will minimize any additional change in controller pressure. an alternative action would be to select a lower cabin altitude with the selector; however, it is important to remember to reselect the proper setting once the cabin pressurizes normally.

12-16

If the safety valve remained open after takeoff, cabin altitude would climb together with the aircraft. This could be caused by: the CaBIN PReSS switch in DUMP, failure of the left main squat switch, failure of the dump solenoid, or a stuck safety valve (Figure 12-9). The pilot should first check to ensure the CaBIN PReSS switch is in PReSS (Figure 12-15). If it is already in PReSS, move the switch to TeST, which will override the squat switch which may have malfunctioned. If the cabin begins to pressurize, hold the switch in TeST until the cabin differential pressure exceeds 0.5 psid, then pull the PReSS CONTROl circuit breaker on the copilot’s circuit breaker panel. Normal pressurization control will be resumed, but the electrical dump functions will not be available. Wi t h t h e P R e S S C O N T R O l c i r c u i t breaker pulled out, the electrical functions related to cabin dump are deenergized. Therefore, pressurization dump with the CaBIN PReSS switch or upon landing touchdown will be disabled until the circuit breaker is reset. If the cabin does not pressurize with the CaBIN PReSS switch in TeST, then the problem is beyond the capability of inflight troubleshooting and must be repaired when the aircraft lands. Symptoms such as unusual or excessive pressure bumps on takeoff or any time the controller is reset indicate sticking outflow/safety valves, possibly due to buildup of tars and nicotine. The filters associated with this system may also be dirty. a symptom of dirty filters is a large difference (over 20 0 feet/minute) in the cabin climb vs. the cabin descent with the rate knob in the same position. These valves and filters should be checked at regular maintenance inspections; however, inspection intervals of these components may be shortened when unusual conditions (heavy smoking, dusty atmosphere) exist.



1. The CaBIN alT gauge indicates cabin _______ and cabin _______ altitude. a. Differential pressure; pressure B. Differential pressure; density C. Rate of climb; pressure D. Rate of climb; density 2. The cabin _______ pressurize on the ground by selecting the _______ position of the CaBIN PReSS switch. a. Will; DUMP B. Will; TeST C. Will not; DUMP D. Will not; RelIeF 3. Th e w h i t e [ C a B I N a lT I T U D e ] status annunciator illuminates when c a b i n p re s s u re a l t i t u d e i n d i c a t e s _______ feet. a. 10,000 B. 10,500 C. 12,000 D. 12,500

4. The red [CaBIN alT HI] warning annunciator illuminates when cabin pressure altitude exceeds _______ feet. a. 10,000 B. 10,500 C. 12,000 D. 12,500 5. Th e f i r s t i m m e d i a t e i t e m f o r PReSSURIZaTION lOSS is: a. Descend ...............aS ReQUIReD B. Bleed air Valves.......eNVIR OFF C. Mic Switch(es) .........................OXY D. Oxygen Mask(s) ......................DON 6.

The first immediate action item for the eMeRGeNCY DeSCeNT is: a. Power levers ..........................FUll FORWaRD B. Prop levers ....FUll FORWaRD C. Power levers ..........................IDle D. Prop levers ...................lOW RPM


12-17

12 PRESSURIZATION

QUESTIONS

13 HYDRAULIC POWER SYSTEM


See Chapter 14, “Landing Gear and Brakes,” for information on the hydraulic power systems.


13-i


CHAPTER 14 LANDING GEAR AND BRAKES CONTENTS INTRODUCTION............................................................................................................. 14-1 GENERAL ......................................................................................................................... 14-1 LANDING GEAR DESCRIPTION .............................................................................. 14-2 Landing Gear Assemblies .......................................................................................... 14-2 Wheel Well Doors....................................................................................................... 14-4 Hydraulic Pack ............................................................................................................ 14-4 Controls and Indicators.............................................................................................. 14-7 Warning System......................................................................................................... 14-12 EXTENSION ................................................................................................................... 14-12 RETRACTION ................................................................................................................ 14-16 MANUAL OPERATION .............................................................................................. 14-19 Extension ................................................................................................................... 14-19

NOSEWHEEL STEERING .......................................................................................... 14-22 BRAKE SYSTEM ........................................................................................................... 14-23 Normal Use of Brakes.............................................................................................. 14-23 Parking Brake............................................................................................................ 14-25 LIMITATIONS................................................................................................................. 14-26 Gear Operating Limits............................................................................................. 14-26 SERVICING..................................................................................................................... 14-26 Shock Struts............................................................................................................... 14-26 Brake Service ............................................................................................................ 14-26


14-i

14 LANDING GEAR AND BRAKES

Retraction .................................................................................................................. 14-21


Hydraulic Service...................................................................................................... 14-27 Brake Wear Limits.................................................................................................... 14-27 QUESTIONS.................................................................................................................... 14-29


14-ii



ILLUSTRATIONS Title

Page

14-1

Main Landing Gear Assembly........................................................................... 14-2

14-2

Nose Landing Gear Assembly ........................................................................... 14-2

14-3

Bulkhead for ER Model..................................................................................... 14-3

14-4

Wheel Well Mechanism ....................................................................................... 14-4

14-5

Hydraulic Pack..................................................................................................... 14-4

14-6

Hydraulic Landing Gear Plumbing Schematic................................................ 14-5

14-7

Hydraulic Fluid Low Caution Annunciator..................................................... 14-6

14-8

SENSOR TEST Button...................................................................................... 14-6

14-9

Safety Switches .................................................................................................... 14-7

14-10

Landing Gear Controls....................................................................................... 14-7

14-11

Gear Position Indicator ...................................................................................... 14-8

14-12

Landing Gear Position Indication—Gear Extended ...................................... 14-9

14-13

Landing Gear Position Indication—Gear in Transit..................................... 14-10

14-14

Landing Gear Position Indication—Gear Up ............................................... 14-11

14-15

Silence Buttons ................................................................................................. 14-12

14-16

Hydraulic Landing Gear Schematic ............................................................... 14-13

14-17

Landing Gear Extension Schematic ............................................................... 14-14

14-18

Landing Gear Extended Schematic................................................................ 14-15

14-19

Landing Gear Retraction Schematic .............................................................. 14-17

14-20

Landing Gear Retracted Schematic................................................................ 14-18

14-21

Landing Gear Alternate Extension Placard .................................................. 14-19

14-22

Landing Gear Circuit Breaker ........................................................................ 14-19

14-23

Hand Pump Extension ..................................................................................... 14-20


14-iii


Figure


14-24

Maintenance Hand Pump Retraction............................................................. 14-21

14-25

Nosewheel Steering Mechanism ..................................................................... 14-22

14-26

Nosewheel Limits.............................................................................................. 14-23

14-27

Brake System Schematic .................................................................................. 14-24

14-28

Parking Brake .................................................................................................... 14-25

14-29

Brake Fluid Reservoir ...................................................................................... 14-26

14-30

Hydraulic Fluid Reservoir ............................................................................... 14-27

14-31

Brake Wear Diagram ........................................................................................ 14-27

TABLES Table

Title

Page

14-1

Landing Gear Warning Horn Operation........................................................ 14-12

14-2

King Air 350 Airspeeds .................................................................................... 14-26


14-iv



CHAPTER 14 LANDING GEAR AND BRAKES

This chapter describes the landing gear and brake systems. A thorough understanding of these systems enables the crew to operate the brakes safely with minimum wear and handle any landing gear abnormal situations that may arise. Operating tips, inspection points, and servicing procedures are also included. Operation of the hydraulic system and nosewheel steering are also part of this chapter.

GENERAL The King Air 350 has a retractable tricycle landing gear system that includes an emergency manual extension pump. A hydraulic power pack powers the extension and retraction cycle.

The brake system with four hydraulically operated brake assemblies includes a brake deice system (refer to Chapter 10).


14-1


INTRODUCTION


LANDING GEAR DESCRIPTION The landing gear system is a retractable, electrically powered, hydraulically actuated, tricycle gear system. When the gear is fully retracted, it is completely enclosed by the gear door assemblies. An alternate means of extension is a manuallyoperated hand pump.

LANDING GEAR ASSEMBLIES The main and nose landing gear assembly consists of a drag leg, shock strut, torque knee (scissors), actuator, wheels, and tires. In addition, the main gear assemblies house the brake assemblies (Figure 14-1). Figure 14-2. Nose Landing Gear Assembly

Shock Strut The air/oil shock struts are filled with both compressed air and hydraulic fluid. The air charge in the shock struts carries the aircraft weight.


At touchdown, the lower portion of each strut is forced into the upper cylinder. This action moves fluid through an orifice to further compress the air charge and absorbing landing shock. Orifice action also reduces bounce during landing.

Figure 14-1. Main Landing Gear Assembly

The nose gear includes the shimmy damper (Figure 14-2).

Drag Leg The upper end of the drag legs and two points on the shock strut assemblies attach to the aircraft structure. When the gear is extended, the drag braces become rigid.

14-2

At takeoff, the lower portion of the strut extends until an internal stop engages on the main gear. On the nose gear, the lower portion of the strut extends until the torque knee prevents further extension.

Torque Knee A torque knee connects the upper and lower portion of the shock strut. It allows strut compression and extension but resists rotational forces. This keeps the wheels aligned with the aircraft longitudinal axis.



On the nose gear assembly, the torque knee a l s o t ra n s m i t s s t e e r i n g m o t i o n t o t h e nosewheel and nosewheel shimmy motion to the shimmy damper.

Actuator One hydraulic actuator is on each landing gear. The actuators extend and retract the landing gear.

The nose landing gear wheel is equipped with a 22 x 6.75 x 10, 8-ply-rated tubeless tire. Each main landing gear wheel is equipped with a 19.5 x 6.75 x 8, 10-ply-rated, tubeless tire. On the King Air 350ER model, the main gear tires are 22 x 6.75 x 10, 10-ply rated tubeless. Check the Pilot’s Operating Handbook for correct tire pressure.

Shimmy Damper Wheels and Tires Each main landing gear and nose gear has a wheel. The main wheels are two forged aluminum 6.50 x 8 wheels. The nose gear has an aluminum 6.50 x 10 wheel.

Bulkhead On the King Air 350ER, a bulkhead has been added to the wheel well to prevent i c i n g o n t h e d ra g b ra c e s t r u t re l e a s e (Figure 14-3).


Each wheel consists of an inner and outer wheel half held together with bolts and nuts. The wheels are sealed against air leakage. The wheel rotates on tapered roller bearings.

The shimmy damper mounted on the right side of the nose gear strut is a balanced hydraulic cylinder. It bleeds fluid through an orifice to dampen the nosewheel shimmy.

Figure 14-3. Bulkhead for ER Model


14-3


WHEEL WELL DOORS

HYDRAULIC PACK

The nose gear door and two sets of main gear doors are mechanically actuated by gear movement during extension and retraction.

A hydraulic power pack (Figure 14-5) in c o n j u n c t i o n w i t h h y d ra u l i c a c t u a t o r s extend and retract the nose and main landing gear assemblies. Each landing gear has one hydraulic actuator. The pack is in the middle of the left wing center section just forward of the main spar.

The doors are hinged at the sides and spring-loaded to the open position. As the landing gear is retracted, a roller on each side of the nose gear engages a cam assembly to draw the doors closed behind the gear (Figure 14-4). For extension, a reverse action occurs as the spring-loading takes effect. When the cam has left the roller, springs pull the linkage over-center to hold the doors open.

Figure 14-5. Hydraulic Pack

14 LANDING GEAR AND BRAKES CAN ASSEMBLY

ROLLERS

Figure 14-4. Wheel Well Mechanism

14-4



The power pack consists of a hydraulic pump with 28VDC motor, a two-section fluid reservoir, filter screens, gear selector valve, up selector solenoid, down selector solenoid, fluid level sensor, and an uplock pressure switch. The reservoir has a dipstick to provide a visual check of fluid level. Three hydraulic lines are routed to the nose and main gear actuators (Figure 14-6). One line is for normal extension and one is for r e t r a c t i o n . Th e s e o r i g i n a t e f r o m t h e power pack.

The power pack pump generates hydraulic fluid under pressure in the accumulator that acts on the piston faces of the actuators attached to the folding drag braces. Power for the pump motor is through the landing gear motor relay and a 60 ampere relay circuit breaker. Both are next to the pump motor in the middle of the left wing center section, just forward of the main spar. The 2-ampere circuit breaker for the landing gear control circuit (see next section) energizes the motor relay.


The third line is for manual extension. It originates from the hand pump.

The normal extension lines and manual extension lines connect to the upper end of each hydraulic actuator. The hydraulic lines for retraction are fitted to the lower ends of the actuators.

Figure 14-6. Hydraulic Landing Gear Plumbing Schematic


14-5


If the pump motor fails to shut off after 13 to 15 seconds, a timer activates the logic relay to open the pump motor relay. This stops the motor, but it also shorts out the 2-ampere control circuit and trips the LANDING GEAR RELAY circuit breaker on the pilot right subpanel.

Hydraulic Fluid Level Indication System A y e l l o w c a u t i o n H Y D F L U I D LOW annunciator in the caution/advisory panel illuminates whenever hydraulic fluid level in the landing gear power pack reservoir is low (Figure 14-7). A fiber-optic sensing unit mounted on the motor end of the power pack provides the necessary switching circuitry to illuminate the low fluid light. Te s t t h e a n n u n c i a t o r a n d a s s o c i a t e d circuitry by pressing the HYD FLUID SENSOR TEST button on the pilot right subpanel (Figure 14-8). The test button sends a signal to the fiber-optic sensor in the reservoir that causes the sensor to sense a low fluid level.

Figure 14-8. SENSOR TEST Button

The sensor then sends a low fluid signal to the annunciator panel. A five-second delay can be expected before the annunciator light illuminates during the test. Release the button to extinguish the annunciator. Expect another five-second delay.


Figure 14-7. Hydraulic Fluid Low Caution Annunciator

14-6



CONTROLS AND INDICATORS Safety Switches Safety switches (Figure 14-9) called “squat” switches are on the main gear torque knees. These switches open control circuits when the oleo strut is compressed to prevent gear retraction on the ground.

The handle must be pulled out of a detent before it can be moved from either the UP or DN position. When the LDG GEAR is moved to the UP position in flight, a control circuit completes to the gear up selector solenoid. This moves the gear selector valve so fluid

When the left squat switch senses the aircraft is on the ground, it opens the control circuit to the landing gear selector valve up solenoid to help prevent inadvertent retraction. When the right squat switch senses the aircraft is on the ground, it opens the control circuit to the landing gear selector valve up solenoid, the landing gear motor control relay, and the landing gear handle lock solenoid.

LDG GEAR CONTROL Handle Figure 14-10. Landing Gear Controls

4

1. Actuator Rod 2. Retaining Nut 3. Switch Arm 4. Locking Screw 5. Adjusting Screw LEFT MAIN

2 1 3

Landing Gear Selector Valve Up Solenoid Ambient Air Modulating Valves Preset Solenoid Dump Solenoid Door Seal Solenoid Stall Warning Heat Control RIGHT MAIN Landing Gear Selector Valve Up Solenoid Landing Gear Motor Control Relay Landing Gear Handle Lock Solenoid Ground Low Pitch Stop System Electric Heat Control Flight Hour Meter

5

Figure 14-9. Safety Switches


14-7


The LDG GEAR CONTROL handle on t h e p i l o t r i g h t s u b p l a n e l c o n t ro l s t h e hydraulic power pack motor (Figure 14-10).


flows to the retraction side of the system. A control circuit is also completed to the pump motor relay to signal the pump motor (or hydraulic power pack) to operate. When the LDG GEAR control is moved to the DN position in flight, a control circuit completes to the gear down selector. The gear selector valve moves so fluid flows to the extension side. A control circuit also completes to signal the pump motor to operate. A 2 ampere circuit breaker on the pilot right subpanel protects the landing gear control circuit.

Control Lock The handle lock prevents the LDG GEAR CONTROL handle from being placed in the UP position when the aircraft is on the ground. It automatically unlocks when the aircraft leaves the ground because the right squat switch closes and completes a circuit through the solenoid that moves the lock. If a malfunction occurs in the solenoid or squat switch circuit, the DOWN LOCK REL button overrides the lock.

Position Indicators 14 LANDING GEAR AND BRAKES

An assembly of three lights in a single unit n e x t t o t h e L D G G E A R C O N T RO L indicate landing gear position (Figure 14-11). One light in each main segment (L and R) and two in the nose segment (NOSE) illuminate green to indicate gear is down and locked. Absence of illumination indicates gear is not down and locked.

Figure 14-11. Gear Position Indicator

Gear DOWN is indicated when all three green gear position indicators illuminate and the red lights in the handle extinguish. The red lights in the handle also illuminate when the landing gear warning horn is actuated. The landing gear in-transit light can indicate one or all of the following conditions: • La n d i n g g e a r h a n d l e u p a n d t h e aircraft on ground with weight on gear; warning horn also sounds • One or both power levers retarded below approximately 86 ±1% N 1 and one or more gear not down and locked; warning horn also sounds • Any one or all gear not in fully retracted or in down-and-locked position. • Warning horn has been silenced and does not operate. Figures 14-12 through 14-14 illustrate the various configurations.

In-Transit Light Two red indicator lights in the LDG GEAR CONTROL handle illuminate to indicate gear is in transit, unsafe, or unlocked. Gear up is indicated when the red lights go out and the gear is obviously retracted.

14-8




Figure 14-12. Landing Gear Position Indication—Gear Extended


14-9



Figure 14-13. Landing Gear Position Indication—Gear in Transit

14-10




Figure 14-14. Landing Gear Position Indication—Gear Up


14-11


With the FLAPS beyond the APPROACH position, the warning horn sounds and intransit lights illuminate regardless of power settings. The horn cannot be silenced.

Indicator Lights Check Check the green position indicator bulbs by depressing on the light housing. Check the red control handle lights by pressing HDL LT TEST button adjacent to the LDG GEAR CONTROL handle (see Figure 14-6). They may also be tested in conjunction with the gear warning horn with the STALL WARN TEST-OFF LDG GEAR WARN TEST switch on the copilot left subpanel.

Landing gear warning horn operational data is shown in Table 14-1.

Silencing Warning Horn

WARNING SYSTEM

Pressing the GEAR WARN SILENCE button adjacent to the LDG GEAR CONTROL handle (and/or the silence button on the left power lever) silences the horn (Figure 14-15).

The landing gear warning system warns the pilot of an unsafe condition within the landing gear system.

The warning horn rearms if power levers are sufficiently advanced.

If the LDG GEAR CONTROL handle is not in the DN position when the aircraft is on the ground, the landing gear warning horn sounds and the red gear-in-transit lights illuminate. Place the LDG GEAR handle DN to silence the horn and extinguish the handle light. The landing gear does not retract on the ground because the squat switches prevent activation of the gear selector valve and power pack motor.

Figure 14-15. Silence Buttons

EXTENSION


In flight, warning modes depend upon the position of the flaps. With the flaps in the UP or APPROACH position and either or both power levers retarded below approximately 86% N1, the warning horn sounds and the LDG GEAR CONTROL handle lights illuminate. The horn can be silenced.

The landing gear is electrically controlled and hydraulically actuated. Folding (drag) braces lock in place when gear is fully extended. An electrically actuated selector valve controls the flow of hydraulic fluid to the

Table 14-1. LANDING GEAR WARNING HORN OPERATION GEAR POSITION

14-12

FLAPS

POWER

HORN

SILENCE MODE

Not down

Up

Above 86% N1

No

N/A

Not down

Up

Below 86% N1

Yes

Silence button

Not down

Approach

Below 86% N1

Yes

Silence button

Not down

Beyond approach

Any

Yes

Lower gear



This action moves the gear selector valve so that fluid can flow to the extension side of the system. After approximately six seconds, the extension cycle is complete.

gear in the down position. In this position, the internal downlock mechanism in the nose gear actuator positions the actuator downlock switch to interrupt current to the nose gear part of the pump motor control circuit. A notched J-hook, lock link, and lock link guide attachments fitted to e a c h m a i n g e a r d ra g b ra c e p ro v i d e a positive downlock action for the main gear. A downlock switch on each J-hook assembly interrupts its part of the pump motor control circuit when the respective main gear is down and locked. The motor continues to run until all three landing gear are down and locked.

When the actuator pistons are positioned to fully extend the landing gear, an internal mechanical lock in the nose gear actuator and the nose gear drag brace lock the nose

Fi g u r e 14 - 18 i s a f t e r a n o r m a l g e a r extension. Electrical power flows from the 2-amp LDG GEAR CONTROL circuit breaker through the control switch and on

individual gear actuators (Figure 14-16). The selector valve receives electrical power through the LDG GEAR CONTROL handle.


Wh e n t h e L D G G E A R C O N T RO L i s moved to the DN position in flight, a c o n t ro l c i rc u i t c o m p l e t e s t o t h e g e a r selector valve down solenoid and energizes the pump motor. The top portion of Figure 14-17 depicts this electrical control circuit.

Figure 14-16. Hydraulic Landing Gear Schematic


14-13



Figure 14-17. Landing Gear Extension Schematic

14-14




Figure 14-18. Landing Gear Extended Schematic


14-15


to the three downlock switches. Each gear is down and locked so these three switches a re o p e n . N o e l e c t r i c a l p o w e r p a s s e s through them. However, power is still provided to the hydraulic selector valve to hold it in the down position.

RETRACTION When the LDG GEAR control is moved to the UP position in flight, a control circuit completes to the gear selector valve up solenoid. This moves the gear selector valve so fluid can flow to the retraction side of the system. A control circuit also energizes the pump motor.

The pressure switch then closes periodically as pressure drops to approximately 2,375 psi (normal system pressure leakdown) to reenergize the pump and restore n e e d e d u p l o c k p r e s s u r e. P r e s s u r e i s maintained between approximately 2,375 to 2,775 psi to keep gear retracted. Figure 14-20 depicts the system after retraction with pressure being maintained. An accumulator in the left wing inboard of the nacelle is precharged to 80 0 psi to aid in maintaining system pressure when the gear is up.

The nose gear actuator unlocks and begins to retract the nose gear when 200 to 400 psi of hydraulic pressure is applied to the retract port of the nose gear actuator. The main gear begins to retract after the main gear actuators unlock the respective Jhooks on the main gear drag braces (Figure 14-19). After approximately six seconds, the retraction cycle is complete. Once the landing gear reaches full-up travel, each actuator physically bottoms out. 14 LANDING GEAR AND BRAKES

The pressure on the retract line builds rapidly until pressure reaches approximately 2,775 psi. The uplock pressure switch opens to break the power circuit to the pump motor; the hydraulic pump stops.

14-16




Figure 14-19. Landing Gear Retraction Schematic


14-17



Figure 14-20. Landing Gear Retracted Schematic

14-18



MANUAL OPERATION A h a n d - p u m p h a n d l e, p l a c a r d e d LA N D I N G G E A R A LT E R NAT E EXTENSION (Figure 14-21), is on the floor between the pilot seat and the pedestal. The pump under the floor below the handle is for emergency extension when normal extension of the gear is incomplete.

Figure 14-22. Landing Gear Circuit Breaker

Stow the pump handle back in the retaining clip when the gear is down and locked. Figure 14-23 illustrates a manual extension.

EXTENSION To engage the system, pull the LANDING GEAR RELAY circuit breaker (Figure 1422) below and to the left of the LDG GEAR CONTROL handle. Ensure the LDG GEAR CONTROL handle is in the DN position. Remove the pump handle from the securing clip. Pump the handle up and down until the green NOSE - L -R gear down indicator lights illuminate. Further resistance should be felt.

If for any reason the green GEAR DOWN lights do not illuminate (e.g., in case of an electrical system failure, or in the event an actuator is not locked down), continue pumping until sufficient resistance is felt to ensure that the gear is down and locked. Do not stow pump handle. The landing gear cannot be manually retracted in flight.

WARNING After an emergency landing gear extension has been made, do not move any landing gear controls or reset any switches or circuit breakers until the aircraft is on jacks. The failure may be in the gear-up circuitry or the hydraulic system and the gear might retract on the ground. The landing gear cannot be retracted manually.


14-19


WARNING Figure 14-21. Landing Gear Alternate Extension Placard


Refer to the Normal Procedures section of the POH for information pertaining to practice manual landing gear extensions.

P u s h t h e LA N D I N G G E A R R E LAY circuit breaker in and move the LDG G E A R C O N T RO L h a n d l e t o t h e U P position.

If the manual extension handle is properly stowed after a practice manual extension, the gear may be retracted hydraulically.


Figure 14-23. Hand Pump Extension

14-20



RETRACTION A service valve forward of the power pack assembly may be used in conjunction with the hand pump to raise the gear for maintenance purposes (Figure 14-24).


With the aircraft on jacks and an external electrical power source attached, unlatch

the hinged retainer and pull up on the red knob located on top of the service valve. The hand pump can then be used to raise the gear to the desired position. After the required maintenance has been performed, push the red knob down. Use the hand pump to lower the gear.

Figure 14-24. Maintenance Hand Pump Retraction


14-21


CAUTION If the red knob on the service valve is pushed down while the landing gear is retracted with electrical power on and the landing gear control handle DN, the landing gear will extend immediately. The service valve must be down for electrohydraulic action (up or down) and for manual extension of the landing gear.

NOSEWHEEL STEERING Direct linkage to the rudder pedals permits nosewheel steering when the nose gear is down (Figure 14-25).

One spring-loaded link in the system absorbs some of the force applied to the interconnected rudder pedals until the nosewheel is rolling. At this time, the resisting force is less; more pedal motion results in more nosewheel deflection. Because motion of the pedals is transmitted via cables and linkage to the rudder, rudder deflection occurs when force is applied to the rudder pedals. With the nose landing gear retracted, some of the force applied to the rudder pedals is absorbed by the springloaded link in the steering system. There is no motion at the nosewheel, but rudder deflection still occurs. Th e n o s e w h e e l upon retraction.

self-centering

When force on the rudder pedal is augmented by a main wheel braking action, the nosewheel deflection can be considerably increased.


Figure 14-25. Nosewheel Steering Mechanism

14-22

is



Nosewheel steering provides 12° to the left and right of center. Castering provides an additional 36°, for a total possible deflection of 48° left and right (Figure 14-26).

Th i s a r ra n g e m e n t a l l o w s d i ff e re n t i a l braking for taxiing and maneuvering on the ground. The dual brakes are plumbed in series. The pilot master cylinders are plumbed through the copilot master cylinders. This allows either set of pedals to perform the braking action and eliminates need for shuttle valves. Neither set of brake pedals can override the other. Proper traction and braking control cannot be expected until the landing gear is carrying the full weight of the aircraft. Use extreme care when braking to prevent skidding. Braking should be smooth and even all the way to the end of ground roll.

BRAKE SYSTEM The King Air 350 uses a non-ass i s t e d hydraulic brake system (Figure 14-27). The main landing gear wheels are equipped with multi-disc dual hydraulic brakes. Toe pressure on the rudder pedals by either pilot actuates the brakes. Depression of either set of pedals compresses the piston rod in the master cylinder attached to each pedal. Hydraulic pressure that results from the piston movement is transmitted through flexible hoses and fixed aluminum tubing to the disc brake assemblies on the main landing gear. This pressure forces the brake pistons to press against the linings and discs of the brake assembly. Each rudder pedal is attached to its own master cylinder. The pilot and copilot right rudder pedals control the brake in the right main landing gear. Similarly, the left rudder pedals control braking in the left main gear.

NORMAL USE OF BRAKES The wheel brakes can be highly effective in stopping the aircraft when the situation requires. Normal operation, however, usually requires minimal brake use. The propellers are more effective than the brakes immediately after touchdown during the landing rollout. As the aircraft decelerates, the brakes become most effective. Since most landing situations are not runway length critical, a smooth, comfortable deceleration can be achieved with almost exclusive use of the propellers until passing approximately 40 knots ground speed. At that time, light wheel brake pressure provides sufficient deceleration for a comfortable transition to taxiing clear of the runway.


14-23


Figure 14-26. Nosewheel Limits

Three automatic brake adjuster assemblies in each brake piston housing maintain the proper brake clearance and compensate for brake lining wear. The automatic brake a d j u s t e r s re d u c e b ra ke d ra g, t h e re b y allowing unhampered roll. They also tend to exhibit a softer pedal and a somewhat longer pedal stroke.


14-24 RESERVOIR

RESERVOIR

COPILOT'S MASTER CYLINDER


LH PARK BRAKE

LH WHEEL CYLINDER

PILOT'S MASTER CYLINDERS

COPILOT'S MASTER CYLINDER

LH PARK BRAKE

RH PARK BRAKE

LH WHEEL CYLINDER

RH WHEEL CYLINDER

PARKING BRAKES SET

RH PARK BRAKE

RH WHEEL CYLINDER

LEFT BRAKE APPLIED

LEGEND FLUID UNDER PRESSURE SUPPLY FLUID STATIC FLUID

Figure 14-27. Brake System Schematic


PILOT'S MASTER CYLINDERS


CAUTION When runway length is critical during landing, maximum braking techniques should be employed. Consult the Performance section of the POH for requirements and landing distances.

Nosewheel steering can accomplish turning with an occasional assist by tapping on one brake to help initiate turns of shorter radius. When minimum radius turns are required, intermittent heavy pressure may be necessary on the inboard brake.

CAUTION Do not hold steady pressure on the inboard brake while turning. It may induce a twisting action on the strut, inflicting serious damage to the strut assembly.

PARKING BRAKE The parking brake uses the regular brakes and a set of valves. Dual parking brake valves are adjacent to the rudder pedals between the master cylinders of the copilot rudder pedals and the wheel brakes.

Figure 14-28. Parking Brake

NOTE When preparing to use the parking brake, ensure the aircraft has come to a complete stop. The parking brake valves should retain the p re s s u re p re v i o u s l y p u m p e d i n t o t h e system. To release the parking brake, depress the brake pedals on either pilot or copilot side to equalize the pressure on both sides of the valves. Simultaneously push the parking brake handle in to allow the parking brake valves to open. To avoid damage to the parking brake system, tires, and landing gear, release the parking brake and install wheel chocks if t h e a i r c r a f t i s t o b e l e f t u n a tt e n d e d . Ambient temperature changes can expand or contract the brake fluid to cause excessive brake pressure or brake release.

The control for the parking brake valves is on the lower left-hand corner of the pilot left subpanel (Figure 14-28). To set the parking brake, depress and hold the brake pedals to maintain pressure in the brake system. Then depress the button in the center of the parking brake control and pull out the control handle. This procedure closes both parking brake valves simultaneously.


14-25


Speed can normally be controlled primarily with the the propellers in the GROUND FINE range during taxi. The brakes should be used when necessary for supplemental control.


LIMITATIONS GEAR OPERATING LIMITS Landing gear cycles (one up/one down) are limited to one every five minutes for a total of six cycles followed by a 15 minute cool-down period. Airspeed limits for the landing gear are summarized in Table 14-2. Table 14-2. KING AIR 350 AIRSPEEDS AIRSPEED

KIAS

REMARKS

Maximum gear operating speed (VLO) Extension Retraction

184 166

Maximum gear extended speed (VLE)

184

Do not extend or retract the landing gear above these speeds. Do not exceed this speed with the landing gear extended.

Figure 14-29. Brake Fluid Reservoir

SERVICING SHOCK STRUTS 14 LANDING GEAR AND BRAKES

Shock struts should always be properly inflated. Do not over- or under-inflate, and never tow or taxi an aircraft when any strut is flat. Correct inflation is placarded on the main and nose struts or refer to the Maintenance Manual.

Brake system servicing is limited primarily to maintaining the hydraulic fluid level in the reservoir. A dipstick is provided for measuring the fluid level. When the reservoir is low on fluid, add a sufficient quantity of MIL-H-5606 hydraulic fluid to fill the reservoir to the full mark on the dipstick. Check all hydraulic connections for signs of seepage and correct if necessary.

BRAKE SERVICE B ra ke f l u i d i s s u p p l i e d t o t h e m a s t e r cylinders from a reservoir accessible through the nose avionics compartment door (Figure 14-29). The brake fluid reservoir is located on the upper corner of the left side of the nose avionics compartment.

14-26



HYDRAULIC SERVICE The hydraulic fill reservoir is just inboard of the left nacelle and forward of the front spar (Figure 14-30). It contains a cap and dipstick assembly to facilitate maintenance of the system fluid level.

A line plumbed to the upper portion of the fill reservoir is routed overboard to act as a vent.

BRAKE WEAR LIMITS To cheek the brakes for wear, measure the distance between the segmented carrier and lining assembly and the piston housing (Figure 14-31). When this distance (with the parking brake set) measures 0.200 inch or more, the brake is ready for a lining inspection per the manufacturer’s maintenance manual.


Figure 14-30. Hydraulic Fluid Reservoir

Figure 14-31. Brake Wear Diagram


14-27




14-28



QUESTIONS 1. The landing gear handle is designed to work airborne: A. Weight off wheels. B. And on the ground. C. If the cabin is pressurized. D. And on the ground if the cabin is pressurized. 2. The green GEAR DOWN annunciators indicate the gear: A. Handle is in the DOWN position. B. Handle is in the UP position. C. Is down and locked. D. Is up and locked. 3. The LDG GEAR CONTROL red light illuminates when the gear position may be unsafe and: A. Can be dimmed. B. Can be extinguished. C. Can be dimmed if low level cabin ambient light is sensed. D. Cannot be dimmed.


4. The alternate landing gear extension system is available for gear: A. Retraction. B. Extension. C. Extension and retraction. D. Extension or retraction. 5. The maximum permitted landing gear extended speed is _______ KIAS. A. 158 B. 184 C. 194 D. 263


14-29


CHAPTER 15 FLIGHT CONTROLS CONTENTS Page INTRODUCTION............................................................................................................. 15-1 GENERAL ......................................................................................................................... 15-1 PRIMARY FLIGHT CONTROLS ................................................................................. 15-2 Trim Tabs...................................................................................................................... 15-2 Electric Elevator Trim ............................................................................................... 15-4 Yaw Damp/Rudder Boost System ............................................................................ 15-4 Stall Warning System.................................................................................................. 15-6 FLAP SYSTEM .................................................................................................................. 15-6 Components................................................................................................................. 15-6 Controls and Indicators.............................................................................................. 15-7 Operation ..................................................................................................................... 15-7 Abnormal Conditions................................................................................................. 15-7 LIMITATIONS................................................................................................................... 15-8 CONTROL LOCKS ......................................................................................................... 15-8 Removal ...................................................................................................................... 15-8 Installation .................................................................................................................. 15-8

15 FLIGHT CONTROLS

QUESTIONS ...................................................................................................................... 15-9


15-i



Title

Page

Elevator and Rudder .......................................................................................... 15-2

15-2

Rudder Pedals...................................................................................................... 15-2

15-3

RUDDER TAB, AILERON TRIM, and ELEVATOR TRIM .................... 15-3

15-4

Electric Elevator Trim ........................................................................................ 15-4

15-5

Yaw Damp/Rudder Boost System..................................................................... 15-5

15-6

RUDDER BOOST Switch ................................................................................ 15-5

15-7

Flap System Components................................................................................... 15-6

15-8

FLAPS Lever ....................................................................................................... 15-7

15 FLIGHT CONTROLS

15-1


15-iii


CHAPTER 15 FLIGHT CONTROLS

INTRODUCTION The flight controls allow the pilot to control the aircraft about the three axes of pitch, roll, and yaw. The flap system helps provide optimum performance in takeoff, approach, and landing modes. This chapter discusses these flight controls.

All flight controls, with the exception of the flaps, are cable-operated conventional surfaces that require no power assistance for normal control by the crew. The flaps and electric elevator trim are e l e c t r i c a l l y p o w e re d . A n e l e c t r i c a l l y

powered rudder boost/yaw dampening system connects directly to the autopilot to aid the pilot during an engine-out condition. It functions automatically when activated as torquemeter oil pressure drops during an engine failure.


15-1

15 FLIGHT CONTROLS

GENERAL


PRIMARY FLIGHT CONTROLS The primary flight controls are a threeaxes control system. The aileron surfaces on each wing provide lateral (roll) control. The elevator provides longitudinal (pitch) control. The rudder provides directional (yaw) control (Figure 15-1).

Figure 15-2. Rudder Pedals

TRIM TABS

Figure 15-1. Elevator and Rudder

All primary flight control surfaces are manually controlled through cable/bellcrank systems. Surface travel stops and linkage a d j u s t m e n t s a re i n e a c h s y s t e m . Du a l controls are available for either the pilot or the copilot. Conventional push-pull control wheels interconnected by a T shaped control column operate the ailerons and elevators. A linkage below the crew compartment floor interconnects the rudder pedals ( Fi g u re 15 - 2 ) . Ru d d e r b e l l c ra n k a r m s adjustable to two positions move the pedals approximately one inch forward or aft.

Tr i m t a b s a r e i n s t a l l e d o n t h e l e f t aileron, rudder and each elevator. The pilot manually controls the tabs through drum-cable systems that use dual jackscrew actuators. The dual actuators drive adjustable doubleend clevis rod assemblies capable of removing joint free play. Positive stops on the primary flight control surfaces limit their travel while traveling stops secured to the cables limit trim tab movement. All trim tab actuators are of a dual configuration to provide additional safety control of trim and tab free play. Tabs are positioned with controls on the pedestal (Figure 15-3). These include the RU D D E R TA B k n o b, t h e A I L E RO N TRIM knob, and the ELEVATOR TRIM wheel.Each knob also contains directional arrows as well as marked settings.

15 FLIGHT CONTROLS

Control locks in the cockpit secure the ailerons, elevators, and rudder when the aircraft is on the ground and out of service.

15-2


15 FLIGHT CONTROLS


Figure 15-3. RUDDER TAB, AILERON TRIM, and ELEVATOR TRIM


15-3


ELECTRIC ELEVATOR TRIM The electric elevator trim system is installed in conjunction with the autopilot system. A dual-element thumb switch on the outboard handle of each control wheel (Figure 15-4) activates the electric motordriven elevator trim tabs. Both elements of the switch must be simultaneously moved forward to achieve a nose-down trim; aft for nose-up trim.

The manual trim control wheel interrupts and resets the electric trim system. Use it anytime the autopilot is off whether or not the electric trim system is in the operative mode. The PITCH TRIM circuit breaker in the FLIGHT group on the right CB panel protects the system.

YAW DAMP/RUDDER BOOST SYSTEM The yaw damp/rudder boost system aids the pilot in directional control associated with asymmetrical thrust in the event of an engine loss or large power difference (Figure 15-5).

Figure 15-4. Electric Elevator Trim

When the elements are released, they return to the center off position. The pilot switch overrides the copilot switch. A bi-level, pushbutton switch is inboard of the thumb switch on the outboard grip of each control wheel. The momentary-on, trim-disconnect switch disconnects the elevator trim system when either of these switches is depressed. Depressing either trim-disconnect switch to the first of the two levels disconnects the autopilot, yaw damp, and rudder boost systems. Depressing the switch to the second level disconnects the electric elevator trim system.

The yaw damp portion senses changes in heading (or yaw rate information) from the rate gyro and uses an electric servo to drive the rudder control cables. Heading changes due to turns are sensed using information from the attitude gyro to allow for turn coordination. On the 350 models, rudder boost senses e n g i n e t o r q u e f r o m b o t h e n g i n e s, a s measured by a torque transducer tapped off of the torque manifold. It is separate from the normal engine torque system. When the difference in these torques exceeds approximately 30%, rudder boost activation begins. An electric servo activates to deflect the rudder, assisting pilot effort. The servo contribution is proportional to the engine torque differential. The pilot trims the rudder.

15 FLIGHT CONTROLS

15-4



LEFT ENGINE TORQUE TRANSDUCER

RUDDER BOOST SWITCH

YAW RATE INFORMATION

YAW DAMP/ RUDDER BOOST COMPUTER

RIGHT ENGINE TORQUE TRANSDUCER

YAW DAMP SWITCH

RUDDER SERVO

LEGEND SYSTEM INPUTS SYSTEM CONTROLS RUDDER

SYSTEM OUTPUTS

KING AIR 350

Figure 15-5. Yaw Damp/Rudder Boost System

The RUDDER BOOST switch is on the pedestal (Figure 15-6). To enable the system, place the switch in RUDDER BOOST and the AP/YD DISC bar on the flight guidance panel up. The rudder boost system is disabled if the switch is in the OFF position. If the DISC TRIM/AP YD switch is depressed, the system is interrupted. A n a m b e r RU D B O O S T O F F o n t h e caution/ advisory/status annunciator panel indicates the switch is inoperative because the switch is in off or the trim disconnect switch on either yoke has been used.

15 FLIGHT CONTROLS

Figure 15-6. RUDDER BOOST Switch


15-5


STALL WARNING SYSTEM

COMPONENTS

Angle of attack is sensed by aerodynamic pressure on the lift transducer vane on the left wing leading edge. When a stall is imminent, an aural stall warning sounds.

The flaps are two segments on each wing. An electric motor drives the flaps through a gearbox mounted on the forward side of the rear spar (Figure 15-7).

The system can be tested during peflight with the STALL WARN TEST switch on the copilot left subpanel. The switch is spring-loaded to the center OFF position. Hold the switch in the STALL WARN TEST position to activate the stall warning system.

The gearbox drives four flexible driveshafts that connect to jackscrew actuators at each flap segment.

FLAP SYSTEM

Protection

The aircraft has a four-segment, Fowlertype flap system. The flaps may be selected t o t h e U P, A P P R OAC H , o r D OW N position. They cannot be stopped at an intermediate point between these positions.

The motor incorporates a dynamic braking system. Two sets of motor windings help prevent overtravel of the flaps.

A 20-amp FLAP MOTOR circuit breaker on the right CB panel protects the flap m o t o r c i rc u i t . A 5 - a m p F LA P I N D & CONTROL circuit breaker protects the control circuit.

Whenever one of the three positions is selected, the flaps move to the selected position. A safety mechanism disconnects power to the electric flap motor if a malfunction occurs that could cause any flaps to be 3° to 6° out of phase with the other flaps.

FLAP ASYMMETRY SWITCHES

LIMIT SWITCHES FLAP INDICATOR TRANSDUCER

15 FLIGHT CONTROLS

20A L GEN BUS

Figure 15-7. Flap System Components

15-6



CONTROLS AND INDICATORS A s l i d i n g F LA P S l e v e r i s j u s t b e l o w the condition levers on the pedestal (Figure 15-8).

• 89 KIAS (91 KIAS for 350ER)— Stalling speed (V S1 ) at maximum weight with flaps approach and idle power • 96 KIAS (99 KIAS for 350ER)— Stalling speed (V S1 ) at maximum weight with flaps up and idle power The white DN arrow marks maximum speed permissible with flaps extended beyond approach: 158 KIAS. The white APP arrow marks maximum speed permissible with flaps in approach position: 202 KIAS.

OPERATION Figure 15-8. FLAPS Lever

Lowering the flaps produces these results: • Attitude—Slight nose up

Flap deflection from 0% (up) to 10 0% (down) displays on an electric indicator on top of the pedestal below the caution/advisory annunciator panel. A potentiometer driven by the right inboard flap segment operates the indicator. The right inboard flap also drives the flap position limit switches.

Airspeed Indicator Display The solid red bar at the bottom of the airspeed scale on the display is the impending stall speed low speed cue (ISS LSC) marker. The top of the marker changes with flap position to reflect the following stall speed: • 81 KIAS (82 KIAS for 350ER)— Stalling speed (V S0 ) at maximum weight with flaps down and idle power

• Airspeed—Reduced • Stall speed—Lowered • Trim—Input required based on the amount of airspeed change

ABNORMAL CONDITIONS Landing Gear Warning System The landing gear warning system is affected by the position of the flaps. Anytime the flaps are in the approach position and the landing gear is not extended, a warning horn sounds until the gear is extended or the flaps retracted with the torque below 40% or 86% N 1 . The landing gear warning system warns the pilot the landing gear is not down and locked during specific flight regimes. The warning horn sounds continuously when the flaps are lowered beyond the APPROACH (40%) position regardless of the power lever setting until the landing gear is extended or the flaps retracted. Refer to Chapter 14 for details.


15-7

15 FLIGHT CONTROLS

The control lever has a position detent TAKEOFF AND APPROACH to select 40% flaps for takeoff or approach. Full flap deflection, or 100%, is equal to approximately 35º of flap travel. Detents also are marked for UP and DOWN positions.


Asymmetrical Flap Protection An asymmetrical flap switch system provides split-flap protection. The switch shuts off the flap motor for any out-ofphase condition of approximately 3° to 6° between the adjacent flap segments on each wing. The switch is spring-loaded to the normallyopen position, but it is rigged so that the roller cam holds the switch in its momentary (closed) position. This provides electrical continuity to the flap motor when the outboard and inboard flap segments on both sides are parallel and in phase with one another.

REMOVAL WARNING Before starting engines, remove control locks. 1. Remove rudder pin 2. Remove control column pin 3. Remove U-shaped power control clamp

Invalid Lever Input If the FLAPS lever input becomes invalid, a default ISS LSC marker displays. This is a checkerboard bar with a yellow bar on top. The junction of the checkerboard bar represents the flaps-down ISS LSC value. The top of the yellow bar represents the flaps-up ISS LSC value.

LIMITATIONS

WARNING Re m o v e c o n t ro l l o c k s b e f o re towing the aircraft. If towed with a tug while the rudder lock is intalled, serious damage to the steering linkage can result.

INSTALLATION

Do not extend flaps or operate with flaps extended above these speeds. Maximum flap extension/extended speeds (V FE ): • Approach—202 KIAS • Full down—158 KIAS

CONTROL LOCKS 15 FLIGHT CONTROLS

The control locks consist of a U-shaped clamp and two pins connected by a chain. The pins lock the primary flight controls. The U-shaped clamp fits around the engine power control levers and serves to warn the pilot not to start the engines with control locks installed.

15-8

It is important that the locks be installed o r re m o v e d t o g e t h e r t o p re c l u d e t h e possibility of an attempt to taxi or fly the aircraft with the power levers released and the pins still installed in the flight controls.

1. Position U-shaped clamp around engine power controls 2. Move control column as necessary to align holes in the control column 3. Insert the L-shaped pin attached to the middle of the chain; holes are aligned when control wheel is full forward and rotated approximately 15° to the left 4. Insert the L-shaped pin attached to the end of the chain through the hole provided in the floor aft of the rudder pedals; rudder pedals must be centered to align the hole in the rudder bellcrank with the hole in the floor 5. I n s e r t p i n u n t i l t h e f l a n g e r e s t s against the floor; this prevents any rudder movement



QUESTIONS 4. The mechanical aileron trim is located _______ of the power quadrant on the _______ side. A. Aft; left B. Aft; right C. Forward; left D. Forward; right

2. Rudder boost aids the pilot in rudder deflection during engine failure operation by sensing: A. Yaw rate. B. Roll rate. C. Torque differential. D. Bleed air differential.

5. The maximum speed permissible with flaps in the approach position is _______ KIAS. A. 160 B. 174 C. 194 D. 202

3. Electric pitch trim is available when: A. Either trim switch is activated. B. Both trim switches are activated simultaneously. C. O p p o s i t e t r i m s w i t c h e s a r e activated. D. Indicated airspeed is greater than 140 knots.

6. Rudder boost: A. Is not required when operating at weights less than 12,500 lbs. B. Is off for takeoff and landing. C. Must be on for takeoff and cruise. D. Must be on and operational for t a k e o f f, c l i m b, a p p r o a c h a n d landing.

15 FLIGHT CONTROLS

1. Secondary flight controls surfaces are: A. Manually-controlled. B. Hydraulically-controlled. C. E l e c t r i c a l l y a n d h y d r a u l i c a l l y controlled. D. M a n u a l l y and electrically controlled.


15-9

16 AVIONICS


CHAPTER 16 AVIONICS INTRODUCTION ................................................................................................................... 16-1 FLIGHT INSTRUMENTS...................................................................................................... 16-1 Adaptive Flight Displays (AFD)..................................................................................... 16-2 Multifunction Display (MFD) ....................................................................................... 16-12 DISPLAY CONTROL PANELS (DCP) ............................................................................. 16-17 INTEGRATED AVIONICS PROCESSOR SYSTEM (IAPS) ....................................... 16-24 AIR DATA COMPUTERS (ADC)..................................................................................... 16-24 ATTITUDE AND HEADING REFERENCE SYSTEM (AHRS)................................ 16-25 REVERSIONARY OPERATIONS.................................................................................... 16-26 OUTSIDE AIR TEMPERATURE .................................................................................... 16-31 STALL WARNING SYSTEM .............................................................................................. 16-33 FLIGHT GUIDANCE SYSTEM (FGS)............................................................................. 16-34 Flight Guidance Computers (FGC).............................................................................. 16-35 Flight Guidance Panel (FGP) ....................................................................................... 16-35 Control Wheel Switches.................................................................................................. 16-43 CONTROL DISPLAY UNIT (CDU).................................................................................. 16-45 FLIGHT MANAGEMENT SYSTEM (FMS) .................................................................... 16-50 Vertical Navigation ......................................................................................................... 16-52 Global Positioning System (GPS) ................................................................................. 16-54 INTEGRATED FLIGHT INFORMATION SYSTEM (IFIS) ........................................ 16-56 Cursor Control Panel (CCP).......................................................................................... 16-58 COMMUNICATION/NAVIGATION SYSTEMS ............................................................ 16-71 Audio System................................................................................................................... 16-75


16-i

16 AVIONICS


Radio Tuning Unit (RTU).............................................................................................. 16-78 CDU Tuning..................................................................................................................... 16-83 ELECTRONIC STANDBY INSTRUMENT SYSTEM (ESIS) ...................................... 16-88 WEATHER RADAR SYSTEM........................................................................................... 16-91 COCKPIT VOICE RECORDER (CVR)........................................................................... 16-96 EMERGENCY LOCATOR TRANSMITTER (ELT)...................................................... 16-96 ENHANCED GROUND PROXIMITY WARNING SYSTEM (EGPWS).................. 16-97 Basic Ground Proximity Warning System (GPWS) .................................................... 16-97 Enhanced Ground Proximity Warning System (GPWS)............................................ 16-99 TRAFFIC COLLISION AND AVOIDANCE SYSTEM (TCAS I) ............................. 16-102 TRAFFIC COLLISION AND AVOIDANCE SYSTEM (TCAS II) (OPTIONAL).. 16-104 APPENDIX A – AVIONICS EQUIPMENT LOCATIONS ......................................... 16-111 APPENDIX B – FLIGHT GUIDANCE MODES.......................................................... 16-113 APPENDIX C – AVIONICS ACRONYMS..................................................................... 16-117 QUESTIONS ........................................................................................................................ 16-121

16-ii


16 AVIONICS



Title

Page

16-1.

Adaptive Flight Displays (AFD)..................................................................... 16-2

16-2.

Primary Flight Display...................................................................................... 16-3

16-3.

Attitude Display ................................................................................................ 16-4

16-4.

Airspeed Display .............................................................................................. 16-4

16-5.

Trend Vector....................................................................................................... 16-4

16-6.

Low Speed Cue.................................................................................................. 16-5

16-7.

High Speed Cue................................................................................................. 16-5

16-8.

Airspeed Speed Bug ......................................................................................... 16-5

16-9.

Acceleration Display......................................................................................... 16-6

16-10.

Altimeter Display.............................................................................................. 16-6

16-11.

Altitude Negative .............................................................................................. 16-6

16-12.

Baro Switch ........................................................................................................ 16-7

16-13.

Vertical Speed Indicator (VSI)........................................................................ 16-7

16-14.

Altitude Preselect Bugs .................................................................................... 16-8

16-15.

Metric Altitude .................................................................................................. 16-8

16-16.

BARO ALT Switch ........................................................................................... 16-8

16-17.

Heading and Navigation Display ................................................................... 16-9

16-18.

DME hold .......................................................................................................... 16-9

16-19.

PFD Compass Rose Format........................................................................... 16-10

16-20.

PFD Arc Format.............................................................................................. 16-10

16-21.

PFD Map Format ............................................................................................ 16-11

16-22.

Terrain and Radar Overlay Section .............................................................. 16-11

16-23.

PFD TCAS Message Area (Non-IFIS)......................................................... 16-11


16-iii

16 AVIONICS


Figure

Title

Page

16-24.

PFD Lower Display Information .................................................................. 16-12

16-25.

Pilot's MFD Display ....................................................................................... 16-12

16-26.

Non-IFIS MFD Checklist............................................................................... 16-14

16-27.

MFD Upper Format (IFIS)............................................................................ 16-14

16-28.

MFD Plan Format ........................................................................................... 16-15

16-29.

MFD TCAS only ............................................................................................. 16-16

16-30.

TCAS ............................................................................................................... 16-16

16-31.

MFD Lower Display Information................................................................. 16-17

16-32.

Display Control Panels ................................................................................... 16-17

16-33.

Display Control Panel (DCP)........................................................................ 16-18

16-34.

Barometric Setting with Yellow Underline .................................................. 16-18

16-35.

IN/hPa Switch .................................................................................................. 16-18

16-36.

Barometric Setting with STD ........................................................................ 16-19

16-37.

PFD REFS Menu Page 1 of 2 ........................................................................ 16-19

16-38.

PFD V-Speeds.................................................................................................. 16-20

16-39.

Radio Altitude Minimum............................................................................... 16-20

16-40.

Barometric Minimum ..................................................................................... 16-21

16-41.

Minimums Annunciator ................................................................................. 16-21

16-42.

PFD REFS Menu Page 2 of 2 ........................................................................ 16-21

16-43.

Metric Altitude ................................................................................................ 16-22

16-44.

Flight Director Formats.................................................................................. 16-22

16-45.

PFD NAV BRG Menu .................................................................................. 16-23

16-46.

Bearing Pointer Information ......................................................................... 16-23

16-47.

IAPS ................................................................................................................. 16-24

16-iv


Figure

Title

Page

16-48.

ADC ................................................................................................................. 16-25

16-49.

AHRS ............................................................................................................... 16-25

16-50.

Heading Slave and Slew ................................................................................. 16-26

16-51.

AFD Reversions ............................................................................................. 16-26

16-52.

Reversionary Modes ....................................................................................... 16-27

16-53.

ADC1 Failure ................................................................................................. 16-28

16-54.

ADC Miscompares ......................................................................................... 16-28

16-55.

ADC Switch - ADC2 Selected....................................................................... 16-29

16-56.

AHRS1 Failure ................................................................................................ 16-29

16-57.

AHRS Miscompares ....................................................................................... 16-30

16-58.

Pitot Tubes........................................................................................................ 16-30

16-59.

Static Ports ..................................................................................................... 16-30

16-60.

Alternate Static Source Selection ................................................................. 16-31

16-61.

System Integration .......................................................................................... 16-32

16-62.

OAT Gauge...................................................................................................... 16-33

16-63.

Rosemont Probe.............................................................................................. 16-33

16-64.

Transducer Vane .............................................................................................. 16-33

16-65.

Stall Warning Test Switch ............................................................................... 16-34

16-66.

Stall Warning Heat .......................................................................................... 16-34

16-67.

Flight Guidance Panel (FGP) ........................................................................ 16-35

16-68.

Flight Guidance System Display ................................................................... 16-35

16-69.

Flight Guidance Panel (FGP) ........................................................................ 16-36

16-70.

Flight Guidance Couple Arrow..................................................................... 16-36

16-71.

Independent Flight Director Operation....................................................... 16-36


16-v

16 AVIONICS


16 AVIONICS


Figure

Title

Page

16-72.

YD/AP Disconnect Bar.................................................................................. 16-37

16-73.

Heading Vector Line....................................................................................... 16-38

16-74.

Half Bank Mode.............................................................................................. 16-38

16-75.

APPR Mode Selection ................................................................................... 16-39

16-76.

Localizer Nav-to-Nav Capture ...................................................................... 16-40

16-77.

VNAV Glidepath (GP) Mode ....................................................................... 16-40

16-78.

Vertical Speed (VS) Mode ............................................................................. 16-41

16-79.

Flight Level Change (FLC) Mode ................................................................ 16-42

16-80.

Left Yoke.......................................................................................................... 16-43

16-81.

Pilot's PFD with SYNC .................................................................................. 16-44

16-82.

Go-Around Button ......................................................................................... 16-44

16-83.

PFD Go-Around (GA) Mode ....................................................................... 16-45

16-84.

Control Display Unit (CDU)......................................................................... 16-45

16-85.

Active Flight Plan Page .................................................................................. 16-47

16-86.

Active Legs Page ............................................................................................. 16-47

16-87.

Direct to Pages................................................................................................. 16-47

16-88.

Hold FPLN Mode ........................................................................................... 16-48

16-89.

MFD Menu Key (CDU)................................................................................. 16-49

16-90.

MFD Advance Key (CDU)............................................................................ 16-50

16-91.

MFD Text Page................................................................................................ 16-50

16-92.

Database Units ................................................................................................ 16-51

16-93.

Active Legs Page with VNAV Altitudes....................................................... 16-53

16-94.

VNAV Top of Descent.................................................................................... 16-54

16-95.

VNAV Modes .................................................................................................. 16-54

16-vi


Figure

Title

Page

16-96.

GPS CONTROL ............................................................................................. 16-55

16-97.

PROGRESS .................................................................................................... 16-55

16-98.

IFIS Block Diagram........................................................................................ 16-57

16-99.

Ethernet Database Unit ................................................................................. 16-58

16-100.

USB Database Unit (DBU-5000)................................................................. 16-58

16-101.

MCDU Menu. ................................................................................................. 16-58

16-102.

IFIS Dataload Block Diagram....................................................................... 16-59

16-103.

CCP................................................................................................................... 16-60

16-104.

MFD Store Complete ..................................................................................... 16-60

16-105.

Geo-Politcal Overlay ...................................................................................... 16-60

16-106.

Airspace Overlay ........................................................................................... 16-61

16-107.

Airways Overlay.............................................................................................. 16-61

16-108.

Database Effectivity (STAT Key) ................................................................. 16-62

16-109.

STAT Menu...................................................................................................... 16-62

16-110.

Chart Subscription (STAT Key).................................................................... 16-62

16-111.

MFD Chart Display ........................................................................................ 16-63

16-112.

MFD Chart Menu ........................................................................................... 16-63

16-113.

MFD Chart Approach Index ......................................................................... 16-64

16-114.

MFD Chart Zoom Box................................................................................... 16-64

16-115.

MFD Chart Geo-Reference Symbols ........................................................... 16-65

16-116.

MFD Chart Menu .......................................................................................... 16-65

16-117.

MFD PLAN Map Weather Overlay ............................................................. 16-66

16-118.

MFD Dedicated Graphical Weather Format (XM Weather) .................... 16-66

16-119.

MFD XM Weather Menu .............................................................................. 16-67


16-vii

16 AVIONICS


16 AVIONICS


Figure

Title

Page

16-120.

MFD Metar Display ....................................................................................... 16-67

16-121.

Overlay Legends ............................................................................................. 16-68

16-122.

MFD Graphical Weather Time Stamps ....................................................... 16-68

16-123.

MCDU Datalink Pages (Universal Weather) .............................................. 16-69

16-124.

Datalink Weather Selections (Universal Weather) ..................................... 16-70

16-125.

MFD_Plan Map Weather Overlay ................................................................ 16-70

16-126.

MFD Dedicated Graphical Weather Format(Universal Weather) ........... 16-71

16-127.

Universal Weather Menu ............................................................................... 16-71

16-128.

RTU / CDU TUNE Switch ............................................................................ 16-72

16-129.

Emergency Frequency Button ....................................................................... 16-72

16-130.

Antennas .......................................................................................................... 16-73

16-131.

RMT Tune Switch............................................................................................ 16-73

16-132.

PFD DME Displays ........................................................................................ 16-74

16-133.

DME Hold Selection and Images ................................................................. 16-74

16-134.

ATC Transponder Switch ............................................................................... 16-75

16-135.

Flight ID Selection ......................................................................................... 16-75

16-136.

Audio Panels .................................................................................................... 16-76

16-137.

Audio System Components............................................................................ 16-76

16-138.

Control Wheel (PTT) Switches...................................................................... 16-78

16-139.

Radio Tuning Unit (RTU).............................................................................. 16-79

16-140.

RTU in Preset Tuning Mode .......................................................................... 16-79

16-141.

RTU COMM Pages ........................................................................................ 16-80

16-142.

RTU NAV Pages ............................................................................................ 16-80

16-143.

RTU ADF Pages ............................................................................................. 16-81

16-viii


Figure

Title

Page

16-144.

RTU ATC Page ............................................................................................... 16-81

16-145.

RTU HF Pages ................................................................................................ 16-82

16-146.

RTU TCAS II Pages ....................................................................................... 16-83

16-147.

CDU Tune with TCAS I ................................................................................. 16-83

16-148.

CDU Frequency Data..................................................................................... 16-84

16-149.

CDU COMM Page ......................................................................................... 16-84

16-150.

CDU NAV Page .............................................................................................. 16-85

16-151.

CDU ATC Page............................................................................................... 16-85

16-152.

CDU ADF Page .............................................................................................. 16-86

16-153.

CDU TUNE With TCAS II............................................................................ 16-86

16-154.

MFD TCAS Display ....................................................................................... 16-87

16-155.

CDU TCAS II Control ................................................................................... 16-87

16-156.

CDU HF Control ............................................................................................ 16-87

16-157.

GND COMM Button ..................................................................................... 16-88

16-158.

Static Wicks...................................................................................................... 16-88

16-159.

ESIS Display ................................................................................................... 16-89

16-160.

ESIS Power Switch.......................................................................................... 16-89

16-161.

ESIS Menu....................................................................................................... 16-91

16-162.

PFD Radar Menu............................................................................................ 16-92

16-163.

Test Mode......................................................................................................... 16-92

16-164.

Radar Ground Map Mode ............................................................................. 16-92

16-165.

Radar Display with Path Attenuation Bar ................................................... 16-93

16-166.

Radar Display Turbulence Mode .................................................................. 16-93

16-167.

Turbulence Only Display ............................................................................... 16-94


16-ix

16 AVIONICS


16 AVIONICS


Figure

Title

Page

16-168.

Radar Gain Display ........................................................................................ 16-94

16-169.

Pilot's PFD with TGT..................................................................................... 16-94

16-170.

Radar Ground Clutter Supression ................................................................ 16-95

16-171.

Radar Tilt Display ........................................................................................... 16-96

16-172.

CVR Controllers ............................................................................................. 16-96

16-173.

ELT Manual Switch ........................................................................................ 16-97

16-174.

PFD GND PROX and PULL UP Annunciators ........................................ 16-97

16-175.

GPWS Failure Annunciators ......................................................................... 16-97

16-176.

EGPWS Buttons ............................................................................................. 16-99

16-177.

Terrain Display.............................................................................................. 16-100

16-178.

Terrain Fail and TERR Annunciators ........................................................ 16-101

16-179.

TCAS I TEST ................................................................................................ 16-102

16-180.

Operating Mode Button............................................................................... 16-103

16-181.

TCAS II Test.................................................................................................. 16-105

16-182.

Overview of Avionics Units ......................................................................... 16-111

16-x


16 AVIONICS


TABLES Table

Title

Page

16-1.

GPWS Cautions and Warnings...................................................................... 16-98

16-2.

EGPWS Buttons ............................................................................................. 16-99

16-3.

EPGWS Cautions and Warnings................................................................. 16-101

16-4.

TCAS Messages ............................................................................................ 16-106

16-5.

TCAS II Annunciators ................................................................................. 16-107

16-6.

TCAS II Traffic Advisory ............................................................................. 16-108

16-7.

TCAS II Resolution Advisories .................................................................. 16-109

16-8.

Flight Guidance Modes................................................................................ 16-113


16-xi

CHAPTER 16 AVIONICS

INTRODUCTION The Super King Air B350 utilizes the Collins Pro Line 21 avionics system. The Pro Line 21 Avionics System is an integrated flight instrument, autopilot, and navigation system. All functions have been combined into a compact, highly reliable system designed for ease of operation, seamless communication between systems, and reduced pilot workload.

FLIGHT INSTRUMENTS ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) The Electronic Flight Instrument System (EFIS) consists of computers and data collectors that, when coupled with other subsystems, result in the display of flight, navigation, and engine indicating on liquid crystal displays (LCD) – these are called Adaptive Flight Dis-

plays (AFD). Compared to conventional instrumentation, an EFIS system permits much more information to be presented to the pilot with a minimum of operating complexity, maintenance, and weight.


16-1

16 AVIONICS


16 AVIONICS


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INCR INCR CABIN CABIN

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15

Figure 16-1. Adaptive Flight Displays (AFD)

ADAPTIVE FLIGHT DISPLAYS (AFD) The liquid crystal (LCD) Adaptive Flight Displays (AFD) contain all the flight and navigation information previously indicated on separate “round dial” instruments. Three AFD’s are installed in the King Air B350 and are all interchangeable . When the IFIS system is installed, the MFD is modified to receive additional information. It is no longer interchangeable and carries a different part number. The left AFD functions as the pilot’s Primary Flight Display (PFD 1) on which airplane attitude, heading, altitude, vertical speed, etc., are shown. The center AFD functions as the multifunction display (MFD) on which engine indications, diagnostic pages, checklists, navigation data, etc., are shown. The MFD receives much of the same data as PFD 1. The right AFD functions as the copilot’s Primary Flight Display (PFD 2) and operates independent of PFD 1. The temperature of LCD displays must stay within appropriate limits to provide normal operation. Should these temperature extremes

16-2

be exceeded each AFD has its own temperature monitor. Depending on what is needed this monitor has control of integral heaters and cooling fans. In the event of a display failure on PFD 1 the MFD can display PFD 1 images in what’s called a reversionary composite mode. However, there is no reversionary backup to PFD 2.

Primary Flight Display (PFD) The PFD displays airplane attitude and dynamic flight data. Flight Director indications, autopilot annunciations, and navigation information are also shown in a centralized location, including during reversionary format. A typical PFD display is shown (Figure 16-2) The PFD has the following controls and indications:

Bright/Dim Rocker Switch The PILOT DISPLAYS rheostat, on the overhead panel, provides primary intensity control.The Bright/Dim Rocker Switch on the


16 AVIONICS


Collins

HDG FMS

PTCH ALTS

6935

1 4 000

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4

20 60

V2 VR V1

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1

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117 110 106

1

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ATC1 4336

RADAR ON UTC 14:41

RAT 15 oC

COM2

125.250

BRT DIM

Figure 16-2. Primary Flight Display

PFD provides secondary intensity control of the PFD. This PILOT DISPLAYS rheostat will control three displays simultaneously; the PFD, MFD and Control Display Unit (CDU) on the pedestal. This allows all three displays to be brightened together. The Bright/Dim Rocker Switch will then allow each display to be fine tuned to make its brightness even with the surrounding displays.

Line Select Keys Four line select keys (LSK) are located on each side of the AFD. These keys are used in conjunction with the information being

viewed on the AFD display. LSKs that are currently active are denoted by carets (< >) displayed adjacent to the LSK.

Attitude Display The primary function of the PFD is to show airplane attitude. The PFD additionally shows the following: flight director steering commands; flight guidance system status/mode annunciations; vertical/lateral deviation; marker beacon annunciations; and radio altitude. A rectangular-shaped slip/skid indicator is located at the base of the “sky-pointer” bank


16-3

16 AVIONICS


index. This is used like the fluid filled slip-skid indicator used in other aircraft (e.g., half of the rectangle to the right equals half ball to the right). See Figure 16-3.

where each knot of airspeed increase or decrease will rollover to show the next digit. The tape and rolling drum will begin indicating as the airspeed is above 40 knots.

Collins

140

HDG FMS

PTCH ALTS

1 4 000

140

80

4

20 60

700

TERM

600 60 6 540 20 1

10 400

0 24

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Figure 16-4. Airspeed Display R T

This display area can also show current Mach, IAS markers (bugs), IAS trend vector, low/high speed cues, and acceleration rates. The trend vector is a magenta line that extends either above or below the pointer to indicate the rate of airspeed increase or decrease. The end of the vector indicates expected airspeed in 10 seconds. A trend vector moving into a warning bar, in V either the overspeed or lowspeed area, will cause the airspeed number to flash yellow (Figure 16-5).

10

HDG FMS

PTCH ALTS

1 20 10

254 280

Figure 16-3. Attitude Display

1

260

25 0 9

Airspeed Display

240

<

The Airspeed Display on the PFD is of a moving tape design (Figure 16-4).

220

A large “pointer” at the center of the display is the current aircraft airspeed. The digital readout at this pointer acts like a rolling drum


<

16-4

Figure 16-5. Trend Vector

The Low Speed Cue / Impending Stall Speed (LSC / ISS) bar is displayed at the AFM value for stall at a maximum gross weight, power idle and no bank condition (Figure 16-6).

160

160 140

120

120

100

100

80

80

60

11 1 0

is reduced to below the red overspeed bar. If the autopilot is engaged during the overspeed, it will begin to pitch the aircraft up until achieving an airspeed just below the current Vmo or Mmo.

16 AVIONICS


91 0

254

254

280

300

260

280

25 0 9

27 0 9

220

240

240

260

200 RA

Low Speed Pre-Warning

200 RA

Low Speed Warning Overspeed Warning

Overpeed Pre-Warning

Displayed above the airspeed tape, is a speed reference that the pilot can set using the speed knob on the Flight Guidance Panel. A bug will appear on the tape next to the selected speed (Figure 16-8). <

This speed is adjusted for flap position as listed here:

Figure 16-7. High Speed Cue

<

Figure 16-6. Low Speed Cue

<

<

• 0% Flaps – 96kts • 40% Flaps – 89kts • 100% Flaps – 81kts It is important to note that these speeds are not adjusted for the current g-forces, power settings or maneuvers. They should be used as reference only and not as the primary indication of a stall. The true indication of a stall will be in the form of a stall horn, or aerodynamic buffet. The autopilot will not stop the aircraft airspeed from getting into the low speed cue but once the stall warning horn sounds the autopilot will disconnect. See the Stall Warning section later in this chapter. The high speed cue consists of a red bar starting at the current Vmo or Mmo whichever is appropriate (Figure 6-7). Should the aircraft actual airspeed enter this red bar area an overspeed warning horn will sound until the speed

Speed Bug Setting

135 160 140

11

Speed Bug

9 8

100 80

Figure 16-8. Airspeed Speed Bug


16-5

Below the airspeed tape two different digital readouts may be displayed. While on the ground the current acceleration rate is displayed in “G's". This can indicate from .00 to + or - .99g. While airborne, the current Mach number is displayed in lieu of the acceleration display (Figure 16-9). The Mach indication will appear only if the current speed is greater than .450 Mach. The display is then removed when the Mach is less than .400.

140

3 000

4

700

2 1

600 60 5 6 40 20

254

80

4000

1

280

60

400

4

260

25 0 9

30. 16I N

240

117 V2 110 VR 106 V1 ACC-.02

220

2

1000

Figure 16-10. Altimeter Display

0

M .471

TERM

1 In Flight Figure 16-9. Acceleration Display On Ground

700 600 60 5 40 20

1

400

2 4

3 0 .1 6 I N

Figure 16-11. Altitude Negative

The Altimeter setting is displayed below the altitude tape . This can be changed between inches and hectopascals. (For IFIS aircraft, see the REFS section of the Display Control Panel (DCP) to see how this is accomplished). For non-IFIS aircraft, this is accomplished by mov<


E

1

R

<

16-6

4 2

<

Should a negative altitude exist, a vertically positioned “NEG” legend will replace the ten thousands position. (Figure 16-11).

RA

<

The Altitude and Vertical Speed Displays indicate the altitude and vertical speed. The altitude data is a moving tape design with a central “pointer”. This pointer contains a digital readout with a rolling drum appearance just like the airspeed display. Each 20 feet of altitude is on a single drum and the hundreds and thousands follow when needed. At lower altitudes, green striped shutters cover the appropriate ten thousand and thousand digits (Figure16-10).

1 4 000

1(*

Altitude and Vertical Speed Displays

<

200 6935

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16 AVIONICS


ing a BARO switch on the overhead panel to the IN (inches) or hPA (hectopascals) position (Figure 16-12). This will affect both pilots PFD’s and cannot be done independently.

16 AVIONICS


4 2 1

Flight Guidance Selected Vertical Speed VNAV Vertical Speed Required

Current Vertical Speed 1 2 4

300

Figure 16-13. Vertical Speed Indicator (VSI) Figure 16-12. Baro Switch

Additionally, this altimeter setting can flash as an advisory of transition altitude / level passage. For IFIS aircraft see the REFS section of the Display Control Panel (DCP) to see how this is accomplished. For non-IFIS aircraft, this is accomplished by moving the FL180 switch on the overhead panel to the ENABLE or DISABLE position depending on whether the advisory flash is desired. This transition point cannot be changed to an altitude other than 18,000’. The vertical speed display consists of a moving green line that will angle up or down depending on the current vertical speed (Figure 16-13). The value of climb or descent will then read at the top of the display for a climb or bottom of the display for a descent,when the value is greater that 300 ft/min. Once the climb or descent decreases below 100 ft/min the digital readout will be removed.

Displayed above the altitude tape is the preselect altitude shown in cyan. This altitude is selected by the pilot using the ALT knob on the Flight Guidance Panel. The selected altitude is then marked with a Fine Preselect Altitude bug that “brackets” the altitude window when captured (Figure 16-14). A smaller Coarse Preselect Altitude bug will appear on the left side of the tape when approximately 1000’ from the selected altitude to indicate proximity to that altitude. An aural tone will sound and the preselected altitude will flash further indicating proximity to the chosen altitude. Once within 200’ of the preselected altitude, the flashing will stop. This flashing can be stopped earlier by pressing the ALT knob on the flight guidance panel. (See the Flight Guidance section later in this chapter). Should the aircraft go +- 200’ from the altitude, an aural tone will sound and the preselected altitude will change to yellow and flash. This flashing will continue until the altitude returns to within 200’ of selected. This flashing can be stopped by pressing the ALT knob on the flight guidance panel.


16 AVIONICS


7500

Preselect Altitude

600 Coarse Preselect Altitude Bug

Fine Preselect Altitude Bug

500 20

7400 80 300

Additionally, a magenta number can be displayed above the VSI (Figure 16-10). This number is FMS generated and indicates the crossing restriction altitude for the current leg (this can come automatically from the FMS database or manually by pilot input into the FMS). If desired, this number, in addition to the preselected altitude, allows the FMS to automatically fly a vertical navigation (VNAV) procedure and comply with all the known step-down fixes.

200

Figure 16-14. Altitude Preselect Bugs

This top display area can also contain the metric altitude and metric altitude preselect (Figure 16-15). For IFIS aircraft see the REFS section of the Display Control Panel (DCP) to see how this is accomplished. For non-IFIS aircraft this is accomplished by moving a BARO ALT switch on the overhead panel to the FT (feet) or meter (M) position (Figure 16-16) . This action will affect both pilots and cannot be done independently. This change does not alter the actual altitude tape; that remains in feet for all phases of flight. METRIC

2450M

6935 4 2 1

100 60 8 0 40 20 1

900

Heading and Navigation Displays The Heading and Navigation Displays at the lower portion of the PFD’s contain heading, FMS navigation display, or ground based navigation display, or radar and terrain imagery (Figure 16-17).

4000M 200

Figure 16-16. BARO ALT Switch

2 4

1018HPA

Figure 16-15. Metric Altitude

16-8

At the top center of this area is the aircraft’s current heading. To the left of that display will appear the cyan heading bug’s current selection when the bug is moved with the Flight Guidance Panel or the heading bug is out of view. Additionally, an open-circle-shaped track pointer will indicate the current aircraft ground track. The difference between the current heading and track pointer indicates drift angle and is helpful in establishing the appropriate crab to maintain course. The track pointer is generated from the FMS and will be green if it is driven from the onside FMS or yellow if it is driven from the cross-side FMS.

FOR TRAINING PURPOSES ONLY <

1

KING AIR 350/350C PRO LINE 21 PILOT TRAINING MANUAL 0

ACC .02

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1 the identifier of the station is removed and a distance will appear with an “H” indicating it is in DME hold (Figure 16-18).

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12

Figure 16-17. Heading and Navigation Display

15

< ET 01:42 COM1 121.800

VOR Active Navigation

Above the active NAV source label is an area reserved for FMS messages and annunciations. Selected messages can appear here. However, the majority of the messages will be displayed on the Control Display Unit (CDU) on the pedestal. These will be prompted by the label “MSG” to instruct the pilots to look down at the CDU and retrieve the message. Immediately below the active NAV source label is a list of related navigation distances and information. When FMS is chosen, this list contains the Desired Track (DTK), name of the next waypoint and distance to that waypoint (Figure 16-17). When LOC or VOR is chosen this list contains the frequency and the current selected course. If DME is collocated with the VOR or LOC, the identifier of the station and DME distance to the station will be displayed. However, if DME hold is selected

2

S

20. 8 H

< PRESET

15

FMS1

12

The upper left corner of the NAV display indicates the active NAV source. This will display in green when the “onside” unit is selected (e.g., NAV1 and FMS1 are green on the pilot’s side; NAV2 and FMS2 are green on the copilot’s side). If the “cross-side” unit is selected, it will display in yellow (e.g., NAV2 and FMS2 are yellow on the pilot’s side; NAV1 and FMS1 are yellow on the copilot’s side). In a single FMS aircraft, the copilot will always have a yellow FMS color and the pilot will have a green FMS color.

VOR Active Navigation With DME Hold

< Figure 16-18. DME hold

Below this list is a PRESET option (Figure 1617). The nav source inside the blue box is on standby. Should the PRESET LSK be pressed, the PRESET nav source will become the active nav source and the active nav source will now be the PRESET (This is the same as course transfer used in other systems). This PRESET option cannot display a secondary CDI and remains in standby. The last LSK on the left side is the Elapsed Timer (ET) (Figure 16-17). Pressing this LSK will start, stop and reset the timer that appears next to the ET label. This is independent of the other pilot’s timer and can only count up and not down.


16-9

16 AVIONICS

1

On the right side of the display there is a FORMAT LSK. This LSK changes the display format of the lower portion of the PFD. This will select one of three options: full compass rose, arc and map (Figure 16-17). The full compass rose is a 360˚ presentation of heading with the ability to display a CDI and two bearing pointers (Figure 16-19). On IFIS aircraft, TCAS traffic can also be displayed in 1 this format by pressing the TFC line select key. When this option is chosen, the range is limited to 50nm. To get a further range, the TCAS traffic must be deselected first. This range is controlled by the DCP and is discussed later.

FMS1

24

DTK 251 (6935) 0. 8NM

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125.250

BRT DIM

Figure 16-19. PFD Compass Rose Format

The arc format can display the same items described for the full compass rose but only presents a 120˚ portion of the compass (Figure 16-20) . In this mode, the display of TCAS traffic does not limit the range to 50nm. The display of any overlays (discussed later in this section) will limit the range to 300nm. If a further range is desired, all overlays must be removed and the arc format can be extended to a 600nm range. This mode cannot display the FMS map. The map format is similar to the arc format but instead of a large CDI image it displays the FMS map (Figure 16-21). This format is only

16-10

144 069

VOR1

N

COM1 121.800

21

< PRESET 33

< ET

1 0 1 8 HPA

W

25

VOR1 V 4.1NM SXW V ----NM SXW

251

50

w 30

< PRESET

24

30

3 0 .1 6 I N

251

TERM

TERM

FMS1

9

0

ACC-.02

DTK 251 (6935) 0. 8NM

0

ACC-.02

available when FMS is the active nav source. This mode will be automatically deselected if a non-FMS source is made active and it will revert to the arc format. Additionally, when map format is chosen on the left PFD it forces the MFD into present position map mode (PPOS) and other MFD map formats are not selec1 table. It is critical to remember that following map lines is not an alternative to CDI displays. For navigation, a lateral deviation display will appear at the bottom of the attitude indicator when map mode is chosen.

<

16 AVIONICS


Figure 16-20. PFD Arc Format

The same range limitations apply in this mode as they did with the arc format. Additional options for display with the FMS map are available through the Control Display Unit on the pedestal (see the CDU section later in this PTM). Below the FORMAT LSK is the TERR/RDR LSK. This key allows for the display of either terrain or radar images. These cannot be displayed simultaneously on the same display or when the compass rose format has been selected. The chosen option will be displayed in cyan and large font. The display of these items does NOT indicate that the unit is active (Terrain and Radar must be turned ON from a different location). Below these labels is an area



0

ACC-.02 TERM

24

FORMAT >

TERR

TERR

RDR

WX T+4.5A

4

30.16IN

251

21

W

TFC >

TFC >

TCAS OFF

TCAS OFF

CUROT JABAN

1

1

1

RALPE

10

1

COM2

FORMAT >

125.250 1

COM2

< PRESET

ONLOE KEGE

TERR RDR

VOR1

ATC1 4336

RADAR ON UTC 14:41

RAT 15 oC

0

>

Figure 16-22. Terrain and Radar Overlay Section

TFC >

F COM2

BRT DIM

BRT DIM

<

5

125.250

125.250

44 69

BRT DIM

FORMAT <

Figure 16-21. PFD Map Format

TERR

For non-IFIS aircraft a TCAS message-only area exists below the TERR/RDR line select key. This has no active caret next to it and therefore has no control over the TCAS display (Figure 16-23). Neither of the PFD’s are

<

RDR TERRAIN

Both can also be deselected from the display and would change the respective label to white. For IFIS aircraft, a TFC line select key allows < 1 the TCAS display to be turned ON or OFF on any of the three formats. When the TCAS display is selected, TFC will be cyan. When deselected, TFC will be white. Below the TFC line is an area reserved for TCAS messages (e.g., TCAS TEST, TA ONLY, etc.) (Figure 16-22). The display of cyan TFC does NOT indicate that TCAS is actually active. TCAS is activated with a different selection discussed later in the TCAS section.

<

<

< ET 01:42 COM1 121.800

>

TERRAIN

30

DTK 251 ONLOE 1. 5NM

FORMAT > RDR >

106

FMS1

44 69

<

V1

0 able to display TCAS traffic unless they are put into a reversionary mode as will be discussed later.

<

0 reserved for detail about the selected option. For instance, if RDR is selected, the display will be cyan and the radar operating mode and F F tilt would be displayed below RDR. If TERR is selected, the display will be cyan and the appropriate operating status for the terrain would be displayed (e.g., “TERRAIN”, “TERRAIN FAIL”, “TERRAIN TEST”, etc.) (Figure 16-22).

TCAS OFF

1

COM2

125.250

BRT DIM

Figure 16-23. PFD TCAS Message Area (Non-IFIS)

Lower Display Information At the bottom of each PFD is a row of information that continuously display these items: COMM1, ATC squawk, UTC, RAT (ram air temperature) and COMM2 (Figure 16-24). Pressing the push-to-talk button on the yoke or microphone will highlight the appropriate COMM frequency label with a blue box. The


16-11

16 AVIONICS

1

1

1

MULTIFUNCTION DISPLAY (MFD)

ATC selection will show which transponder is chosen and whether that transponder is on STBY or active. It does not display the difference between ON and ALT. The RAT is derived from the currently selected air data computer. <

< ET 01:42 COM1 121.800

ATC1 4336

RADAR ON UTC 14:41

RAT 15 oC

COM2

The MFD displays engine indications, diagnostic pages, weather radar, two formats of navigation information, and terrain information. A typical MFD display is shown in Figure 16-25.

125.250

BRT DIM

The MFD has the following controls and indications:

Figure 16-24. PFD Lower Display Information

Collins

10500

ITT

ITT 516

26

TORQ TORQ

3.40

62.2 0.0

FIRE

PROP

PROP

N1 NI

1740 1980 106.0 98.5

ITT 830

734

0

TORQ TORQ

FIRE AFX

110.0 2000

0 . 0NM 0 . 8NM 4 . 4NM 198NM

RW25 ( 6 9 3 5) SXW152 KBJC

FMS

PRESS OIL

120

OIL 49 TEMP C 112 46 TEMP°C 73 o

: - : - - : CL I MB - : - - : ( 6 9 3 5) 6 9 3 5 A - :- - : - : - - / 0 . 8NM

24

DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

130 FF 750 FF 0 430 122 PRESS 80

ITT

251

W ABOVE

21

30

50

<

<

<

16 AVIONICS


25 TERR

SXW152 ( 6 9 3 5) KEGE /6935A

RDR < WX T+5.7

TFC <

F

GS

0

TAS

0

RLG /14000A SAT 15 oC

ISA +13 oC

BRT DIM

Figure 16-25. Pilot's MFD Display

16-12


Bright/Dim Rocker Switch The Bright/Dim Rocker Switch provides secondary intensity control of the MFD. The PILOT DISPLAYS rheostat, on the overhead panel, provides primary intensity control. This PILOT DISPLAYS rheostat will control all three displays: the PFD; MFD; and Control Display Unit (CDU) on the pedestal, simultaneously. Each display does not have to be individually dimmed or brightened but can be operated together. The Bright/Dim Rocker Switch will then allow each individual display to be fine tuned to make its brightness compatible with the surrounding displays.

Line Select Keys Four line select keys (LSK) are located on each side of the AFD. The keys are used in coordination with the information being viewed on the individual AFD display. LSKs that are currently active are denoted by carets (< >) displayed adjacent to the LSK.

Engine Display The engine instrument display is shown at the top of the MFD. This is called the Engine Indicating System (EIS). The EIS is always visible with aircraft power on. Refer to Chapter 7, Powerplant, of this Pilot Training Manual for more information.

MFD Window The MFD Window can display the following items: specific FMS waypoint and/or Vertical Navigation (VNAV) information; or a checklist.

Non-IFIS equipped aircraft

on the pedestal (see the CDU section for more information). The checklist information is turned ON or OFF using buttons mounted on the backside of both yokes. Once the checklist appears, the pages are advanced using the LSK’s on the left side and chosen with the SELECT LSK on the right side of the MFD. Each individual item is then “checked off” using LINE ADV buttons on the back of either yoke, or the caret line select keys on the MFD. To return to a higher level menu, press the INDEX key on the MFD (Figure 16-26). The checklist is reset when the avionics are shut down. However, if there is a need to reset the checklist without turning the avionics OFF, there is a line item on the main checklist menu page that will reset all previously “checked off” items.

IFIS equipped aircraft The FMS waypoint information must be turned ON by the left LSK on the MFD. When pressed, the UPPER FORMAT menu will appear that allows selection of the checklist, FMS-TXT or OFF (Figure 16-27) . Each repeated press of the UPPER FORMAT LSK will cycle through the options. Once the FMSTXT is chosen, the information presented is changed with the Control Display Unit (CDU) (see the CDU section for more information). The checklist can be selected either by using the UPPER FORMAT LSK described above and choosing “CHKLST”, or by using the checklist ON / OFF button on the back of either yoke. The pages are advanced using the Cursor Control Panel (CCP). For IFIS checklist operation details see the CCP section in this PTM.

The FMS waypoint information is turned ON or OFF with the Control Display Unit (CDU)


16-13

16 AVIONICS


16 AVIONICS


Collins

10500

ITT ITT

516 26

TORQ TORQ

3.40

62.2 0.0

FIRE

PROP 1900 ITT PROP 1740 ITT

830 800

N1 N1

106.0 98.5

Collins

130 FF 750 FF 0 430 122 PRESS 80 PRESS 120 0 OIL OIL 49 TEMP C 112 46 TEMP°C 73

10500

ITT

ITT 516

26

62.2 0.0

o

TORQ TORQ

AFX FIRE

110.0 2000

TORQ TORQ

3.40

FIRE

PROP

PROP

1740 1980

N1 NI

106.0 98.5

130 FF 750 FF 0 430 122 PRESS 80

ITT

ITT 830

734

0

TORQ TORQ

FIRE AFX

110.0 2000

PRESS OIL

120


FORMAT

CHECKLIST INDEX

CHKLIST

FMS-TXT

NORMAL CHECKLIST MENU ABNORMAL CHECKLIST MENU EMERGENCY CHECKLIST MENU USER CHECKLIST MENU

OFF

0 . 0NM 0 . 8NM 4 . 4NM 198NM

RW25 ( 6 9 3 5) SXW152 KBJC

: - : - - : CL I MB - : - - : ( 6 9 3 5) 6 9 3 5 A - :- - : - : - - / 0 . 8NM

RESET CHECKLIST COMPLETE HISTORY - RESET

-----------------------------------------24

DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

FMS

251

W

DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

ABOVE

21

24

251

W ABOVE

21

30

30

>

50 < UPPER FORMAT

<

50

<

LOWER FORMAT >

25 F

SXW152 ( 6 9 3 5) KEGE /6935A

RDR > WX T +5 .7

TFC >

GS

0

TAS

FORMAT PPOS

25 TERR

SXW152 ( 6 9 3 5) KEGE /6935A


0

GS

TERR

RDR < WX T+5.7

TFC <

F

ISA +13 oC

PLAN TCAS GWX

<

FMS

0

TAS

0


ISA +13 oC

BRT DIM

MFD Checklist Index

Checklist Line Advance

Collins

10500

ITT ITT

516 26

TORQ TORQ

3.40

62.2 0.0

FIRE

PROP 1980 ITT PROP 1740 ITT

106.0 98.5

Checklist ON/OFF

130 FF 750 FF 0 430 122 PRESS 80 PRESS 120 0

830 734

N1 NI

BRT DIM

OIL OIL

49 TEMP C 112 46 TEMP°C 73 o

TORQ TORQ

AFX FIRE

110.0 2000

NORMAL CHECKLIST MENU BEFORE ENGINE START ENGINE STARTING (BATTERY) BEFORE TAXI BEFORE TAKEOFF (RUNUP) BEFORE TAKEOFF (FINAL ITEMS) TAKEOFF CLIMB

1/3

----------------------------------------------FMS DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

24

251

W ABOVE

21

30

FORMAT <

Figure 16-27. MFD Upper Format (IFIS) <

50 25 SXW152 ( 6 9 3 5) KEGE /6935A

V

> F

RDR < WX T+5.7

NAVIGATION Information

TFC <

>V > SELECT GS

TERR

0

TAS

0


INDEX <

ISA +13 oC

BRT DIM

MFD Cecklist Normal Menu Checklist Line Advance

Checklist ON/OFF

Figure 16-26. Non-IFIS MFD Checklist

16-14

The following formats can be chosen for display on the MFD by pressing the top right line select key (labeled FORMAT in non-IFIS aircraft):

Plan Map Format The Plan Map Format (MAP) is used for planning/verifying the entered FMS information. It is displayed as a true north up, waypoint centered display (Figure 16-28). The Plan Map format is not intended to be used for primary navigation nor for the duration of the flight. In this mode the aircraft position may fly “off” the map since it is waypoint centered not aircraft centered. Additionally the following overlays cannot be displayed: terrain; radar; or


TCAS. For IFIS equipped aircraft with the XM weather option, this format can also overlay downloaded Nexrad radar for the 48 continuous states. Collins

10500

ITT

ITT 516

26

3.40

FIRE

PROP

PROP

1740 1980

N1 NI

106.0 98.5

ITT 830

734

TORQ TORQ

0 . 0NM 1 . 5 NM 4 . 0NM 540NM

0

FIRE AFX

110.0 2000

RW25 ONLOE RALPE KLAS

130 FF 750 FF 0 430 122 PRESS 80

ITT

PRESS OIL

The currently selected range is displayed on the edge of the range circle. This is controlled by the DCP and will be discussed later. This range will always be equal to the range displayed on the left PFD. This will limit to the following; 50nm if TCAS traffic has been selected on the left PFD; 300nm if TCAS display is OFF and overlays have been selected on the left PFD or MFD; or 600nm if no overlays or TCAS are selected on the left PFD or MFD.

120


16:24 - :- - - - :- - :- - - - :- - : - - - - : - - - - - - - LB - - . - GW

FMS

N

STORE COMPLETE

10

<

<

<

TORQ TORQ

62.2 0.0

OBLOE KEGE

CUROT

RALPE

Further display options for the FMS map display are controlled by the Control Display Unit on the pedestal (see the CDU section later in this PTM).

TERR

RDR

JABAN

TERRAIN

TFC < FATPO GS

0

TAS

0

To see an extended image beyond the range arc on the MFD, the MFD window option previously discussed can be turned OFF by using the UPPER FORMAT key (IFIS aircraft) or the CDU (non-IFIS aircraft). This will provide 50% more range above the normal navigation display.

ISA +13 oC

SAT 15 oC

BRT DIM

MFD Window ON

FMS Present Position Map Format

Collins

10500

ITT

ITT

516

26

TORQ TORQ

3.40

62.2 0.0

FIRE

PROP

PROP

1740 1980

N1 NI

106.0 98.5

130 FF 750 FF 0 430 122 PRESS 80

ITT

ITT 830

734

TORQ TORQ

0

FIRE AFX

110.0 2000

The FMS Present Position (PPOS) map is a moving pictorial of the flight. The map is centered on the airplane present position with the current heading at the top of the display.

PRESS OIL OIL

120

49 TEMP C 112 46 TEMP°C 73 o

JESIE

FMS

To see an extended image beyond the range arc, the MFD window previously discussed can be turned OFF, either by using the UPPER FORMAT key (IFIS aircraft), or the CDU (non-IFIS aircraft). This provides 50% more range above the normal navigation display similar to the Plan Map Format discussed earlier.

N

10

<

<

< OBLOE KEGE

CUROT

RALPE

TERR

RDR

JABAN

TERRAIN

TFC < FATPO GS

0

TAS

0

SAT 15 oC

ISA +13 oC

BRT DIM

MFD WIndow OFF

The current range is displayed on the two concentric range arcs, controlled by the DCP. The displayed range will always be equal to the ranges displayed on the left PFD. This will be limited to 50nm if TCAS traffic has been selected on the left PFD; 300nm if TCAS display is OFF and overlays have been selected on the

Figure 16-28. MFD Plan Format


16-15

16 AVIONICS


left PFD or MFD; or 600nm if no overlays or TCAS are selected on the left PFD or MFD.

Collins

10500

ITT

ITT

51626

TCAS Information

TORQ TORQ

FIRE

3.40

PROP PROP

1740 1980

N1 NI

106.0 98.5

0 . 0NM 0 . 8NM 4 . 4NM 198 NM

0

FIRE AFX

24

251

W

ABOVE BELOW

SXW152

21

120

OIL OIL

49 TEMP C 112 46 TEMP°C 73 o

TORQ TORQ

: - : - - : CL I MB - : - - : ( 6 9 3 5) 6 9 3 5 A - :- - : - : - - / 0 . 8NM

FM S DTK 25 1 ( 6 9 3 5) TT G - - : - 0 . 8N M

830 734

110.0 2000

RW25 ( 6 9 3 5) SXW152 KBJC

130 FF 750 FF 0 430 122 PRESS 80 PRESS

ITT ITT

30

TCAS traffic may be displayed on a TCASonly format, or overlayed on the PPOS format. To overlay TCAS on the PPOS format, simply press the TFC line select key to turn it cyan. A TCAS message-only area will be present below this TFC key (e.g., TCAS TEST, TA ONLY, etc.).

62.2 0.0

+10 -10

The TCAS-only format can be selected by the LOWER FORMAT key or by pressing and holding the traffic (TFC) key for more than 2 seconds (Figure 16-29). The display is a 360˚ , heading up image that only shows traffic and initially displays with a 10nm scale. It does not show the weather radar, terrain, or FMS map.

5

<

< 2.5

( 6 9 3 5) /6935A KEGE

-02

<

TERR

RDR < WX T+5 .7

TFC <

F

TCAS TEST

GS

0

TAS

0

ISA +13 oC

o

SAT 15 C

BRT DIM

Collins Collins 10500

ITT

ITT 516

10500

ITT

ITT 516

26

TORQ TORQ

3.40

62.2 0.0

FIRE

PROP

PROP

N1 NI

1740 1980 106.0 98.5

ITT 830

734

TORQ TORQ

110.0 2000

26

130 FF 750 FF 0 430 122 PRESS 80

ITT

0

FIRE AFX

49 46

PRESS OIL OIL TEMP°C TEMP oC

TORQ TORQ

120

FIRE

3.40

112 73

62.2 0.0

PROP PROP

1740 1980

N1 NI

106.0 98.5

0 . 0NM 0 . 8NM 4 . 4NM 198 NM

24

DTK 25 1 ( 6 9 3 5) TT G - - : - 0 . 8N M

251

734

TORQ TORQ

0

FIRE AFX

o

251

W

ABOVE BELOW

SXW152

21

120

PRESS OIL

OIL 49 TEMP C 112 46 TEMP°C 73

: - : - - : CL I MB - : - - : ( 6 9 3 5) 6 9 3 5 A - :- - : - : - - / 0 . 8NM

FM S FMS

ITT 830

110.0 2000

RW25 ( 6 9 3 5) SXW152 KBJC

130 FF 750 FF 0 430 122 PRESS 80

ITT

30

+10 +10

-10 -10

5

<

<

<

< 2.5

<

10

-02

TERR RDR

( 6 9 3 5) / 6 9 3 5 A +02 KEGE

TERR

RDR < WX T+5 .7

TFC <

F

TERRAIN

<

16 AVIONICS


TCAS TEST

TFC < GS GS

0

TAS

0

SAT 15 oC

0

TAS

0

SAT 15 oC

ISA +13 oC

BRT DIM

ISA +13 oC

BRT DIM

Figure 16-30. TCAS Figure 16-29. MFD TCAS only

Either selection will depict nearby transponder-equipped airplanes who are in close proximity or who are predicted collision threats (Figure 16-30). There can be up to 30 traffic indications on the display at one time.

16-16

The TFC line select key is only a display selection and does not actually turn ON the TCAS unit. This must be accomplished with a separate procedure (see the TCAS section of this PTM).


Graphical Weather (IFIS equipped aircraft only)

16 AVIONICS


DISPLAY CONTROL PANELS (DCP)

Another possible format is the dedicated graphical weather page. The options available here depend on the chosen weather provider. See the aircraft documentation and the IFIS section of this manual for more information.

Display control panels are vertical panels located adjacent to each PFD (Figure 16-32). The DCP and the bezel mounted line select keys on each PFD provide the primary pilot interface to control the flight displays. The left display control panel (DCP 1) provides control for PFD 1 and the MFD. DCP 2 controls only PFD 2. All menus and pages controlled < by the DCP will “time out” after 10 seconds if there is no activity. This will return the PFD to the main display.

Lower Display Information At the bottom of the MFD is a line of information that always contains the following items: GS, TAS, SAT, ISA (Figure 16-31). The Ground Speed (GS) indication is derived from the FMS. Should the FMS fail, the GS indication will be removed. True Airspeed (TAS), Static Air Temperature (SAT) and ISA deviation (ISA) are all derived from the ADC. Should the ADC fail, these indications will be removed.

<

GS

0

TAS


0

ISA +13 oC

BRT DIM

Figure 16-31. MFD Lower Display Information


CABIN DIFF HI

MASTER CAUTION

DISCHARGED

L BLEED FAIL

R FUEL PRES LO

ES PR S

R OIL PRES LO

O T S E

R BLEED FAIL

ENG FIRE

EXTINGUISHER PUSH


DISCHARGED

MASTER CAUTION


MASTER WARNING

PRESS PRES PR ESS TO TO R RESET ESET

SH

H

SY C N

AP

CRS2

YD/AP DISC

CPL

PPRESS RES E S TO TO RE RESET R SET

FD

PUSH

IR EC

T

T

ALT A LT

PUSH

MAC

YD

D

IA

D

S/

UP

A LT ALT

1/2 BANK

PUSH

IR EC

APPR

HDG

SPEED

VNAV VNA AV

PUSH

HDG

NAV N AV

CAN

CRS1

FLC

DOWN

PU

VS

EL

FD

C

ESET RESET OR SS TTO RESS RE PPRESS

Collins

Collins


W

20:03

R RAT AT

1 °C

COM2

F TIL TILT LT

121.90 BRT DIM

GS

0

TAS TAS

0

ON

ON OFF

LEFT

12 °C

ISA

TILT TILT

TAXI TAXI

ICE

NAV NAV

RECOG

BAT B AT

ARM

PROP TEST GND IDLE STOP ST OP GOV

PILOT BRAKE DEICE

HI

OFF

COM1

SURF FACE SURFACE DEICE SINGLE SINGLE

FUEL

MANUAL

DN

VENT DOWN LOCK REL

LEFT STALL STALL WARN WARN

BEACON

STROBE

PITOT PITOT

RIGHT LANDING GEAR

GEAR DOWN


NOSE E

HD LT LT TEST

L

EMER FREQ

ADC 2

1

RMT TUNE NORM

TUNE CDU

2

RTU

NORM

121 121.5 21.5

DG FREE

SLEW –

NORM

AT ATC C

OFF

L GEN TIE OPEN

HYD FLUID LOW

RVS NOT READY

R GEN TIE OPEN

R DC GEN

L NO FUEL XFR

BAT TIE OPEN

DUCT OVERTEMP

R NO FUEL XFR

R CHIP DETECT

L ENG ICE FAIL

L FUEL QTY

ELEC HEAT ON

EXT PWR

R FUEL QTY

R ENG ICE FAIL

L BL AIR OFF

AUTOFTHER OFF

HYD FLUID SENSOR

LEFT

BLOWER

RIGHT

FUEL CROSSFEED

L BK DEICE ON

MAN TIES CLOSE

CABIN ALTITUDE

LDG/TAXI LIGHT

UP

RELAY RELA Y

TEST

FLAPS

RUD BOOST OFF

R BL AIR OFF

R ENG ANTI-ICE

R IGNITION ON

PROP GND SOL L ENG ANTI-ICE

WING DEICE

2 OFF

OXY NOT ARMED

L PITOT HEAT

L PROP PITCH

20

1


60

DOWN

80

.5

PASS OXYGEN ON

2

4


0

6

.5

1

2

2

DME

2

1

ADF

TE RR TERR RD RDRR

UTC

SPKR

2

MKR

TE RR TERR

R RAT AT

o

C


COM2


IINPH NPH

NORM

GND COM

O ON N

+

DISABLE DISABLE

+

TEMP

4


AUT O AUTO

40 35 30 25

100 0F

0

T AL

5

20

3 4


MAN HEAT HEAT

ELEC HEAT HEAT

DECR


LOW

2

6

MODE

IINCR NCR TEMP

PSI

0 VACUUM V ACUUM

OFF


OFF




ENG FIRE TEST DET

OFF

0

50 80 ÛÛ) ) 100

FLIGHT HOURS 1/10

CABIN AIR

500 0 USE NO OIL

1000

1500 2000 PSI


OFF

T

5

1 7


BLOWER

10

3456


MAN TEMP INCR INCR WINDOW WINDOW DEFOG

OFF AUT O AUTO

35k

MAN COOL

R PITOT HEAT


15 5k 15k

+

NORM


L DC GEN L CHIP DETECT

L IGNITION ON

R

AUT O AUTO COMM

BRT DIM

UP


COPILOT COPILOT

OFF TEST

AHRS

NORM

$87 2


AUT OFEATTHER AUTOFEATHER

RIGHT


MAIN

OPEN

2))

OFF

PARKING P ARKING BRAKE

LEFT ACTUATORS ACTUAT TORS STANDBY ST ANDBY

2))

TEST IGNITION AND ENGINE START START ENGINE LEFT RIGHT ON

LDG GEAR CONTROL

NAV NAV

1

Collins

1

MFD NORM

1

2))

L GEN R GEN BUS SENSE GEN TIES RESET MAN CLOSE 1250

ESIS ON

OFF TEST

OFF

OFF

2

< ET

RANGE USH A AUTO UTO TILT

PILOT DISPLAY Y PFD

STBY

+

FORMAT FORMAT <

FMS F


ON EMER OFF

< PRESET

+13 °C

ATC A TC 1

SLEW –

5

2))

ENGI NE ANTI-ICE ENGINE LEFT RIGHT ON

350.0

TFC

SSAT AT

NORM

+

/,*+76

GEN RESET

RIGHT

/,*+76

BATT BUS NORM

DG FREE

ARM

ENG AUT O AUTO IGN OFF

RADAR

BRT DIM

PROP SYNC


OFF–RESET

1/2

GCS

Collins

EXT PWR

DME-H

TA TA ONLY ONLY

J10-1


P

UTC

PA

29.88IN 29.88I N

3

12

051 05111

2

E

ONLY TTA A ONL Y

ATC1 AT TC1

T

COMM

N

33

108.50

TERR RDR

J JNETT JN

/82 /8215A

GCS

1

NORM

G GS

TCAS F AIL FAIL

HDG

HDG HDG

LOC1

DATA DATA

ELEC

NA NAV/BRG AV/BRG

ADF

11 051 1

( NTC) (INTC) C) (8700) 87 ((8215) 82 KASE K AS ASE ))

TERRAIN

1118.85 18.85

ATC A TC 1

S MIC MIC OXY

LOC

6

ET COM1

T IDENT H

12.5

11 1 13.00

GLENO G LENO E LI ONDZ LINDZ DBL /16000A 0 0A 000

RA

PUSH

IDENT

123.80

NAV1 NAV1

3

50

TERR RDR

VOR1

VOICE B O

DC P DCP

COM1

11 11 8 . 8 5

MENU ADV

B BRT

Collins

N

30

RADAR

H

NORM

((8215)

TTG --:-1.4NM

FORMA AT FORMAT

12.5 12.5

V 13.6NM DBL F

AC ACC C .–– CRS 057

3 9 329

HDG 329

VS

PULL UP GND PROX

ENG2 EN G2 REFS

S

S

AUDIO ALTN ALTN

INPH

FMS

DATA DATA PUSH

ELEC

ALT A LT G GPWS

ENG1 EN G1

XMIT XMIT PA

1

VOL

IAS

XAHS XA HS XADC XA DC

N

NAV/BRG NA AV/BRG

SPKR

XTLK

STD

7500 0MIN

33 3 M

MENU ADV

PRESET MKR

78 20 0 0 00

10

10

30 0

2 ALT A LT

ATT AT T

FD

PUSH

8000

10

10

1 40 9

8215A

-:--/ 1.4NM

0KTS

138 38 082 82 2

25 25

AUTO AUTO COMM

2

0.6NM 1.4NM -:-- : CLIMB 2.6NM -:-- : (8215) 169NM -:-- :

REFS

3

ADF

60

KASE ((8215)) (8700) KCOS

4

A TT V ATT N V

IAS

BARO

30

N

ADC2 13.6NM 29.92 in

VOR1 057CRS

80

w

300

AHS2

OIL

49 TEMP°C 11 1122

FIRE

24

329

TERM

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830

PUSH

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2

600

HDG 329 (8215) 0.8NM

130 FF 750 122 PRESS 80

ITT

TORQ TORQ

110.0 110.0

T

PA

106.0

A

P

2

1740

N1

J206

2

PROP

62.2 3.4

W

NAV NAV

DME

1050

ITT

1

700

AP AP

TRIM TR IM

T

1

1

10

0

N350KA

516

1

40

V2 107 VR 103 V1 100 AC ACC–.03 C –. 03

RADIO CALL

TTORQ ORQ

800 00

MIC OXY

2

Collins

BARO

7820 COMM

NORM

1

FLAP OVRD D

ACTIVE A CTIVE

2

900

S 1

TERR INHIB B

A ACTIVE

STEEP APPR PR

ACTIVE A CTIVE

21

VOL

10

G/S INHIB IB

A ACTIVE

4

8 000

20 60

T

16000

170

80

Collins Colli ns

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S

1

PTCH ALTS ALTS

15

HDG FMS

XMIT 2 PA

<

MASTER WARNING


T

L OIL PRES LO

ENG FIRE

T

L FUEL PRES LO

EXTINGUISHER PUSH

AUTO AUTO


INCR INCR

EXT

10

15

Figure 16-32. Display Control Panels


16-17

The two versions of the DCP (IFIS and nonIFIS) are shown in Figure 16-33. (Information for Weather Radar controls are found in this chapter).

low underline will appear when the altimeter settings are different by more than 1 hPa. The range for this mode is 745 to 1100hPa.

1 4 000

1

6935

4

700

2 1

600 60 6 5 40 20 1

400

2 4

3 0 .1 6 I N

Figure 16-34. Barometric Setting with Yellow Underline

non-IFIS

IFIS

Figure 16-33. Display Control Panel (DCP)

<

16 AVIONICS


BARO Knob Rotating the BARO knob adjusts the altimeter setting for the on-side altimeter. The current altimeter setting is displayed below the PFD altitude scale. Altimeter settings are independent for each side and a yellow underE line will appear below the altimeter setting when they are different by more than .02”Hg (Figure 16-34). Single pilot operations will require a manual setting of each DCP barometric knob. The altimeter setting has the range of 22.00 to 32.50”Hg. In flight regions where the barometric setting is given in hPa this setting can be changed. For IFIS aircraft, the DCP is used to change the units for the barometric setting using the REFS button. In non-IFIS aircraft a switch labeled IN/hPa located on the overhead panel, and can select between inches of Hg and hPa (Figure 16-35). When using hPa units, the yel-

16-18

Figure 16-35. IN/hPa Switch

BARO PUSH STD Button When pushed, the standard altimeter setting QNE is selected and “STD” will be displayed in lieu of the pressure setting. The cyan preselect altitude above the altitude display will display a flight level (FL) format when this button is pushed (e.g., 22,000 will be displayed as FL220; 8,000 will be FL80) (Figure 16-36 ) . To return the setting to normal units, turn the Baro Knob and select the new altimeter setting.


minimums (RA MINS), and MDA/DA minimums (BARO MINS) shown on the PFD.

FL250 4

700

Menus are controlled with the knob at the center of the DCP (Figure 16-33). For IFIS aircraft, there are two concentric knobs labeled MENU ADV and DATA. The PUSH SELECT feature of the DATA knob will enter data or choose items from the avionics selections.

2 1

600 60 6 5 40 20 1

400

2 4

STD

Figure 16-36. Barometric Setting with STD

REFS Button The REFS button will bring up a menu on the respective PFD (Figure 16-37). <

Collins

HDG FMS

1 4 000

80

6935

VT is a general purpose “target” speed that is not affected by the takeoff related V-speeds.

4

20

700

2 1

10

600 60

6 540 20 V2

TERM

24

FMS1

251

REFS 1/2 RA MIN < <

<

30

<

30.16IN

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50

160 <

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144 069

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21

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400

0

106 V1 ACC-.02

200

25

V2

BARO MIN <

117

2980

VR

VREF <

110

V1

106

For IFIS equipped aircraft, the values are set by placing the cyan box cursor around the desired label. This can be accomplished by pressing the adjacent line select key on the PFD or by rotating the MENU ADV knob until the cursor covers the desired value. Once the cursor is set, rotate the DATA knob to set the desired value. To move to the next item, repeat the steps listed above.

1

10

117

160 ATC1 4336

RADAR ON UTC 14:41

The left side of the menu contains V-speeds. Beginning from the bottom, the pilot’s can set V1, VR, V2 and VT. Speeds will show up on both PFD’s so only one pilot needs to set the values. Additionally, the setting of one value will affect the remaining values in this relationship: V2 ≥ VR ≥ V1.

PTCH ALTS

140

60

For non-IFIS aircraft, there is a single knob labeled MENU ADV. This knob has a button labeled PUSH MENU SET that will enter data or choose items from the avionics selections.

RAT 15 oC

BRT DIM

Figure 16-37. PFD REFS Menu Page 1 of 2

REFS Page 1 With this menu, it is possible to control the display of selected V-speeds, radio altitude height

For non-IFIS equipped aircraft, the values are set by placing the cyan box cursor around the desired value to be changed. This can be moved by pressing the adjacent line select key or by rotating the MENU ADV knob. This cursor must flash to indicate the value is settable. If the cursor was moved by pressing the adjacent line select key on the PFD the cursor will automatically begin flashing. If the cursor was moved with the MENU ADV knob then the PUSH MENU SET button must be


16-19

16 AVIONICS


pressed to get the cursor to flash. Once it is flashing, the MENU ADV knob can be used to change the value inside the cursor instead of moving the cursor. To move on to the next V-speed press the line select key next to the subsequent V-speed and rotate the MENU ADV knob to change the value. Alternatively, press the PUSH MENU SET button to stop the cursor from flashing and move the cursor to the desired value with the MENU ADV knob. For both aircraft installations, these speeds must be cyan in order to be shown on the airspeed display. They will turn white (deselected) by pressing the PUSH SELECT feature of the DATA knob (or by pressing and holding the PUSH MENU SET button for non-IFIS aircraft). Once they are cyan, a list appears below the airspeed display while on the ground. The display contains all but the VT setting. V-speed settings will also appear as reference bugs on the airspeed display (Figure 16-38).

mum (BARO MIN) value and the radio altimeter minimum (RA MIN) value will be identical on both pilot’s displays. Only one pilot needs to set the values. Setting RA MIN will create a hollow bar on the altitude tape the length of the value chosen. For instance, setting 200 feet will create a bar starting from radio altitude “Zero” up 200’ on the altitude tape. Radio altitude “Zero” is the point where the altimeter changes from blue to brown (Figure 16-39) .

1 4 000 700 RAD Minimum Altitude

600 60 6 5 40 20

Radio Altitude Zero

400 Radio Altitude Minimum Setting

140

30. 16I N M I N 200 R A

80 60

V2 VR V1

Figure 16-39. Radio Altitude Minimum

The change of altimeter color is solely based off of the radio altimeter. It is not dependent on putting in the RA MIN number and will always display when the radio altimeter is operational. It would not display if the radio altimeter were inoperative. The RA MIN reference is not used as a desired minimum reference since the King Air B350 is certified only to CAT I minimums.

117 110 106

ACC-.02 TERM

Figure 16-38. PFD V-Speeds

Setting BARO MIN is the desired minimum reference altitude. This will create a cyan bar across the altitude tape at the altitude selected (Figure 16-40).

The right side of the menu contains the numbers used for landing. The barometric mini<

16 AVIONICS


16-20

FOR TRAINING PURPOSES ONLY T

16 AVIONICS


6720 GS

700 BARO Minimum Altitude

barometric minimum setting

6935

1 4 000

1

4

20

600 60 6 5 40 20

700

2 1

10

MIN

600 60 6 5 40 20 1

10

400

400

350

4

3 0 . 1 6 IN M I N 6 6 0 0 BARO

251

3 0 .1 6 IN MIN 6 7 2 0 B A R O

4

w

Figure 16-40. Barometric Minimum

Figure 16-41. Minimums Annunciator <

An additional benefit of setting BARO MIN is that the altitude preselector can be set to the exact BARO MIN value. For example, if BARO MIN is set to 1830, the preselected altitude can now be set to 1830 to allow for autopilot capture at the desired MDA. The BARO MIN can be set to the nearest ten feet of altitude.

2

REFS Page 2 (IFIS-equipped aircraft) For IFIS equipped aircraft, there is a second page to the REFS menu (Figure 16-42). This is accessed by pressing the REFS key a second time.

E

HDG FMS

PTCH ALTS

1 4 000

140

80

4

20 60

6935

700

2 1

10

600 60

6 540 20 V2

DTK 251 (6935) 0. 8NM

24

251

2 4

30.16IN

W

21

144 069

REFS 2/2 50 < PRESSURE HPA IN

400

0

TERM

FMS1

1

10

117

106 V1 ACC-.02

REFS 2/2 <

The last option on the right side of the menu is VREF. This acts just like the V-speeds discussed earlier. Once one pilot adjusts the value it will turn cyan for both pilots and will place a bug on both airspeed tapes.

Collins

30

Both RA MIN and BARO MIN will generate a “MINIMUMS” aural callout and flashing MIN annunciator on the PFDs (Figure 16-41). If the aircraft continues below the values, the RA MIN hollow bar will turn yellow or the BARO MIN altitude bar will turn yellow. The minimum reference displayed is the last one adjusted (e.g., if RA was set first and then BARO, the BARO minimums are the only ones displayed). For non-IFIS equipped aircraft each pilot can choose to display BARO MIN or RA MIN independent of the other pilot. However, if each pilot sets a different reference (one shows BARO MIN and the other RA MIN) the “MINIMUMS” aural callout will occur at the first value achieved.

25

< METRIC ALT ON OFF < FL ALERT ON OFF < FLT DIR

V-BAR X-PTR ATC1 4336

RADAR ON UTC 14:41

RAT 15 oC

BRT DIM

Figure 16-42. PFD REFS Menu Page 2 of 2


16-21

The PRESSURE option allows the altimeter setting units to change from HPA (hectopascals) to IN (inches of mercury). This will affect both pilots and cannot be set independently. It does not affect the standby unit which will have to be adjusted separately. The METRIC ALT selects the display of metric altitudes ON or OFF above the altimeter display (Figure 16-43). This setting does not change the feet presentation on the actual altimeter tape. This action will affect both pilots displays and cannot be set independently. METRIC

2450M

6935

4000M 4

200

2

0

10

10

10

10

20

V-BAR

0

20

X-PTR

Figure 16-44. Flight Director Formats

For non-IFIS equipped aircraft this page 2 does not exist but most of the features are accessed with external switches located on the overhead panel. How they affect the PL21 system is discussed in the altimeter section of the PFD.

1

MENU ADV Knob (IFIS)

100 60 8 0 40 20

The MENU ADV knob moves the menu cursor around the displays.

1

900

2

DATA Knob (IFIS)

4

The DATA knob will change the value inside the menu cursor.

1018HPA

Figure 16-43. Metric Altitude

The FL ALERT turns the advisory flashing of altimeter setting ON or OFF. The setting will flash when passing through transition altitude 18,000’, or transition level FL180. A change of the altimeter setting or pressing the center STD button will stop the advisory flashing. This transition level trigger cannot be changed to a value other than 18,000’.

PUSH MENU SET (IFIS) The PUSH MENU SET feature will enter or accept selected items in the menu cursor.

<

16 AVIONICS


Finally, the FLT DIR line will change the flight director image changing it from a v-bar presentation to a cross-pointer (X-PTR) presentation (Figure 16-44). This change will affect both pilots and cannot be set independently.

MENU ADV Knob (non-IFIS) The MENU ADV knob accomplishes two tasks. When the menu cursor is flashing, this knob is used to change the value inside. When the menu cursor is not flashing, this knob is used to move the cursor around the display to position it on another item.

PUSH MENU SET (non-IFIS) The PUSH MENU SET button will start the menu cursor flashing on the first press. The sec-

16-22


ond press will enter the information and stop the cursor from flashing. This will also change the value of items where they are just two options inside the cursor (e.g., ON / OFF).

NAV/BRG Button Pressing the NAV/BRG button displays the NAV SOURCE and BRG SOURCE menus on the PFD (Figure 16-45). The navigation source (NAV SOURCE) section is on the left side of the menu and allows selection of the appropriate active navigation source. Each press of the left line select key will cycle the options. For IFIS aircraft the DATA knob on the DCP will also cycle the options. On nonIFIS aircraft the cursor can be placed with the MENU ADV knob and then press the PUSH MENU SET button to select the appropriate navigation source. Caution must be used when manipulating this NAV SOURCE because it will immediately change the active navigation display. Collins

VT

20

300

DN

10

200

4 2

14 1 0

1

20

5100 80

120

10

1

000

2

100

900

4

300

33

W

S

DTK 301

24

FMS1 FMS2 LOC1 VOR2

N

<

OFF FMS ADF1

25

ATC1 1200

UTC 16:42

RAT - 4 oC

F

BRG SOURCE <

<

ICT 4.1NM NAV SOURCE

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FMS 1

30.16IN MIN 200 RA

301

17

VOR1

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V 4.1NM SXW V ----NM SXW

< ET

BRT DIM

COM1 121.800

12

160

1 Once the bearing pointers are chosen, an information area will appear on the bottom left corner of the PFD (Figure 16-46). The following labels are possible: V (VOR); F (FMS); A (ADF). Below the “V” will appear the frequency of the VOR. If DME is available, the station identifier will replace the frequency once the identification is received from the DME. Additionally, the DME to the station will appear next to the “V”.DME information will not display if the radio is on DME hold or the active navigation source is the same VOR. In both cases the DME will appear up by the active navigation source.

E

180

3200

5000

The bearing source (BRG SOURCE) section is on the right side of the menu and allows selection of the appropriate bearing pointers. Two pointers can be displayed; a magenta single-needle pointer; and a cyan double-needle pointer. The magenta needle will only point to the #1 navigation systems (e.g., VOR1, ADF1, FMS1). The cyan needle will only point to the #2 navigation systems (e.g., VOR2, ADF2, FMS2). The exception is when there is only one FMS installed. In this case, both needles can be selected to that single FMS. Selection is accomplished by pressing the appropriate line select keys. These selections are independent for each pilot. For IFIS aircraft, the DATA knob will also cycle the options. For non-IFIS aircraft the cursor can be placed with the MENU ADV knob and then press the PUSH MENU SET button to select the appropriate bearing source.

FMS1 AP VPTCH A LT S

185

16 AVIONICS


ATC1 4336

UTC 1

Figure 16-45. PFD NAV BRG Menu Figure 16-46. Bearing Pointer Information


16-23

RAT

COM2

12

16 AVIONICS


The active FMS fix name and distance to that fix will appear next to the “F”. The ADF frequency will appear next to the “A”.

RADAR Button The RADAR button displays the weather radar menus on the PFD. See the Weather section of this manual.

GCS Button

temperature of each unit to eliminate sustained overheating which would cause an automatic shutdown of the respective power supply. Additionally, the power supply operation is inhibited in extreme cold temperatures below -40˚C. Each IAPS section contains the Flight Guidance Computers (FGC’s) and the Flight Management Computers (FMC’s) for the respective side.

The GCS button controls the ground clutter suppression selection of the weather radar. See the Weather section of this manual.

TILT Control The TILT knob controls the weather radar antenna tilt angle. See the Weather section of this manual.

RANGE Knob The RANGE knob controls the display range shown on the PPOS map, North-up Planning Map, and TCAS only Display. The selected range annunciations are shown on the PFD and MFD as discussed above.

INTEGRATED AVIONICS PROCESSOR SYSTEM (IAPS) The Integrated Avionics Processor System (IAPS) provides system integration and operating logic for most systems that make up the ProLine 21 avionics. This unit is installed in the nose of the aircraft in the avionics bay (Figure 16-47). It consists of two sections; the No. 1 (left) section monitors the No. 1 aircraft systems while the No. 2 (right) section monitors the No. 2 systems. Each section is powered by a dedicated power supply. Fans control the

16-24

Figure 16-47. IAPS

AIR DATA COMPUTERS (ADC) Two digital Air Data Computers (ADC 1 and ADC 2) convert raw dynamic flight data into electronic signals for use by various airplane systems (Figure 16-48). The ADC’s generate independently and are supplied with the following inputs: • Ram air pressure from the onside pitot mast • Static pressure from the static ports • Air tempe rature


ATTITUDE AND HEADING REFERENCE SYSTEM (AHRS) The Attitude and Heading Reference System (AHRS) provides pitch, bank, and magnetic heading data to the onside displays (Figure 1649).

Figure 16-48. ADC

Each ADC supplies its onside systems (the MFD is supplied from ADC 1). Reversionary switching allows use of the cross-side ADC as a backup. In the reversionary ADC mode, the selected ADC supplies all systems. Each ADC processes the data and provides electronic signals to the following systems and components: • EFIS • Displays the following information • Uncorrected Pressure Altitude • Baro-Corrected Altitude • Vertical Speed • Airspeed (KIAS & KCAS) • Indicated Airspeed Trend Vector • Mach Number • Maximum Airspeed (VMO/MMO) • True Airspeed • Ram Air Temperature (RAT) • Static Air Temperature (SAT) • ISA Deviation Temperature • Wind Direction and Speed Vector • Attitude and Heading Reference Systems (AHRS) • Integrated Avionics Processor System (IAPS)

Figure 16-49. AHRS

Magnetic heading information is obtained from separate magnetic sensors located in opposite sides of the horizontal stabilizer. Compensator units automatically correct for magnetic interference within the airplane or due to sensor error. Attitude information is obtained from two attitude and heading computers (AHC). Each system includes an inertial measurement unit (IMU) that monitors angular rates and accelerations about the airplane axes. The IMU does not provide self generated navigation position. The AHC processes IMU data to determine airplane pitch and bank attitude. Each AHC is provided with a primary and secondary power supply for redundancy. If the secondary power supply should fail, the primary power supply will continue powering the AHC. After 10 minutes of operation on primary power only,the primary power supply will cease operating. The power loss to the AHC will result in a total failure of that AHC. There will be no indication, except from a pos-


16-25

16 AVIONICS


16 AVIONICS


sible tripped circuit breaker. This indicates a failure of the secondary power supply. If the primary power supply should fail, the AHC will immediately fail. In either case, the crossside AHC may then be selected using the AHRS reversionary switch to regain AHRS information on the affected side. The output of each AHRS is supplied to the integrated avionics processor system (IAPS) for distribution to the appropriate display or component. AHRS 1 data is displayed on the pilot displays while AHRS 2 data is displayed on the copilot display. Each AHRS can provide reversionary support to the other. The AHRS switch on the reversionary control panel controls reversionary operation. Compass controls are provided for control of the slaving operations for the pilot and copilot compass systems. The controls are labeled DG–FREE–NORM and SLEW + / – (Figure 16-50) . The DG switch selects whether the respective heading is “slaved” to the compass (NORM) or acting as an unslaved, free unit (FREE). When the FREE Mode is selected, the pilot can manually adjust the heading by moving the SLEW switch to either the + or – position.

Figure 16-50. Heading Slave and Slew

REVERSIONARY OPERATIONS AFD Reversion The pilot’s PFD and the MFD are designed to provide reversionary support to each other in the event of a single display failure. Reversionary display switching for the pilot’s PFD or the MFD is accomplished via the PILOT DISPLAY switch on the reversionary control panel (Figure 16-51). Selecting the remaining AFD will display a composite image. When an AFD fails a XTLK annunciator will appear on the remaining display. This indicates that the other displays have lost communication with the failed display. This helps identify

Figure 16-51. AFD Reversions

16-26


that an actual display failure has occurred, not a brightness control problem.

functions should remain normal and unaffected.

The selection of PFD or MFD is always made toward the unit that is still functional. (e.g., if the PFD is still operating, select PFD) If the PFD position of the PILOT DISPLAY switch is selected, the composite display will appear on both the pilot and copilot PFDs. Selecting the MFD position of the switch will result in the composite display appearing on only the MFD (Figure 16-52). When selecting reversionary modes, all flight director and autopilot

ADC Reversion The Air Data Computer (ADC) switch on the reversionary control panel provides reversion capabilities for the ADCs. If a single ADC fails, the red IAS, ALT, and VS failure flags will appear on the affected PFD and a white XADC flag will appear on the cross-side PFD (Figure 16-53). The ADC switch should be

Collins

1050 0

ITT ITT

516 26

TORQ TORQ

62.2 0.0

FIRE

0 3.4

PROP PROP

1740 1980

N1 NI

106.0 98.5

830 734

TORQ TORQ

AFX FIRE

2000 110.0 PTCH ALTS

80

516 26

TORQ TORQ

700

FIRE

1740 1980

N1 NI

106.0 98.5

130 FF FF 750 0 430 122 PRESS PRESS 120 80 0 OIL OIL TEMP oC 112 49 TEMP°C 46 73

ITT ITT

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TORQ TORQ

AFX FIRE

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PTCH ALTS

4

20 60

700

600 60

600 60

30. 16I N

DTK 251 (6935) 0. 8NM

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< PRESET

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TERR RDR

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144 069

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FORMAT <

2 4

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400

0

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6 540 20 1

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0

2 1

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6 540 20 117 V2 110 VR 106 V1 ACC-.02

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140

80

2 1

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PROP PROP

62.2 0.0

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20 60

1050 0

ITT ITT

0 3.4

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Collins


ITT ITT

HDG FMS

Collins

N

< ET

COM1 121.800

BRT DIM

12

ATC1 4336

UTC 14:41

BRT DIM

BRT DIM

PILOT DISPLAY Switch - PFD Selected Collins

Collins

1050 0

0 3.4

1740 1980

N1 NI

106.0 98.5


ITT ITT

830 734

TORQ TORQ

AFX FIRE

2000 110.0 HDG FMS

1 4 000

80

700

60

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4

2 4

30. 16 I N

251 w

24 21

S

21

FORMAT < TERR

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ABOVE

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ATC1 4336

UTC 14:41

RAT 15 oC

FORMAT <

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144 069

TERR RDR

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400

33

COM1 121.800

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DTK 251 (6935) 0. 8NM

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< ET

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< PRESET

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1

30

DTK 251 (6935) 0. 8NM

251

1 2

2

600 60 6 540 20

30

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400

0 24

700

10

117 V2 110 VR 106 V1 ACC-.02

600 60

6 540 20

TERM

4

2 1

10

6935

1 4 000

4

20

117 V2 110 VR 106 V1 ACC-.02

PTCH ALTS

20

PTCH ALTS

140

60

HDG FMS

140

80


TFC <

SAT 15 oC ISA +15 oC COM2 125.250

< ET

COM1 121.800

BRT DIM

12

FIRE

PROP PROP

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TORQ TORQ

62.2 0.0

ATC1 4336

6

516 26

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ITT ITT

Collins

<

UTC 14:41

TFC >

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<

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COM1 121.800


TFC <

SAT 15 oC ISA +15 oC RAT 15 oC COM2 125.250

6

< ET

ABOVE

25

3

F

UTC 14:41

RAT 15 oC

COM2

125.250

BRT DIM

PILOT DISPLAY Switch - MFD Selected

Figure 16-52. Reversionary Modes


16-27

16 AVIONICS


16 AVIONICS


Collins

HDG FMS

Collins

PTCH ALTS

HDG FMS

PTCH ALTS

6935

14 000

140

80

20 FD

60

10

IAS

ALT

XADC

0

TERM

24

FMS

251

600 60 6 540 20 1

10

117 V2 110 VR 106 V1 ACC-.02

400

0

4

24

w 30

21

S

30

DTK 251 (6935) 0. 8NM

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FMS 144 069

2 1

10

TERM

W

21

DTK 251 (6935) 0. 8NM

700

VS

10 ACC-.--

4

20 FD

<

RADAR ON UTC 14:41

RAT oC

COM2

E

ATC1 4336

12


TCAS OFF

< ET

125.250

COM1 121.800

ATC1 4336

TFC >

6

COM1 121.800

TERR RDR

VOR1

3

F

< ET

< PRESET

<

RDR TERRAIN

VOR1

N

TERR

< PRESET

FORMAT >

15

25

<

FORMAT <

33

50

UTC 14:41

RAT 15 oC

COM2

125.250

BRT DIM

BRT DIM

Pilot’s PFD

Copilot’s PFD

Figure 16-53. ADC1 Failure

moved to the operating ADC (e.g., if ADC1 is still working, choose ADC1). Miscompare indications also require the use of ADC reversion. This occurs when the pilot and copilot systems are still functional but have different values displayed on the PFD’s. Yellow IAS, ALT and VS flags will appear on both PFD’s (Figure 16-54). The pilots must determine which system is correct and choose the operating ADC. Once the operative ADC has been selected, a yellow-boxed ADC1 or ADC2 flag will appear on both PFDs indicating they are both using the same ADC. (Figure 16-55). When using the reversionary mode, normal flight director and autopilot functions will return when the flight

guidance computer is coupled to the operating ADC. See the Flight Guidance section of this manual for the method of coupling to each side. Collins

HDG FMS

PTCH ALTS

140 80 IAS

20

60

V2 VR V1

1 4 000 ALT

HDG 010

1

1

400

0 24

251

2

600 60 6 540 20

10

ACC-.02 TERM

4

700

10

117 110 106

6935

2 4

30.16IN

W

Figure 16-54. ADC Miscompares <

R

T

16-28

F


16 AVIONICS


AHRS Reversion The Attitude Heading Reference System (AHRS) switch on the reversionary control panel provides reversion capabilities for the AHRS. If a single AHRS fails, the red HDG and ATT flags will appear on the affected PFD and a white XAHS flag will appear on the cross-side PFD (Figure 16-56). The AHRS switch should then be moved to the operating AHRS (e.g., if AHRS2 is still working, choose AHRS2).

140

ADC2

80 60

117 V2 110 VR 106 V1 ACC-.02 TERM

Figure 16-55. ADC Switch - ADC2 Selected

Miscompare indications also require the use of AHRS reversion. This occurs when the pilot and copilot systems are still functional but have different values displayed on the PFD’s. Yellow HDG and ATT flags will appear on < both PFD’s (Figure 16-57). The pilots must de<

T

Collins

HDG FMS

PTCH ALTS

HDG FMS

6935

14 000

140

Collins

80 60

0 24

FMS

4

W

1

400

0

4

24

w 30

21

S

30

DTK 251 (6935) 0. 8NM

2

30.16 I N

251

FMS

21

DTK 251 (6935) 0. 8NM

600 60 6 540 20

10

TERM

2 1

10

117 V2 110 VR 106 V1 ACC-.02

2

30.16IN

HDG

700

XAHS 1

TERM

FD

60

1

400

4

20

2

600 60 6 540 20 117 V2 110 VR 106 V1 ACC-.--

6935

14 000

80

4

700

ATT

PTCH ALTS

140

<

ATC1 4336

RADAR ON UTC 14:41

RAT 15 oC

COM2

< ET

125.250

COM1 121.800

12

E


TCAS OFF

ATC1 4336

UTC 14:41

RAT 15 oC

BRT DIM

Pilot’s PFD

TFC >

6

COM1 121.800

TERR RDR

VOR1

3

F

< ET

< PRESET

<

RDR TERRAIN

VOR1

N

TERR

< PRESET

FORMAT >

15

25

<

FORMAT <

33

50

COM2

125.250

BRT DIM

Copilot’s PFD

Figure 16-56. AHRS1 Failure


16-29

termine which system is correct and choose the operating AHRS. Collins

HDG FMS

PTCH ALTS

140

ATT

80

14 000

4

20 60

6935

700

2

Figure 16-58. Pitot Tubes

1

10

600 60

6 540 20 V2 VR V1

117 110 106

ACC-.02

FMS1

2 4

30.16IN

W

21

144 069

30

DTK 251 (6935) 0. 8NM

400

0 HDG 24 251

TERM

Each heated mast provides ram air pressure to its respective Air Data Computer (ADC). The copilot’s mast also provides ram air pressure to the Electronic Standby Instrument System (ESIS) ADC.

1

10

Figure 16-57. AHRS Miscompares <

16 AVIONICS


R

Once the operating AHRS has been selected, a yellow-boxed AHS1 or AHS2 flag will appear on both PFDs indicating they are both using the same AHRS. T

If the Attitude portion of the AHRS fails, then the autopilot will automatically disengage and cannot be reengaged until the AHRS is repaired by maintenance. If only the heading portion has failed, the autopilot will remain engaged. If the heading failed on the side that is coupled to the flight director or autopilot, there will be limited lateral control and it is recommended to select the operating AHRS or couple to the unaffected side. See the Flight Guidance section of this manual for the method of coupling to each side.

Dual static ports are located on each side of the aft fuselage in a vertical arrangement (Figure 16-59). The top port on the left side is connected to the bottom port on the right side and the resulting average pressure is supplied to the pilot’s static air source valve, located just below the right side circuit breaker panel. The other two static ports are also connected and the resulting average pressure is supplied to the copilot’s ADC. The copilot does not have an alternate static source selection. The copilot’s static source is also attached to the ESIS ADC. The static ports are not heated as they are in a position that does not accumulate ice.

PITOT AND STATIC SYSTEM Independent pitot and static systems are provided for the pilot and copilot flight indications. The pilot and copilot pitot masts (Figure 1658) are located on the forward lower nose section of the airplane.

16-30

Figure 16-59. Static Ports

In addition, an alternate static air source is provided to the pilot’s static air source valve from the aft side of the rear pressure bulkhead. The output from the pilot’s static air source valve is manually selected by the crew and


provides either normal static air pressure or alternate static air pressure to the pilot’s ADC. During preflight, the pilot should ensure the PILOT’S STATIC AIR SOURCE valve switch is held in the NORMAL (forward) position by the spring-clip retainer (Figure 1660). See Figure 16-61 to see the connections from pitot-static lines to the ADC’s for pilot and copilot and the ADC for the ESIS. I

ENVIRONMENTAL

WSHLD

10

25

EQMT COOLING FLT INSTR

RIGHT

DEICE CONTROL

WIPER

COM

ANT

LEFT

LEFT

SURF

SELCAL

PILOT PFD

5

5

5

1

OXY

NOSE

71 2

5

5

5

5

1

BUS TIE POWER

BLEED AIR CONTROL

71 2 BUS TIE POWER

71 2

CABIN ALT

5

5

5

2

RIGHT

HIGH

HIGH

PNL COOLING

FURNISHING MASTER

CIGAR

DC

ANTI ICE

5

PRESS

CONTROL

P

5

TEMP

CABIN DIFF

MN ENG ANTI ICE

PROP

LEFT

DC

FUEL VENT

STBY ENG ANTI ICE

HF

DEICE IEC

71 2

10

5

15

15

5

5

5

LIGHTER

CONV2

CONV2

RIGHT

RIGHT

DEICE

10 HEATER

BRAKE

CONTROL

ADC2

5

2

M

1 HEA

COPILOT PFD

10 HEATER

PILOT'S STATIC AIR SOURCE

NORMAL ALTERNATE

SEE FLIGHT MANUAL PERFORMANCE SECTION FOR INSTR CAL ERROR

Figure 16-60. Alternate Static Source Selection

Selecting the alternate static source will induce errors in altitude and airspeed indications and should only be selected when the normal static source is blocked.

WARNING The pilot’s airspeed and altimeter normal indications are changed when the alternate static air source is in use. Refer to the Airspeed Calibration – Alternate System, and the Altimeter Correction – Alternate System graphs in the POH/AFM (PERFORMANCE Section) for operations when the alternate static air source is selected.

16 AVIONICS


OUTSIDE AIR TEMPERATURE The digital outside air temperature (OAT) gage is located on the left sidewall, and displays Indicated Outside Air Temperature (IOAT) in Celsius (Figure 16-62). When the adjacent button is depressed, Fahrenheit is displayed. The probe is located on the lower fuselage under the pilot’s position. Indicated Outside Air Temperature (IOAT) is a combination of Static Air Temperature (SAT) and temperature due to air friction across the probe. This is referred to as Ram Air Temperature (RAT) or Total Air Temperature (TAT). For determination of actual OAT, refer to the Indicated Outside Air Temperature Correction – ISA chart in the Performance section of the POH/AFM. This sidewall OAT gage must be used for performance computations. The Ram Air Temperature (RAT) and Static Air Temperature (SAT) indications are located at the bottom of the PFD and MFD respectively. Information is derived from the Air Data Computers. This input comes from a Rosemont probe located behind the nose gear well area on the underside of the fuselage. This is an unheated probe as is the OAT gauge probe (Figure 16-62). The term ambient temperature, when used for Engine Anti-ice operations, refers to IOAT corrected for ram air temperature as found in the above listed correction chart in the POH.


16-31

16 AVIONICS


RAT TEMPERATURE PROBE

L PITOT MAST

R PITOT MAST

NUMBER 1 UNITS

NUMBER 2 UNITS

FGC FMC

FGC FMC

AHRS

(Optional)

IAPS

IAPS

ADC DRAIN

AHRS

ADC DRAIN

FWD PRESSURE BULKHEAD STANDBY UNIT PILOT PFD

COPILOT PFD

PILOT MFD

DRAIN DRAIN

CABIN PRESSURE CABIN DIFFERENTIAL PRESSURE GAGE

LEFT STATIC PORTS

PNEUMATIC PRESSURE

DRAIN

PNEUMATIC PRESSURE GAGE

ALTERNATE STATIC SOURCE

RIGHT STATIC PORTS TOP

TOP

BOTTOM

BOTTOM

Figure 16-61. System Integration

16-32

PILOT'S STATIC SOURCE SELECTOR


to 14 knots above stall with the flaps fully extended.

Figure 16-62. OAT Gauge Figure 16-64. Transducer Vane

The left main-gear squat switch disconnects the stall warning system when the aircraft is on the ground. The system has preflight test capability through the use of the STALL WARN TEST switch mounted on the copilot’s left subpanel (Figure 16-65). The STALL WARN TEST switch, when held in the TEST position, raises the transducer vane and actuates the warning horn. Figure 16-63. Rosemont Probe

STALL WARNING SYSTEM The stall warning system consists of a transducer, a lift computer, a warning horn, and a test switch. Angle of attack is sensed by air pressure on the transducer vane located on the left wing leading edge (Figure 16-64). When a stall is imminent, the transducer output is sent to a lift computer. The Lift Computer activates a stall warning horn at approximately 5 to 12 knots above stall with flaps in the 40% (Approach) position, and at 8

In the ICE group of switches on the pilot’s right subpanel, a STALL WARN switch controls electrical heating of the mounting plate (Figure 16-66). With the squat switch in the Ground Mode, power is limited on the mounting plate to one-half the system voltage. Full system voltage is applied to the plate with the squat switch in the Airborne Mode. The transducer vane is heated to system voltage anytime power is applied to the aircraft.

WARNING The formation of ice at the transducer vane, or on the wing leading edge, results in erroneous indications in flight


16-33

16 AVIONICS


16 AVIONICS


The airspeed tape on the PFDs incorporates an Impending Stall Speed/Low Speed Cue (ISS/LSC) to visually indicate when the airspeed is nearing AFM published stall speeds.. It has no connection or input from the stall warning transducer vane. See the Airspeed Display section of the PFD earlier in this chapter.

Figure 16-66. Stall Warning Heat

FLIGHT GUIDANCE SYSTEM (FGS) B300 with Early Environmental System

DG FREE

GND COM

ON

+

+

NORM 2))

OVERSPEED STALL WARN TEST WARN TEST

MAN COOL AUTO

MAN HEAT

MAN TEMP INCR

OFF

ELEC HEAT

DECR

OFF

LDG GEAR WARN TEST

MODE

$872

ENVIR BLEED AIR NORMAL

BLEED AIR VALVES LEFT OPEN RIGHT ENVIR OFF

LOW

PNEU & ENVIR OFF

OFF

CABIN ALT WARN TEST SILENCE 2))

WINDOW DEFOG

CABIN DIFF WARN TEST

OFF

ENG FIRE TEST DET OFF EXT

B300 with New Environmental System

Figure 16-65. Stall Warning Test Switch

16-34

The Flight Guidance System (FGS) consists of an integrated flight director (FD) and autopilot (AP) system. It includes yaw damping and pitch trim functions. The Flight Guidance Panel (FGP), the SYNC and YD/AP DISC buttons are on the control wheels, with the GA button on the left power lever. These inputs control the FGS (Figure 16-67). The FGS consists of two flight guidance channels with independent computers, related hardware, and control circuits. This provides independent output for flight director and autopilot functions. AP/FD indications are displayed along the top of the PFDs (Figure 16-68 ). Active modes are displayed in green and armed modes are displayed in white, below the active modes.


16 AVIONICS



CABIN DIFF HI


L BLEED FAIL

R FUEL PRES LO

ES PR S

R OIL PRES LO

O T S E

R BLEED FAIL

ENG FIRE

EXTINGUISHER PUSH


DISCHARGED

MASTER CAUTION


MASTER WARNING

PRESS PRES PR ESS TTO OR RESET ESET

SH

PPRESS RES E S TO TO RE RESET R SET

FD

CRS2

YD/AP DISC

CPL

PUSH

IR EC

T

T

AP

D

IA

D

UP

SY C N

MAC

YD

ALT A LT

PUSH

PUSH

S/

A LT ALT

1/2 BANK

H

PUSH

IR EC

APPR

HDG

SPEED

VNAV VNA AV

HDG

NAV N AV

CAN

CRS1

FLC

DOWN

PU

VS

EL

FD

C

ESET R RESET O RE RESSSS TTO RE PPRESS

Collins

Collins

516

AC ACC–.03 C –. 03

PA

329

TERM

3300

FMS

N

W

051 05111

UTC

20:03

R RAT AT

1 °C

COM2

F

0

GS

TAS TAS

0

ON

ON OFF

LEFT

12 °C

ISA

TAXI TAXI

ICE

NAV NAV

RECOG

LEFT ACTUATORS ACTUAT TORS STANDBY ST ANDBY

ARM

PILOT


GOV

HI

TEST

PARKING P A ARKING BRAKE

OFF

FUEL


MANUAL

STROBE

L DC GEN

PITOT PIT OT

LEFT

L NO FUEL XFR

L ENG ICE FAIL

L FUEL QTY

L BL AIR OFF

AUTOFTHER OFF

RIGHT


NOSE E

HD LLTT TEST

L

HYD FLUID LOW

RVS NOT READY

BAT TIE OPEN

DUCT OVERTEMP

ELEC HEAT ON

1

RMT TUNE NORM

TUNE CDU

2

RTU

NORM

121 121.5 21.5

OFF

L ENG ANTI-ICE

FUEL CROSSFEED

WING DEICE

L BK DEICE ON

MAN TIES CLOSE

L PROP PITCH

CABIN ALTITUDE

LDG/TAXI LIGHT

UP

RIGHT

RELAY RELA Y

TEST

FLAPS

20

1


60

DOWN

80

.5

R GEN TIE OPEN

COM1

AT ATC C

FORMAT FORMAT <

AUT O AUTO COMM

TE RR TERR RD RDRR

UTC

SPKR

TE RR TERR

R RAT AT

o

C


COM2 BRT DIM

O ON N

+

DISABLE DISABLE

+

R NO FUEL XFR

BLOWER

TEMP

R FUEL QTY

R ENG ICE FAIL

RUD BOOST OFF

R BL AIR OFF

R ENG ANTI-ICE

R IGNITION ON

PASS OXYGEN ON

4

CABIN CLIMB

0

6

.5

2

ADF

2

MKR


IINPH NPH

NORM

4


MAN TEMP INCR INCR

40 35 30 25

100 0F

0

T AL

OFF

7 5

20

3 4

MODE

IINCR NCR


TEMP

ELEC HEAT HEAT

PSI

0 VACUUM V ACUUM


OFF



DECR


LOW


ENG FIRE TEST DET

OFF

0

50 80 ÛÛ) ) 100

FLIGHT HOURS 1/10

CABIN AIR

500 0 USE NO OIL

1000

1500 2000 PSI


OFF

T

2

6


BLOWER

5

1


AUT O AUTO

10

3456


MAN HEAT HEAT

AUT O AUTO

R PITOT HEAT


35k

MAN COOL

R CHIP DETECT

EXT PWR

THDS FT PER MIN

1

2

1

15 5k 15k

+

NORM


R DC GEN

OXY NOT ARMED

2

2

DME

GND COM

DG FREE

SLEW –

NORM

PROP GND SOL


R

HYD FLUID SENSOR

LANDING GEAR

2 OFF

L GEN TIE OPEN

L CHIP DETECT

GEAR DOWN

DN

VENT DOWN LOCK REL

LEFT

COPILOT COPI LOT


2))

BRAKE DEICE

OFF


EMER FREQ

ADC 2

NORM

$872


OFF


MAIN

OPEN

< ET

RANGE

UP


2))

TEST IGNITION AND ENGINE ENGINE START START LEFT RIGHT ON

RIGHT

BEACON

LDG GEAR CONTROL

NAV NAV

1

2))


BAT B AT

OFF TEST

OFF

1250

ESIS ON

AHRS 1

MFD NORM

1

Collins


ON EMER OFF

2

5

2))


PILOT DISPLAY Y PFD

STBY

+

PA

3

FMS F

USH A AUTO UTO TILT

+13 °C

ATC A TC 1

SLEW –

/,*+76

GEN RESET

< PRESET

TILT TILT

G GS

TCAS F AIL FAIL

29.88IN 29.88I N

TFC

SSAT AT

NORM

+

RIGHT

/,*+76

BATT BUS NORM

DG FREE

ARM


350.0

BRT DIM

PROP SYNC


OFF–RESET

RADAR

GCS

Collins

EXT PWR

NA NAV/BRG AV/BRG

1/2

TA TA ONLY ONLY

J10-1


BRT DIM

DME-H

TERR RDR

J JNETT JN

/82 /8215A

TIL TILT LT

121.90

P

ATC1 AT TC1

11 051 1

(8700) 87 ((8215) 82 KASE AS 5))

GCS

2

N

33

108.50 ADF

LOC

HDG

HDG HDG

LOC1

DATA DATA

ELEC

12

1118.85 18.85

ATC A TC 1

T

COMM

E

ET COM1

12.5

11 1 13.00

GLENO G LENO E LINDZ ONDZ DBL /16000A 0 0A 000 ( NTC) (INTC) C)

1

NORM

6

TERR RDR

ONLY TTA A ONL Y

H

VOICE B O T IDENT H

DC P DCP

RA

PUSH

IDENT

123.80

NAV1 NAV1

3

50

S MIC MIC OXY

PULL UP GND PROX

AC ACC C .––

MENU ADV

B BRT

Collins

COM1

11 11 8 . 8 5

VS

ENG2 EN G2

N

30

RADAR

G GPWS

ENG1 EN G1 REFS

S

S

NORM

((8215)

TTG --:-1.4NM

NAV/BRG NA AV/BRG

TERRAIN

XAHS XA HS

ALT A LT

XADC XA DC

XMIT XMIT PA

1

VOL

IAS

CRS 057

3 9 329

HDG 329

FORMA AT FORMAT

12.5 12.5 VOR1

STD

7500

N

M

FMS

DATA DATA PUSH

ELEC

V 13.6NM DBL F

78 20 0 0 00

0MIN

33 3

2 ALT A LT

ATT T AT

FD

XTLK

PUSH

10

10 0KTS

30 0

MENU ADV

SPKR

AUDIO ALTN ALTN

INPH

1 40 9

8215A

A TT V ATT N V

IAS

BARO 8000

10

10

-:--/ 1.4NM

REFS


PRESET MKR

0.6NM 1.4NM -:-- : CLIMB 2.6NM -:-- : (8215) 169NM -:-- :

4

138 38 082 82 2

25 25

AUTO AUTO COMM

2

ADC2 13.6NM 29.92 in

VOR1 057CRS

80

3

ADF

KASE ((8215)) (8700) KCOS

STD

A

AHS2

OIL

49 TEMP°C 11 1122

FIRE

60

2

600

HDG 329 (8215) 0.8NM

830

30

2

130 FF 750 122 PRESS 80

ITT

TORQ TORQ

w

2

106.0

P

NAV NAV

DME

1740

N1

110.0 110.0

1

700

PROP

62.2 3.4

J206

1

1

10

0

1050

ITT

TTORQ ORQ

24

ACTIVE A CTIVE

T

V2 107 VR 103 V1 100

AP AP

TRIM TR IM

PUSH

800 00

MIC OXY

2

N350KA

W

COMM

RADIO CALL

BARO 1

40

NORM

1

ACTIVE

Collins

2

900

7820

S 1

FLAP OVRD D

T

VOL

10

TERR INHIB B

A ACTIVE

STEEP APPR PR

4

8 000

20 60

G/S INHIB IB

A ACTIVE

21

16000

170

80

1

T

Collins Colli ns

8215

S

XMIT PA

2

PTCH ALTS ALTS

15

HDG FMS

<

MASTER WARNING

ENG FIRE

DISCHARGED

T

L OIL PRES LO EXTINGUISHER PUSH

T

L FUEL PRES LO

MASTER CAUTION

AUTO AUTO


INCR INCR

EXT

10

15

Figure 16-67. Flight Guidance Panel (FGP) Collins

HDG FMS

AP

PTCH ALTS

140

80

20

14 000

coupled FGC. The autopilot automatically disengages when autopilot control discrepancies are detected.

6935 4

Figure 16-68. Flight Guidance System Display 1

FLIGHT GUIDANCE COMPUTERS (FGC) Each FGC is supplied with input from the AHRS, navigation data, FGP selections, servo, and ADC computers. The coupled FGC produces control signals for yaw damping, AP/FD, and pitch trim functions. Each FGC is supplied data from the onside ADC, EFIS, and AHRS. The autopilot and flight director require both attitude portions of the AHRS to be operational. <

Each FGC produces an independent AP control signal. Only one FGC may be coupled to the autopilot at any time. AP control computations from the other FGC are continuously compared with AP control signals from the

FLIGHT GUIDANCE PANEL (FGP) The Flight Guidance Panel (FGP) controls both FGC’s. The coupled FGC then controls the Flight Guidance System (Figure 16-69. The FGP is centered at the top of the instrument panel. All AP/FD mode selections are made on this panel. The FGP has the following controls:

AP Button The AP button controls autopilot engagement. The autopilot engages if the following conditions are met: (1) YD/AP DISC switch-bar is raised; (2) no unusual attitudes/rates exist; (3) and the flight guidance computer does not detect any autopilot faults. The yaw damper is automatically engaged when the AP button is pushed.


16-35

NAV

T

ALT SH

H

PUSH SY C N

MAC

YD

AP

FD

CRS2

YD/AP DISC

CPL

PUSH

IREC

D

IA

D

UP

S/

ALT

1/2 BANK

PUSH

IREC

APPR

HDG

SPEED

VNAV

PUSH

HDG

T

FLC

CAN

CRS1

DOWN

PU

VS

EL

FD

C

16 AVIONICS


Collins

Figure 16-69. Flight Guidance Panel (FGP)

YD Button The YD button controls yaw damper engagement. The yaw damper may be engaged without engaging the autopilot. Disengaging the yaw damper with the autopilot ON will also disengage the autopilot.

CPL Button The CPL button controls which flight guidance computer (FGC), right or left side, supplies flight director commands and attitude data to the autopilot. With the autopilot on, a green arrow on the PFD indicates the coupled FGC (Figure 16-70). With the autopilot off, a white arrow on the PFD indicates which FGC is generating the flight director commands. The cross-side flight director will be a duplicate of coupled side. Flight director modes will default to ROLL and PTCH modes each time the CPL button is pushed.

couple arrow will always point to the left after avionics power-up. Each PFD will display AP/FD commands from the coupled side. They do not normally operate independently. There are two exceptions: go-around mode; full-ILS approach mode. When GA and full-ILS modes are active, each Flight Guidance Computer (FGC) provides independent guidance to the onside PFD flight director. When either of these conditions exist, the single pointer arrow adds another barb to show that the flight directors are now independent (Figure 16-71). For this condition to exist in the full-ILS approach mode, the same localizer frequency must be tuned on both radios (e.g., LOC1 and LOC2) and the glideslope must be captured. If independent operation can not be accomplished an annunciator will appear on the non-coupled side showing that an independent mode was attempted but unsuccessful. Collins

Collins

HDG FMS

AP

PTCH ALTS

140

APPR LOC1

14 000

80

6935

AP

1 4 000

80 4

20

GS

140 20

Left Side Couple

6935 4

700

2

6

Successful Independent Operation Collins

Collins

1

4

2

HDG FMS

PTCH ALTS

140

80

20

14 000

4 6935

APPR LOC1

140

4

80

AP

GS

1 4 000

FD1

20

RIght Side Couple

Figure 16-70. Flight Guidance Couple Arrow

6935 4

700

2

6 Unuccessful Independent Operation

Figure 16-71. Independent Flight Director 1 Operation <

<

2

At power-up, the left side FGC is automatically chosen as the computer to supply the flight director. Autopilot commands and the

4

4

FOR TRAINING PURPOSES ONLY <

<

16-36

R

The coupled FGC provides automatic pitch trimming with the autopilot engaged. Pitch

R

trimming is disabled if a pitch trim fault occurs. If a pitch trim fault is detected before the autopilot is selected ON, the autopilot will be prevented from engaging. A pitch trim fault detected after autopilot engagement will not disengage the autopilot. Failures are indicated by the appearance of a red TRIM annunciation on the PFDs (see the Flight Controls section of this PTM).

YD/AP Disconnect Switch-Bar The YD/AP Disconnect switch-bar removes power from the autopilot and yaw damper causing both to disengage. When pulled down, a red and white band is visible to indicate the disengage position (Figure 16-72). Raise the switch-bar to permit autopilot/yaw damper engagement. LT

YD

ALT

YD/AP DISC

H

S

CPL

AP

CR PU

IREC

D

CE L

FD

Figure 16-72. YD/AP Disconnect Bar

FD Mode Buttons All mode buttons on the FGC are ON / OFF buttons. Caution should be exercised when selecting each mode, as the buttons do not indicate which one is already engaged. A scan of the mode selection area on each PFD is required first to verify current mode. When a mode is then selected, incompatible modes are automatically removed. Lateral modes include HDG, ROLL, ½ BANK, APPR, and NAV. Vertical modes include VS, ALT, VNAV, PTCH, FLC (or IAS), and altitude select (ALTS).

FD Buttons The left and right side FD buttons control display of the flight director command bars on the respective PFD. At power-up, both flight

directors are off. Both flight directors are automatically activated when the autopilot is engaged or when a flight director mode is selected. Pushing the FD button will initially display both flight directors in the PTCH and ROLL modes. Either pilot can independently remove their command bars from view by pressing the respective FD button. The command bars will be removed from view but the mode selections and opposite pilot’s command bars will remain in view. If both pilots remove the command bars from view, the flight director will be completely turned off. This includes all mode selections. For IFIS equipped aircraft the flight director image can be a v-bar or cross pointer (x-ptr). See the REFS section of the DCP in this chapter.

UP/DOWN Pitch Wheel The pitch wheel controls reference values used to set the vertical speed in the VS mode, or pitch angle in the pitch mode. Caution must be taken when using this control because it will override or change active vertical modes. There are two exceptions: glideslope (GS) captured; GPS Vertical Glidepath (VGP) captured. This override is active during altitude capture so care should be taken not to manipulate the pitch control wheel during the display of ALT CAP on the PFD.

ROLL Mode The ROLL mode is the basic lateral mode and is activated automatically if no other lateral mode is selected when the flight director is on, or when the CPL button is pressed. ROLL annunciates on the PFD when the mode is selected. In the ROLL mode, the FGC maintains the current bank angle at engagement if the bank angle is more than 5 degrees. The current heading is maintained, with a bank angle limit of 5 degrees, if the bank angle is 5 degrees or less when the ROLL mode is activated.


16-37

16 AVIONICS


HDG Button

1/2 BANK Button

The HDG button controls selection of heading mode. HDG annunciates on the PFD when active. The FGC maintains the heading selected by the heading bug.

The 1/2 BANK button limits the maximum bank angle to 15˚. While in this mode, a white arc appears bellow the roll scale that spans ±15 degrees either side of level (Figure 16-74).

HDG Knob The HDG knob simultaneously controls the heading bugs shown on both PFDs and the MFD. If the bug is out of view on a display, a cyan dashed line will extend from the airplane symbol to indicate its location. A digital readout of the selected heading will be displayed to the left of the current heading display (Figure 16-73). The commanded turn will take the 1 shortest distance to the selected heading unless the heading bug was rotated beyond 180˚ from the current heading. When rotated beyond 180˚, the turn will continue in the direction the bug was moved. 0 HDG 010

TERM

FMS1

24

W

21

144 069

30

DTK 251 (6935) 0. 8NM 50

FORMAT > <

25

TERR

< PRESET

RDR

VOR1

>

TERRAIN

TFC >

F

TCAS OFF

< ET 01:42 COM1 121.800

ATC1 4336

RADAR ON UTC 14:41

RAT 15 oC

COM2

125.250

BRT DIM

Figure 16-73. Heading Vector Line

PUSH SYNC Button The PUSH SYNC button within the HDG knob resets the heading bugs to the current heading.

16-38

PTCH ALTS

140

14 000

80

20 60

6 4

700

2 1

Figure 16-74. Half Bank Mode

The half-bank mode is automatically selected when climbing through 18,500 feet and deselected when descending through 18,500 feet. This mode is also deselected with the following; localizer capture; go-around mode selection; or onside FMS navigation capture.

APPR Button

30.16IN

251

HDG FMS

The APPR button controls selection of the approach mode. The type of approach is determined by the active navigation source shown on the PFD (APPR LOC1, APPR VOR2, APPR FMS2, etc.). The mode also arms the glideslope capture after the front course localizer has captured if GS is valid. At glideslope capture, the FGC will descend on the glideslope and disregard any preselected altitudes. The FGC will not capture an altitude after the glideslope is captured. The displayed position of the CDI course is significant when APPR is pressed. If the head of the needle is more than 110 degrees from the present heading, then the approach mode will assume a localizer back-course is desired and the annunciation APPR B/C1 or APPR B/C2 will appear. This position of the CDI will also suppress any glideslope indications. If the course is less than 110 degrees from the present heading the approach mode assumes a


<

16 AVIONICS


normal localizer based approach and the annunciation APPR LOC1 or APPR LOC2 will appear and the GS will arm and capture normally (Figure 16-75).

called NAV-to-NAV capture as the pilot does not have to manually change navigation sources or change flight guidance modes. It is accomplished automatically.

Additionally, this mode will allow the FMS to accomplish what is called a NAV-to-NAV capture. When FMS is the current active NAV source and has been loaded with a localizerbased procedure (ILS, LOC, LOC BC, LDA, SDF) the FMS will automatically tune that localizer and set up a preselected course when within 30nm of the airport. The preselected course will appear as a cyan dual line, dashed CDI on the PFD. This preselected course must become the active navigation source when on final for the localizer procedure as it is required by limitation. This transfer will happen automatically only if the APPR mode has been pressed and the preselected course is trending toward center (Figure 16-76) . This is

The APPR button is also used when flying a non-localizer-based approach to a DA (Decision Altitude). When established on final for an appropriate RNAV (GPS) approach, the APPR button will activate the approach mode (APPR FMS1 or APPR FMS2). When VNAV is then pressed, it will arm the vertical glidepath (GP) mode (Figure 16-77). This allows the FMS to follow a glidepath down to a published decision altitude (DA) minimum. This approach descent is based on barometric altitudes and does not consider a ground based antenna. Like the ILS glideslope, however, the GPS GP will disregard any preselected altitudes. Reference the VNAV section of this chapter for more information.

6 Collins

Collins

1

APPR B/C1

ALTS

140

700

LOC1 109.75

24

251 LOC11 109.75

w

24

CRS 235 IEJC 1 0. 8NM

S 15

< PRESET

E

3

<

6 COM2

< ET

125.250

COM1 121.800

ATC1 4336

TFC > TCAS OFF

6

RAT 15 oC

12

12

V 4.1NM SXW

TFC > TCAS OFF

3

E

UTC 14:41

TERR RDR

FMS1

<

15

<

ATC1 4336

<

COM1 121.800

FORMAT >

N

< ET

1

33

N

TERR RDR

VOR1 V 4.1NM SXW

4 600 60 5 30. 16 IN

4

FORMAT >

33

< PRESET

2

w

21

S

30

21

700

30

B/C 055 IESJ 0. 8NM

10

DN

1

3 0 .1 6 I N

251

TERM

160

600 60

4

20

2 1

10

UTC 14:41

RAT 15 oC

COM2

BRT DIM E

6935

1 4 000

180

4

20

GS

185

80 60

APPR LOC1

6935

14 000

Localizer Back Course

125.250

BRT DIM E

Localizer Front Course

Figure 16-75. APPR Mode Selection


16-39

16 AVIONICS


16 AVIONICS


Collins

APPR FMS1 APPR LOC1

ALTS GS

700

400

0

24

w

S

S

3

E

6 COM2

15 12

E

RAT 15 oC

< ET

125.250

COM1 121.800

ATC1 4336

TFC > TCAS OFF

6

12

V 4.1NM SXW

TFC > TCAS OFF

UTC 14:41

TERR RDR

FMS1

3

15

<

ATC1 4336

<

< PRESET

N

COM1 121.800

FORMAT >

33

TERR RDR

N

< ET

w

21

FORMAT >

33

V 4.1NM SXW

4

24

CRS 235 IEJC 0. 8NM

2

30. 16 IN

251

30

LOC1 109.75

400

850

100

LOC1 109.75

21

1

10

120

30

DTK 235 CHARL 0. 8NM

600 60 6 540 20

2 4

2 1

10

DN

3 0 .1 6 I N

251

TERM

FMS1

700

14 1 0 1

10

117 V2 110 VR 106 V1 ACC-.02

160

600 60 6 540 20

4

20

2 1

10

6935

1 4 000

180

4

20

GS

185

80 60

APPR LOC1

6935

14 000

140

Collins

UTC 14:41

RAT 15 oC

COM2

125.250

BRT DIM

BRT DIM

FMS with Localizer Preselect

Localizer Capture

Figure 16-76. Localizer Nav-to-Nav Capture

Collins

APPR FMS

VPTCH GP

185 180 160

700

160

24

30.16IN

251

700 600 60 6 540 20

1

10

120

400

2

24

TOD

W

GP Armed

2 4

100

1000

2 1

10

DN

4000 4

20

14 1 0

4

100

TOD

180

1

400

3 000

2

600 60 6 540 20

10

VGP

185

1

10

14 1 0 120

APPR FMS

4000 4

20 DN

3 000

Collins

251

30.16IN

1000

W

GP Active

Figure 16-77. VNAV Glidepath (GP) Mode < <

V

16-40


NAV Button The NAV button controls selection of the navigation mode. Heading mode remains active until course intercept. After intercept, the FGC maintains the selected course. The active NAV identifier annunciates on the PFD (FMS, VOR1, LOC2, etc.). The NAV mode should be used during the enroute phase of flight, for appropriate terminal procedures and when flying an approach to an MDA. This excludes an FMS NAV-to-NAV capture as referenced in the APPR section. Refer to the VNAV section of this chapter for more information on how this mode interacts with FMS vertical navigation.

CRS Knobs The CRS knobs select the course to be flown on the respective PFD. This knob is not active when FMS is the active navigational source.

is not engaged, pushing the SYNC button on the control wheel synchronizes the pitch reference to the current attitude.

VS Button The VS button controls selection of the vertical speed mode. When VS is activated, the FGC initially maintains the current aircraft vertical speed when the mode is selected. Rotating the UP/DOWN pitch wheel changes the vertical speed reference value. When the autopilot is not engaged, pressing the SYNC button on the control wheel synchronizes the VS reference to the current vertical speed. VS and the vertical speed reference value appear on the PFD (Figure 16-78). An up arrow appears for climbs and a down arrow appears for descents. A reference arrow (bug) appears on the vertical speed scale adjacent to the selected vertical speed.

PUSH DIRECT Button

Collins

The PUSH DIRECT button within the CRS knob automatically selects a direct course to the active VOR, and centers the CDI on the respective PFD. This button is not active when either FMS or LOC is the active navigational source.

HDG FMS

1 4 000

80

700

2 1

10

117 1

6935 4

20 60

V2

Pitch Mode

VS 1100 ALTS

140

600 60 6 540 20 1

10 4

Figure 16-78. Vertical Speed (VS) Mode 4

3

Rotating the UP/DOWN pitch wheel changes the pitch reference value. When the autopilot

VNAV Mode The VNAV button controls Vertical Navigation mode selection and is annunciated on the PFD as a “V” located in front of the active vertical mode (e.g., VPTCH, VVS, VALTS, etc…). The flight management computer (FMC) determines the VNAV capture point and provides vertical steering commands to waypoints R that contain altitude restraints in the FMS. See the VNAV section and the Flight Guidance Mode Annunciations table for more information.


<

Pitch mode is a basic vertical operating mode. It activates when no other vertical mode is active and the flight director is on. The annunciation PTCH displays on the PFD. When active, the FGC maintains the pitch attitude which existed when the pitch mode was engaged. This will occur when the previously selected vertical mode is pressed again (deselected) or when the UP/DOWN Pitch Wheel is moved and VS mode is not active.

R

16-41

16 AVIONICS


FLC Button The FLC button controls the Flight Level Change mode. The FLC mode will climb or descend the airplane towards the preselected altitude at the IAS or Mach speed reference located above the airspeed display. FLC indications are modified by the SPEED Knob (Figure 16-79). It is important to note that when the autopilot is engaged after the FLC mode is selected, the present speed of the aircraft will be indicated as the active speed, not the one dialed in with the SPEED knob. The pilot can reset the desired speed by rotating the SPEED knob. Collins

HDG FMS

FLC 160 ALTS

160

14 000

80

20

2

FLC M.31 ALTS

M.31

4

14 000

80

20

6935 4

700

The SPEED knob selects the IAS or Mach reference value, as appropriate, to be used by the FLC mode. This value displays at the top of the Airspeed Tape. When the FLC mode is selected, the selected speed will also be annunciated adjacent to the FLC mode annunciation at the top of the attitude display.

IAS/MACH Button

2

Collins

HDG FMS

SPEED Knob

6935 4

700

the selected speed. This same procedure will occur if a lower altitude is preselected but the power is left too high. In this situation the aircraft will initially pitch to achieve the selected speed. If this results in a speed faster than selected, the aircraft will begin to pitch back up until it maintains a descent of approximately 100 ft/min, regardless of what speed that generates.

The IAS/MACH button within the SPEED knob, when pushed, selects Mach mode or IAS mode for the FLC Speed Bug and FLC reference. The system automatically changes from IAS to Mach or Mach to IAS when climbing or descending through 20,517 feet.

2

Figure 16-79. Flight Level Change (FLC) Mode

ALT Button

<

2

The FLC mode controls the pitch of the aircraft and requires pilot manipulation of power to establish a climb or descent. If the power is set inappropriately or the speed is unachievable, the aircraft will not be allowed to deviate further from the preselected altitude to achieve the selected speed. As an example, if an altitude of 5000’ is preselected and FLC mode is chosen for a 160kt climb and the power is not increased, the aircraft will initially begin to pitch up. If this results in a speed below 160kts, the aircraft will then lower the pitch until the VSI indicates a climb of approximately 100 ft/min and stay there regardless of what speed that generates. It will not allow the aircraft to pitch down and deviate away from the preselected altitude to achieve 4

<

16 AVIONICS


16-42

The ALT button is used to hold the aircraft at the current barometric altitude. The ALT button is used to level at an altitude other than a preselected altitude. ALT will annunciate on the PFD when this is pressed. If the autopilot is not engaged, pressing the SYNC button on the control wheel synchronizes the altitude reference to the current altitude. As with all flight guidance modes, pressing the ALT button when “ALT” is already annunciated on the PFD will remove the altitude capture.

Altitude Preselect Mode The altitude preselect mode permits the pilot to select a target altitude for automatic level off by the autopilot or FD command. The


ALTS armed mode annunciates in white on the PFD. The altitude preselect mode is automatically selected with the following: the ALT knob is turned; go-around mode is cleared or the flight director is turned on. Altitude preselect is automatically deselected when glideslope approach mode becomes active, the VNAV glidepath approach mode (VGP) becomes active, altitude hold mode is selected, or the altitude capture mode (ALT CAP) is annunciated. If a descent or climb is desired, a new altitude must be preselected. The appropriate vertical mode must then be selected to climb or descend. Changing the altitude preselector alone does not cause the aircraft to climb or descend. If the ALT knob is turned while ALT CAP is annunciated, the pitch mode is selected and the altitude preselect mode rearms. Altitude capture (ALT CAP) occurs when the airplane altitude approaches the selected altitude. The capture point depends on the closure rate. When within 1000’ of the selected altitude a single aural tone will sound and the preselected altitude will flash. The flashing will stop when within 200’ of the selected altitude. Should the aircraft subsequently deviate by more than 200’ from the selected altitude the single aural tone will sound and the preselected altitude will flash yellow. The flashing will stop with an input by the pilot (pressing the altitude selector knob) or the aircraft returns to within 200’ of selected altitude. In either case the number will stop flashing and return cyan in color.

16 AVIONICS


ALT Preselect Knob The ALT knob selects the desired altitude for level off (displayed on the PFD). Rotating the knob while in its default position will select thousands of feet. Pressing the knob IN while rotating will select hundreds of feet. See the Altitude Display section of the PFD for more information on the bugs that appear on the altitude tape.

PUSH CANCEL Button The PUSH CANCEL button within the ALT knob cancels the flashing visual altitude alerts on the Altitude Display section of the PFD as described earlier.

CONTROL WHEEL SWITCHES The following control wheel switches affect FGS operation:

DISC TRIM AP/YD Button The DISC TRIM AP/YD button is located on the outboard horn of each control wheel. It is used for disengagement of the autopilot and yaw damper (Figure 16-80). Pushing the button to the first detent will disconnect the autopilot and/or yaw damper. Pushing the button to the second detent will interrupt electric trim operation. Releasing the button will reset the trim and allow continued operation.

ALTS shows in yellow if the capture is inhibited due to invalid data and ALTS CAP shows in yellow if the capture is cleared without a subsequent selection of altitude hold or glideslope/glidepath capture. Figure 16-80. Left Yoke


16-43

SYNC Button

GA Button

The SYNC button is located on the outboard horn of each control wheel. It is used to synchronize the PTCH, FLC, VS, ALT and ROLL modes of the flight director to the current parameters if the autopilot is not engaged (Figure 16-81). Inputs known as Control Wheel Steering (CWS) or Touch Control Steering (TCS) features are not installed on this system.

The GA button is located on the outboard side, in the center, of the left power lever (Figure 16-82). The G/A button selects the goaround (GA) mode of the flight director. Selecting GA mode will disengage the autopilot, but not yaw damper and clear all other flight director modes. The flight director will display approximately +7 degree pitch up attitude. ROLL mode will be selected and heading will be held if bank angle is less than 5 degrees. (Figure 16-83). The heading being held is independent of the heading bug. This mode will not follow any lateral or vertical commands and will not capture the preselected altitude. During go-around mode, the flight directors are independent and the failure of one will not affect the other. This allows for redundancy during a critical flight maneuver. The independent flight director capability also occurs during a full ILS and provides the same redundancy.

Collins

SYNC

HDG FMS

PTCH ALTS

14 000

140

80

4

20 60

6935

700

2 1

10

600 60

6 540 20 V2 VR V1

117 110 106

400

0

ACC-.02 TERM

1

10

24

251

2 4

30.16IN

W

Figure 16-81. Pilot's PFD with SYNC

Electric Pitch Trim Switches

<

16 AVIONICS


The electric pitch trim switch is comprised of two segments. The trim switch is located on the outboard horn of each control wheel. The trim switch applies electric pitch trim commands. Both segments of the switch must be actuated to operate the electric pitch trim. The segmented pitch trim switch reduces the potential of trim runaway or inadvertent activation. T

When moved in either direction, the electric pitch trim switches will disconnect the autopilot while leaving the yaw damper engaged. See the Flight Controls section of this PTM for further discussion of electric pitch trim and its annunciations.

Figure 16-82. Go-Around Button

16-44


It is necessary to reselect a desired mode after the aircraft is configured in the go-around to regain full flight director control.

CONTROL DISPLAY UNIT (CDU)

See the Flight Guidance Mode Annunciations table at the end of this chapter.

The Control Display Unit (CDU-3000) serves as a control of the communication and navigation radios, Flight Management System (FMS) and limited display control for the PFDs and MFD (Figure 16-84). The pedestal can contain either one or two CDUs. The second CDU is an option. If two are installed, each CDU will communicate only with the respective FMS. In the optional two CDU installation, reversionary mode is not available should one fail. The remaining CDU will be capable of communicating with the on-side FMS only.

Collins

GA

GA

14 000

140

80

4

20 60

6935

700

2 1

10

600 60

6 540 20 V2 VR V1

117 110 106

400

0

ACC-.02 TERM

1

10

24

251

2 4

30.16IN

The CDU has a normal operating temperature range of -20˚C to +70˚C. Should the unit temperature get below -20˚C the CDU will turn ON but the LCD display will delay indications

W

Figure 16-83. PFD Go-Around (GA) Mode <

T

ACT FPLN ORIGIN

1/4 DIST

KICT

DEST

452

KDEN

ROUTE

ALTN

PLANT2

KAPA ORIG RWY

VIA

TO

DIRECT ICT ------------------
PERF INIT>

[

[

MSG

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-84. Control Display Unit (CDU)


16-45

16 AVIONICS


16 AVIONICS


by a power-up timer. During this time the CDU will monitor its internal temperature. With extreme unit temperatures of -30˚C and colder, this timer can take as much as 10 minutes to illuminate the display.

Scratchpad Line

This button provides secondary control of the display intensity. The PILOT DISPLAYS rheostat on the overhead panel provides primary control.

The scratchpad line displays data entered by the alphanumeric keys, or data selected for transfer by a line key. Brackets identify this line and it is the only place where the operator can input information from the keypad. Once input data is displayed on this line it should be verified before transferring to a selected field. Should an entry occur that is not compatible with the selected item, the scratchpad will momentarily display a message to indicate details about the error. This message will time out and the previously entered information will return, so that it may be corrected.

Title Line

Message Line

This line displays the page title and page number. The page number is formatted as the current page number followed by a slash and the total number of pages.

A single message line is reserved along the bottom line of every page to annunciate conditions requiring operator attention or simply to provide information. If more than one message is active the message key (MSG) may be used to display additional messages as discussed later in this section.

The CDU has the following controls and displays:

BRIGHT/DIM Button

Line Select Keys These keys activate functions displayed on the CDU adjacent to the line select key. The line functions depend on which page is displayed.

Label/Data Line Pairs Two display lines are associated with each line select key. The top line is normally a label for the information that is shown on the data line Displayed on the second (bottom) line. The data line can display large or small characters. When the system has entered information the text will be in a smaller size. When the operator has entered information the text will be larger in size.

Alphanumeric Keys These keys enter data in the scratchpad line of the display. The data entry keys are as follows; the 0-9 number keys; the A-Z letter keys; the period key; the +/- (plus/minus) key; the SP (space) key; the / (slash) key; and the CLR/DEL (clear/delete) key. The compass cardinal headings of N, E, S, and W are highlighted with a white box to ease entry of items requiring direction inputs. Care must be exercised not to confuse the letter “O” with the number “0” on the keypad.

IDX Key The IDX (index) key controls display of items that do not have a dedicated function key. It also is a central location for setup and configuration pages for FMS and GPS operations.

16-46


16 AVIONICS


FPLN Key ACT LEGS

ACT FPLN ORIGIN

1/4 DIST

KICT

DEST

452

KDEN

ROUTE

MSG

ALTN

PLANT2

KAPA ORIG RWY

VIA

TO

DIRECT ICT ------------------
DIR

FPLN

DEP ARR

LEGS

PERF

MFD MENU

MFD ADV

MFD DATA

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

J K L M N

7 8 9 O P Q R S T U SP

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

/ +/ - V W X Y Z 0 NEXT

/ +/ - V W X Y Z 0

PERF

MFD MENU

/

SP

EXEC

PREV

I

DEP ARR

7 8 9 O P Q R S T U

[ LEGS

KICT / o 309 12NM ---/----ICT o / 307 9.2NM MUGER ---/----/ o 3.3NM 307 WUKOL ---/----/ o 0.5NM / 307 WUKUS ---/--------------------------LEG WIND> [ [

PERF INIT>

[

MSG

1/6 SEQUENCE AUTO/INHIBIT

The FPLN (flight plan) key controls display of the active flight plan (Figure 16-85). This page will give an overview of the entered flight plan, not each individual waypoint.

/

Figure 16-85. Active Flight Plan Page

Figure 16-86. Active Legs Page

DIR Key The DIR (direct) key controls display of the active direct-to page. Navigating backward through these pages will lead to a HISTORY page of all the previous waypoints in the flight plan (Figure 16-87).

LEGS Key The LEGS key controls display of the waypoint-to-waypoint detail contained in the active flight plan. The display includes the lateral information from waypoint-to-waypoint and vertical information when applicable. Page 1 always contains the current FROM waypoint in cyan at the top and the current TO waypoint in green (Figure 16-86). Page 1 also contains the selection of AUTO sequencing or INHIBIT sequencing when the progression of waypoints is desired (AUTO) or not desired (INHIBIT).

ACT DIRECT-TO

1/1

HISTORY

/o 250

<(6935) 215o

[

MSG

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-87. Direct to Pages


16-47

16 AVIONICS


DEP ARR Key

NEXT Key

The DEP ARR key controls display of the departure/arrival pages. The selectable procedures are those related to the current active flight plan ORIGIN and DESTination airports or the current secondary flight plan ORIGIN and DESTination airports. If diversion to a different airport is desired, the identifier for that airport must be placed in the DEST slot on the FPLN page to retrieve departures / arrivals for that airport.

The NEXT key is used to display the next page when the current CDU function has more than one page.

PERF Key The PERF key controls display of the performance menu page. These pages contain manually entered loading data, fuel advisory pages, and some VNAV advisory pages.

EXEC Key The EXEC (execute) key activates modifications made to the active flight plan. The label EXEC annunciates on the CDU when the active flight plan has been modified and the changes have not been activated (Figure 16-88). Pushing the EXEC key activates the modified flight plan. If this key is not pressed the changes will not take effect. A CANCEL MOD option is available when the modification to the flight plan has not yet been executed. It will erase the modification and return the FMS to the original flight plan.

MSG Key The MSG (message) key controls display of the system message page. This is necessary when more than one message is active. Should multiple messages be active pressing the MSG key will allow additional messages to be viewed. To return to the last viewed page simply press the MSG key again.

TUN Key

MOD FLPN HOLD ENTRY

DIRECT

1/1

INBD CRS/DIR

HOLD SPD FAA/ICAO MAX KIAS // 200 FIX ETA

LEG TIME

EFC TIME

QUAD/RADIAL

--/---o

/ o/R TURN 307

14:16

/ MIN 1.0

// 02:00 /

LEG DIST

MSG

The TUN (tune) key controls display of the radio tuning page. These pages are used to tune the communication, navigation and ATC transponder equipment in conjunction with the Radio Tuning Unit (RTU). If two CDU’s are installed, the right CDU will not have this page active.

FIX

WUKOL

/ NM NEW HOLD> 3.0 ----------------------
DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-88. Hold FPLN Mode

PREV Key The PREV (previous) key is used to display the previous page when the current CDU function has more than one page.

16-48

MFD MENU Key The MFD MENU key opens the display of the MFD menu page on the CDU (Figure 16-89). The MFD menu page displays a menu of the


point on the FMS plan map display on the MFD. It will also control advancing through the pages within a selected MFD DATA text page.

possible MFD display options, or available text pages for display on the MFD when the MFD Data Key has been pressed. A “L/R” is displayed on the lower right corner of this page. The left (L) selection will be all the options for the left PFD and the MFD; the right (R) selection will be all the options for the right PFD only. For each menu the items in green are selected and the items in white are not selected.

MFD DATA Key The MFD DATA key controls the display of text data pages on the MFD (Figure 16-91). The text data page displayed is the last one selected from the MFD menu page. Other pages can be accessed through the MFD MENU Key.

MFD ADV Key The MFD ADV key controls display of the MFD Advance page on the CDU (Figure 1690). The MFD advance page displays a menu enabling a move to the next or previous way-

SPEED

LO NAVAIDS

ALTITUDE

INTERS

APTS

RNG: ALT SEL

NAV STATUS

LRN POS

POS SUMMARY

ALTN FPLN

POS REPORT

MISS APPR SIDE L/R>

TERM WPTS

WINDOW OFF/ON/VNAV

VOR STATUS

SIDE L/R>

[

[ DEP ARR

LEFT DISPLAY MENU TEXT DISPLAY FPLN PROG

MFD MENU

MFD ADV

MFD DATA

2/2

LEFT DISPLAY MENU MAP DISPLAY

ETA

HI NAVAIDS

MSG

NDBS

1/2

LEFT DISPLAY MENU MAP DISPLAY NEAREST APTS

EXEC

MSG

DEP ARR

MFD MENU

MFD ADV

MFD DATA

SIDE L/R>

LRN STATUS

[

[

EXEC

[

[

MSG

MFD ADV

MFD DATA

EXEC

FPLN

PREV

NEXT

DIR

FPLN

PREV

NEXT

DIR

FPLN

IDX

1 2 3 A B C D E F G

CLR DEL

IDX

1 2 3 A B C D E F G

CLR DEL

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

TUN

4 5 6 H

BRT DIM

TUN

4 5 6 H

BRT DIM

PERF

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

LEGS

PERF

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Pages with Map on MFD

PERF

MFD MENU

DIR

LEGS

LEGS

DEP ARR

I

PREV

J K L M N

NEXT

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

/

SP

Page with Text on MFD

Figure 16-89. MFD Menu Key (CDU)


16-49

16 AVIONICS


16 AVIONICS


LEFT DISPLAY ADVANCE ACT PLAN MAP CENTER
LEFT DISPLAY ADVANCE TEXT DISPLAY

[

[

MSG

SIDE L/R>

LEGS

DEP ARR

MFD MENU

PERF

MFD ADV

MFD DATA

EXEC

[

[

MSG

MFD ADV

MFD DATA

EXEC

FPLN

PREV

NEXT

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

TUN

4 5 6 H

BRT DIM

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

I

J K L M N

7 8 9 O P Q R S T U

/

SP

PERF

MFD MENU

DIR

I

LEGS

DEP ARR

/ +/ - V W X Y Z 0

With Map Displayed on MFD

SP

/

With Text Displayed on MFD

Figure 16-90. MFD Advance Key (CDU) Collins

26

PROP

PROP

62.2 0.0

TORQ TORQ

N1 NI

1740 1980 106.0 98.5

130 FF 750 FF 0 430 122 PRESS 80

ITT

ITT 830

734

0

TORQ TORQ

ICT

DIST 3.6NM

ETA 20:07

7.4NM

20:11

1/3

2450

16.6NM

20:15

2390

WUKOL

19.8NM

20:16

2370

WUKUS

20.4NM

HUT

29.8NM

20:20

20:16 2310

2370

OATHE

218NM

21:32

1170

SELLS

283NM

21:56

760

----------------------------------------

GS

0

TAS

415NM

22:47

0

435NM

22:54 RESERVE EXTRA

0 0 0

0

The FMS provides multiple flight management functions. These functions include lateral navigation, (LNAV) using multiple navigation receivers, and vertical navigation (VNAV). Navigation input includes GPS, DME and VOR receivers. Vertical navigation (VNAV) is provided by a computed vertical output from the FMS using these receivers. The system also provides course-tracking signals to the flight guidance system. The Flight Management Computers (FMCs) are housed in the IAPS unit located in the nose avionics bay.

FUEL (LB) 0

MUGER

DEST KDEN ALTN KAPA

120

o

FMS ACT PROGRESS WPT KICT

PRESS OIL

OIL 49 TEMP C 112 46 TEMP°C 73

FIRE

110.0 2000

3.40

SAT 15 oC

FORMAT <

<

10500

ITT

ITT 516

FLIGHT MANAGEMENT SYSTEM (FMS)

TFC <

ISA +13 oC

BRT DIM

Figure 16-91. MFD Text Page

16-50

The FMS uses a blended combination of GPS and VOR/DME data to construct a three dimensional position of the aircraft in space. To achieve this blend, the NAV1 radio and NAV2 radio must be receiving a valid signal. This can be accomplished by manually tuning the receiver or setting a feature called “auto-tuning” which will be discussed later.


The CDU is the primary interface with the FMS. Each CDU will communicate with the “on-side” FMS (e.g., Left CDU for No.1 FMS, Right CDU for No.2 FMS). The FMS’s can be synchronized so that selected operations on one CDU (and its related FMS) will automatically be transferred to the cross-side CDU (and its related FMS). (See FMS quick reference guides and other handouts for information on how to synchronize the units).

Rockwell Collins software titled “PCD Software” is required for this operation.

The FMS database is updated using the Database Unit (DBU). This system can consist of a 3.5-inch high-density floppy disk drive (DBU4100) located on the center pedestal and either an additional computer port located on the lower right sidewall of the center pedestal or an Ethernet port on top of the pedestal on IFIS equipped aircraft. Optionally, a unit called the DBU-5000, which consists solely of two USB ports on top of the pedestal may be installed. The installed system is used to upload data to the aircraft or download data from the aircraft. This can include avionics malfunction reports (Figure 16-92). All update methods require the aircraft battery and avionics to be ON. It is strongly recommended that a ground power unit be applied to the aircraft for this operation. If a laptop is required during the update, make sure the battery has sufficient power to last the whole process or have it connected to an external power supply.

Figure 16-92. Database Units

To use the floppy disk drive, the FMS database must first be loaded onto a computer and then written onto the disks. These disks are then inserted into the disk drive and prompts and the CDU will provide the necessary prompts for the update.

To use the Ethernet port located on top of the pedestal, the FMS database must first be loaded onto a laptop computer. This port will also accept IFIS information such as Jeppesen charts, airways, airspace, etc. This information must be on the laptop computer or in the laptop CD-ROM drive. The laptop computer and standard Ethernet cable are then connected through this port to either upload or download information. The use of Rockwell Collins software is required for this operation (CPAS3000).

To use the computer port (PCD-3000) located on the sidewall of the pedestal, the FMS database must first be loaded onto a laptop computer. The laptop computer and a special cable are then connected through this port to either upload or download information. The use of

To use the USB port (DBU-5000), the FMS data and IFIS data must first be loaded onto a computer and then moved to a USB drive. The USB device must not have preinstalled software which manages passwords or security, as this can interfere with the proper loading of


16-51

16 AVIONICS


16 AVIONICS


the database. If Jeppesen charts are involved, it is recommended to have a device at least 1GB in size. This drive is then plugged into the USB port in the aircraft. The generated prompts are displayed on the CDU. In this case the laptop does not need to be connected to the aircraft.

PLAN The flight plan will be loaded on the FPLN page. ORIGIN, DESTination, and fixes along the route of flight may be entered. Instrument Departures or Arrivals may be loaded as necessary.

PERFORMANCE INITIALIZATION

FMS INITIALIZATION The FMS must be initialized prior to each flight. The initialization may be accomplished using the following acronym: V – Verify FMS database coverage and effective dates I – Initialize FMS position P – Plan the flight (build the flight plan) P – Performance initialization For further explanation of these steps, refer to the FMS quick reference guides and FMS manuals.

VERIFY Verify the coverage of the database and verify the currency of the database. Flight with an out of date database is allowed, but the use of FMS / GPS dependent procedures are not authorized.

INITIALIZE Initialize the FMS position, or verify that the current position is correct. This position needs to be in a latitude / longitude format and can be retrieved / verified using airport reference point (ARP), a pilot defined point or the GPS.The GPS should be able to update the system quickly unless the aircraft was moved a significant distance (>40nm) with the FMS inoperative or the FMS was removed and replaced. This step will consist primarily of verifying the known position as opposed to actively entering the position.

16-52

Performance is initialized by entering the desired weights for passengers, cargo, fuel, etc. The CRZ ALT is an optional entry and helps the unit forecast a descent point later in the flight. CRZ ALT does not change any fuel calculations when changed or updated.

VERTICAL NAVIGATION The FMS-3000 is capable of creating and displaying a descent profile or a glidepath to comply with crossing altitude restrictions issued by ATC, or an associated instrument procedure. The Flight Guidance System is able to use this information to capture and track the computed glidepath. VNAV altitude restrictions are displayed in magenta along the right side of the LEGS page (Figure 16-93). A VNAV altitude will be automatically entered if it is part of a database derived procedure. The pilot can manually insert an altitude associated with any waypoint. Once an altitude restriction is inserted either automatically or manually, the FMS will generate the associated glidepath. The glidepath will be displayed at the appropriate point. As long as the proper conditions are met, the FGS will capture and track the vertical glidepath. The conditions are as follows: • The altitude must be entered into the LEGS page • The VNAV mode of the FGS must be selected (indicated by a “V” prior to the active vertical mode) • The Preselected Altitude must be set at, or beyond, the VNAV altitude


MSG

ACT LEGS 2/6 / o 9.5NM 307 HUT ---/----THEN - DISCONTINUITY /o 3.0 FEBIT // ---/ 3600A / /o / o 6.0NM / 054 0.0 CEPGA // ---/ 3600A /o / 144o 6.0NM 3.0 FAXIM // ---/ 3100A -----------------------LEG WIND> [ [ LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-93. Active Legs Page with VNAV Altitudes

The default VNAV glidepath is a 3.0˚ descent angle unless otherwise published in an instrument procedure. The pilot has the ability to modify this angle on every leg except for the final approach segment between the Final Approach Fix (FAF) and the Missed Approach Point (MAP). The FMS may create an angle other than 3.0˚, if required. The glidepath is based on aircraft position relative to the associated waypoint, a commanded vertical directto, or the associated waypoints position relative to a prior waypoint with an altitude restriction. When two or more waypoints in a flight plan have altitude restrictions, and they are sufficiently close in proximity to each other the FMS will compute the best glidepath to meet the requirements of all altitude restrictions. Instead of flying a 3.0˚ path to a waypoint, leveling off, and then flying another 3.0˚ path to the next waypoint, the FMS will adjust the paths to varying angles resulting in a continuous descent. This is sometimes called “smoothing” the descent.

A magenta Top Of Descent (TOD) circle will appear on the display maps to indicate the projected point where this descent will occur. The TOD point will indicate when the vertical deviation indicator nears the center position on the vertical deviation scale (Figure 16-94). This indicator is sometimes called the “snowflake” or “star”. As with Glideslope operations, these GPS Glidepath operations will only capture VNAV when initially below the projected angle. If the aircraft is already passed the descent point, manual intervention is required to place the aircraft in a position where the FGS can capture the glidepath. When the FGS captures a glidepath, the vertical mode will be annunciated as VPATH when NAV is selected or VGP when APPR is selected (Figure 16-95). VPATH will allow the FGS to level at either the preselected altitude or VNAV altitude, whichever it encounters first. It is necessary to be aware of the armed altitude mode when accomplishing this maneuver. ALTS indicates that VNAV will reach and level off at the preselected altitude even though there may be multiple step downs in between. This indicates that smoothing the descent is possible and an intermediate level off is not required. ALTV indicates that VNAV will reach and level off at the next VNAV altitude posted in magenta above the VSI. This indicates that smoothing the descent is not possible and the aircraft must accomplish an intermediate level off. Another TOD will appear indicating where the descent will begin if there is another altitude in the FMS. The use of NAV and VNAV should be used when flying enroute VNAV and when flying an approach to MDA This selection does not include localizer based procedures which are flown with a NAV-to-NAV capture function of the FMS. These approaches require the APPR mode for the NAV-to-NAV function to operate correctly.


16-53

16 AVIONICS


16 AVIONICS


any manually chosen vertical mode (VS or FLC). When VNAV is selected, the altitude preselector is then placed at the highest authorized altitude and the FGS will level off at each intermediate VNAV altitude. Once leveled off at the intermediate altitude, FLC will arm indicating there is another climb. Passing the altitude restricted fix, FLC will become the active vertical mode at the aircraft’s current indicated speed. The pilot must now change the FLC speed and aircraft power for the climb. The aircraft will level off at the next altitude restricted fix and FLC will arm again. This process will be repeated until the aircraft levels at the altitude shown on thepreselector. The aircraft is not allowed to go beyond the preselector setting.

Collins

FMS

VPATH ALTS

180 160

4

20

700

2 1

10

DN

4000

3 000

185

600 60

14 1 0

6 540 20 1

10

120

400

2 4

100 24

TOD

FMS1

W

JABAN

21

30

DTK 251 RALPE 5. 2NM

30.16IN

251

RALPE 10

FORMAT >

TOD

<

5 TERR RDR

< PRESET VOR1

>

TFC > < ET 01:42 COM1 121.800

ATC1 4336

UTC 14:41

RAT 15 oC

125.250

COM2

Collins

BRT DIM

Collins

FMS

VPATH ALTS

185 FMS

700

2 1

10

600 60

14 1 0

6 540 20

APPR FMS

400

4

3 000

185

2

24

TOD

30.16IN

251

4000 4

4

100

APPR + VNAV

1000

W

Figure 16-95. VNAV Modes

21

30


VGP

1

10

120

FMS1

Collins

4

20 DN

NAV +VNAV

4000

3 000

185

160

4000 4

VPATH ALTS

180

3 000

JABAN

GLOBAL POSITIONING SYSTEM (GPS)

<

< PRESET

RALPE TERR RDR

TOD

VOR1

>

TFC > < ET 01:42 COM1 121.800

ATC1 4336

UTC 14:41

RAT 15 oC

COM2

125.250

BRT DIM

Figure 16-94. VNAV Top of Descent

Additionally, VNAV can be used during an altitude restricted climb. The FGS will be in NAV and VNAV modes and never in APPR mode. The same three conditions mentioned for a VNAV descent apply here too. The initial climb from the airport will be accomplished by

16-54

4

R

The global positioning system (GPS) provides worldwide navigation via signals received from orbiting satellites. The GPS receiver is located in the nose avionics bay and is labeled GPS-4000( ) (The parentheses will contain either an “A” for standard GPS or an “S” for WAAS GPS). Using an antenna mounted on the top of the fuselage, it will track and monitor up to 12 satellites to provide a three dimensional position for the FMS and the Terrain Awareness and Warning System (TAWS). The GPS 1 and optional GPS 2 systems are controlled by the CDU(s).


<

FORMAT > 5

<

10

R

The FMS’s will default to GPS navigation sources as the primary reference for their position. Whether they are still enabled and part of the navigation can be seen with a few pages in the CDU Index (IDX) page. The GPS Control page will indicate whether the GPS sensors are enabled for navigation use, and will indicate the difference between the GPS position and the calculated FMS position (Figure 16-96). The PROGESS page on the CDU displays the current navigation sources used by the FMS to determine current position (Figure 16-97). The PROGRESS page shows a label on the bottom titled NAVIGATION. In this example the NAVIGATION area indicates that the system is using VOR, DME and GPS. Should the GPS malfunction or lose its Receiver Autonomous Integrity Monitoring (RAIM) the GPS label would be removed from the NAVIGATION line. If the GPS portion of the position begins to malfunction, a message will appear on the CDU. Some examples of GPS messages are as follows:

16 AVIONICS


GPS CONTROL GPS1

POS DIFF / 322 /0.4

GPS2

/ 322 /0.3

SAT DESELECT

-DEST KDEN

MSG

APPR RAIM AVAILABLE

ETA 12:16

-----------------------
DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

[

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

/

SP

Figure 16-96. GPS CONTROL

PROGRESS LAST DIST KIRLE 2.3

GPS – FMS Disagree (indicates the computed FMS position is different than the GPS position by a selected amount)

1/2 FUEL-LB / 0

ETE

TO

DBL

19.6

/ / 0:04

/ 2440

72.3

/ 0:16

// 2300

193

/ 0:42

/ 1980

NEXT

BASEE DEST

KDEN

GPS Not Available (indicates the FMS is not using the GPS for position information)

ALTN

----

---

-:--

-----

NAVIGATION

VOR/DME1 GPS

NO GPS RAIM (indicates the FMS is using the GPS but the GPS position is degraded) As with any approved GPS navigation receiver, this system allows the check of integrity and accuracy through certain pages in the CDU. For a RAIM prediction it is necessary to navigate to the Index page of the CDU and choose GPS CNTL. On this page it is possible to enter a desired airport and ETA. The RAIM system will then indicate RAIM availability 15 minutes before to 15 minutes after that entered time. The default entry for the airport line will automatically contain the DESTination airport. ETA will be an active number based on the loaded flight plan and current ground speed.

MSG

[

[ LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-97. PROGRESS


16-55

16 AVIONICS


INTEGRATED FLIGHT INFORMATION SYSTEM (IFIS)

or via the original floppy disk drive located at the aft end of the pedestal. If DATABASE UNIT is selected, the floppy disk drive can now accept diskettes and the CDU is used to update the FMS.

The Integrated Flight Information System IFIS-5000 is a part of the ProLine 21 architecture to provide extra information storage, increasing the available display features. The added items known as Enhanced Maps (EMaps) are displayed only on the MFD and include geographic / political boundaries, airways (high and low), and airspace. Optionally, the IFIS system can also display downloaded graphical weather (GWX), and Electronic Charts (E-Charts).

If IFIS DATA is chosen, then the disk drive is not active and the database loads must occur through the Ethernet port. Once IFIS DATA is selected, the switch labeled ENABLE / DISABLE must be enabled to prepare the aircraft avionics to accept data. A laptop is used and an Ethernet cable is connected between the computer and the covered IFIS DATA port. Software from Rockwell Collins will organize and coordinate the upload of data from the laptop to the aircraft. Refer to the CPAS-3000 Collins software manual for appropriate dataload order and instructions.

The main storage unit is the File Server Unit (FSU-5000) located in the empennage avionics shelf. This contains the memory needed for all the display options and outputs information only to the MFD via a fast Ethernet bus. This unit also receives inputs from a graphical weather system, FMS(s), database update unit and the pilot’s Cursor Control Panel (CCP) (Figure 16-98). There are two kinds of database update units. One form of database update allows for the update of all Collins related items including FMS(s), E-charts, E-maps, graphical weather and maintenance items. Alternately, a disk drive may be used to update only the FMS(s). This DBU-4100 contains an Ethernet port and two switches that are used to prepare the aircraft to accept the data (Figure 16-99). The switch labeled FMC Load will chose whether to update the FMS data via this Ethernet port

16-56

Another type of database unit, the DBU-5000, uses two USB ports located at the aft end of the pedestal (Figure 16-100). Either port is used to update the FMS(s), E-charts, E-maps, graphical weather and/or maintenance items. Once the databases are loaded onto the USB device from a computer it is connected to one of these ports. The remainder of the database load is controlled through the MCDU MENU line key on the CDU Index (Figure 16-101). Pressing the DBU option will allow the CDU to query the aircraft and the USB device to see what files are available for loading. After the load is complete the CDU can be exited to the main Index page and the USB device can be disconnected and used for the next database cycle. The two USB ports are to be used only for database loading and will not support external USB devices.


16 AVIONICS


Collins

ITT

26

TORQ

0

FIRE

KEGE 11-1

CCP

MENU ADV

0

PROP

0.0 0

NI

1980 98.5

0 0

ITT

734

TORQ

2000

AFX

46

FF PRESS OIL TEMP oC

430 120 73

AIRPORT

DATA

MFD

P U SH

T

S

ELEC

Collins

<

<

<

ACT FPLN ORIGIN

KICT

1/4 DIST

452

ACT FPLN

BRT DIM

DEST

ORIGIN

KDEN

ROUTE

KAPA

ORIG RWY

ORIG RWY

TO

FMC 1

FMC 2

VIA

TO

DIRECT ICT ------------------
PERF INIT>

[

[

KDEN ALTN

PLANT2

DIRECT ICT ------------------
DEST

452

ROUTE

KAPA

VIA

1/4 DIST

KICT

ALTN

PLANT2

PERF INIT>

[

[

ETHERNET CDU

CDU ETHERNET

FSU-5010 E-CHARTS E-MAPS GWX

DA ATA LOADER LOAD

CMU-4000 OR RIU-40X0

OR

XMWR-1000

COMMUNICA ICA ATION SYSTEM (VHF F, HF F,, ETC.)

XM Satellite Antenna

RF LINK DA ATALINK PROVIDER P (ARINC) INFORMA ATION PROVIDER (Universal)

ATHER THER UNIVERSAL WEA (GWX-5000)

XM WEA ATHER TH THER (GWX-3000)

Figure 16-98. IFIS Block Diagram


16-57

16 AVIONICS


INDEX
Figure 16-99. Ethernet Database Unit

Regardless of which dataloader is installed, the available subscriptions are listed in Figure 16102. Collins will provide the FMS and Enhanced Map (E-Map) databases through internet download or a shipment of CD’s. Jeppesen will provide the Electronic Chart (EChart) database through a shipment of CD’s only (no internet download). Finally, Hawker Beechcraft will provide the electronic checklist through an internet download. Although not specifically a part of the IFIS system, the electronic checklist will be uploaded through the same dataloader units discussed earlier. With each revision of the aircraft AFM that affects the checklist, it is the operator’s responsibility to update the electronic checklist manually or download a new version from Hawker Beechcraft.

Figure 16-100. USB Database Unit (DBU-5000)

16-58

MSG

1/2 GPS POS> FREQUENCY>

FIX>

HOLD>

PROG>

SEC FPLN> [

EXEC

MCDU MENU
GPS POS>

MSG

[

[

EXEC

Figure 16-101. MCDU Menu.

CURSOR CONTROL PANEL (CCP) The primary pilot interface with the IFIS system is accessed through the Cursor Control Panel (CCP) located on the pedestal (Figure 16-103). The left most section is used to enter and manipulate menus that appear on the MFD. The center section is used to store MFD display options to more quickly retrieve a desired display setup. The right most section contains a joystick and input buttons to control the E-Charts and downloaded weather. The memory keys are used to store the main MFD line select key format options. They do not store IFIS related map selections such as E-Maps or E-Charts. The selected Upper Format, Lower Format, Terrain or Radar, and TCAS options are stored. When the appropriate selections are made, press and hold the desired memory key until STORE is indicated on the MFD. Releasing the memory key will display a STORE COMPLETE (Figure 16104). This can be repeated for each of the three memory keys. To retrieve the selected options


16 AVIONICS


FILE SERVER UNIT (FSU) JEPPESEN E-CHARTS (CD) - 14 DAYS COLLINS E-MAPS (DOWNLOAD) - 28 DAYS GEO-POLITICAL (DOWNLOAD) - AS REQUIRED GRAPHICAL WX DATABASE (DOWNLOAD) - AS REQUIRED

ETHERNET BUS DATALOADER FLIGHT MANAGEMENT COMPUTER (FMC)

MAINTENANCE DIAGNOSTIC COMPUTER (MDC)

COLLINS FMS NAV DATABASE (DOWNLOAD) 28 DAYS

HAWKER BEECHCRAFT MFD CHECKLIST (DOWNLOAD) AS REQUIRED

SIMULTANEOUS

FMC 1

FMC 2

INDIVIDUALLY DATABASE UNIT (DBU) diskettes

Not Applicable with DBU-5000

Figure 16-102. IFIS Dataload Block Diagram


16-59

press and release the desired memory key and the MFD will change to the stored settings. MENU

ESC

MENU ADV

STAT

MEM 1

MENU ADV knob on the CCP. After the cursor is at the desired position, another press of F the line select key will change the value as will rotating the DATA knob on the CCP.

CHART Collins

DATA

MEM 2

P US H

516 26

T

S

1050 0

ITT

ITT

ELEC

MEM 3

ZOOM

Collins

TORQ TORQ

3.4 0

62.2 0.0

FIRE

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE

130 FF FF 430 750 0 122 PRESS PRESS 80 0 OIL 120 OIL o 49 TEMP°C 46 TEMP C 112 73

Figure 16-103. CCP FMS FMS1 DTK 275 ICT TTG --:-17.5NM

S

191

21 ABOVE

15 24

STORE COMPLETE

PPOS

PPOS

< GEO-POL 50 ON OFF

10

MAP SRC

<

<

FMS1 25

<

FMS2

KBEC <

16 AVIONICS


ICT

< AIRSPACE ON OFF < AIRWAYS HI LO OFF

TAS

0

SAT 25 oC

ISA +13 oC BRT DIM

Figure 16-104. MFD Store Complete Figure 16-105. Geo-Politcal Overlay

Enhanced Maps (E-MAPS) The IFIS system contains Collins provided data with certain enhanced map features. These include geographic / political boundaries, airspace and airways (high and low). The geographic / political option (GEO POL) will overlay state and country boundaries on the MFD display. The location of international boundaries on the overlay must not be used as an accurate representation of true boundary position. The GEO-POL overlay should only be used for information. This overlay is accessed by pressing the MENU button on the CCP when a PPOS map or PLAN map is in view on the MFD. Moving the cursor to the GEO-POL option will allow turning the overlay ON or OFF (Figure 16105) The cursor can be moved by pressing the adjacent line select key or by rotating the

16-60

The airspace option will overlay certain airspace boundaries. The airspace boundaries include Class A and B airspace along with CTA and TMA/TCA airspace. Airport related boundaries are shown with a solid magenta outline. Additionally, restricted and prohibited airspace is shown with a dashed magenta outline. The vertical limits and identifying marks of the airport or restricted/prohibited areas are not shown on the MFD. They must be used as information only and not to navigate or stay clear of these areas. The overlay is accessed with the MENU button on the CCP with the PPOS map or PLAN map displayed on the MFD. As discussed earlier, moving and manipulating the cursor to the Airspace option will allow turning the overlay ON or OFF. (Figure 16-106).


Collins

1050 0

ITT

ITT

516 26

TORQ TORQ

3.4 0

62.2 0.0

FIRE

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE

16 AVIONICS


Collins

130 FF FF 430 750 0 122 PRESS PRESS 80 0 OIL 120 OIL o 49 TEMP TEMP°C C 112 46 73

1050 0

ITT

ITT

516 26

TORQ TORQ

3.4 0

62.2 0.0

FIRE

PROP PROP

1740 1980

N1I N

106.0 98.5


ITT ITT

830 734

TORQ TORQ

AFX FIRE

110.0 2000

V3

V140 V53 2

V7 4

V73-2

4

V7

V35 V7

7 2

V77-53

ABOVE

2

4-7

PPOS MAP SRC

54

V35

<

FMS1 V12

0

FMS2

V77-53

2

25

V77

<

V3

FMS2

4 35

90

V7

FMS1 25

V

PPOS

< GEO-POL 50 ON OFF

V1

24

<

21

V73

PPOS MAP SRC

56

15

24

PPOS

< GEO-POL 50 ON OFF

0-2

V77-53

ABOVE

15

V19

191

S

<

FMS1 DTK 275 ICT TTG --:-17.5NM

21

2

191

V516

V53

ICT TTG --:-17.5NM

6

4

S

DTK 275

V25

V7

V190

FMS1

7

56

V190

0

SAT 25 oC

< AIRWAYS HI LO OFF

ISA +13 oC

V12 V77

V12

ICT

V350

V5

V10-23 4 V502

V73

TAS

< AIRSPACE ON OFF

V73

< AIRWAYS HI LO OFF

KBEC ICT

< AIRSPACE ON OFF

V1 3 V12 2

32

KBEC

V77

TAS

0

V280

V280

V280 V234

SAT 25 oC

ISA +13 oC

BRT

BRT

DIM

DIM

Figure 16-106. Airspace Overlay

The airway feature will superimpose all the selected airways on top of the current MFD map to help orient their positions. Only the airway is labeled and not the intersections. Once the airway is loaded in the FMS the intersection names will appear for that airway only. This overlay is accessed by pressing the MENU button on the CCP when a PPOS map or PLAN map is in view on the MFD (Figure 16107). As discussed earlier, moving and manipulating the cursor to the Airway option will allow selection of HI / LO / OFF. The overlay selections are the same for the PLAN map with the exception of a Graphical Weather (GWX) option. The GWX overlay will be discussed later.

Figure 16-107. Airways Overlay

STATUS PAGES The File Server Unit (FSU) contains status pages that indicate settings and configurations for the IFIS system. Pressing the STAT key on the CCP will display the last viewed page (Figure 16-108). The DATABASE EFFECTIVITY page indicates the current dates of each installed item. If a database is out of date the affected line will be yellow. The CCP MENU ADV and PUSH SELECT knobs are used to move the cursor and display more information for the selected database in the lower box. Pressing the CCP MENU key will display the status menu options (Figure 16-109). Using the CCP MENU ADV and PUSH SELECT knobs allows for the selection of another status page. One example, is the optional Electronic Chart subscription page (Figure 16-110). On this page the pilot can enter a Jeppesen provided Access Code and be able to instantly retrieve more charts. This capability can be used when a one-time flight is planned outside the current chart coverage. It is important to note that electronic chart coverage is a sepa-


16-61

Collins

rate subscription than the FMS database and may not cover the same regions. 1050 0

ITT

ITT

516 26

Collins

TORQ TORQ

3.4 0

FIRE

62.2 0.0

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE


CHART SUBSCRIPTION

1050 0

ITT

ITT

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TORQ TORQ

3.4 0

FIRE

62.2 0.0

PROP PROP

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N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE

130 FF FF 430 750 0 122 PRESS PRESS 80 0 OIL 120 OIL o 49 TEMP°C 46 TEMP C 112 73

SUBSCRIPTION NUMBER VNX12ABCD5AB3A1C

REGIONS ENABLED LATIN AMERICA SOUTH AMERICA USA - 48 STATES EUROPE

DATABASE EFFECTIVITY DATABASE

BEGIN

END

FMS 1 NAV 11 MAY 06 07 JUN 06 CHARTS 16 JUN 06 05 JUL 06 AIRSPACE 08 JUN 06 05 JUL 06 AIRWAYS 08 JUN 06 05 JUL 06 GEOGRAPHIC 25 JUL 05 30 SEP 06 POLITICAL 25 JUL 05 30 SEP 06 GRAPHICAL WX 01 MAY 06 N/A

STATUS

NOT CURRENT CURRENT CURRENT CURRENT CURRENT CURRENT CURRENT

< CHARTS

ADD REGIONS

DATABASE NAME: RCPL0612

ACCESS CODES

<

<

COVERAGE REGIONS: LATIN AMERICA SOUTH AMERICA USA - 48 STATES EUROPE

<

<

<

----------------------------------------------------------------

TFC <

TERR RDR < BRT

TFC <

DIM

DATE 27 JUN 06 BRT DIM

Collins

1050 0

ITT

ITT

516 26

TORQ TORQ

3.4 0

FIRE

62.2 0.0

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE


To return to an MFD map display press the STAT key again or one of the line select keys on the MFD bezel.

DATABASE EFFECTIVITY DATABASE

BEGIN

END

STATUS

FMS 1 NAV 11 MAY 06 07 JUN 06 NOT CURRENT CHARTS 16 JUN 06 05 JUL 06 CURRENT STAT MENU AIRSPACE 08 JUN 06 05 JUL 06 CURRENT AIRWAYS 08 JUN 06 05 JUL 06 CURRENT GEOGRAPHIC 25 JULDATABASE 05 30 SEP 06 CURRENT EFFECTIVITY POLITICAL 25 JUL 05CHART 30 SEP 06 CURRENT SUBSCRIPTION GRAPHICAL WX 01 MAY N/A CURRENT FCS06 DIAGNOSTICS

ELECTRONIC CHARTS (E-CHARTS) [Optional]

MAINTENANCE MAIN MENU FILE SERVER CONFIGURATION CHARTS DATABASE NAME: RCPL0612

<

COVERAGE REGIONS: LATIN AMERICA SOUTH AMERICA USA - 48 STATES EUROPE

<

TERR RDR <

TFC <

DATE 27 JUN 06 BRT DIM

Figure 16-109. STAT Menu

16-62

Figure 16-110. Chart Subscription (STAT Key)

Other STAT pages are the Flight Control System (FCS) Diagnostics, Maintenance Main Menu, and File Server Configuration. These pages mainly contain maintenance related information and are not necessary to be accessed by the pilot.

Figure 16-108. Database Effectivity (STAT Key)

<

16 AVIONICS


The IFIS system can optionally contain Jeppesen created instrument charts. These charts are loaded to the FSU through the dataloader discussed earlier. It is important to note that the chart coverage chosen is a different subscription than the FMS coverage. The charts will come from Jeppesen while the FMS database will come from Collins. See the dataloader section for more database information. Once a flight plan is entered in the FMS, the E-Chart feature will automatically be linked


to the airports in the Origin, Destination, and Alternate airport fields. To retrieve the desired charts, press the CHART key on the CCP (Figure 16-111). The MFD stores the last viewed image and will display that chart every time the CHART key is pressed until manually changed with the MFD chart menu . There are two items to note for this process. Even if the FMS procedure has changed, pressing the CHART key will display the last viewed chart not the new procedure’s chart. The pilot must change the chart manually to agree with the procedure in the FMS. Secondly, if the avionics have just been turned on, no chart will appear (the MFD does not have a chart stored in memory yet) and the pilot will have to choose the desired chart. Collins

1050 0

ITT

ITT

516

26

TORQ TTORQ ORQ

3.4 0

FIRE

KBJC 21-1

62.2 0.0

PROP 1740 1980 ITT ITT

PROP

830

734

N1I N

106.0 98.5

TTORQ ORQ TORQ

1110.0 10.0 2000

FIRE AFX

can be changed from this page. All other airport identifiers are retrieved from the FMS flight plan. Procedures loaded in the FMS will automatically link to this menu and the shortcut field will update with the new procedure and will show in magenta. There are airports where multiple charts exist for one runway (e.g., ILS Rwy 01 and Converging ILS Rwy 01). For these airports the shortcut field will be a white “SELECT CHART” and the pilot must press the PUSH SELECT key and choose the appropriate chart. It is important to note that the FMS will only contain one approach type for each runway. Even though the Converging ILS Rwy 01 may be chosen for chart display, that procedure will not be in the FMS database. Charts that have been manually selected will show in cyan. To exit out of the menu press the CCP ESC key. Collins

130 FF FF 750 0122 PRESS 430 80 PRESS 120 0 OIL OIL 49 TEMP°C 11732 46 TEMP oC 112 1050 PROP 1740 0 PROP 1980

ITT

ITT

516 26

ILS R RWY WY 29R

TORQ TORQ

3.4 0

FIRE

KEGE 11-1

62.2 0.0

NN1 I

106.0 98.5

ITT

ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE


AIRPORT

CHART MAIN INDEX FMS1 ORIGIN - KEGE AIRPORT [AIRPORT] DEPARTURE [GYPSUM 3 DEP]

<

<

<

TFC < TFC

BRT BRT DIM M DIM

ARRIVAL -APPROACH [] ANY CHART [] CHART NOTAMS -DESTINATION - KBJC ARRIVAL [RAMMS 5, TOMSN 4 ARRS] APPROACH [ILS RWY 29R] AIRPORT [AIRPORT, AIRPORT INFO, TAKEOFF MNMS] DEPARTURE [] ANY CHART [] CHART NOTAMS -ALTERNATE - KDEN ANY CHART [] ANY CHART [] CHART NOTAMS ALL OTHER AIRPORT - KHUT ANY CHART [GPS RWY 4] ANY CHART [] CHART NOTAMS ALL CHART DIMMING DAY NIGHT

<

TFC <

Figure 16-111. MFD Chart Display

Choosing the desired chart is accomplished by first pressing the CHART key and then the MENU key on the CCP (Figure 16-112). The CHART Main index is divided into the following areas; Origin; Destination; Alternate; Other airport. Only the OTHER AIRPORT

BRT DIM

Figure 16-112. MFD Chart Menu

The cursor is moved with the CCP MENU ADV knob. Once the cursor is over the desired entry two actions are possible with the PUSH SELECT feature on the CCP DATA


16-63

16 AVIONICS


16 AVIONICS


knob. A single press will choose the indicated chart for display on the MFD (e.g., the ILS Rwy 29R in the previous figure). Secondly, pressing and holding the PUSH SELECT feature will bring up a selection menu allowing the choice of every chart in that category. (e.g., all airport diagram charts, or all departure procedure charts, or all instrument approach charts, etc.) (Figure 16-113).

Collins

1050 PROP 1740 0 PROP 1980

ITT

ITT

516 26

TORQ TORQ

3.4 0

FIRE

KEGE 11-1

62.2 0.0

NN1 I

106.0 98.5

ITT

ITT

830 734

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110.0 2000

AFX FIRE


AIRPORT

Collins

1050

ITT

516

62.2

PROP

1740

N1

106.0

TORQ

830

TORQ

110.0

3.4

KBJC 21-1

130 FF 750 122 PRESS 80

ITT

<

OIL

FIRE

<

49 TEMP°C 112

AIRPORT

APPR OACH - KBJC

TFC <

ALL PRECISION APPR OACHES 21-1 ILS R WY 29R ALL NON-PRECISION APPR OACHES 23-1 28-1 28-2 29-1

VOR DME R WY 29L/R GPS R WY 29R GPS R WY 29L VOR DME RNAV R WY 29R

BRT DIM

Figure 16-114. MFD Chart Zoom Box <

<

TFF <

BRT DIM

Figure 16-113. MFD Chart Approach Index

After the chart is displayed, it is moved as needed using the CCP joystick to display areas that may be off the screen. An orientation button on the CCP will turn the chart clockwise 90 degrees. Pressing the orientation key again will return the chart to its original state. Additionally, there are two levels of zoom using the CCP ZOOM key. The first press will zoom into the area bounded by the green box (Figure 16114). Another press of the ZOOM key will return the chart to the original size. To return to the MFD map imagery, press the CHART key again or press one of the line select keys on the MFD bezel.

16-64

If the chart is geo-referenced, the aircraft position and orientation will be displayed using a magenta aircraft icon. (Figure 16-115). This indicates that the latitude / longitude positions on the chart agree with the GPS coordinate system, known as WGS-84. When the aircraft icon does not appear, two possible symbols will appear at the upper right corner of the chart. A magenta crossed-out aircraft symbol indicates the chart is not geo-referenced. A yellow crossed-out aircraft symbol indicates the chart is geo-referenced but GPS1 present position daa is not available. Chart NOTAMS are also available from the Chart Main Index when applicable. Caution should be exercised since these NOTAMS were loaded at the last database update which may have been 14 days earlier. This information does not receive updates from an active datalink. To enter the OTHER AIRPORT information, the cursor must be moved to that airport and then press PUSH SELECT. This allows for manual entry of the identifier by turning the


CCP DATA knob and advancing the cursor to the next letter with the MENU ADV knob. After the identifier is entered, pressing PUSH SELECT will enter the airport and allow the use of ANY CHART fields to retrieve the desired charts. This feature can be used to view airport or airport chart information when it is not part of the FMS flight plan or when the link between FMS and FSU has failed.

clockwise will display the Airport diagram. This is useful after landing where a single click clockwise from the approach chart will display the airport diagram and help with taxiway orientation. Collins

1050 PROP 1740 0 PROP 1980

ITT

ITT

516 26

TORQ TORQ

3.4 0

FIRE

KEGE 11-1

62.2 0.0

NN1 I

106.0 98.5

ITT

ITT

830 734

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110.0 2000

AFX FIRE


AIRPORT

CHART MAIN INDEX FMS1 ORIGIN - KEGE AIRPORT [AIRPORT] DEPARTURE [GYPSUM 3 DEP] Collins

1050 1740 40 PROP 17 0 PROP 1980

ITT

ITT

516 26

TORQ TORQ TORQ

3.4 0

FIRE

KEGE 11-1

62.2 0.0

NN1 I

106.0 98.5

ITT

ITT

830 734

TORQ TORQ TORQ

1110.0 10.0 2000

AFX FIRE

130 FF FF 430 750 0 122 PRESS PRESS 80 0 OIL 120 OIL o TEMP°C 112 49 TEMP C 11 46 732

AIRPOR AIRPORTT

<

<

ARRIVAL -APPROACH [] ANY CHART [] CHART NOTAMS -DESTINATION - KBJC ARRIVAL [RAMMS 5, TOMSN 4 ARRS] APPROACH [ILS RWY 29R] AIRPORT [AIRPORT, AIRPORT INFO, TAKEOFF MNMS] DEPARTURE [] ANY CHART [] CHART NOTAMS -ALTERNATE - KDEN ANY CHART [] ANY CHART [] CHART NOTAMS ALL OTHER AIRPORT - KHUT ANY CHART [GPS RWY 4] ANY CHART [] CHART NOTAMS ALL CHART DIMMING DAY NIGHT

<

TFC <

<

BRT DIM

TFC <

Figure 16-116. MFD Chart Menu

BRT DIM

Figure 16-115. MFD Chart Geo-Reference Symbols

Graphical Weather (GWX) [Optional]

At the bottom of the Chart Main Index is a two level Chart Dimming control. Setting the DAY option will display charts in a standard white background color. Setting the NIGHT option will change the white background to a cyan hue reducing the intensity of the MFD image during dark conditions.

There are two weather providers that will allow for the display of select weather maps. These two providers are not compatible and the aircraft will be configured for only one version. The XM weather provider uses a satellite downlink system and is available only for weather images within the US 48 Contiguous States. The Universal weather provider uses a COMM3 VHF datalink and is available for weather images for many parts of the world.

After a chart is displayed it can be changed using the procedures described earlier or using the DATA knob shortcut. By rotating the DATA knob clockwise or counterclockwise all the charts linked for the current airport can be viewed without having to navigate to the Chart Main Index. For instance, if the ILS Rwy 29R for KBJC is in view from Figure 16-116 one click counterclockwise will display the RAMMS 5, TOMSN 4 ARR chart or one click

As with all satellite or radio-based weather, the data provided should be used only with reference to onboard radar and appropriate preflight planning. All downloaded information is a view of past weather conditions and is not instantaneous. Some information may be


16-65

16 AVIONICS


more than 15 minutes old and unusable for appropriate weather avoidance.

Collins

1050 0

ITT

ITT

516 26

XM WEATHER (GWX-3000) The XM weather provider is labeled as the GWX-3000 system for the Collins IFIS. XM weather uses a satellite antenna collocated within the GPS antenna housing on top of the aircraft. The antenna is then connected to the XMWR-1000 unit located in the empennage avionics shelf. The XMWR-1000 receives the XM provided weather data and images on a continuous basis and sends the information to the File Server Unit (FSU) for potential display on the MFD. Refer to the IFIS-5000 Operator’s Guide for more detailed information.

TORQ TORQ

FIRE

3.4 0

The dedicated weather format is chosen from the FORMAT line select key on the MFD by choosing the GWX selection (Figure 16-118 ). This format is used for NEXRAD and all other XM weather images and information. The CCP is used to control all the overlays and position of this format.

16-66

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE


STORE COMPLETE PLAN

<

PLAN

GEO-POL 50 ON OFF

MAP SRC

<

FMS1 FMS2

< AIRSPACE ON OFF

GWX

<

ON OFF

< AIRWAYS GS HI LO0 OFF

Once images are available they are displayed in two MFD formats. For NEXRAD radar, weather returns can be displayed on a dedicated weather format or overlayed with the PLAN Map format. All other images can be displayed only on the dedicated weather format. To overlay NEXRAD on the PLAN Map format, first choose the PLAN Map format, then press MENU on the CCP (Figure 16-117). The lower right option allows for graphical weather (GWX) to be turned ON or OFF. This overlay depicts the FMS course along with NEXRAD returns to help anticipate radar returns along the route of flight. The age of NEXRAD information is displayed at the upper right portion of the PLAN map and should update every time a new NEXRAD download is received. Changing the range is accomplished with the DCP range knob. Changing the position of the map is accomplished using the MFD ADV key on the CDU to advance the map to each FMS waypoint.

PROP PROP

62.2 0.0

<

16 AVIONICS


0

TAS

SAT 15 oC

ISA +13 oC BRT DIM

Figure 16-117. MFD PLAN Map Weather Overlay Collins

1050 0

ITT

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N1I N

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830 734

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AFX FIRE


LOWER FORMAT > FORMAT PPOS PLAN

GWX

GS

0

TAS

0

SAT 15 oC

ISA +13 oC BRT DIM

Figure 16-118. MFD Dedicated Graphical Weather Format (XM Weather)


Pressing the CCP MENU key will display the XM graphical weather menu (Figure 16-119). The MENU ADV, DATA and PUSH SELECT knobs on the CCP are used to choose the applicable options. The TAF/METAR reports are textual only and are chosen by pressing the PUSH SELECT knob (Figure 16-120). Rotating the DATA knob will cycle through multiple pages, if they exist, as indicated by “Page 1 of 2” in the figure. The Origin, Destination, and Alternate airports are automatically retrieved from the FMS flight plan. The Other airport can be manually inserted as described earlier in the Chart Main Index. To exit out of the textual pages press the CCP ESC key.

indicating that for the first 18 minutes of flight the NEXRAD cannot be animated on the display. Once the animation is possible the AVAILABLE message will appear on the menu. Collins

516 26

TORQ TORQ

3.4 0

1050 0

ITT

516 26

TORQ TORQ

3.4 0

FIRE

PROP PROP

1740 1980

N1I N

106.0 98.5

62.2 0.0

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE

FIRE

62.2 0.0

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE


GRAPHICAL WEATHER OTHER - KICT

PAGE 1 OF 2 METAR METAR KICT O71456Z COR 15O1OKT 1OSM FEWO5O OVC25O 26/18 A3OO5 RMK AO2 SLP163 TO256O178 51O1O METAR KICT O71356Z 13OO8KT 1OSM FEWO41 OVC25O 24/18 A3OO4 RMK AO2 SLP16O TO244O183 METAR KICT O71256Z 13OO7KT 1OSM FEW25O 22/18 A3OO3 RMK AOK SLP156 TO222O178

Collins

ITT

1050 0

ITT

ITT


>

TFC >

GS

0

TAS

0

SAT 15 oC

ISA +13 oC BRT DIM

GRAPHICAL WEATHER TAF/METAR REPORTS ORIGIN KBEC DESTINATION KDAB ALTERNATE KICT OTHER [ KHUT ] NATIONAL MET REPORTS SIGMET AIRMET

Figure 16-120. MFD Metar Display

ANIMATED NEXRAD - AVAILABLE OVERLAYS NEXRAD OFF ON OFF ON ECHO TOPS OFF ON METAR AIRPORT IDENTS OFF ON SIGMETS OFF ON A/C FLIGHT INFO OFF ON OVERLAY LEGENDS GS

0

TAS

0

TFC <

SAT 15 oC

ISA +13 oC BRT DIM

Figure 16-119. MFD XM Weather Menu

The NATIONAL METerological REPORTS are also text only and are chosen with the PUSH SELECT knob. The Animated NEXRAD selection is available only after the XM system has downloaded at least three NEXRAD images. These are delivered approximately every 6 minutes

The available Overlays have ON or OFF selections that are controlled with the CCP. The METAR overlay will change the airport symbols to visually indicate weather conditions. The SIGMET overlay will indicate areas of SIGMET coverage with different colored boxes corresponding with the coordinates affected. The A/C FLIGHT INFO will display or remove the aircraft icon to help orient present position with displayed weather. The FMS course line is not viewable on the dedicated weather page. The last item, OVERLAY LEGENDS, defines what the colors and symbols represent on the dedicated weather page (Figure 16-121). Additionally, the ECHO TOPS overlay will include textual descriptions of storm intensity that are defined on the LEGENDS page.


16-67

16 AVIONICS


16 AVIONICS


age, the time below the label will turn yellow with a yellow box. The pilot cannot request a specific update since XM weather is designed to continuously receive weather information. Caution should be exercised when referencing the affected overlay for weather information. If an overlay is selected OFF then the label and time stamp are removed.

Collins

1050 0

ITT

ITT

516 26

TORQ TORQ

FIRE

3.4 0

62.2 0.0

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

GRAPHICAL WEATHER OVERLAY LEGENDS METAR NO DATA VFR MARGINAL VFR IFR LOW IFR

ECHO TOPS HAIL PROBABLE HAIL MESO MESOCYCLONIC TVS TORNADIC

AFX FIRE

SIGMET VOLCANIC ASH CONVECTIVE TURBULENCE ICING DUST STORMS OTHER NEXRAD RAIN PRECIP MIXED PRECIP SNOW PRECIP

2O

0

TAS

Figure 16-122. MFD Graphical Weather Time Stamps

TFC <

35O HAIL

GS


0

RADIO ID OQ8N5OCU SAT 15 oC

ISA +13 oC BRT DIM

Figure 16-121. Overlay Legends

Finally, the RADIO ID field is the XM subscription number. This is needed when the XM feature needs to be turned ON initially or reinstated after it fails to communicate with the satellite system. Each press of the CCP ESC key will remove one submenu at a time until all menus are removed and the dedicated graphical weather page is in view. The graphical weather page can be moved using the CCP joystick to the full extent of the US borders and is not limited by aircraft position or FMS waypoints. Additionally, each press of the CCP ZOOM key will provide three levels of zoom. Each level of zoom is indicated above the weather map (Figure 16122). The zoom levels are indicated with these labels: x1=Entire CONUS; x4 = ¼ of CONUS; x16 = 1/16 of CONUS. Time entries are also displayed above the weather map. The current UTC time is used to provide a reference for the age of each chosen overlay. Once an affected overlay exceeds a set

16-68

UNIVERSAL WEATHER (GWX-5000) The Universal weather provider is labeled as the GWX-5000 system for the Collins IFIS. Universal weather uses an additional VHF COM3 radio and an additional VHF antenna. The antenna is located under the empennage of the aircraft and is attached to a Collins Communications Management Unit (CMU4000) in the aft avionics shelf. The CMU handles all outbound and inbound COM3 VHF transmissions that are requested from the pilot through an additional CDU page. The COM3 system is not connected to the audio panels or audio controls in the cockpit. Optionally, the CMU unit is capable of datalink communications (e.g., ACARS or AFIS) using an HF, SATCOM and/or VHF radio. The Universal weather provider is a request only system. Each weather image or weather data is first requested by the pilot through the CDU datalink page. If the aircraft is within radio coverage of an appropriate groundbased station, the image or information is sent via VHF communication to the CMU unit. A CDU and MFD message will appear when the image is available for view. To access the CDU graphical weather page, press IDX MCDU MENU. On this page, a


message will remain active until all new images are viewed.

Datalink (DL) option is available that will show the Graphical Weather request page (Figure 16-123). The images shown only contain the graphical weather selection, but each page may contain other optional items such as textual weather, digital ATIS, received ATC messages, etc. Selecting the REQ field for GRAPHICAL WX, will display the available weather products (Figure 16-124). Navigating between the two available pages allows selection of the desired weather image. Pressing the left side keys will select the main image and turn it green. Pressing the right side keys will display a new page where the desired Region, Altitude, or Forecast time options can be set for the selected image. Once the selections are complete pressing the SEND line select key will initiate the CMU communication with an available VHF datalink station. The REQUEST STATUS option can be used to identify which images are still downloading and which images have been received. If the CDU is used for other functions while the information is downloading a “GWX RCVD” message will appear on the CDU message line. This

MCDU MENU
Once images are available they are displayed in two MFD formats. For U.S. NEXRAD radar, weather returns can be displayed on a dedicated weather format or overlayed with the PLAN Map format. All other images can be displayed only on the dedicated weather format. To overlay NEXRAD on the PLAN Map format, first choose the PLAN Map format and then press MENU on the CCP (Figure 16-125) The lower right option allows for graphical weather (GWX) to be turned ON or OFF. This overlay depicts the FMS course along with NEXRAD returns to help anticipate radar returns along the route of flight. The age of NEXRAD information is displayed at the upper right portion of the PLAN map and should update every time a new NEXRAD download is requested. Changing the range is accomplished with the DCP range knob. Changing the position of the map is accomplished using the MFD ADV key on the CDU to advance the map to each FMS waypoint.

DL DATALINK

WEATHER

DL

GPS POS>

EXEC LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

MSG

EXEC MFD ADV

MFD DATA

MSG

EXEC MFD ADV

MFD DATA

PREV

NEXT

DIR

FPLN

PREV

NEXT

DIR

FPLN

PREV

NEXT

1 2 3 A B C D E F G

CLR DEL

IDX

1 2 3 A B C D E F G

CLR DEL

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

TUN

4 5 6 H

BRT DIM

TUN

4 5 6 H

BRT DIM

/ +/ - V W X Y Z 0

SP

/

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

PERF

MFD MENU

FPLN

J K L M N

LEGS

DEP ARR

IDX

I

PERF

MFD MENU

DIR

7 8 9 O P Q R S T U

LEGS

DEP ARR

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-123. MCDU Datalink Pages (Universal Weather)


16-69

16 AVIONICS


16 AVIONICS


REQ GWX

DL

2/2

REQ GWX

DL

1/2

N AMERICA

WINDS/TEMPS

NE US

REGION>

REGION>

NEXRAD

FL340 / ICING

TOPS/MOVE

ALTITUDE> 42HR

FORECAST>

TURBULENCE

WX DEPICTION

RCVD

RCVD

SEND*

MSG

EXEC

DIR

FPLN

DEP ARR

LEGS

SEND*

PERF

MFD MENU

MFD ADV

MFD DATA

MSG

EXEC MFD ADV

MFD DATA

NEXT

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

TUN

4 5 6 H

BRT DIM

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

I

J K L M N

7 8 9 O P Q R S T U

/

SP

PERF

MFD MENU

PREV

I

LEGS

DEP ARR

/ +/ - V W X Y Z 0

SP

/

Figure 16-124. Datalink Weather Selections (Universal Weather)

The dedicated weather format is chosen from the FORMAT line select key on the MFD by choosing the GWX selection (Figure 16-126). This format is used for NEXRAD and all other Universal weather images. The image that appears will be the last viewed weather image. To change the selection, press the CCP MENU key to display the Universal weather menu page (Figure 16-127). The menu is organized with the most recently received image at the top. Older items may be on the next page with up to 50 total stored images. Once an image is past a selected effective time the entry will turn yellow to better indicate its age.

Collins

1050 0

ITT

ITT

516 26

TORQ TORQ

3.4 0

FIRE

62.2 0.0

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE


STORE COMPLETE

<

PLAN

GEO-POL 50 ON OFF

MAP SRC

<

<

PLAN

FMS1 FMS2

< AIRSPACE ON OFF < AIRWAYS GS HI LO0 OFF

GWX

<

ON OFF TAS

0

SAT 15 oC

ISA +13 oC BRT DIM

Figure 16-125. MFD_Plan Map Weather Overlay

16-70

Use the CCP MENU ADV and PUSH SELECT knobs to move the cursor and select the desired weather image from the menu. The displayed image and corresponding time of effectiveness will appear on the MFD. The image is static and cannot be zoomed in or moved around. If weather from an adjacent area is desired the appropriate image needs to be requested from the CDU and then viewed when received.


COMMUNICATION/ NAVIGATION SYSTEMS

Collins

1050 0

ITT

ITT

516 26

TORQ TORQ

3.4 0

FIRE

62.2 0.0

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

AFX FIRE


The Pro Line 21 avionics system uses either the Control Display Unit (CDU), or the Radio Tuning Unit (RTU) to tune the communication and navigation radios and the transponder. The CDU and RTU provide redundant control of all devices. Reversionary control is provided should one unit fail.

< UPPER FORMAT

RADIO SENSOR SYSTEM

LOWER FORMAT > FORMAT PPOS PLAN

GWX

TFC > UK GS

0

NEW GWX VAL 30JUL/1200Z ISSUED 30JUL/1000Z SAT 15 oC ISA +13 oC

WINDS/TEMPS TAS 0

BRT DIM

Figure 16-126. MFD Dedicated Graphical Weather Format(Universal Weather) Collins

1050 0

ITT

ITT

516 26

TORQ TORQ

3.4 0

FIRE

62.2 0.0

PROP PROP

1740 1980

N1I N

106.0 98.5

ITT ITT

830 734

TORQ TORQ

110.0 2000

GRAPHICAL WEATHER

AFX FIRE


2/3

IMAGES VALID NEXT AVAIL <-- PREVIOUS IMAGES <-N-PAC WINDS/TEMPS FL15O 3OHR 31OCT 2OOOZ O1NOV O1OOZ USA TURBULENCE FL15O 12Z 31OCT 12OOZ 31OCT 23OOZ NW-US NEXRAD 3OOCT 15O6Z 3OOCT 1512Z SW-US NEXRAD 3OOCT 1454Z 3OOCT 15OOZ NW-US TOPS/MOVE 3OOCT 143OZ 3OOCT 144OZ < USA ICING FL15O 3O HR 29OCT O95OZ 29OCT 12OOZ < W-PAC WINDS/TEMPS FL15O 3OHR 28OCT 2OOOZ 29OCT O1OOZ W-PAC TURBULENCE FL15O 12Z 28OCT 12OOZ 28OCT 23OOZ NE-US NEXRAD 27OCT 15O6Z 27OCT 1512Z SE-US NEXRAD 27OCT 1454Z 27OCT 15OOZ NE-US TOPS/MOVE 27OCT 143OZ 27OCT 144OZ NE-US ICING FL15O 3OHR 26OCT O95OZ 26OCT 12OOZ S-PAC WINDS/TEMPS FL15O 3OHR 25OCT 2OOOZ 26OCT O1OOZ NC-US NEXRAD 24OCT 15O6Z 24OCT 1512Z TFC < SC-US NEXRAD 24OCT 1454Z 24OCT 15OOZ NC-US TOPS/MOVE 24OCT 143OZ 24OCT 144OZ NC-US ICING FL15O 3OHR 23OCT O95OZ 23OCT 12OOZ --> MORE IMAGES --> GS 0 TAS 0 SAT 15 oC ISA +13 oC BRT DIM

Figure 16-127. Universal Weather Menu

The Radio Sensor System provides the control, displays, and sensors for VHF voice communication, HF voice communication (if installed), VOR/ILS/DME, ADF, transponder tuning, and TCAS II (if installed). The system consists of the Radio Tuning Unit (RTU-4220) located in the center instrument panel, and the Control Display Unit (CDU) which is located in the pedestal. The RTU is considered to be the primary method of tuning, with the CDU functioning as the secondary method of tuning. The tuning capabilities of the CDU are accessed by using the TUNE page as described earlier. If Dual CDUs are installed, only the left CDU (CDU 1) has radio tuning capabilities. A RTU/CDU TUNE switch is located on the reversionary panel (Figure-128). When this switch is in the NORM position, radios may be tuned using either the RTU or the CDU. Should the RTU become inoperable, tuning the #1 radios (COM1, NAV1, ADF1, etc) will not be possible. If the CDU should become inoperable, tuning the #2 radios (COM2, NAV2, ADF2, etc.) will not be possible. Moving the RTU / CDU TUNE switch to the operating unit (CDU or RTU) will return full tuning capability. If the RTU is the only unit still operating, selecting RTU will allow that unit to tune both the #1 and #2 radios. If the CDU is the only unit still operating, selecting CDU will allow that unit to tune both the #1 and #2 radios.


16-71

16 AVIONICS


16 AVIONICS


mounted on the lower fuselage (Figure 16130).

VHF Navigation System

Figure 16-128. RTU / CDU TUNE switch

If radio tuning capability is lost from both the RTU and the CDU, the EMER TUNE annunciator-switch, located on the reversionary panel, may be pushed to tune the No. 1 COM to the emergency frequency 121.5 MHz (Figure 16-29). Activation of the switch is indicated by the illumination of the annunciator, 121.5, located on the switch.

One NAV-4000 and one NAV-4500 navigation receivers (NAV 1 and NAV 2) provide VOR and Localizer navigation capabilities in the frequency range of 108.00 through 117.95 MHz in 25 kHz increments. The NAV-4000 also contains the ADF receiver. As an option, the aircraft may be equipped with two NAV-4000 units for a dual – ADF installation. The NAV 1 and NAV 2 antennas are located on either side of the vertical stabilizer. The CDU has the capability of automatically tuning the VHF NAV receivers in order to improve the calculation of airplane position by the FMS. This feature has no effect on current procedural navigation aids and will choose only those VORs or ILSs that provide the best signal reception and position information. This auto tune function is selected from the navigation portion of the CDU TUNE page. The auto tune function is automatically cancelled if any of the following occur. • DME HOLD is selected • A NAV receiver is manually tuned using either the RTU or the CDU • The FMS is deselected as a NAV source • A NAV receiver fails

Figure 16-129. Emergency Frequency Button

VHF Communications System Two VHF-4000 communication transceivers (COM 1 and COM 2) provide two-way communications in the frequency range of 118.000 through 136.975 MHz in 25 or 8.33 kHz increments. These units are located in the forward avionics compartment (See Appendix A). The COM 1 antenna is mounted on the top of the fuselage while the COM 2 antenna is

16-72

If a malfunction occurs when the auto tune function is active, it may be manually disabled using the RMT TUNE switch located on the reversionary panel (Figure 16-131). Moving this switch from the NORMAL position to the DISABLE position will disable the auto tuning function of the CDU. This includes the auto tune feature discussed here and localizer auto tuning after loading an approach. In other words, having the RMT TUNE switch selected to DISABLE requires the pilot to tune the NAV radios manually for all subsequent operations.


TCAS II (OPT) / TRANSPONDER (OPT)

SATELLITE SKYWATCH ANTENNA

16 AVIONICS


COMM 1 ANT ELT ANTENNA (RIGHT SIDE OF FIN)

PHONE

GPS ANTENNA

LH, RH NAV ANTENNA

GLIDESLOPE ANTENNA

DME NO. 1 ANT TRANSPONDER ANTENNAS

RADIO ALTIMETER

COMM 3 (OPT)

ADF ANTENNA

MKR ANTENNA TCAS II ANT (OPT)

COMM 2 ANT

DME NO. 2 ANT

Figure 16-130. Antennas

The ADF antenna is mounted on the lower fuselage. A second ADF receiver is optional.

Distance Measuring Equipment (DME)

Figure 16-131. RMT Tune Switch

Automatic Direction Finder (ADF) The automatic direction finder (ADF) allows navigation using non-directional beacons (NDBs). As mentioned in the VHF Navigation section, the ADF is part of the NAV-4000 unit and does not have a separate line replaceable unit (LRU). Magnetic bearing to NDB stations is displayed on the PFD and MFD with selectable bearing pointers. ADF receivers are tuned using the CDU tune page or the RTU.

The DME-4000 receiver determines slantrange distance, groundspeed, and time-to-station for the navaid tuned on the respective Nav receiver. A single DME-4000 is standard but it contains three channels. Channel 1 is the DME for NAV 1, Channel 2 is the DME for NAV 2 and Channel 3 is a “blind” channel that the FMS can use to tune any frequency it chooses. Should the optional second DME4000 be installed, Channel 1 for each unit will be the DME for NAV 1 and NAV 2. Channels 2 and 3 for each DME-4000 will be “blind” channels that the FMS can use to tune any frequency it selects. DME information is shown on the PFD (Figure 16-132) when the ground-based navigation source is selected for display. If only FMS is selected, then DME will not be displayed in the active NAV location. In that case, a bearing


16-73

1

wheels. The Mode S does provide an “onground” or “in-air” message for other TCAS operators and ground based ATC radar, but this does not control the actual mode of the transponder. Additionally, Elementary or En1 hanced surveillance transponders are available as options including Flight ID which can be entered with the RTU or CDU (Figure 16135). The antenna is located on the lower fuselage. In the optional TCAS II installations, Dual TDR-94D Diversity Mode S transponders are installed indicating that they have an antenna on the top and bottom for each transponder.

pointer will have to be displayed to 1 get ground-based DME. The DME receivers are tuned using the CDU tune page or RTU. Each DME receiver can also be automatically tuned by the FMS as described in the VHF Navigation section. The DME antenna is mounted on 1 the lower fuselage. 1 Localizer DME

LOC1 109.75

VOR Bearing Pointer DME

CRS 235 IESJ 0. 8NM

TERM

FMS1

<

S

S

DTK 251 (6935) 0. 8NM

COM1 121.800

A

4336

1


< ET

125.250 A

DME With FMS

Figure 16-132. PFD DME Displays

A DME hold function allows retention of the currently tuned DME frequency after changing the active frequency on the respective VHF Nav radio (Figure 16-133). This can be selected by the DME HOLD button on the RTU or the DME HOLD option in the CDU.

4336

1

C RECALL / 134.250 NAV1 / 113.80/ICT DME1 / HOLD 116.80 ATC1

121.700 // RECALL

123.875 MK-HI NAV2 110.30 / / DME2

HOLD

3144 ADF

412.5 [

[ BRT

Collins

IDENT

COM 1

126 . 700

ATC 2

Dual TDR-94 Mode S transponders provide ATC secondary radar returns. The transponder code selection is done through either the CDU tune page or the RTU. To activate the transponder the ATC switch must be moved to either 1 or 2 as desired (Figure 16-134). This switch must be moved prior to departure since this operation is not controlled by weight on

125.250

122.875

113 . 80

ATC Transponder

16-74

w

21

20. 8 H

COM1 121.800

DME Without FMS

24

CRS 251

3

< ET

VOR1 113.80

S

V 4.1NM SXW

< PRESET

30

VOR Bearing Pointer DME Not Received

15

VOR1

<

< PRESET

15

<

16 AVIONICS


4176

118 . 200

NAV 1 MK-HI 116.80H

25 SEL

DME--H

110 . 20 ADF

1/2

3 3 2 .0 ANT BFO

Figure 16-133. DME Hold Selection and Images


16 AVIONICS


AUDIO SYSTEM The all-digital audio system manages the communication and navigation systems. An audio control panel, adjacent to each pilot’s PFD, enables individual audio control (Figure 16-136).

IDENT

A press-to-transmit (PTT) button on the outboard horn of each control wheel facilitates communication transmissions. A microphone jack on each sidewall allows connection of headset microphones. Two speakers in the cockpit ceiling repeat audio heard through the headphones (Figure 16-137). The speaker volume for audible warnings cannot be muted. Additionally, each pilot’s oxygen mask contains a microphone.

DME--H

Passenger Address System

Figure 16-134. ATC Transponder Switch BRT

Collins

COM 1

118 . 200

126 . 700

25 SEL

NAV 1

113 . 80

MK-HI 116.80H

1/2

ADF

ATC 2

4176

110 . 20 3 3 2 .0

ID

ALT OFF N218KA

Audio Control Panels

ATC CONTROL ATC1

5211

RPLY /// ALT 14000FT IDENT ADC1 FLIGHT ID

The passenger address (PA) system facilitates amplified broadcasts to the cabin for passenger announcements, and seat belt and no smoking chimes. The XMIT knob on the respective audio panel controls PA broadcasts from the crew.

ALT REPORT ON/OFF

The audio control panels contain the following controls:

TEST

N218KA

XMIT Selects the transmitter to be use and its associated audio if the AUTO COMM switch is on. [

[

MSG

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-135. Flight ID Selection

1 – Selects COM 1 transceiver 2 – Selects COM 2 transceiver PA – Selects the PA system TEL – Selects the optional AirCell Phone HF – Selects the optional HF transceiver

Audio Control Knobs The audio control knobs control the volume of the associated radio. Pushing the knob in turns the audio off and pulling it out turns it on. These controls are independent of AUTO


16-75

16 AVIONICS



CABIN DIFF HI

MASTER CAUTION

L BLEED FAIL

R FUEL PRES LO

ES PR S

R OIL PRES LO

O T S E

R BLEED FAIL

ENG FIRE

EXTINGUISHER PUSH


DISCHARGED

MASTER CAUTION


MASTER WARNING

PRESS PRES PR ESS TO TO RESET RESET

SH


FD

CRS2

YD/AP DISC

CPL

PUSH

IR EC

T

T

AP

D

IA

D

UP

SY C N

MAC

YD

ALT A LT

PUSH

PUSH

S/

A LT ALT

1/2 BANK

H

PUSH

IR EC

APPR

HDG

SPEED

VNAV VNA AV

HDG

NAV N AV

CAN

CRS1

FLC

DOWN

PU

VS

EL

FD

C

ESET R RESET O RE SS TTO RESS RE PPRESS

Collins

Collins

N

3 30

MENU ADV

RADAR

GLENO G LENO E LINDZ ONDZ DBL /16000A 0 0A 000A

W

051 05111

UTC

20:03

R RAT AT

1 °C

F

GS

0

TAS TAS

0

12 °C

ISA

TILT TILT

ON OFF

LEFT

TTAXI AXI

ICE

N NAV AV

RECOG

ACTUATORS ACTUAT TORS STANDBY ST ANDBY


MAIN

OPEN

ARM


GOV

PILOT BRAKE DEICE

HI

OFF

FUEL

MANUAL

DN

VENT DOWN LOCK REL

LEFT

COPILOT COPI LOT


OFF TEST

EMER FREQ

ADC 2

NORM

1

RMT TUNE NORM

TUNE CDU

2

RTU

NORM

121.5 21.5


BEACON

STROBE

AT ATC C

L GEN TIE OPEN

HYD FLUID LOW

RVS NOT READY

R GEN TIE OPEN

R DC GEN

L CHIP DETECT

L NO FUEL XFR

BAT TIE OPEN

DUCT OVERTEMP

R NO FUEL XFR

R CHIP DETECT

L FUEL QTY

PITOT PITOT

RIGHT LANDING GEAR

GEAR DOWN


NOSE E

HD LT LT TEST

L

L BL AIR OFF

AUTOFTHER OFF


LEFT

EXT PWR

RIGHT

R FUEL QTY

O ON N

+

DISABLE DISABLE

TEST

MAN TIES CLOSE LDG/TAXI LIGHT

FLAPS

20

1


60

DOWN

80

UTC

SPKR

TE RR TERR

R RAT AT A

o

C


COM2

+

.5

PASS OXYGEN ON

2

4


0

6

.5

1

2

2

ADF

2

MKR


IINPH NPH

NORM

BLOWER

RUD BOOST OFF

R BL AIR OFF

R ENG ANTI-ICE

R IGNITION ON

TEMP

4


MAN HEAT HEAT

OFF

40 35 30 25

100 0F

0

T AL

5

20

3 4


PSI

0 VACUUM V ACUUM A

MAN TEMP INCR INCR

OFF

ELEC HEAT HEAT

CABIN CABIN ALT ALT WARN WARN TEST SILENCE SILENCE

DECR

BLEED AIR VALVES VA ALVES LEFT OPEN RIGHT ENVIR ENVIR OFF PNEU & ENVIR ENVIR OFF

OFF


LDG GEAR WARN TEST WARN


ENG FIRE TEST DET

OFF

0

50 80 ÛÛ) ) 100

FLIGHT HOURS 1/10

CABIN AIR

500 0 USE NO OIL

1000

1500 2000 PSI


OFF

T

2

6

MODE

IINCR NCR TEMP

LOW 5

1 7


BLOWER

10

3456


WINDOW WINDOW DEFOG

AUTO AUTO

35k

MAN COOL AUTO AUTO

R PITOT HEAT


5k 15k

+

NORM

R ENG ICE FAIL

PROP GND SOL FUEL CROSSFEED

L BK DEICE ON CABIN ALTITUDE

UP

RELAY RELA Y

OXY NOT ARMED

L ENG ANTI-ICE

WING DEICE L PROP PITCH

2 OFF

ELEC HEAT ON


R

TE RR TERR RD RDRR

BRT DIM

ENVIRONMENTAL ENVIRONMENTAL

L DC GEN

L ENG ICE FAIL

DME

1

GND COM

DG FREE

SLEW –

NORM

$872


COM1

UP


2))


RIGHT

2))

TEST IGNITION AND ENGINE ENGINE START START LEFT RIGHT ON OFF

PARKING P ARKING BRAKE

LEFT

LDG GEAR CONTROL

1

Collins

AHRS 1

MFD NORM

2

2))


BAT B AT

OFF TEST

OFF

1250

ESIS ON

2

< ET

RANGE


ON EMER OFF

PFD

STBY

+

AUTO AUTO COMM

2))


PILOT DISPLAY Y

ATC A TC 1

SLEW –

NORM

+

/,*+76

GEN RESET

RIGHT

/,*+76

BATT BUS NORM

DG FREE

ARM


FORMAT FORMAT <

FMS F

USH A AUTO UTO TILT

+13 °C BRT DIM

PROP SYNC


OFF–RESET

< PRESET

TFC

SSAT AT

3

RADAR

GCS

Collins

ON EXT PWR

1/2

ADF

350.0

TA TA ONLY ONLY

J10-1


BRT DIM

108.50

TERR RDR

J JNETT

/82 /8215A

TIL TILT LT

121.90

COM2

P

ATC1 AT TC1

ATC A TC 1

11 051 1

((INTC) (IN NTC) (8700) 87 ((8215) 82 KASE K ASE AS ))

2

NAV NAV

N

33

5

12

1118.85 18.85

12.5 GCS

COMM

29.88IN 29.88I N

LOC1

DATA DATA

ELEC

NA NAV/BRG AV/BRG

DME-H

PA

E

ONLYY TTA A ONL

T

1

1

NORM

G GS

TCAS F AIL FAIL

6

TERRAIN

11 1 13.00

S MIC MI C OXY

LOC

HDG

PUSH

IDENT

123.80

NAV1 NAV1

RA

HDG HDG

CRS 057 MENU ADV

B BRT

Collins

COM1

VS

PULL UP GND PROX

AC ACC C .–– DC P DCP

11 11 8 . 8 5

VOL

ENG2 EN G2

N

30

--:--

50

TERR RDR

VOR1 V 13.6NM DBL F ET COM1

T IDENT H

TTG

XMIT XMIT PA

1

ALT A LT G GPWS

ENG1 EN G1 REFS

S

VOICE B O

1.4NM

H

NORM

((8215)

2 ALT A LT

IAS

XAHS XA HS XADC XA DC

N

3 9 329

HDG 329

3

AUDIO ALTN ALTN

INPH

FMS

DATA DATA PUSH

ELEC

FORMA AT FORMAT

12.5 1122.5

SPKR

STD

7500 0MIN

33 3

NAV/BRG NA AV/BRG

PRESET MKR

78 20 0 0 00

10

10

30 0

M

138 38 082 82 2

25 25

AUTO AUTO COMM

2

1 40 9

8215A

-:--/ 1.4NM

0KTS

S

ADF

0.6NM 1.4NM -:-- : CLIMB 2.6NM -:-- : (8215) 169NM -:-- :

REFS

3

1

KASE ((8215)) (8700) KCOS

STD

2

XTLK

PUSH

8000

10

10

ATT T AT

FD

BARO

60


A TT V ATT N V

IAS

13.6NM 29.92 in

VOR1 057CRS

80

30

TERM

FMS

ADC2

112 49 TEMP°C 112

FIRE

110.0 110.0

4

600

HDG 329 (8215) 0.8NM

OIL

TORQ TORQ

3.4

A

AHS2

PUSH

1

700

106.0

830

w

2

10

0 329

N1

62.2 TORQ TORQ

130 FF 750 122 PRESS 80

ITT

24

DME

V2 107 VR 103 V1 100 AC ACC–.03 C –. 03

PA

1740

P

1

MIC OXY

AP AP

TRIM TR IM

BARO 1

40 7820 800 00

NORM

N350KA

PROP

J206

2

RADIO CALL

516

W

2

NAV NAV

Collins

1050

ITT

S

COMM

1

FLAP OVRD D

ACTIVE A CTIVE

2

900

S 1

TERR INHIB B

A ACTIVE

STEEP APPR PR

ACTIVE AC

T

VOL

10

G/S INHIB IB

A ACTIVE

4

8 000

20 60

Collins Colli ns

8215

T

170

80

1

T

16000

21

PTCH ALTS ALTS

15

HDG FMS

XMIT PA

2

<

MASTER WARNING


T

L OIL PRES LO

ENG FIRE

DISCHARGED

T

L FUEL PRES LO

EXTINGUISHER PUSH

AUTO AUTO


EXT

INCR INCR

10

15

Figure 16-136. Audio Panels

Speakers (one on each side)

Push to Talk Button

Hand Mic and Headset Connection

Figure 16-137. Audio System Components

16-76


COMM operation. Rotating the knob adjusts the volume.

COMM 1 – Controls the COM 1 audio volume 2 – Controls the COM 2 audio volume

NAV 1 – Controls the NAV 1 audio volume 2 – Controls the NAV 2 audio volume

DME 1 – Controls the DME 1 audio volume 2 – Controls the DME 2 audio volume

ADF 1 – Controls the ADF 1 audio volume 2 – Controls the ADF 2 audio volume (this knob exists only if the optional 2nd ADF is installed)

16 AVIONICS


MIC OXY – Selects the microphone in the associated oxygen mask as the active microphone. Automatically turns ON the on-side cockpit overhead speaker. NORM – Selects the headset or hand microphone as the active microphone

AUTO COMM Controls operation of the auto comm system. On – Allows audio from the selected transmitter on the XMIT knob to automatically be received without having to pull ON the respective control knob . Off – Inhibits auto comm control and requires the desired control knob to be pulled ON to receive the audio.

SPKR Controls the on-side cockpit overhead speaker.

HF Controls HF radio audio volume

VOICE/BOTH/IDENT

MKR

Controls the NAV audio filter.

Controls the marker beacon audio volume

VOICE – Removes morse code identification and allows only voice communications on the NAV audios.

TEL Controls the AirCell telephone volume

INPH Controls interphone communications. The knob on the pilot’s audio panel can be pulled out and pushed in to turn on and off the interphone system and then rotated to control the pilot’s side interphone volume. The copilot’s INPH knob is a volume control only.

BOTH – Voice communications and Morse code identification are both heard on the NAV audios. IDENT – Only Morse code identifications are audible on the NAV audios.

AUDIO Controls reversionary operation of the on-side audio control panel.


16-77

16 AVIONICS


NORM – Places the on-side audio control panel in normal mode. ALTN – Places the on-side audio control panel in reversionary operation. This bypasses the on-side audio amplifier and utilizes the pre-set amplifier associated with each COM and the PA. The pilot can transmit and receive on COMM 1 using a hand mic or boom mic, and cockpit speaker or headphones. The volume of radio receptions is not controllable. Transmissions may be made on COMM 2 and the PA, but COMM 2 receptions are not possible.

Control Wheel (PTT) Switches Each control wheel has the following PTT switches and functions (Figure 16-138): MIC Button – Controls COM radio and PA transmissions. IDENT – Controls the transponder identification function.

to the CDU is that all green frequencies are the active frequencies and all white frequencies are the standby or unused frequencies (Figure 16-139).

RTU Tuning There are three methods of RTU radio tuning: direct tuning, recall tuning, and tuning from the preset pages.

Direct Tuning The radios are directly tuned by changing the active frequency. This is accomplished when the white cursor (hollow white box) is over the green active frequency.

Recall Tuning Recall tuning is accomplished by tuning a frequency in the recall position (white color frequencies) and then swapping the active and recall frequencies by pressing the recall line select key.

RADIO TUNING UNIT (RTU) As with the CDU, the radio tuning unit (RTU) can be used for all radio tuning. Also similar

Figure 16-138. Control Wheel (PTT) Switches

16-78


16 AVIONICS



CABIN DIFF HI


MASTER CAUTION

R FUEL PRES LO

ES PR S

R OIL PRES LO

O T S E

R BLEED FAIL

ENG FIRE

EXTINGUISHER PUSH


DISCHARGED

MASTER CAUTION

ESET RESET OR RESS TTO PRESS PPR

MASTER WARNING


T

H

SY C N

MAC

CPL


FD

CRS2

YD/AP DISC

PUSH

IR EC

T

PUSH

AP

D

IA

D

S/

UP

YD

ALT A LT

1/2 BANK

PUSH

IR EC

ALT ALT

SH

VNAV VNA AV

APPR

HDG

SPEED

PUSH

HDG

NAV N AV

CAN

FLC

PU

CRS1

DOWN

EL

VS

FD

C

ESET R RESET O RE RESS TTO RE PPRESS

Collins

Collins

FLAP OVRD OVR RD

ACTIVE A CTIVE

62.2

MENU ADV

W

TERRAIN ONLYY TTA A ONL

GLENO G LENO E LINDZ ONDZ DBL /16000A 0 0A 000

1118.85 18.85

ATC1 A TTC1

051 05111

UTC

20:03

R RAT AT

1 °C

COM2

F

GS

0

TA TAS S

0

ON

PROP SYNC

ON

ON

OFF

OFF–RESET

OFF

LEFT

12 °C

ISA

TILT TILT

N NAV AV

RECOG

TEST

MAIN

OPEN AUTOFEATHER AUTOFEA ATTHER ARM

HI

PILOT

PROP TEST GND IDLE STOP STOP

GOV

BRAKE DEICE

TEST

STARTER STARTER ONLY ONLY PARKING P ARKING BRAKE

OFF

FUEL

MANUAL

ST ALL STALL ARN W WARN

PIT OT PITOT

VENT

RIGHT LANDING GEAR

GEAR DOWN

DN DOWN LOCK REL

LEFT

COPILOT COPI LOT

SURFA ACE SURFACE DEICE SI NGLE SINGLE

OFF

2))

OFF

AHRS 1

MFD NORM

EMER FREQ

ADC 2

NORM

1

RMT TUNE NORM

TUNE CDU

2

RTU

NORM

121.5 21.5

DG FREE

SLEW –

NORM

COM1

ATC ATC


NOSE E

HD LT LT TEST

L

L DC GEN

L GEN TIE OPEN

HYD FLUID LOW

RVS NOT READY

R GEN TIE OPEN

R DC GEN

L NO FUEL XFR

BAT TIE OPEN

DUCT OVERTEMP

R NO FUEL XFR

R CHIP DETECT

L ENG ICE FAIL

L FUEL QTY

ELEC HEAT ON

EXT PWR

R FUEL QTY

R ENG ICE FAIL

L BL AIR OFF

AUTOFTHER OFF

OXY NOT ARMED

RUD BOOST OFF



UP

OFF

LEFT

RIGHT

RELAY RELAY

TEST

FLAPS

L ENG ANTI-ICE

FUEL CROSSFEED

L BK DEICE ON

MAN TIES CLOSE

CABIN ALTITUDE

LDG/TAXI LIGHT

20

BLOWER

TEMP

1


60

DOWN

80

o

C


COM2

+

.5


6

.5

2

2

ADF

2

MKR


IINPH NPH

NORM

MAN TEMP INCR INCR

4

OFF

40 35 30 25

AUTO AUTO


BLOWER

100 0F

0

2

6 5

20

3 4

10

3456

PSI

0 VACUUM V ACUUM

ELEC HEATT HEA

CABIN ALTT W WARN CABI N AL ARN TEST SILENCE SILENCE

DECR

ENVIR ENVI R BLEED AIR NORMAL

BLEED AIR VALVES VAL VALVES LEFT OPEN RIGHT

LOW


ENVIR ENVIR OFF

OFF


LDG GEAR WARN WARN TEST


OFF

ENG FIRE TEST DET

0

50 80 ÛÛ) ) 100

FLIGHT HOURS 1/10

CABIN AIR

500 0 USE NO OIL

1000

1500 2000 PSI


OFF

T

5

1 7

MODE

IINCR NCR TEMP

R PROP PITCH

T AL


TAIL DEICE

AIR COND N1 LOW

35k


MAN HEAT HEAT

AUTO AUT O

R IGNITION ON

R BK DEICE ON

4

0 1

DME

1

15k

+

MAN COOL

R PITOT HEAT R ENG ANTI-ICE

PASS OXYGEN ON

2

1

GND COM

R BL AIR OFF

PROP GND SOL


R

TE RR TERR

RA RAT AT

NORM


L CHIP DETECT

2

UTC

SPKR

BRT DIM

ON O

+

DISABLE DISABLE

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IGNITION AND ENGINE START ENGI NE ST ART LEFT RIGHT ON

< ET

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OFF ICE PROTECTION PROP AUTO MANUAL AUTO


AUT O AUTO COMM

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ACTUATORS ACTUA ATORS STANDBY STANDBY

RIGHT

LDG GEAR CONTROL

STROBE

FORMAT FORMAT <

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ICE

/,*+76

TTAXI AXI

2))

OFF TEST

LEFT

1250

ESIS ON

2

BEACON

2

Collins

OVERSPEED STALL STALL WARN WARN WARN TEST W ARN TEST LANDI LANDING NG

OFF


BAT B AT

PILOT DISPLAY DISPLA AY PFD

STBY

+

3

5

FMS F

USH A AUTO UTO TILT

+13 °C

ATC A TC 1

SLEW –

ON EMER OFF

< PRESET

TFC

SSAT AT

NORM

+

RIGHT

ENGINE ENGINE ANTI-ICE LEFT RIGHT ON

GEN RESET

/,*+76

BAT BA AT BUS

RADAR

BRT DIM

DG FREE

ARM

ENG AUTO AUTO IGN

AVIONICS AVIONICS MASTER POW WER POWER

EXT PWR

1/2

ADF

350.0 GCS

Collins

NORM

NA NAV/BRG AV V/BRG

TTA A ONL ONLYY

J10-1


BRT DIM

2

NAV NAV

N

33

ELEC

DME-H

108.50

TERR RDR

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TIL TILT LT

121.90

ATC A TC 1

0 5 11 11

( NTC) (INTC) C) (8700) 87 ((8215) 82 KASE AS 5))

GCS

P

ET COM1

12.5

COMM

29.88IN 29.88I N

LOC1

DATA DATA PUSH

IDENT

123.80

NAV1 NAV1

PA

12

VOR1 V 13.6NM DBL F

SPKR

VOICE B O T IDENT H

RADAR

TERR RDR

T

1

1

NORM

G GS

TCAS F AIL FAIL

P

12.5 12.55

H

NORM

COM1

11 1 1 8 .8 5 1 13.00 11

MENU ADV

B BRT

Collins

N

30

50

LOC

HDG

E

AUDIO ALTN A LTN

INPH

((8215)

TTG --:-1.4NM

RA

HDG HDG

CRS 057

329

HDG 329

3

MKR

FMS

DATA DATA PUSH

ELEC

S MIC MIC OXY

PULL UP GND PROX

AC ACC C .––

REFS

DC P DCP

NAV/BRG NA AV V/BRG

VOL

VS

ENG2 EN G2

N

M

FORMAT FORMA AT

G GPWS

XMIT XMIT PA

1

ALT A LT

ENG1 EN G1

7500 0MIN

33 3

30 0

2

IAS

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10

10 0KTS

138 38 082 82 2

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XTLK XAHS XA HS

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25 25

PUSH

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ALT A LT

ATT T AT

FD

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TERM

FMS

8215A

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1 40 9

4

600

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AUT O AUTO COMM

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A TT V ATT N V

13.6NM 29.92 in

VOR1 057CRS

80 60

1

700

ADC2

49 TEMP°C 11 1122

FIRE

110.0 110.0

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W

2

TORQ TORQ

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ADF

10

0 329

OIL

106.0 06.0

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PUSH

STD

A

AHS2

S

DME

1

V2 107 VR 103 V1 100 AC ACC–.03 C –. 03

PA

830

6

1

MIC OXY

NORM

N1

130 FF 750 122 PRESS 80

ITT

S

2

1740

PROP

3

2

NAV NAV

AP AP

TR IM TRIM

BARO 1

800 00

S

COMM

1

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2

900 40

1

RADIO CALL

516

7820

S

Collins

1050

ITT

24

TERR INHIB IB

A ACTIVE

STEEP APPR APP PR

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VOL

10

G/S IB INHIB

A ACTIVE

4

8 000

20 60

Collins Colli ns

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T

80

1

T

16000

21

PTCH ALTS ALTS

17 170

15

HDG FMS

XMIT PA

2

<

MASTER WARNING

L BLEED FAIL

T

L OIL PRES LO

ENG FIRE

DISCHARGED

T

L FUEL PRES LO

EXTINGUISHER PUSH

AUTO AUTO


INCR INCR

EXT

10

15

Figure 16-139. Radio Tuning Unit (RTU)

Preset Tuning

Line Select Keys

Preset tuning (i.e. stored frequencies) is enabled when the TUNE MODE on the COM PRESET PAGE is set to PRESET. The tuning knobs are then used to select the desired preset memory number instead of tuning a frequency (Figure 16-140).

The line select keys (LSK) are used to place the cursor, navigate to a subpage, and make selections. Pressing the line select keys once places the cursor (a hollow white box) around the frequency at that location. Pressing the LSK next to active frequencies twice navigates to the appropriate menu display page. Pressing the LSK next to standby frequencies twice swaps the active and recall frequencies.

BRT

Collins

IDENT

COM 1

126 . 700 3

113 . 80 4

MK-HI 116.80H

ATC 2

4176

118 . 200

NAV 1

ID

2

25 SEL

DME--H

110 . 20 1

ADF

1/2

3 3 2 .0

ALT OFF

Figure 16-140. RTU in Preset Tuning Mode

COM Operation The COM section of the RTU top-level page provides tuning functions for the COM radio. Other COM control functions are handled on the dedicated COM main page and COM preset page. The active and recall frequency can be tuned from either the COM section of the top-level page or the COM main display page. The COM squelch, 8.33 and 25 kHz tuning, COM self-test and COM preset page access are controlled from the COM main display page (Figure 16-141).


16-79

16 AVIONICS


BRT

Collins

IDENT

COM 1

126 . 700

118 . 200

TX

KNOB SEL 8.33 25

SQUELCH ON OFF

beacon sensitivity, NAV self-test and NAV preset page access are controlled from the NAV main display page (Figure 16-142).

DME--H

1/2

PRESET PAGE

BRT

Collins

IDENT

NAV1

113.80

TEST

116.80 DMEH MKR SENS LO HI

RETURN

1/2

PRESET PAGE BRT

Collins

DME--H

TEST

RETURN

IDENT

COM 1

17

126 . 725

118 . 250 18

19

118 . 275 121.500 EMER

DME--H

TUNE MODE FREQ PRESET

PAGE

BRT

Collins

1/2

IDENT

NAV 1

5

1

116.80

109.50 2

25 SEL

3

110.50

110.80 4

DME--H

ACTIVE

RETURN

125.500

SQ OFF


Figure 16-141. RTU COMM Pages

PAGE

1/2

1

ACTIVE

RETURN

108.80

MK-HI

AUTO

The COM preset page allows for storing known frequencies. Once they are entered, the RTU preset tuning option can be activated and frequencies are chosen simply by selecting the memory number rather than tuning the frequency. In this preset tuning mode however, only the active frequency on the RTU top level page can be tuned directly if ATC gives a different frequency to contact.

NAV Operation The NAV section on the RTU top-level page provides tuning functions for the NAV radios. Other NAV control functions are handled on the NAV main display page and NAV preset page. The active and recall frequency can be tuned from either the NAV section of the top-level page or the NAV main display page. Marker

16-80

Figure 16-142. RTU NAV Pages

The NAV preset page allows for storing known frequencies. Once they are entered, the RTU preset tuning option can be activated and frequencies are chosen simply by selecting the memory number rather than tuning the frequency. In this preset tuning mode however, only the active frequency on the RTU top level page can be tuned directly if a different navigation source is required.

ADF OPERATION The ADF section on the RTU top-level page provides tuning functions for the ADF radio. Other ADF control functions are handled on the ADF main display page and ADF preset page.


The active frequency can be tuned from the ADF section of the top-level page and both the active and the recall frequencies can be tuned from the ADF main display page. The ADF or ANT modes, BFO feature, ADF selftest and ADF preset page access are controlled from the ADF main display page (Figure 16-143).

BRT

Collins

1200 5322

IDENT

ATC1 ID RPLY

ALT ON OFF ADC1 3000FT

DME--H

1/2

TEST XPNDR FAIL

RETURN

BRT

Collins

404.0

IDENT

ADF 1

320.0 BFO ON OFF

MODE ADF ANT

DME--H

1/2

PRESET PAGE

TEST

Figure 16-144. RTU ATC Page

ATC CONTROL Page The ATC CONTROL page annunciations are shown below:

RETURN

ATC Source Annunciation BRT

Collins

IDENT

ADF 1

1

404.0

390.0

2

3

566.0

304.0

4

DME--H


PAGE

1/2

1

Transponder Code Display This display shows the selected transponder code.

ACTIVE

RETURN

The ATC source annunciation indicates which transponder the CDU and RTU are controlling. Only one transponder is active at a time.

404.0

ANT BFO

Figure 16-143. RTU ADF Pages

ATC OPERATION The ATC section on the RTU top-level page provides the setting functions for the ATC code. Other ATC control functions are handled on the ATC main display page. The active code can be selected from the ATC section of the top-level page and both the active and the recall codes can be set from the ATC main display page. The Mode-C operation and self-test initiation are also controlled on the ATC main page display (Figure 16-144).

IDENT Line Select Key and Annunciation The IDENT line select key controls the transponder IDENT function. The IDENT annunciation enlarges and changes to cyan during ident functions (approximately 18 seconds).

Altitude Source Annunciator When Mode-C is enabled, the altitude data source (ADC 1 or ADC 2) is shown in cyan below the altitude readout.


16-81

16 AVIONICS


16 AVIONICS


Mode-C Control

BRT

Collins

The ALT line select key controls altitude reporting. ALT is shown in larger cyan when altitude reporting is selected. When selected off, only mode A replies are transmitted.

2.0000

IDENT

HF

5.0000 AM

SQ3 UV

410.0

Reporting Altitude Display

1/2

ADF

1330.5

RETURN

The Mode-C pressure altitude readout is shown in green when altitude reporting is selected.

BRT

Collins

Flight ID Display

2.0000

The Flight ID, if option is installed, is displayed and adjusted on the RTU top-level page and the ATC Control page.

TEST Function

IDENT

HF

5.0000

SQ2 AM

AM

POWER LO MED HI

SIMPLEX

DME--H

DUPLEX

1/2

PRESET PAGE

TEST

RETURN

The TEST line select key initiates the transponder self-test. The TEST annunciator enlarges in cyan while the test is active (approximately 10 seconds).

XPDR FAIL Annunciator XPDR FAIL appears in yellow to the right of the ATC legend when a transponder fails.

HF Operation (OPTIONAL) The Rockwell Collins HF-9000 is an option that can be installed in the aircraft. This creates a second page in the RTU. Pressing the “NEXT PAGE” LSK on the top-level page accesses the optional HF sub-display (Figure 16145). This display provides tuning functions for the HF radio. Refer to the Aircraft Flight Manual and HF-9000 operators guide for more information.

16-82

DME--H

BRT

Collins

R 10.0000 4 T 10.0000 AM 6

20.0000

IDENT

HF

15.0000 5 UV SIMPLEX

DME--H

DUPLEX

AM TUNE MODE FREQ PRESET EMER MAR ACTIVE RETURN 10.0000 SQ1 AM

PAGE

1/2

2

Figure 16-145. RTU HF Pages

TCAS II OPERATION (OPTIONAL) The Rockwell Collins TCAS-4000 TCAS II is an option that can be installed in the aircraft.


This option will replace the standard ADF frequency on page 1 and moves it to page 2 (Figure 16-146). This allows for quick selection of the desired TCAS mode from the main level page. Additional control is available on the TCAS main page. See the TCAS section later in this chapter for more information. BRT

Collins

IDENT

COM 1

118 . 200

126 . 700 113 . 80

25 SEL

NAV 1

CDU TUNING TUNE PAGE Display The TUNE PAGE has the following controls/displays. Similar to the RTU all green frequencies are the active frequencies and all white frequencies are the standby or unused frequencies (Figure 16-147). For installations that have a second CDU this TUNE feature is not active on the right CDU.

DME--H

110 . 20

MK-HI 116.80H

TCAS

ATC 2

4176

16 AVIONICS


TUNE

1/2

TA ONLY

ID

COM1

COM2

122.875

121.700 //

RECALL

ALT OFF

RECALL

/ 134.250

NEXT PAGE

ADF

410.0

123.875 MK-HI NAV2 110.30 / / DME2

NAV1 / 113.80/ICT DME1 / HOLD 116.80

HOLD

ATC1

3144 BRT

Collins

2.0000

HF

410.0

IDENT

ADF

MSG

5.0000 AM

SQ3 UV

ADF

DME--H

1/2

1330.5

412.5 [ LEGS

[ DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

RETURN

SP

/

Figure 16-147. CDU Tune with TCAS I BRT

Collins

TCAS ABOVE STBY TA ONLY NORM ALT REL ABS BELOW

IDENT

TA/ RA

TRAFFIC ON OFF

DME--H

1/2

TEST

RETURN

Figure 16-146. RTU TCAS II Pages

COM Display COM radio tuning is accomplished by entering the desired frequency in the scratchpad and then touching either the first or second line select keys on either side. The second position serves as the RECALL or PRESET frequency (i.e. standby frequency) and is the standard method of entry. Pressing the RECALL or PRESET key again will then swap the frequencies. If a frequency is inserted in the first line it will immediately be the active frequency and the previous one will move to the second line. For all frequencies, the decimal is assumed and does not need to be inserted (e.g., 123.4 can be entered as 1234).


16-83

16 AVIONICS


Additionally, the active frequencies are always identical between the RTU and CDU. Use caution when working with the standby frequencies as they are handled differently between the CDU and RTU.

display allows for turning the squelch ON or OFF and for testing the COM radio.

For IFIS equipped aircraft there is another option for tuning. The CDU contains a FREQUENCY selection under the IDX (index) page (Figure 16-148). This page contains frequencies for those airports entered into the flight plan. Press the line select key next to the desired frequency and it will enter into the scratchpad. The pilot can then navigate to the TUNe page and the frequency will still be in the scratchpad for use. FREQUENCY DATA

1/1

ATIS

TWR

1

118.200 // ICT TWR

2

// ICT DEP 126.700

3

119.500 // [

4 [

MSG

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N SP

/

DEP

MULTIPLE> APP

MULTIPLE> -----------------------
MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-148. CDU Frequency Data

The SQ OFF annunciation beside the COM legend appears when squelch has been disabled. TX annunciates when the radio is transmitting.

COM CONTROL Page The COM 1 or COM2 CONTROL page is selected by pushing the respective COM1 or COM2 line select key (the scratch pad must be empty) (Figure 16-149). The top portion of this

16-84

TEST

------ COM PRESETS -----## / ICT GND 121.750

Figure 16-149. CDU COMM Page

// 118.200

UNICOM / 122.950 CLNC DEL // 125.700

DEP ARR

RECALL / 134.250

/ +/ - V W X Y Z 0

GND

FSS

LEGS

122.875

// 121.900

MSG

SQUELCH ON/OFF

7 8 9 O P Q R S T U

SEL APT

KICT/KSLN/KHUT/ / 125.150

1/5

COM1 CONTROL COM1

The lower section of this display contains numbered COM PRESETS. This can contain up to 20 preset COM frequencies. Push the NEXT or PREV function keys to select the next or previous preset page. To create or modify a COM PRESETS frequency, enter the desired frequency into the scratchpad. Then push the appropriate left line select key to transfer this frequency to the numbered preset frequency field. If the frequency is valid, it displays in the data field. Once this is done, a label can be applied by simply typing in the desired name and pressing the left line select key again. To use these stored frequencies press either the left or right line select key from the COM PRESETS page and it will immediately become the active frequency. Another method is to simply enter the corresponding memory number (1 thru 20) into the scratchpad and then insert that into a COM tuning line. The associated frequency will be entered automatically.


NAV Display NAV radio tuning is accomplished by inserting the nav frequency in the scratchpad and then touching the appropriate NAV1 or NAV2 line select key. Additionally, the nav radio identifier can be typed into the scratch pad and selected by touching the NAV line select key. The CDU tuning will search the nearest frequency associated with that identifier and enter it along with the nav frequency. Additionally, the active frequencies are always identical between the RTU and CDU.

NAV CONTROL Page The NAV1 or NAV2 CONTROL page is selected by pressing the respective NAV1 or NAV2 line select key (the scratchpad must be empty) (Figure 16-150). The NAV CONTROL page will then allow for auto or manual tuning, DME hold, testing the radio, and changing marker beacon sensitivity (NAV1 CONTROL page only). See the VHF Navigation System section discussed earlier for more information on AUTO vs MANual tuning. NAV1 CONTROL

/ HUT 116.70

2

111.50 / [

3 [

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

ATC CONTROL ATC1

5211

RPLY /// ALT 14000FT IDENT ADC1

ALT REPORT ON/OFF

TEST

MFD DATA

EXEC

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

J K L M N

7 8 9 O P Q R S T U SP

[

[ LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

FPLN

/ +/ - V W X Y Z 0

The ATC CONTROL page is selected by pressing the ATC line select key (the scratchpad must be empty). (Figure 16-151). This page allows for transponder code entry, altitude reporting selection, testing the transponder and optionally entering a Flight ID. With the altitude reporting turned ON the automatically selected ADC will be displayed along with its corrected barometric pressure. Should an ADC fail the opposite ADC will automatically be selected. Additionally, the selected code is always identical between the RTU and CDU

MSG

DIR

I

ATC CONTROL Page

1/7

NAV1 NAV TUNING // AUTO/MAN 113.00/DBL DME1 TEST HOLD MKR SENS LO/HI ------ NAV PRESETS -----## 113.80 / ICT 1

MSG

The lower section of this display contains the NAV PRESETS. This section operates exactly like the COM PRESETS discussed earlier.

/

Figure 16-150. CDU NAV Page

/ +/ - V W X Y Z 0

SP

/

Figure 16-151. CDU ATC Page

The Flight ID field should contain only the ATC given identifier or the aircraft registration as appropriate. To turn the transponder ON or OFF and to select STBY, a separate switch on the reversionary panel must be moved. See the ATC Transponder section earlier in this chapter.


16-85

16 AVIONICS


16 AVIONICS


ADF CONTROL Page The ADF control page is selected by pressing the ADF line select key (the scratchpad must be empty) (Figure 16-152). From here the ADF can be tuned, Beat Frequency Oscillator (BFO) can be turned ON or OFF, the mode selected, or the ADF can be tested. The BFO selection should only be used for an NDB that cannot produce a typical Morse code identifier. The ANT mode provides only an audio output and does not create bearing-to-the-station signals. The bearing pointer will “park” at the 3 o’clock position. Both of these selections are abnormal and the CDU will annunciate on the main level TUNe page when chosen. 1/5

ADF CONTROL

ADF BFO / / 404.0 ON/OFF MODE ADF/ANT TEST ------ ADF PRESETS -----## / / 390.0 1

304.0 / /

TCAS Display and CONTROL Page (only with Optional TCAS II) The TCAS display and control pages allow for manipulation of the Rockwell Collins TCAS4000 TCAS II. When this option is installed, the external TCAS buttons on the reversionary panel are removed and all control is accomplished through the RTU or CDU. The TCAS display allows for TCAS mode selection without having to enter a menu (Figure 16-153). Each press of the TCAS MODE line select key will cycle through the available modes. The selected mode is then displayed on the PFD and MFD on the lower right corner advisory section (Figure 16-154). This selection works together with the RTU and either unit can change the TCAS mode. TUNE

1/2

COM1

COM2 121.700 // RECALL

122.875

2

RECALL / 134.250

MSG DIR

/ / 404.0

3

/ / 280.0 [

4 [

FPLN

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

HOLD EXEC NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

ATC1

TA/RA/STBY

3144 PREV

I

123.875 MK-HI NAV2 110.30 / / DME2 111.70 / HOLD TCAS MODE

NAV1 / 116.30 DME1

/

Figure 16-152. CDU ADF Page

ADF

REL

412.5

TCAS> [

[

MSG

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

16-86

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

The lower section of the display contains the ADF PRESETS display. Just like the COM and NAV radios this can contain up to 20 preset ADF frequencies. This section operates exactly like the COM PRESETS discussed earlier.

MFD DATA

SP

/

Figure 16-153. CDU TUNE With TCAS II

The TCAS CONROL page is selected by pressing the TCAS line select key (Figure 16155). This page allows for mode selection, altitude tag selection (relative or absolute), turning “other” traffic on or off, testing the TCAS system or doing an extended test and choosing the altitude volume. See the TCAS section later in this chapter for more detail.


HF CONTROL POWER LV
<

<

HF1 //// 2.0000

-10

16 AVIONICS


TERR RDR <

SQ 3

TFC <

TA ONLY

VOL 7

MODE

FREQ/EMER/MAR C

ISA +13 oC

BRT DIM

MSG

Figure 16-154. MFD TCAS Display

DEP ARR

[ PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

TCAS CONTROL MODE TA/RA/STBY

TEST EXT TEST ON/OFF ALT LIMITS

TRAFFIC ON/OFF

NORM BELOW

[

[ DEP ARR

MFD MENU

MFD ADV

MFD DATA

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

4 5 6 H

PERF

EXEC

DIR

TUN

LEGS

I

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BRT DIM

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

HF1 PRESETS 1/5 POWER LV
//// 2.0000

ABOVE

MSG

/

SP

ALT TAG REL/ABS

/

Figure 16-155. CDU TCAS II Control

HF Display and CONTROL PAGE When the optional HF system is installed, the CDU HF display and HF control page allow for selection of frequencies, emission modes, power output selections and squelch selections (Figure 16-156). Refer to the Aircraft Flight Manual and HF-9000 operators guide for more information.

MSG

LEGS

DEP ARR

PERF

MFD MENU

MFD ADV

MFD DATA

EXEC

DIR

FPLN

PREV

NEXT

IDX

1 2 3 A B C D E F G

CLR DEL

TUN

4 5 6 H

BRT DIM

I

J K L M N

7 8 9 O P Q R S T U

/ +/ - V W X Y Z 0

SP

/

Figure 16-156. CDU HF Control

HF Communication Systems One HF communication radio, available as an option, provides worldwide communications capability. The HF radio operates in the HF band of 2.0000 to 29.9999 MHz in 100-Hz steps. Operating emission modes include upper sideband voice (UV), lower sideband voice (LV), and amplitude modulation equivalent (AM). The AM Emission Mode has a


16-87

16 AVIONICS


frequency bandwidth of 15 KHz. Thus, radio stations with frequencies separated by 15 KHz or less may be received simultaneously. Both Simplex and Half-Duplex Tuning Modes are available. Refer to the AFM and HF-9000 operators guide for more information.

Ground Communications Power When the Battery Bus switch is in the normal position, the ground communications electric bus provides electric power directly from the main aircraft battery when selected by the pilot. Control of the system consists of a push on/push off solenoid-held annunciator switch labeled GND COM and is located on the reversionary panel (Figure 16-157). Selection provides operation of COM 1 through the RTU utilizing the headsets or the hand mic and cockpit speakers. No other radios are available during ground comm operations. An “ON” annunciation will illuminate when ground comm has been selected and extinguish when deselected.

Static Discharging A static electrical charge builds up on the surface of an airplane while in flight and causes interference in radio and avionics equipment operation. The charge is also dangerous to persons disembarking after landing, as well as to persons performing maintenance on the airplane. Static wicks (Figure 16-158) are installed on the training edges of the flight surfaces and the wing tips and assist discharging of the static electrical charge.

Figure 16-158. Static Wicks Figure 16-157. GND COMM Button

Subsequent activation of the main battery switch will result in an automatic disconnect of the ground communications bus from the com system; however, the normal method for deactivation of the system is accomplished by pressing the GND COM switch. This switch does not have a timer and will remain selected unless turned off, or the battery is turned on, or the Battery Bus switch is turned off.

16-88

ELECTRONIC STANDBY INSTRUMENT SYSTEM (ESIS) The L3 Avionics GH-3100 Electronic Standby Instrument System (ESIS) provides backup attitude, heading, airspeed and altitude information in a single display should a failure with the ProLine 21 system occur (Figure 16-159).


ered from the aircraft electrical system. A 30minute backup battery is provided to power the ESIS should the aircraft electrical input fail.

The ESIS can also provide lateral and vertical deviation information from NAV 1, with some limitations as discussed later in this section.

The TEST position tests the charge of the backup battery located in the avionics nose section. A green light adjacent to the switch illuminates if a sufficient charge is indicated. The ON position powers the ESIS from either the aircraft electrical system or the ESIS battery. An amber light adjacent to the switch illuminates if only the ESIS battery is powering the unit. The ESIS battery will not provide backup power to NAV 1 if it has lost power from the aircraft electrical system. Loss of aircraft electrical, will prevent its display on the ESIS. Figure 16-159. ESIS Display

Adjustment Knob The Adjustment knob on the bezel of the ESIS is used to set the barometric pressure setting or make selections within a menu. Pushing the knob selects standard pressure or selects the highlighted item on the menu when the menu is displayed.

The ESIS has the following controls:

ESIS Switch The ESIS switch on the pilot’s left subpanel controls power to the unit (Figure 16-160). During normal operations, the ESIS is pow-


CABIN DIFF HI


MASTER CAUTION

R FUEL PRES LO

ES PR S

R OIL PRES LO

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R BLEED FAIL

ENG FIRE

EXTINGUISHER PUSH


DISCHARGED

MASTER CAUTION

ESET RESET OR RESS TTO PRESS PPR

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SH

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FD

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LEFT

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AUTOFEATHER AUTOFEA ATHER ARM

PROP TEST GND IDLE STOP STOP

GOV

PILOT BRAKE DEICE

HI

OFF

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MANUAL

DN

VENT DOWN LOCK REL

LEFT

COPILOT COPI LOT

SURFACE SURFA ACE DEICE

OFF TEST

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121.5 121 21.5

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L GEN TIE OPEN

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RVS NOT READY

R GEN TIE OPEN

R DC GEN

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R NO FUEL XFR

R CHIP DETECT

L ENG ICE FAIL

L FUEL QTY

ELEC HEAT ON

EXT PWR

R FUEL QTY

R ENG ICE FAIL

L BL AIR OFF

AUTOFTHER OFF

OXY NOT ARMED

RUD BOOST OFF

R BL AIR OFF

OFF

LEFT

RIGHT

RELAY RELA Y

PROP GND SOL


R



UP TEST

FLAPS

L ENG ANTI-ICE

FUEL CROSSFEED

L BK DEICE ON

MAN TIES CLOSE

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TAIL DEICE

AIR COND N1 LOW

AUT O AUTO

MAN HEAT HEAT

OFF

ELEC HEAT HEAT

100 0F

0

2

6 5

20

3 4

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3456

PSI

0 VACUUM V ACUUM

MAN TEMP INCR INCR

OFF

CABIN CABIN ALT ALT WARN WARN TEST SILENCE SILENCE

DECR


BLEED AIR VALVES VAL VALVES LEFT OPEN RIGHT

LOW


ENVIR ENVIR OFF

OFF


LDG GEAR WARN WARN TEST


ENG FIRE TEST DET

OFF

0

50 80 ÛÛ) ) 100

FLIGHT HOURS 1/10

CABIN AIR

500 0 USE NO OIL

1000

1500 2000 PSI


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T

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1 7

MODE

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R PROP PITCH

T AL

35k


WINDOW WINDOW DEFOG

R IGNITION ON

R BK DEICE ON

THDS FT PER MIN

1

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1

15k

+

MAN COOL

R PITOT HEAT R ENG ANTI-ICE

PASS OXYGEN ON

2

1

GND COM

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L CHIP DETECT

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AUT O AUTO COMM

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DISABLE DISABLE

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LEFT ACTUATORS ACTUA ATORS STANDBY STANDBY

1250

ESIS ON

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BAT BAT

2

3

5

FMS

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PILOT DISPLAY DISPLA AY PFD

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29.88IN 29.88I N

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ENGINE ENGINE ANTI-ICE LEFT RIGHT ON

/,*+76

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/,*+76

BAT BA AT BUS

350.0

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ENG AUTO AUTO IGN

AVIONICS AVIONICS MASTER

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TA TA ONLY ONLY

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TCAS F AIL FAIL

12

TERRAIN

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S MIC MIC OXY

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123.80

NAV1 NAV1

RA

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VOR1 V 13.6NM DBL F ET COM1

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11 1 1 8 .8 5 1 13.00 11

MENU ADV

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TTG --:-1.4NM

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AC ACC C .–– CRS 057

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8000

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329

TERM

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HDG 329 (8215) 0.8NM

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KASE ((8215)) (8700) KCOS

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NAV NAV

AP AP

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STD

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900

7820

S

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1050

ITT

24

TERR TE INHIB IB

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STEEP APPR PR

ACTIVE A CTIVE

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G/S S INHIB I

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4

8 000

20

Collins Colli ns

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21

17 170

80 60

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PTCH ALTS ALTS

15

HDG FMS

XMIT 2 PA

<

MASTER WARNING

L BLEED FAIL

T

L OIL PRES LO

ENG FIRE

DISCHARGED

T

L FUEL PRES LO

EXTINGUISHER PUSH

AUTO AUTO


INCR INCR

EXT

10

15

Figure 16-160. ESIS Power Switch FOR TRAINING PURPOSES ONLY

16-89

16 AVIONICS


16 AVIONICS


ESIS Display The ESIS display incorporates aircraft heading, altitude, airspeed, pitch, and roll data into a compact display. Nav data from NAV 1 is also capable of being displayed provided NAV 1 is receiving power from the aircraft’s electrical system. A dedicated internal AHRS and an internal ADC provide data to the ESIS. HEADING – The aircraft heading is displayed along the bottom in a tape format. The compass “slides” horizontally with a lubber line placed in the center denoting the current heading. This reference comes from the internal AHRS and from a magnetometer located at the base of the aircraft T-tail, dedicated to the ESIS AHRS. ALTITUDE – The aircraft altitude is displayed in a tape format along the right hand side. The present altitude is depicted in a digital format within a box in the center of the altitude tape. The barometric pressure (shown at the top of the altitude tape) is adjusted with the Adjustment knob. The ESIS ADC generates this information. However the ADC retrieves air input from the copilot’s static source and does not have an independent port. This ESIS altitude is not RVSM certified. AIRSPEED – The aircraft airspeed is displayed in a tape format along the left hand side. The present airspeed is displayed in a digital format within a box in the center of the airspeed tape. A red band is displayed at VMO/MMO and VSO. These indications are not associated with any aural alerts. The ESIS ADC generates this information. The ADC receives air input from the copilot’s pitot source and does not have an independent input. PITCH – Aircraft pitch is displayed on the attitude display through the use of a pitch ‘ladder” and an Aircraft Reference Symbol. An ”Excessive Attitude” display provides assistance in determining the direction the pilot needs to pitch the aircraft to return to a level

16-90

pitch attitude. The Excessive Attitude display consists of red chevrons located within the pitch ladder. During an excessive attitude condition, the NAV data will be removed to declutter the display. The data will be removed when roll attitude exceeds 65˚ left or right bank or the pitch attitude exceeds 20˚ nosedown or 30˚ nose-up. The ESIS AHRS generates this information. ROLL – Aircraft roll attitude is depicted through the use of a sky pointer-type roll pointer and roll scale. A rectangular shaped slip/skid indicator is located below the roll pointer similar to the main ProLine 21 displays. The indicator moves with the roll pointer and “slides” left and right to depict slip/skid information. The ESIS AHRS generates this information. See the Pitot and Static System discussed earlier in this chapter for the air source connections.

MENU Button The MENU button on the bezel of the ESIS is used to configure the display. Once the button is pressed use the adjustment knob to move the cursor up or down the display (Figure 16161). At the appropriate item press the adjustment knob to enter that selection. A “…” placed at the end of the menu selection indicates the presence of a sub-menu. A period placed at the end of the menu selection indicates an action selection with no sub-menu. The following are available on the ESIS menu: Set Brightness Offset – Provides for manual adjustment of the display intensity. Rotating the Adjustment knob will adjust the brightness. Fast Align – Provides for realignment of the ESIS AHRS system and will initiate another 90 second count down timer.


Menu item selections will be restored to the last selected values after power is cycled

180 29.88 in 160

2 500

1407

10

120

10

10

136 5

40 22 20 00

10

2 000

100

Fast Erect

Set Brightness Offset... Fast Align Set 33Heading... N 03 M Figure 16-161. ESIS Menu

Set Heading – Provides for manual control of the compass. This places the compass in the “Free” mode. Nav On or Nav Off – Displays or removes from the display the nav data derived from NAV 1. Set Crs… – Provides selection of the course to be displayed for the nav data. Rotating the Adjustment knob adjusts the course. SILS BC or ILS Normal – Provides for normal or back course sensing of the course needle in reference to the type of approach being flown. Nav Displays… – Submenu allows selection of the DME Speed (On or Off) and DME Time (On or Off) to be displayed (Note that this information is from DME only and not the FMS).

WEATHER RADAR SYSTEM The WXR-852 radar system is installed in the Pro Line 21 King Air B350. The WXR-852 provides precipitation-based turbulence detection and has sector scan and auto-tilt functions. Weather radar controls are located on the display control panels (DCP). Weather radar display is shown on the MFD or PFD, depending on display selections. The weather radar is operated in a split mode with independent radar scans shown on each PFD. The following weather radar controls are located on the display control panel:

Radar Button The RADAR line select key controls display of the weather radar menus on the PFD (Figure 16-162). The following modes are selected with the MODE line select key and are displayed on the PFD’s weather radar status field.

Standby Mode (STBY) The STBY (standby) mode inhibits the radar transmitter and antenna scan drive. Selecting STBY or TEST will affect both pilot’s radar displays. The other three modes (WX, WX+T, or MAP) can be independently chosen. This STBY mode will automatically be selected 60 seconds after weight on wheels. However, once on the ground the radar can be turned ON again by reselecting a desired mode.

Baro Type… – Allows selection of the barometric pressure to be displayed in inHg or HPa or MB.


16-91

16 AVIONICS


300

FMS 1

Collins

30.16IN MIN 200 RA

301

17

33

W

DTK 301

ICT 4.1NM RADAR MODE

24

N

STBY TEST MAP WX WX+T

NORM <

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TARGET ATC1 1200

UTC 16:42

RAT - 4 oC

ARM OFF

V2 VR V1

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117 110 106

600 60

<

VOR1 113.85 CRS 229

1

10 400

0 24

30.16IN

251

W

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300

TERR

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301

RDR

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RADAR GAIN

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ATC1 4336

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125.250

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STBY WX WX+T TURB MAP TEST

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FMS 1

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6 540 20 SEC SCAN

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6935

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140

RADAR GAIN

25

Figure 16-163. Test Mode SEC SCAN ON OFF

TARGET ATC1 1200

UTC 16:42

RAT - 4 oC

ARM OFF

<

Collins <

BRT DIM

26

TORQ TORQ

IFIS

3.40

Figure 16-162. PFD Radar Menu

Test Mode (TEST)

10500

ITT

ITT 516

62.2 0.0

FIRE

PROP

PROP

N1 NI

1740 1980 106.0 98.5

0 . 0NM 0 . 8NM 4 . 4NM 198NM

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734

TORQ TORQ

0

FIRE AFX

PRESS OIL

120


: 0 . 0NM - : - - : 0 . 8NM - : - - : 4 . 4NM 6 9 3 5 A - : - - : 198NM - : - - / 0 . 8NM

FMS 24

251

W ABOVE

21 30

The system self-test is initiated by selecting the TEST mode of operation. A test pattern made up of six rainbow-like arcs show on the display(s) when the TEST mode is active (Figure 16-163).

ITT 830

110.0 2000

KTCT RW!$ ---KBJC

130 FF 750 FF 0 430 122 PRESS 80

ITT

50

<

<

<

16 AVIONICS


25 TERR

RDR < MAP T+5.7

Map Mode (MAP)

TFC <

The MAP mode allows the weather radar to provide the most detailed ground returns. The signal processing and target display colors are changed to accentuate ground features. Ground targets show in cyan, green, yellow, and magenta (Figure 16-164). This mode should not be used for weather avoidance.

16-92

GS

0

TAS

0

SAT 15 oC

ISA +13 oC

BRT DIM

Figure 16-164. Radar Ground Map Mode


Weather Mode (WX) Puts the weather radar in the basic weather detection mode. The weather mode displays precipitation-based returns in one of four colors: green, yellow, red, or magenta. The highest precipitation rates show in red (Figure 16-165). Should a significant return cause a potential masking of the radar image a path attenuation bar will appear on the display. This indicates a potential radar “shadow” and flight should not be conducted into that region until the pilot is assured it is clear of precipitation.

mode, with the addition of turbulence being displayed as magenta (Figure 16-166). The WX+T mode is only active out to 50 NM. When a display range greater than 50 NM is selected, the turbulence feature is automatically disabled. The turbulence detection is reactivated once the selected range is 50nm or less. Collins

10500

ITT

ITT 516

26

TORQ TORQ

3.40

PROP

62.2

0.0

FIRE

PROP

N1 NI

1740 1980 98.5

130 FF 750 FF 0 430 122 PRESS 80

ITT

ITT 830

734

106.0

TORQ TORQ

110.0 2000

0

FIRE AFX

PRESS OIL

120


Collins

ITT 516

26

TORQ TORQ

3.40

62.2 0.0

FIRE

1740 PROP 1980

PROP N1 NI

106.0 98.5

ITT 830

734

0

TORQ TORQ

FIRE AFX

110.0 2000

49 46

120

FMS DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

24

24

251

W ABOVE

21

50

<

DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

6935A - : - - / 0 . 8NM

112 73

: - : - - : CL I MB - : - - : ( 6 9 3 5) 6 9 3 5 A - :- - : - : - - / 0 . 8NM

FMS

:: 0 . 0NM - : - - :: 0 . 8NM - : - - :: 4 . 4NM - : - - :: 198NM

30

0 . 0NM 0 . 8NM 4 . 4NM 198NM

RW25 ( 6 9 3 5) SXW152 KBJC

PRESS OIL OIL TEMP°C TEMP oC

0 . 0NM 0 . 8NM 4 . 4NM 198NM

RW25 ( 6 9 3 5) SXW152 KBJC

130 FF 750 FF 0 430 122 PRESS 80

ITT

< 25

251

TERR

SXW152 KEGE

W

RDR < WX+T T+5.7

ABOVE

21

30

TFC <

<

GS

0

TAS

0

<

50

<

25

TAS

0

ISA +13 oC DIM

TERR

RDR < WX T+5.7

Figure 16-166. Radar Display Turbulence Mode

TFC <

F

0

SAT 15 oC

BRT

SXW152 ( 6 9 3 5) KEGE /6935A

GS

<

10500

ITT


ISA +13 oC

BRT DIM

Figure 16-165. Radar Display with Path Attenuation Bar

On IFIS equipped aircraft a small cyan indicator sweeps across the display helping assure that radar is ON even though the display may remain black (e.g., no returns).

Turbulence Only Mode (TURB) The turbulence only mode shows precipitation-related turbulence targets only in their magenta color (Figure 16-167). This is useful for closely analyzing areas of precipitation-related turbulence that have been detected while in the WX+T mode. TURB mode is automatically deselected in ranges greater than 50 NM.

Weather + Turbulence Mode (WX + T)

Gain Control

Detects precipitation and precipitation-related turbulence targets. The colors of the displays remain unchanged from those of the weather

The current GAIN setting is displayed in a box next to the GAIN legend (Figure 16-168). Turn the DATA knob ( the MENU SET knob for non-IFIS aircraft) on the DCP to set the gain


16-93

16 AVIONICS


at NORM, ±1, ±2, or ±3. Use caution when selecting a setting other than NORM as this will change the purpose of the standard radar colors. (i.e. a green area may actually be yellow or red in NORM setting and should be avoided) Collins

10500

ITT

ITT 516

26

TORQ TORQ

3.40

PROP

PROP

62.2

0.0

FIRE

N1 NI

1740 1980 98.5

0 . 0NM 0 . 8NM 4 . 4NM 198NM

734

TORQ TORQ

:: 0 . 0NM - : - - :: 0 . 8NM - : - - :: 4 . 4NM - : - - :: 198NM

FMS 24

DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

ITT 830

110.0 2000

RW25 ( 6 9 3 5) SXW152 KBJC

130 FF 750 FF 0 430 122 PRESS 80

ITT

106.0

251

0

FIRE AFX

PRESS OIL

120


6935A - : - - / 0 . 8NM

W

ABOVE

21

for faster updates. When this is not selected, the standard sweep is +/- 60˚(120˚total).

Antenna Stabilization This selection is only available on non-IFIS aircraft. The antenna stabilization function enables or disables automatic stabilization of the radar antenna. When enabled, the antenna sweep will maintain a constant angle relative to the earth’s surface as the aircraft’s pitch and bank change. This eliminates ground returns when banking the aircraft and allows for a precise left and right sweep. For IFIS equipped aircraft this feature is always selected and cannot be manually deselected.

1 30

50

<

<

The target alert function allows radar display to be deselected while the system continues monitoring the intensity of radar returns. The following annunciations on the PFD indicate how this feature is working (Figure 16-169).

<

25 TERR

SXW152 KEGE

RDR < TURB T+5.7

TFC <

GS

0

TAS

0

SAT 15 oC

ISA +13 oC BRT

4

DIM

Figure 16-167. Turbulence Only Display

W

TGT

30.16IN

F 144 069

30

>

<

TERR

RDR >

FORMAT > Figure 16-169. Pilot's PFD with TGT

WX G+3

<

16 AVIONICS


T+ 5 .7

TFC >

<

Cyan TGT: indicates the target function is selected when the PFD’s and MFD are not displaying radar. This indicates the system is working appropriately.

ISA +13 oC

Figure 16-168. Radar Gain Display

Sector Scan Function The sector scan function limits the sweep of the radar to +/- 30˚ sweep (60˚ total) providing

16-94

White TGT: indicates the target function is selected but both PFD’s are displaying TERRain. In this orientation the target function does not work. At least one display must have terrain deselected.


Yellow TGT: indicates the target function has detected a significant return and radar should be selected for display to see the area of interest. This does not cause the radar display to auto “pop up”.

Collins

10500

ITT

ITT 516

26

TORQ TORQ

3.40

62.2 0.0

FIRE

PROP

PROP

1740 1980 106.0 98.5

N1 NI

0 . 0NM 0 . 8NM 4 . 4NM 198NM

ITT 830

734

TORQ TORQ

0

FIRE AFX

PRESS OIL

120


: 0 . 0NM - : - - : 0 . 8NM - : - - : 4 . 4NM - : - - : 198NM - : - - / 0 . 8NM

FMS DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

130 FF 750 FF 0 430 122 PRESS 80

ITT

110.0 2000

RW25 ( 6 9 3 5) SXW152 KBJC

The target alert function searches in a ±15˚ sector in front of the aircraft within a range of 7 to 200 NM.

24

251

W

FMS DR ABOVE

21

30

GCS Button

16 AVIONICS


50

<

<

<

The GCS button controls ground clutter suppression. When selected, the system suppresses ground returns (clutter) in the WX and WX+T modes to help identify precipitation targets. GCS is only active for 30 seconds. GCS annunciates on the PFD and MFD when the radar mode is on and the GCS button has been pressed (Figure 16-170).

25 TERR

RDR < WX T+5.7

TFC < RLG GS

0

TAS

0

ISA +13 oC

SAT 15 oC

BRT DIM

Collins

TILT Control 10500

ITT

The TILT knob controls the antenna tilt angle. The selected angle (-15 to +15 degrees) is displayed with the letter T on the displays (Figure 16-171). Since each pilot has a tilt control the radar produces an image on only one sweep. This enables the pilot’s tilt to be shown on the clockwise sweep while the copilot’s tilt can be shown on the counterclockwise sweep.

ITT 516

26

TORQ TORQ

3.40

62.2 0.0

FIRE

PROP

PROP

1740 1980

N1 NI

106.0 98.5

0 . 0NM 0 . 8NM 4 . 4NM 198NM

734

TORQ TORQ

0

FIRE AFX

PRESS OIL

120


: 0 . 0NM - : - - : 0 . 8NM - : - - : 4 . 4NM - : - - : 198NM - : - - / 0 . 8NM

FMS DTK 251 ( 6 9 3 5) TTG -- : -0. 8NM

ITT 830

110.0 2000

RW25 ( 6 9 3 5) SXW152 KBJC

130 FF 750 FF 0 430 122 PRESS 80

ITT

24

251

W

FMS DR ABOVE

21

30

The PUSH AUTO TILT button located in the center of the TILT / RANGE knob selects automatic antenna tilt control. The letter "A" adjacent to the tilt angle indicates that auto-tilt is selected. The auto tilt function compensates for airplane attitude changes and range changes by adjusting the tilt angle to maintain the selected reference to ground. This will cause the tilt number to change when pitching up, pitching down, or changing the range.

50

<

<

<

PUSH AUTO TILT Button

25 TERR

RDR < GCS T+5.7

TFC < RLG GS

0

TAS

0

SAT 15 oC

ISA +13 oC

BRT DIM

Figure 16-170. Radar Ground Clutter Supression


16-95

<

activate during a crash and transmit a sweeping tone on 121.5 MHz, 243 MHz, and 406 MHz, through a system of satellites. This activation is independent of the remote switch setting or availability of aircraft power. The ability of the ELT to transmit on 406 MHz requires that the ELT be activated with the National Oceanic and Atmospheric Association (NOAA) as the beacon provides a unique identifier code traceable to a specific aircraft and operator. The registration is free, good for two years, and can be done on-line at www.beaconregistration.noaa.gov.

<

16 AVIONICS


TERR

RDR < WX T+5.7

TFC <

ISA +13 oC

BRT DIM

Figure 16-171. Radar Tilt Display

RANGE Knob The RANGE knob controls the scanning range shown on the MFD map and radar pictorial. Range annunciations are shown on the displays as discussed earlier.

COCKPIT VOICE RECORDER (CVR) The typical CVR is the Fairchild FA2100 which simultaneously records audio from each audio panel, PA system, and the cockpit area microphone. Depending on the selected option this can be a recording of 30 minutes or 2 hours on the solid-state recorder. An impact switch stops further recording when sufficient G-force is encountered. There are 2 styles of controllers installed on the pedestal of the aircraft (Figure 16-172). Refer to the Aircraft Flight Manual supplement for necessary test procedures of the installed CVR.

EMERGENCY LOCATOR TRANSMITTER (ELT) The Emergency Locator Transmitter (ELT) is designed to provide beacon location to the aircraft after a crash. The ELT will automatically

16-96

Figure 16-172. CVR Controllers

The remote switch located on the left-hand sidewall of the cockpit, is installed to perform the following functions (Figure 16-173): • Test the ELT • Deactivate the ELT if it has been inadvertently activated by the “G” switch • Activate the ELT in an in-flight emergency if an off-airport landing is anticipated • Activate the ELT after an off-airport landing, if the impact did not automatically activate it


An amber light is located adjacent to the switch that will illuminate any time the ELT has been activated, either manually or automatically. The ELT will automatically activate, with the “G” switch, regardless of the position of the remote switch.

timeter. If the radio altimeter were to fail an appropriate GPWS annunciator would appear on the PFDs indicating that all the following modes are inoperative. (Figure 16-175). Collins

APPR FMS

VGP

3 000

185 180 160

4

20

700

2 1

10

DN

14 1 0

4000

600 60

6 540 20 GND PROX 10

120

1

400

700

100 24

TOD

2 4

30.16IN

251

1000

W

Collins

Figure 16-173. ELT Manual Switch

ENHANCED GROUND PROXIMITY WARNING SYSTEM (EGPWS)

3 000

180 160

14 1 0

1

400

550 24

TOD

2 4

30.16IN

251

1000

W

Figure 16-174. PFD GND PROX and PULL UP Annunciators Collins R

APPR FMS

VGP

3 000

185 180 160

The following operating modes generate cautions and warnings that are part of the basic GPWS. The cautions will generate a “GND PROX” PFD message while the warning will generate a “PULL UP” PFD message (Figure 16-174). Each caution and warning is also accompanied by an aural command as shown in the following table. This portion of the MK VIII system is solely related to the radio al-

1

600 60 6 540 20

PULL UP 10

100

2R

<

BASIC GROUND PROXIMITY WARNING SYSTEM (GPWS)

700

10

DN

4000 4

20

120

The Honeywell Mark VIII Enhanced Ground Proximity Warning System (EGPWS) provides visual and aural cautions and warnings to the pilot of potential collision with terrain or obstructions, other potentially unsafe conditions, as well as altitude awareness callouts. The EGPWS has two versions of operation: basic GPWS and Enhanced GPWS (EGPWS).

VGP

<

APPR FMS

185

4

20

700 600 60 6 540 20

14 1 0

400

GPWS

RA

100

TOD

1

10

120

24

251

2 1

10

DN

4000

2 4

30.16IN

1000

W

Figure 16-175. GPWS Failure Annunciators

<

R


16-97

16 AVIONICS


16 AVIONICS


Table 16-1. GPWS CAUTIONS AND WARNINGS Mode

Function

PFD Caution Message

Aural Caution

PFD Warning Message

Aural Warning

1

Excessive Descent Rate

GND PROX

Sink Rate

PULL UP

Pull Up

2

Excessive Closure on Terrain

GND PROX

Sink Rate

PULL UP

Pull Up

3

Altitude Loss After Takeoff

GND PROX

Don’t Sink, Don’t Sink

4a

Unsafe Terrain Clearance

GND PROX

Too Low, Gear

4b


GND PROX

Too Low, Flaps

4c


GND PROX

Too Low, Terrain

5

Excessive Glideslope Deviation

GND PROX

Glideslope

GND PROX

Glideslope, Glideslope

6

Bank Angle

Bank Angle

Altitude Callouts

Smart 500, 200, 100, 50, 40, 30, 20, 10

Minimums

Minimums, Minimums

The following equipment is required to be operational for the proper function of Modes 1 through 6 of the Mark VIII system: 1. Enhanced Ground Proximity Warning Computer (EGPWC) 2. Radio Altimeter 3. Vertical Speed from the Air Data Computer 4. Airspeed from the Air Data Computer 5. Glideslope Deviation 6. Landing Gear Position 7. Flap Position 8. Roll Attitude from Pilot’s Attitude System (for BANK ANGLE voice message) 9. Decision Height System (for MINIMUMS voice message)

16-98

The following Mode 6 advisory callouts are enabled for altitude awareness: 1. Five Hundred (classified as a Smart Callout) 2. Two Hundred 3. One Hundred 4. Fifty 5. Forty 6. Thirty 7. Twenty 8. Ten 9. Minimums Three push-button switch annunciators are located directly in front of the pilot between the pilot’s PFD and the MFD (Figure 16-176).


ENHANCED GROUND PROXIMITY WARNING SYSTEM (GPWS)

Figure 16-176. EGPWS Buttons

These push-buttons allow the pilot to desensitize modes 1, 4b, and 5 and the Enhanced modes for abnormal operations.

The enhanced features of the EGPWS include Terrain (or Obstacle) Alerting and Display (TAD) and the Terrain Clearance Floor (TCF). Terrain for the entire world and obstacles of 100 feet or more are contained in a database that covers the United States, parts of Canada, the Caribbean and gradually the rest of the world. These functions require GPS1 latitude / longitude, airplane altitude, and the terrain / airport database. Note that the database is Honeywell specific and contained within the ground proximity unit located in the nose of the aircraft. It is not mandatory to update this database however it will help eliminate nuisance alerts by updating airport and obstacle information. The update procedure requires access to the aircraft nose avionics section and must be accomplished by qualified personnel. A Honeywell specific cable and a PCMCIA card will be attached to

Table 16-2. EGPWS BUTTONS

Switch/ Annunciator

Color

Function

GPWS FLAP OVRD

AMBER

Pressing the switch disables the TOO LOW FLAPS portion of the GPWS Mode 4b alert and desensitizes the Mode 1 alert boundaries. The annunciator illuminates when the switch is pressed.

G/S INHIBIT

AMBER

Illuminates to indicate the GPWS Mode 5 glideslope alert has been inhibited. While the airplane is on the ground, this switch is used to initiate the EGPWS system self-test. The annunciator illuminates when the switch is pressed.

TERR INHIBIT

GREEN

Pressing the switch deselects all enhanced functions of the EGPWS system. The annunciator illuminates when the switch is pressed


16-99

16 AVIONICS


the EGPWS unit. A series of lights on the unit will indicate successful or unsuccessful loading.

The TAD algorithms continuously compute terrain clearance envelopes ahead of the airplane. Two envelopes are computed, one corresponding to a Terrain Caution Alert (roughly 60 seconds prior to impact) and the other corresponding to a Terrain Warning Alert (roughly 30 seconds prior to impact). If the boundaries of these envelopes conflict with terrain or obstacle elevation data, alerts are issued. The Caution and Warning envelops use the terrain clearance floor as a baseline, and look ahead of the airplane in a volume that is calculated as a function of groundspeed, flight path angle, and track. If terrain or obstacle data penetrates the caution or warning envelopes, then the corresponding aural and visual alerts are generated. Additionally, the terrain display will automatically pop up on the MFD and display any terrain penetrating the warning envelope in solid red with a 10 nm range. If the display has been automatically changed to terrain by the pop up feature, the original display will need to be manually reselected after the terrain conflict has been resolved. It is important to note that the EGPWS system does not account for performance degradation or actual climb capability of the aircraft. This requires good situational awareness of

16-100

Terrain display can be selected manually at any time. Areas of terrain sufficiently close to the airplane that do not penetrate the terrain caution or warning envelopes are depicted by 1 areas of red, yellow or green dot patterns (Figure 16-177). The color and dot density vary based on terrain elevation relative to the airplane. Magenta coloring is used to indicate 1 areas where terrain information is unavailable. 0

ACC-.02 TERM

24

FMS1

251

W

3 0 .1 6 IN

T GT

21

DTK 251 (6935) 0. 8NM

144 069

30

The TCF creates an increasing terrain clearance envelope around the nearest airport runway and generates alerts based on current airplane location, the nearest runway center point and radio altitude. TCF protection is provided in all airplane configurations and protects from those conditions where the airport may be located on higher terrain than what is currently under the aircraft. The nominal airport altitude is extended outward from the airport area and a caution will alert the aircraft even though it is not close to the immediate terrain.

the surrounding terrain to help define the best escape route should it become necessary.

50

FORMAT > <

16 AVIONICS


25

TERR

< PRESET

RDR

VOR1

>

TERRAIN

TFC >

F

TCAS OFF

< ET 01:42 COM1 121.800

ATC1 4336

RADAR ON UTC 14:41

RAT 15 oC

COM2

125.250

BRT DIM

Figure 16-177. Terrain Display

EGPWS Terrain Display Overlay is available only on Present Position Map and Arc formats. Selection of weather radar and Terrain Display are mutually exclusive. The following equipment is required to be operational for the proper functioning of the enhanced features of the Mark VIII EGPWS System: 1. Enhanced Ground Proximity Warning Computer (EGPWC) 2. Heading from the No. 1 Compass System 3. GPS position from the Flight Management System (if GPS position is not available/reliable, the TERR INHIB switch/annunciator must be pushed) 4. Terrain and Obstacle Data Base


16 AVIONICS


Table 16-3. EPGWS CAUTIONS AND WARNINGS PFD Caution Message

Mode

Function

TAD

Terrain Alerting and Display

Aural Caution

GND PROX

Caution Terrain, Caution Terrain

Or Obstacle Alerting

Aural Warning

PULL UP

Terrain, Terrain, Pull Up

Or Or Caution Obstacle, Caution Obstacle

and Display

TCF

PFD Warning Message

Terrain Clearance Floor

GND PROX

Too Low, Terrain

Obstacle, Obstacle, Pull Up

PULL UP

Terrain, Terrain, Pull Up

Should a failure of one of these items occur a TERR and TERRAIN FAIL annunciator will appear on the AFD’s and the terrain / obstacle display will be removed (Figure 16-178). Once 4 1 the accuracy of the Enhanced features is reduced or has failed the TERR INHIB switch should be pushed to eliminate any misleading information. This causes the ground proximity system to revert to a basic GPWS and use only the radio altimeter for further callouts.

The following enhanced features are available:

251

The auto-ranging feature will affect the pilot’s PFD and MFD

24

TOD

FMS1

W

21

NOTE

30


1. A visual display of terrain on the PFD’s and/or the MFD which is conFigured for: a. A Peaks Display b. Pop-Up feature 10nm range (MFD only)

JABAN 10

FORMAT >

TERRAIN FAIL

<

5

< PRESET

RALPE

TERR > RDR

TOD

VOR1

TFC > < ET 01:42 COM1 121.800

ATC1 4336

UTC 14:41

TERR RAT 15 oC

COM2

125.250

BRT DIM

2. Forward Looking Terrain and Obstacle Cautions and Warnings 3. Envelope Modulation of GPWS Modes 1, 2, 3, 4, 5, and 6. 4. Runway Field Clearance Floor (RFCF) 5. Terrain Clearance Floor (TCF) (Also requires a radio altitude input)

Figure 16-178. Terrain Fail and TERR Annunciators


16-101 101

TRAFFIC COLLISION AND AVOIDANCE SYSTEM (TCAS I)

Collins

HDG FMS

PTCH ALTS

1

600 60

400

0

TERM

24

FMS1

1

10

117 V2 110 VR 106 V1 ACC-.02

251

2 4

TRAFFIC W

30.16IN

21

DTK 251 (6935) 0. 8NM

30

144 069 50

FORMAT > 25

TERR

< PRESET

>

RDR TERRAIN

VOR1

TFC >

F

TCAS TEST

< ET

COM1 121.800

RADAR ON UTC 14:41

ATC1 4336

RAT 15 oC

COM2

125.250

BRT DIM

Collins

10500

51626

TORQ TORQ

62.2 0.0

FIRE

3.40

PROP PROP

1740 1980

N1 NI

106.0 98.5

0 . 0NM 0 . 8NM 4 . 4NM 198 NM

RW25 ( 6 9 3 5) SXW152 KBJC

ITT 830 734

TORQ TORQ

0

FIRE AFX

o

24

251

W

ABOVE BELOW

SXW152

21

120

PRESS OIL OIL

49 TEMP C 112 46 TEMP°C 73

: - : - - : CL I MB - : - - : ( 6 9 3 5) 6 9 3 5 A - :- - : - : - - / 0 . 8NM

FM S DTK 25 1 ( 6 9 3 5) TT G - - : - 0 . 8N M

130 FF 750 FF 0 430 122 PRESS 80

ITT

110.0 2000

30

+10 -10

< 2.5

( 6 9 3 5) /6935A KEGE

<

5

<

-02

16-102

2

6 540 20

ITT

The display of traffic can selected on the MFD by pressing and holding the TFC line key for more than 1 second or by navigating through the lower format key (Figure 16-179). For IFIS installed aircraft, TCAS is also available for display on the PFD’s by using the TFC line key. However, if TCAS is selected for display on the HSI format this will limit the range to 50nm. The TCAS must be deselected from the PFD or the PFD must be placed in the ARC or MAP formats for the range to extend beyond 50nm.

700

10

ITT

The SKY899 TAS is an active system that operates as an aircraft-to-aircraft interrogation device. The system can interrogate up to 35 different aircraft transponders in a 35 nm radius in the same way ground based radar interrogates aircraft transponders. When the SKY899 receives replies to its interrogations, it computes the responding aircraft’s range, relative bearing, relative altitude, and closure rate. The SKY899 then predicts collision threats and plots the eight most threatening aircraft locations.

4

20 60

The L3 Communications SKYWATCH HP Traffic Collision and Avoidance System (TCAS), Model SKY899, is to be used for aiding visual acquisition of conflicting traffic. The system includes a transmitter-receiver computer (TRC), and a directional antenna mounted on the top of the fuselage. The installation receives pressure altitude information from the pilot’s or copilot’s encoding altimeter through the No. 1 or No. 2 transponder. The system also receives inputs from the right weight-on-wheels switch, the right landing gear downlock switch, and heading input from the No. 1 compass. The system is powered from the Left Generator Avionics Bus, and is protected by a 5-amp circuit breaker, placarded TCAS.

6935

1 4 000

140

80

<

16 AVIONICS


TERR

RDR < WX T+5 .7

TFC <

F

TCAS TEST

GS

0

TAS

0

SAT 15 oC

ISA +13 oC

BRT DIM

Figure 16-179. TCAS I TEST

The SKY899 has the following controls:

Operating Mode Button This switch/light is placarded ON/STBY (Figure 16-180). ON is illuminated when the sys-


tem is in the operating mode. The switch/light will be blank when the system is in the standby mode. On the ground, this switch can be used to change the operating mode between ON and STBY. In flight, this switch is inactive and the system is continuously ON due to inputs from the squat switch.

16 AVIONICS


Solid Cyan Diamond This is the Proximate Traffic symbol that is generated when intruder traffic is detected within 6 nm and 1200 feet, but does not pose a threat.

Open Cyan Diamond This is the symbol for Other Traffic and is generated to represent an intruder aircraft that has been detected but it outside of the Proximate Traffic boundary.

Solid Yellow Semicircle This is a Traffic Advisory (TA) symbol that is generated when an intruder aircraft may pose a collision threat but is out of the current display range.

Vertical Trend Arrow Figure 16-180. Operating Mode Button

Display Range Knob The display range is controlled through the range knob on the Display Control Panel (DCP).

Vertical Display Mode/Test Button This push-button is placarded TEST/ALT. On the ground, pressing this button will initiate an internal self-test. This test should be conducted before the first flight of the day. When the TCAS is turned ON, this button acts as a Vertical Display Mode control, allowing the pilot to toggle the display between ABOVE, BELOW, ABOVE/BELOW AND Normal. The SKY899 will display the following features:

Solid Yellow Circle

The vertical trend arrow appears to the right of the traffic symbol to indicate that the intruder aircraft is climbing or descending at a rate greater than 500 fpm. The arrow will be pointing up or down as appropriate for the climb or descent. The vertical trend arrow will not be displayed for non-altitude reporting aircraft.

Data Tag (Example +04) A two-digit number representing the relative altitude, in hundreds of feet, of the intruder aircraft is shown above or below the traffic symbol. A positive data tag will be shown above the traffic symbol representing that the intruder is located above your aircraft. A negative data tag will be shown below the traffic symbol representing that the intruder is located below your aircraft. If the intruder is located at the same altitude as your aircraft, 00 is displayed above the traffic symbol. Four altitude display modes are available:

This is the Traffic Advisory (TA) symbol that depicts an intruder aircraft that may pose a collision threat. This is accompanied by the aural alert “TRAFFIC, TRAFFIC”. Additionally, the PFD will annunciate a flashing TRAFFIC below the attitude indicator.

Look-up Mode (ABOVE) Displays traffic detected within +9,000 feet to –2,700 feet of your airplane.


16-103

16 AVIONICS


Normal Mode (blank) Displays traffic detected within ±2,700 feet of your airplane.

Look-down Mode (BELOW) Displays traffic detected within +2,700 feet to –9,000 feet of your airplane.

Unrestricted Mode (ABOVE/BELOW) Displays traffic detected within ±9,000 feet of your airplane

TCAS Self-Test Mode When the TCAS self-test is conducted, the following test pattern will be displayed on the MFD: Traffic Advisory (solid yellow circle) will appear at 9 o’clock, range 2 miles, 200 feet below and climbing. Proximate Traffic (solid cyan diamond) will appear at 1 o’clock, range 3.6 miles, 1000 feet below and descending. Other Traffic (open cyan diamond) will appear at 11 o’clock, range 3. 6 miles, flying level 1000 feet above, and in level flight. The SKY899 has the following automatic features: Using the right weight-on-wheels switch, the system will automatically switch from the STBY mode to the ON mode in the 6 nm range and ABOVE mode approximately 8 to 10 seconds after takeoff. Using the right weight-on-wheels switch, the system will automatically switch from the ON mode to the STBY mode approximately 24 seconds after landing.

16-104

Using the radio altimeter, the system will inhibit aural traffic alerts below 400 feet AGL to minimize pilot distraction.

TRAFFIC COLLISION AND AVOIDANCE SYSTEM (TCAS II) (OPTIONAL) The Collins TCAS-4000 is a TCAS II system designed to protect a volume of airspace around the TCAS II-equipped airplane by warning the pilot of the threat of other transponder equipped airplanes penetrating that airspace. The system interrogates Mode C and Mode S transponders in nearby airplanes and analyzes their replies to identify potential and predicted collision threats. The system advises the pilot when to climb, descend, or maintain altitude to avoid passing too close to, or colliding with, the threat airplane. When an intruder airplane is equipped with TCAS II, the system coordinates avoidance maneuvers with this airplane using data link capability of the Mode S transponders. If traffic gets within 25 to 45 seconds (depending on altitude) of the projected Closest Point of Approach (CPA), it is considered an intruder and a Traffic Advisory (TA) is issued. This TA calls attention to what may develop into a collision threat using visual and aural alerts. The visual alert consists of a solid yellow circle depicting the intruder on the traffic map and a yellow flashing TRAFFIC message on the PFDs. The aural alert consists of the voice message, TRAFFIC, TRAFFIC. These alerts promote mental and physical preparation for a possible maneuver that may follow, and assists the pilot in achieving visual acquisition of the intruding aircraft (Figure 16-181).


sory (RA) is issued. This RA provides a recommended vertical maneuver using modified instantaneous vertical speed indicators (IVSIs) and voice messages to provide adequate vertical separation from the threat aircraft (a Corrective RA) or prevents initiation of a maneuver that would place the TCAS II aircraft in jeopardy (a Preventive RA). In addition to the voice messages, e.g., CLIMB, CLIMB, the threat aircraft is depicted as a solid red square on the Traffic Map, and a flashing red TRAFFIC message is displayed on the PFDs.

Collins

HDG FMS

PTCH ALTS

6935

1 4 000

140

80

4

20 60

700

2 1

10

600 60

6 540 20 400

0

TERM

24

FMS1

1

10

117 V2 110 VR 106 V1 ACC-.02

251

2 4

TRAFFIC W

30.16IN

21

144 069

30

DTK 251 (6935) 0. 8NM 5

FORMAT > <

2.5

TERR

< PRESET

>

RDR TERRAIN

VOR1

TFC >

F

TCAS TEST

< ET

COM1 121.800

RADAR ON UTC 14:41

ATC1 4336

RAT 15 oC

COM2

125.250

BRT DIM

Collins

10500

ITT

ITT

51626

TORQ TORQ

62.2 0.0

FIRE

3.40

PROP PROP

1740 1980

N1 NI

106.0 98.5

830 734

0 . 0NM 0 . 8NM 4 . 4NM 198 NM

130 FF 750 FF 0 430 122 PRESS 80 PRESS 0 120 OIL OIL 49 TEMP C 112 46 TEMP°C 73 o

TORQ TORQ

110.0 2000

RW25 ( 6 9 3 5) SXW152 KBJC

FIRE AFX

: - : - - : CL I MB - : - - : ( 6 9 3 5) 6 9 3 5 A - :- - : - : - - / 0 . 8NM

FM S DTK 25 1 ( 6 9 3 5) TT G - - : - 0 . 8N M

ITT ITT

24

251

W

ABOVE BELOW

SXW152

21

The TCAS II system consists of a TCAS II receiver-transmitter, and two mode S transponders. The TCAS II transponders contain dual-element antennas and are called diversity transponders. One element is on top of the fuselage and one element is on the bottom of the fuselage to help reduce the chance of losing aircraft targets while maneuvering. The system also receives altitude and vertical speed information from the pilot’s Air Data Computer (ADC1). If that system fails, the copilot’s Air Data Computer (ADC2) automatically provides information. Radio altitude information is provided from the radio altimeter, and heading information from the pilot’s AHRS. The system also receives inputs from the right weight-on-wheels switch and right landing gear downlock switch.

30

+10 -10

FORM AT < 2.5

-02

( 6 9 3 5) / 6 9 3 5 A +02 KEGE

<

5

TERR

RDR < WX T+5 .7

TFC <

F

TCAS TEST

GS

0

TAS

0

SAT 15 oC

ISA +13 oC

BRT DIM

Figure 16-181. TCAS II Test

If the intruder gets within 20 to 35 seconds (depending upon altitude) of the CPA, it is considered a threat, and a Resolution Advi-

The TCAS II system generates vertical guidance commands that are displayed on the pilots and copilots IVSIs in the form of vertical red and green bands. Vertical speeds located next to the red band are to be avoided. The vertical speed associated with the green band (either descending or climbing) is the vertical speed the pilot should attain. Intruder targets are displayed on the MFD on the TCAS Only Map, or may be overlaid on the Present Position Map. Aural alerts are sounded over the speakers, whether or not they are selected on, and also over the headsets. Controls for the


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TCAS II system are integrated into the RTU and the CDU. Either unit may be used to control the TCAS system. The TCAS II system is powered by the Left Generator Avionics Bus and is protected by a 5-amp TCAS circuit breaker located on the right circuit breaker panel. Power is applied to the system when the Avionics Master switch is turned on.

MFD Displays and Controls The TCAS Traffic Only Map may be selected by pressing the TFC line select key for more than 1 second. The TCAS Traffic Only Map will be displayed in the 10 nm range. The range

of the display may be adjusted from 5 nm to 50 nm using the RANGE knob on the Display Control Panel (DCP). The TFC key may also be used to select the TCAS Traffic Display on or off. Once the Traffic Only Map has been selected using the TFC key, the FORMAT key may be used to select the Plan Map, the Present Position Map, or the TCAS Only Map. The following messages appear along the right side of the display when appropriate. They are listed, as they would appear from top to bottom:

Table 16-4. TCAS MESSAGES ABS INOP (white)

If the Absolute Altitude Mode is selected and the airplane is below 18,000 feet P.A. this display is presented

ALT XXX (cyan)

If the Absolute Altitude Mode is selected and the airplane is above 18,000 feet P.A. this display will show airplane altitude in thousands and hundreds of feet Example: 23,000 feet = 230.

ABOVE/BELOW (white)

◊ OFF (cyan)

These messages indicate the operating altitude volume of the TCAS system. These messages will be shown as ABOVE, ABOVE BELOW, BELOW, or will be blank. The operating volume of each display is as follows: ABOVE = -2700 ft to +9900 ft BELOW = -9900 ft to +2700 ft ABOVE/BELOW = -9900 ft to +9900 ft Blank = -2700 ft to +2700 ft This message indicates that the OTHER TRAFFIC symbol has been selected off.

TFC (white or cyan)

This legend indicates that the TCAS II system has been selected for display (cyan), or has been selected off (white)

TCAS TEST (cyan)

This message indicates that the TCAS II is in the Test Mode. (Color is white if TCAS has not been selected.)

TCAS OFF (cyan)

This message indicates that the Standby Mode of the TCAS system has been selected, the standby mode of the transponder has been selected, or that the Mode C has been selected Off. (Color is white if TCAS display has not been selected.)

TA ONLY (cyan)

This message indicates that the TA Only Mode has been selected. It will always be displayed on the ground. The message will change color from cyan to yellow and flash when a TA is issued by the TCAS. (Color is white if TCAS display has not been selected.)

TCAS FAIL (yellow)

This message indicates a TCAS fault has been detected.

TA or RA with no bearing data

Two lines are provided for the first two detected TAs or RAs without valid bearing data. Each line of data will include the range of the intruder followed by the relative or absolute altitude, if available, and a rate-of-climb or descent direction arrow if applicable

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When the TCAS self-test is conducted, the following test pattern will be displayed on the MFD. ·• Traffic Advisory (solid yellow circle)will appear at 9 o’clock, range 2 miles, 200 feet below and climbing. • Proximate Traffic (solid cyan diamond) will appear at 1 o’clock, range 3.6 miles, 1000 feet below and descending. • Other Traffic (open cyan diamond) will appear at 11 o’clock, range 3. 6 miles, flying level 1000 feet above, and in level flight. • Resolution Advisory Traffic (solid red square) will appear at 3 o’clock, range 2 miles, 200 feet above, and in level flight.

PFD Displays For non-IFIS aircraft, the PFD does not display traffic unless in the reversionary mode. For IFIS aircraft the PFD can show traffic any time by selecting the TFC line key. The following TCAS messages and displays are provided just below the lower right corner of the EADIs. Table 16-5. TCAS II ANNUNCIATORS TRAFFIC (yellow or red)

TCAS FAIL (yellow)

This message will be yellow for a TA and red for an RA. It will flash approximately 6 times and then become steady.

This message is identical to the one shown on the MFD.

The following messages will be displayed on the right side of the PFD opposite the third Line Select Key. They are identical to those shown on the MFD. • TCAS TEST (white) • TCAS OFF (white) • TA ONLY (white) During a Resolution Advisory, red or red and green bands will be displayed on the IVSIs . There are two types of RAs; corrective and preventive. If a corrective RA is issued, red and green bands will be displayed. The green band indicates the rate-of-climb or descent required for the pilot to obtain in response to the RA. The red bands indicate the rate-of-climb or descent required for the pilot to obtain in response to the RA. The red bands indicate the rate-ofclimb and descent the pilot is to avoid during the response to the RA. If a preventive RA is issued, normally only a single red band will be displayed indicating the vertical speeds to be avoided. If intruders exist above and below the airplane, it is possible to have a green band covering the lower rates-ofclimb and/or descent followed by two red bands indicating the higher rate-of-climb and descent to avoid. During the TCAS self-test, the IVSIs will display the following test pattern. • A red band will extend from 0 fpm to the bottom of the display. • A green band will extend from 0 fpm to +300 fpm. • A red band will extend from +2000 fpm to the top of the display.


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System Characteristics Only the TA ONLY Mode is available during ground operations. The RA Mode is available after takeoff above approximately 1150 feet. The traffic Display may be overlaid on the radar or EGPWS display on the MFD. EGPWS and radar displays are not available on the TCAS Traffic Only Map. EGPWS voice alerts have priority over TCAS II voice messages. During such occasions, the TCAS II will automatically switch to the TA Only Mode with no TCAS voice messages. The TCAS II surveillance may not function at distances less than 900 feet. CLIMB and INCREASE CLIMB RAs are inhibited with flaps extended beyond the Approach position.

CLIMB and INCREASE CLIMB RAs are inhibited above 32,000 feet P.A. When below approximately 1000 feet, the TCAS II will automatically revert to the TA Only Mode. All RA and TA voice messages are inhibited below 600 feet AGL while climbing and 400 feet AGL while descending. DESCEND RAs are inhibited below 1200 feet AGL while climbing and below 1000 feet AGL while descending. INCREASE DESCENT RAs are inhibited below 1450 feet AGL.

Voice Messages The following voice message accompanies a TCAS II Traffic Advisory (TA).

Table 16-6. TCAS II TRAFFIC ADVISORY VOICE MESSAGE TRAFFIC, TRAFFIC

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PILOT RESPONSE Gain visual contact with traffic. Check the TCAS II display for range and bearing of the traffic if necessary. Assess the threat and prepare to execute the evasive maneuver if a Resolution Advisory is subsequently issued.


16 AVIONICS


The following voice messages accompany TCAS II Resolution Advisory Traffic (RAs). Table 16-7. TCAS II RESOLUTION ADVISORIES VOICE MESSAGE

PILOT RESPONSE

CLIMB, CLIMB, CLIMB (corrective)

Change vertical speed to 1500 fpm climbing, or as indicated by the green band on the IVSI.

CLIMB, CROSSING CLIMB, CLIMB, CROSSING CLIMB (corrective)

Same as previous except that this message indicates that flight paths will cross at some altitude.

INCREASE CLIMB, INCREASE CLIMB (corrective)

This follows a CLIMB voice message. The climbing vertical speed is typically increased to 2500 fpm as shown by the green band on the IVSI.

ADJUST VERTICAL SPEED, ADJUST (corrective)

Reduce climbing vertical speed to that shown on the IVSI.

DESCEND, DESCEND NOW (corrective)

This follows a CLIMB voice message. This message indicates that a reversal of vertical speed from a climb to a descent is needed to provide adequate separation.

DESCEND, DESCEND, DESCEND (corrective)

Change vertical speed to 1500 feet descending, or as indicated by the green band on the IVSI.

DESCEND, CROSSING DESCEND, DESCEND, CROSSING DESCEND (corrective)

Same as previous except that this message indicates that flight paths will cross at some altitude.

INCREASE DESCENT, INCREASE DESCENT (corrective)

This follows a DESCENT voice message. The descending vertical speed is typically increased to 2500 fpm as shown by the green band on the IVSI.

ADJUST VERTICAL SPEED, ADJUST (corrective)

Reduce descending vertical speed to that shown on the IVSI.

CLIMB, CLIMB NOW (corrective)

This follows a DESCEND voice message. This message indicates a reversal of vertical speed from a descent to a climb is needed to provide adequate separation.

CLEAR OF CONFLICT

Resume normal flight. Apparent conflict of airspace has been resolved.

MONITOR VERTICAL SPEED (preventive)

Be alert for approaching traffic. Ensure that the IVSI needle does not enter the area of the red band.

MAINTAIN VERTICAL SPEED (preventive)

Maintain present vertical speed and direction. Ensure that the IVSI needle does not enter the area of the red band.

MAINTAIN VERTICAL SPEED, CROSSING, MAINTAIN (preventive)

A flight path crossing is predicted, but being monitored by the TCAS II. Maintain present vertical speed and direction. Ensure that the IVSI needle does not enter the area of the red band.

ADJUST VERTICAL SPEED, ADJUST (preventive)

Indicates a weakening of the RA. This allows the pilot to start returning to an assigned altitude.


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16 AVIONICS



16-110


APPENDIX A – AVIONICS EQUIPMENT LOCATIONS

Nose Avionics: ADC 1 / 2 AHRS 1 / 2 COMM, NAV, DME: 1 / 2 EGPWS GPS 1 / 2 IAPS Standby Battery Weather Radar

Aft Avionics: Air Cell Satellite Phone CVR ELT FSU HF (and HF SelCal, if installed) TCAS I or II Transponder 1/2 Universal Weather (COMM 3 and CMU) XM Weather

Figure 16-182. Overview of Avionics Units


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16 AVIONICS


APPENDIX B – FLIGHT GUIDANCE MODES Table 16-8. FLIGHT GUIDANCE MODES

MODE (FGP Mode Button)

PFD ANNUNCIATION DEFINITION ARMED

ACTIVE LATERAL MODES

Roll Hold FD

N/A

ROLL

Holds bank angle present at the time it is selected or holds existing heading if the bank angle is 5˚ or less without reference to the heading bug. Default mode for the flight director if no other modes are selected, if flight guidance is transferred or if current lateral mode is deselected.

Heading Hold HDG

N/A

HDG

Holds the heading as selected by the Heading Bug. HDG is automatically selected when no other lateral mode is active and any other lateral or vertical mode is selected.

FMS Lateral Navigation NAV

FMS FMS1, FMS2

FMS FMS1, FMS2

Tracks the active course generated by the selected FMS. A single-FMS installation annunciates FMS. A dual-FMS installation annunciates FMS1 or FMS2, as appropriate.

VOR Lateral Navigation NAV

VOR1, VOR2

VOR1, VOR2

Tracks the selected VOR course from the selected NAV radio with a VOR frequency tuned. Annunciates VOR1 or VOR2 as appropriate to the selected radio.

Localizer Lateral Navigation NAV

LOC1, LOC2

LOC1, LOC2

Tracks the selected Localizer course from the selected NAV radio with a localizer frequency tuned. Annunciates LOC1 or LOC2 as appropriate to the selected radio.

FMS Approach APPR

APPR FMS, APPR FMS1, APPR FMS2

APPR FMS, APPR FMS1, APPR FMS2

Tracks the active course generated by the selected FMS. A single-FMS installation annunciates FMS. A dual-FMS installation annunciates FMS1 or FMS2, as appropriate.

VOR Approach APPR

APPR VOR1, APPR VOR2

APPR VOR1, APPR VOR2

Tracks the selected VOR course from the selected NAV radio with a VOR frequency tuned. Annunciates VOR1 or VOR2 as appropriate to the selected radio.

Localizer Approach APPR

APPR LOC1, APPR LOC2

APPR LOC1, APPR LOC2

Tracks the selected Localizer course from the selected NAV radio with a localizer frequency tuned and enables GS mode. Annunciates LOC1 or LOC2 as appropriate to the selected radio.

Go Around

N/A

GA

Go Around button on the left power lever pressed. Maintains the existing heading with a 5˚ bank limit. Does not reference the heading bug.


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Table 16-8. FLIGHT GUIDANCE MODES (Cont)



ACTIVE VERTICAL MODES

Pitch Hold FD

N/A

PTCH

Maintains the pitch present at the time the mode is selected. Default mode for the flight director if no other modes are selected, if flight guidance is transferred, or if current vertical mode is deselected. Can be adjusted with the UP/DN Wheel or the SYNC button.

Vertical Speed Hold VS

N/A

VS 1500

Maintains the vertical speed present at the time the mode is selected. Can be adjusted with the UP/DN Wheel or the SYNC button. Selected vertical speed is annunciated adjacent to VS.

Flight Level Change FLC

FMS FMS1, FMS2

FLC 160

Maintains the Indicated Airspeed at the time the mode is selected. Can be adjusted with the SPEED Knob or the SYNC button. Selected speed is annunciated adjacent to FLC.

Altitude Hold ALT

VOR1, VOR2

ALT

Maintaining an altitude other than the Preselected or VNAV altitude. Maintains the altitude present at the time the mode is selected. Can be adjusted with the SYNC button.

Preselect Altitude Hold

ALTS

ALTS

Preselected altitude is being maintained or will be maintained (if armed).

Glide Slope APPR

GS

GS

The APPR LOC mode has been selected and the flight director will, or has, intercepted the localizer glide slope. This mode will not recognize any Preselected or FMS generated altitudes.

Go Around

N/A

GA

Commands a +7o pitch attitude. Selected with the Go Around button on the left power lever.

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16 AVIONICS


Table 16-8. FLIGHT GUIDANCE MODES (Cont)



ACTIVE VNAV MODES

VNAV – Pitch Hold VNAV

PTCH

VPTCH

Pitch Hold Mode has been selected with VNAV enabled. Can be adjusted with the SYNC button. Armed mode exists if next leg does not have a VNAV path.

VNAV – Vertical Speed Hold VS + VNAV

N/A

VVS 1500

Vertical Speed Hold Mode has been selected with VNAV enabled. Selected vertical speed is shown adjacent to VVS. Can be adjusted with the UP/DN Wheel or the SYNC button.

VNAV – Flight Level Change FLC + VNAV

FLC

VFLC 160

Flight Level Change Mode has been selected (or armed by the FMS during a VNAV climb) with VNAV pressed. Selected speed is annunciated adjacent to VFLC. Can be adjusted with the SPEED Knob or the SYNC button.

VNAV – Altitude Hold ALT + VNAV

N/A

VALT

Maintaining an altitude other than the Preselected or VNAV altitude. Maintains the altitude present at the time the mode is selected. Can be adjusted with the SYNC button.

VNAV – Preselected Altitude Hold VNAV

ALTS

VALTS

Preselected altitude is being maintained or will be maintained (if armed) with VNAV enabled.

VNAV – FMS VNAV Altitude Hold VNAV

ALTV

VALTV

FMS VNAV altitude is being maintained or will be maintained with the altitude preselector set at a different altitude.

VNAV – PATH VNAV

PATH

VPATH

FMS has captured the manually or automatically generated descent angle to the next waypoint. Aircraft must stay within lateral deviation limits (cross-track error or track angle error) to remain active.

VNAV – Glide Path APPR + VNAV

GP

VGP

The APPR Mode has been selected and the FMS generated VNAV Glide Path is, or will be, captured. Ignores the Preselected altitude or FMS altitudes.


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16 AVIONICS


APPENDIX C – AVIONICS ACRONYMS A

E

ACP

Audio Control Panel

E-Chart

Electronic Charts

ADC

Air Data Computer

E-Maps

Enhanced Maps

ADF

Automatic Direction Finder

EDC

Engine Data Concentrator

ADI

Attitude Direction Indicator

EFIS

Electronic Flight Instrument System

AFD

Adaptive Flight Display

EGPWS

AFCS

Automatic Flight Control System

Enhanced Ground Proximity Warning System

AHC

Attitude Heading Computer

EIS

Engine Indicating System

AHRS

Attitude and Heading Reference System

AHS

Attitude Heading System

AM

Amplitude Modulation

FD

Flight Director

AP

Autopilot

FGC

Flight Guidance Computer

FGP

Flight Guidance Panel

FGS

Flight Guidance System

FMC

Flight Management Computer

FMS

Flight Management System

FSA

File Server Application

FSU

File Server Unit

B BFO

Beat Frequency Oscillator

C

F

CCW

Counterclockwise

CDU

Control Display Unit

CMU

Communication Management Unit

GCS

Ground Clutter Suppression

CPL

Couple

GPS

Global Positioning System

CVR

Cockpit Voice Recorder

GPWS

Ground Proximity Warning System

CW

Clockwise

GWX

Graphical Weather

G

D

H

DBU

Database Unit

HF

DCP

Display Control Panel

DCU

Data Concentrator Unit

High Frequency Radio


16-117

16 AVIONICS


I

O

IAPS

Integrated Avionics Processor System

IEC

IAPS Environmental Controller

IFIS

Integrated Flight Information System

IMU

Inertial Measurement Unit

IND

Indicators

IOC

Input / Output Concentrator

P PA

Passenger Address

PFD

Primary Flight Display

PTT

Press-to-Talk

Q J R K L LCD

Liquid Crystal Display

LSC/ISS

Low Speed Cue/Impending Stall Speed

LSK

Line Select Keys

LV

Lower Sideband Voice

RA

Resolution Advisory

RAT

Ram Air Temperature

RIU

Radio Interface Unit

RSS

Radio Sensor System

RTU

Radio Tuning Unit

S SAT

Static Air Temperature

SELCAL Selective Call SFDS

M

Secondary Flight Display System

MCDU

Maintenance Control Display Unit

MDC

Maintenance Diagnostic Computer

MFD(1)

Multifunction Display

TA

Traffic Advisory

MFD(2)

Multi-Function Display

TAWS

MFD(3)

Multifunctional Flight Display

Terrain Awareness and Warning System

TCAS

Traffic Alert Collision Avoidance System

TFC

Traffic

N NDB

16-118

T

Non-Directional Beacon


16 AVIONICS


U USTB

Unstabilized (Weather Radar)

UV

Upper Sideband Voice

V W X Y Z


16-119

16 AVIONICS



16-120


16 AVIONICS


QUESTIONS 1. The minimum autopilot use height during an approach is _______ feet. A. 79 B. 100 C. 400 D. 1000 2. A copilot side heading failure can be corrected by placing the: A. AHRS switch to No. 1. B. ADC switch to No. 2 C. PILOT DISPLAY to MFD. D. PILOT DISPLAY to PFD. 3. The active No. 2 bearing pointers are: A. Magenta. B. Cyan. C. Amber. D. Green. 4. ISA deviation can be found on the: A. RTU. B. Pilot PFD. C. Copilot PFD. D. MFD. 5. In order for the BARO MINS to be displayed, the values on the REFS page must be: A. White. B. Magenta. C. Cyan. D. Amber. 6. The composite mode is activated by selecting the _______ reversion switch. A. PILOT DISPLAY B. AHRS C. ADC D. RMT TUNE

7. P r e s s i n g t h e BA R O k n o b o n t h e display control panel (DCP) will: A. Display the on-side barometric minimums. B. Display the on-side and off-side barometric minmums. C. Set Flight Level altitudes on the altitude preselector display. D. Cycle the altimeter setting between inches of mercury and hectopascals. 8. The color of the to-waypoint on the CDU is: A. White. B. Blue. C. Magenta. D. Green. 9. A i r s p e e d t r e n d i n f o r m a t i o n i s available: A. On the MFD. B. On the CDU. C. From a magenta indicator on the airspeed indicator. D. Fr o m a c y a n i n d i c a t o r o n t h e airspeed indicator. 10. The minimum autopilot use height during an approach is _______ feet. A. 79 B. 10 0 C. 40 0 D. 10 00 11. An FMS preflight includes: A. Loading a SID. B. Checking the Database. C. Designating an alternate departure airport. D. Testing the TCAS I or TCAS II (as installed).


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16 AVIONICS


12. Using the FMS for guidance is not authorized: A. For WASS approaches. B. For a missed approach procedure. C. O u t s i d e t h e f i n a l a p p ro a c h f i x (FAF) on a localizer approach. D. I n s i d e t h e FA F o n a l o c a l i z e r approach. 13. For an FMS preflight procedure, the I in VIPP stands for: A. Initialize. B. Instrument. C. IFIS. D. Integrated. 14. For an FMS preflight procedure, the second P in VIPP stands for: A. Preflight. B. Performance. C. Plan. D. Precision. 15. APPR must be pressed: A. When cleared for any approach. B. Before cleared for any approach. C. When VNAV to a decision altitude is desired. D. To sequence the waypoints on a missed approach procedure.

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16. VNAV guidance is: A. Required inside the final approach fix (FAF). B. Prohibited inside the FAF. C. Required to fly a missed approach. D. P r o h i b i t e d d u r i n g a m i s s e d approach. 17. If GPS APPR is not displayed inside the final approach fix: A. The GPS must not be used for flight guidance. B. A d i f f e r e n t a p p r o a c h m u s t b e selected. C. The approach may be continued if the aircraft is in instrument meteorological conditions (IMC). D. The approach must be flown to a minimum descent altitude. 18. Magenta color text on the CDU LEGs page indicates _______ information. A. Bearing and distance B. Airspeed and altitude C. Course and wind D. Airspeed and distance


CHAPTER 16A WIDE AREA AUGMENTATION SYSTEM (WAAS) CONTENTS Page INTRODUCTION............................................................................................................. 16A-1 GENERAL ........................................................................................................................ 16A-1 OPERATION .................................................................................................................... 16A-3 Integrity ..................................................................................................................... 16A-3 Departures ................................................................................................................. 16A-3 Enroute ...................................................................................................................... 16A-4 Arrivals ...................................................................................................................... 16A-4 Approaches ................................................................................................................ 16A-4 Degraded SBAS Integrity During LPV Approach.................................................... 16A-9 Missed Approach..................................................................................................... 16A-10 Lateral Guidance ..................................................................................................... 16A-11 QUICK REFERENCE ROCKWELL COLLINS WAAS FMS (Version 4.0) ............... 16A-12 Select SBAS Provider ............................................................................................. 16A-13 Load LPV Approach ............................................................................................... 16A-13 Failure of SBAS During LPV Approach................................................................. 16A-15 Load LNAV/VNAV or LNAV Approach ................................................................ 16A-17 Failure of SBAS During LNAV/VNAV Approach.................................................. 16A-18 Load LNAV/VNAV Approach with WAAS (Rare)................................................. 16A-19 Load Non-GPS Approach ....................................................................................... 16A-20 Navigation Integrity ................................................................................................ 16A-21 RAIM Prediction..................................................................................................... 16A-22 ROCKWELL COLLINS FMS DIFFERENCES............................................................ 16A-23


16A-i

16A WIDE AREA AUGMENTATION SYSTEM



Title

Page

16A-1

Worldwide SBAS Providers............................................................................. 16A-2

16A-2

SBAS Service Providers .................................................................................. 16A-5

16A-3

Check SBAS Provider...................................................................................... 16A-5

16A-4

Approach Loading ........................................................................................... 16A-5

16A-5

Approach Selection.......................................................................................... 16A-6

16A-6

Arrival Data ..................................................................................................... 16A-6

16A-7

Non-WGS-84 Airport ...................................................................................... 16A-7

16A-8

WAAS Channel Number.................................................................................. 16A-7

16A-9

PFD Annunciations LPV Approach................................................................. 16A-8

16A-10 Course to Final Approach Message ................................................................. 16A-8 16A-11 SBAS Failure Messages................................................................................... 16A-9 16A-12 VNAV Flag ...................................................................................................... 16A-9 16A-13 Changing VNAV Guidance............................................................................ 16A-10 16A-14 PFD Annunciations Non-SBAS..................................................................... 16A-11 16A-15 Loss of Nonprecision Approach RAIM ........................................................ 16A-11 16A-16 Select SBAS Provider.................................................................................... 16A-13 16A-17 LPV Approach ............................................................................................... 16A-13 16A-18 Failure of SBAS During LPV Approach ....................................................... 16A-15 16A-19 Load LNAV/VNAV or LNAV Approach ....................................................... 16A-17 16A-20 RAIM Failure after SBAS Failure ................................................................. 16A-18 16A-21 LNAV/VNAV Approach with WAAS............................................................ 16A-19 16A-22 Load Non-GPS Approach.............................................................................. 16A-20 16A-23 Navigation Integrity....................................................................................... 16A-21 16A-24 RAIM Prediction ........................................................................................... 16A-22


16A-iii



TABLE Table 16A-1

Title

Page

Loss of Integrity................................................................................................. 16A-3


16A-v



CHAPTER 16A WIDE AREA AUGMENTATION SYSTEM (WAAS)

INTRODUCTION For the standard GPS system to provide lower minimums on an approach the GPS signal needed to be corrected. The correction was primarily needed to increase the accuracy of vertical navigation but lateral navigation was also improved.

GENERAL Two forms of correction have been implemented to achieve this goal: Ground-based Augmentation Systems (GBAS) and Satellitebased Augmentation Systems (SBAS). GBAS uses towers in the vicinity of an airport that correct the GPS signal locally and send the correction message back to the aircraft using

VHF radios. The special equipment requirements for this system have limited its implementation to a small number of airports and operators [the FAA has termed this as a Local Area Augmentation System (LAAS)].


16A-1





SBAS is much more widely implemented. In the US, over 2,000 runway ends are served by SBAS approaches. The FAA has termed this as a Wide Area Augmentation System (WAAS) because it does not rely on airport specific towers to correct the signal and send the correction message. Instead, it uses data from stations throughout North America and a correction signal from geostationary satellites. SBAS approved units are able to receive correction messages from these satellites and create a very accurate vertical and lateral navigation unit. (See gps.faa.gov and the Aeronautical Information Manual (AIM) for more information). Other countries will label SBAS differently when it is implemented as shown in Figure 16A-1.

The Rockwell Collins FMS version 4.0 is the unit needed to use the SBAS system in Collins equipped aircraft. This FMS is used with a SBAS capable receiver labeled GPS-4000S. The FMS uses the corrected signal to create appropriate vertical and lateral navigation displays during all phases of flight to include WAAS approaches. SBAS and other software/ equipment upgrades are included with FMS v4.0 and this addendum will highlight the most critical. Refer to the appropriate Collins FMS user guide, AFM or AFM supplement for a more complete listing of limitations. The FMS v4.0 upgrade includes a new Flight Management Computer (FMC) and processor. This allows for the increased rate of error checking and position updates that occur during WAAS flight and approaches. Additionally, updating the FMS database should be faster through the DBU-5000 since the communication speed has increased.

Figure 16A-1. Worldwide SBAS Providers

16A-2


OPERATION INTEGRITY WAAS geo-stationary satellites provide integrity messages for the FMS v4.0. When the FMS detects a navigational problem “LOSS OF INTEGRITY” will show on the CDU and MFD. The PFD will also show an “LOI” or “LOI TERM” message depending on the phase of flight (see Table 16A-1). When the “LOSS OF INTEGRITY” message i s a c t iv e t h e F M S m u s t n o t b e u s e d a s primary navigation. If only the WAAS signal is degraded but the GPS signal is unaffected (for instance, a loss of geo-stationary satellites or being outside of WAAS ground station coverage) no messages will appear for non-SBAS procedures since they do not require WAAS. The FMS will automatically begin using what is called Receiver Autonomous Integrity Monitoring (RAIM). RAIM is the error checking technique used by all non-SBAS units or in SBAS units after SBAS has failed.

The aircraft position will not be as accurate but is still well within the boundaries of standard RNAV operations. If the RAIM error gets too large, the FMS will post the “LOSS OF INTEGRITY” message as previously discussed.

DEPARTURES During RNAV departures CDI deflection values will match the navigational performance requirements of the procedure. US RNAV departures and Europe P-RNAV departures are labeled RNAV 1 and the CDI will be ± 1nm for the entire procedure. This will be annunciated as “TERM” on the PFD. CDI deflection values will change according to the following: • ± 1 nm: On a departure procedure OR within 31nm of an airport • ± 2 nm: Outside of 31nm from an airport AND not on a departure

Table 16A-1. LOSS OF INTEGRITY CDU

PFD

MFD

Terminal (within 31nm of origin airport or on a RNAV departure)

Enroute (outside of 31nm of origin and not on a RNAV departure)


16A-3





ENROUTE

APPROACHES

During the enroute phase of flight CDI deflection values will be ± 2nm unless on a RNAV departure or RNAV arrival. If those procedures are active the CDI deflection will be ± 1nm as discussed earlier.

The most signif icant changes for the Collins FMS v4.0 will be in the approach phase of flight. The FMS is now capable of flying RNAV (GPS) or RNAV (GNSS) approaches to the Localizer Performance with Vertical (LPV) guidance minimums. If airport marking and approach lighting standards are met, some LPV DA minimums can be 200 feet above the runway surface. However, LPV approaches are part of the group labeled Approaches with Vertical Guidance (APV) and are not considered Precision approaches.

US RNAV airways labeled “Q” and “T”-routes are labeled as RNAV 2 procedures. Once the RNAV departure is f inished, the CDI deflection will be ± 2nm on these airways and remain that way until joining an RNAV arrival or arriving within a 31nm ring around the destination air por t. Europe B-RNAV routes are labeled as RNAV 5 procedures but the CDI will remain at ± 2nm as discussed. The PFD will not show an annunciator when in the enroute scale. When the aircraft is beyond ground-based navaid services volumes, CDI deflection will change. Deflection values will be ±4nm and the label “OCEANIC” will annunciate on the PFD. This will continue until the aircraft is back inside navaid service volumes and the enroute or terminal mode is automatically reselected, as appropriate.

ARRIVALS During RNAV arrivals CDI deflection values will match the navigational performance requirements of the procedure. US RNAV arrivals and Europe P-RNAV arrivals are labeled RNAV 1 and the CDI will be ± 1nm for the entire procedure. This will be annunciated as “TERM” on the PFD. Navigational integrity and messages on the CDU, PFD, and MFD are the same as discussed in the Departures section.

16A-4

SBAS Provider The appropriate SBAS providers are chosen on the “SBAS SERVICE PROVIDERS” CDU page. This can be found on the GNSS Control page under the main index [IDX]. The GNSS control page will show how many are enabled as shown on the Figure 16A-2. Each provider on the SBAS Service Providers page can be manually enabled or disabled by pressing the appropriate left line select key. The following providers are on this page: 1. Wi d e A r e a A u g m e n t a t i o n S y s t e m (WAAS) for the US; 2. European Geostationary Navigational Overlay System (EGNOS) for Europe; 3. MTSAT Satellite based Augmentation System (MSAS) for Japan; and 4. GPS-Aided GEO Augmented Naviga tion (GAGAN) for India. Enabling an SBAS provider will allow the FMS to use it should the aircraft fly into that region of the world.


quired provider is enabled. The approach can still be flown to LNAV/VNAV or LNAV minimums since these do not require SBAS. The SBAS Service Providers page does not have a default selection and once the appropriate SBAS is enabled it will remain that way for every flight.

Loading The Approach The DEP/ARR key is used to load a SBAS approach. The instrument approach listing is labeled “APPROACHES” and the visuals are labeled “RUNWAYS” (Figure 16A-4). The FMS is able to load multiple named approaches such as the RNAV (GPS) Y 10L and RNAV (GPS) Z 10L as shown in the f igure.

Figure 16A-2. SBAS Service Providers

As each area develops LPV minimum approaches, the FMS database will contain the required SBAS provider for that approach (only one SBAS provider is actively used by the FMS at any one time). If the appropriate SBAS provider is not enabled once the approach is loaded, a “CHK SBAS SVC PRVDR” message will appear on the CDU when within the terminal area (Figure 16A-3). The approach cannot be continued to LPV minimums until the re-

Figure 16A-3. Check SBAS Provider

Figure 16A-4. Approach Loading


16A-5





Pressing next to the desired approach will turn the label green and display available transitions (Figure 16A-5). The VECTORS option is always chosen by default and will initially display in green. Selecting another transition will turn its label green and change VECTORS to white.

Pressing the Execute key will load the approach into the active flight plan. Colors for the selected approach are the same before and after the execute key is pressed.

Arrival Data Page The ARR DATA line select key is a shortcut to the Active Arrival Data page. This page can also be accessed from the main index [IDX] (Figure 16A-6).

SBAS APPROACH

Figure 16A-5. Approach Selection

Additionally, VNAV guidance for the selected approach and the required SBAS provider (if appropriate) will display at the 5R key. In the example, “WAAS LPV” indicates the US WAAS system is required and the approach will use LPV vertical guidance. It must be understood that this label does not indicate the actual navigation integrity available but is only database information.

16A-6

NON-SBAS APPROACH

Figure 16A-6. Arrival Data

For non-SBAS approaches this page is only informational and not required to be viewed. For SBAS approaches it provides information for the approach and is the only page where the pilot can change approach VNAV guidance: LPV or BARO (discussed later in this section).




The following paragraphs provide a brief description of the Arrival Data page. The GNSS label indicates whether the approach can be flown as a GPS overlay.If NO, ground-based navaids that define the approach must be tuned, in view during the approach, and must be used as final authority to determine whether to continue or execute a missed approach. If YES, the procedure may be flown using only the FMS. The World Geodetic System (WGS-84) will indicate if the airport is referenced to standard GNSS coordinates. If the WGS-84 label is NO, the FMS must not be used as primary navigation or reference navigation when it is using GPS. The location of f ixes and airports could be very different than their actual positions. If an approach is loaded at an airport not referenced to WGS-84, a CDU message “NON-WGS-84 AIRPORT” will indicate the need to rely on ground based navigation (Figure 16A-7).

Figure 16A-7. NON-WGS-84 Airport Figure 16A-8. WAAS Channel Number

The Channel number will only display on approaches with SBAS guidance. This number is a unique identifier for that approach and can be referenced from the approach chart. Every SBAS approach will have a Channel number assigned (Figure 16A-8). (Used with permission from Jeppesen.) The Required Provider label is derived from the FMS database and indicates which SBAS provider must be enabled as discussed earlier in this section.

Approach VNAV Selection Before discussing approaches it is necessary to review Collins vertical navigation. Non-SBAS FMS units accomplish VNAV by using barometric inputs (“baro-VNAV”) from the altimeter system. This is used during enroute and terminal operations. It is also used on

LNAV/VNAV approaches to DA minimums. Baro-VNAV, however, is only as accurate as the altimeter system on board the aircraft and is affected by normal barometric errors (temperatures colder and hotter than ISA, inappropriate barometric settings, etc.) SBAS FMS’s will use two forms of VNAV; BaroVNAV and GPS altitude VNAV (LPV VNAV). Baro-VNAV will be used for select procedures where highly accurate vertical navigation is not required. GPS altitude VNAV will be used where highly accurate vertical navigation is required. GPS altitude VNAV does not rely on altimeter indications and is not affected by altimeter errors because it is created by the SBAS signal. This vertical navigation is similar to an ILS glideslope because it is unaffected by temperatures or inappropriate barometric settings. SBAS FMS units will use baro-VNAV for enroute proce-


16A-7



dures, terminal procedures and non-LPV approaches. GPS altitude VNAV will only be used for LPV approaches.

Flying the LPV Approach Once an LPV approach is loaded in the CDU the integrity of SBAS is monitored continuously. Within 31nm of the destination airport “LPV TERM” will annunciate in white on the PFD (Figure 16A-9). During this phase of flight CDI deflection will be ± 1nm. Baro-VNAV will be used with a Vertical Deviation Indicator (VDI) deflection of ± 500 ft. When the aircraft is past the Final Approach Course Fix (FACF), the SBAS integrity is appropriate for the approach, and the course leg to the FAF is within 45 degrees of the inbound course, “LPV APPR” will annunciate in green on the PFD (Figure 16A-9). The FACF is the f ix immediately prior to the FAF. The change from LPV TERM to LPV APPR occurs at the FACF because the aircraft will transition from baro-VNAV to LPV VNAV. Baro-VNAV will be affected by the surrounding temperature and the two glidepaths may not coincide. The glidepath indicator (“snowflake”) may appear to move suddenly when transitioning from baro-VNAV to LPV VNAV and more time is needed to be established on glidepath before

crossing the Final Approach Fix (FAF). If VNAV is already selected on the flight guidance panel the aircraft will smoothly increase or decrease the rate of descent as required to center the new LPV glidepath. Once LPV APPR is annunciated, lateral and vertical guidance is angular and will get more and more sensitive to course deviations during the approach descent. (This is similar to ILS and glideslope guidance). Lateral CDI deflections start at ± 1nm and will decrease to approximately ± 350 ft at the runway end. Vertical VDI deflections start at ± 500 ft and will decrease to the appropriate scale needed for that approach. The amber message “CRS TO FAF>45 DEG” will appear on the CDU if a “Direct-to” the FAF creates a leg more than 45 degrees to the inbound (Figure 16A-10). Sequencing to LPV APPR will be delayed until the “Direct-to” leg is fixed.

Figure 16A-10. Course To Final Approach Message

Figure 16A-9. PFD Annunciations LPV Approach

16A-8


Descent on the LPV approach is accomplished using the APPR and VNAV modes on the flight guidance panel. FMS APPR and VGP will be annunciated on the PFD. Missed approach operations are the same as non-LPV approaches.

DEGRADED SBAS INTEGRITY DURING LPV APPROACH The following messages will appear any time SBAS integrity degrades during an LPV approach. “LPV NOT AVAILABLE” will display on the CDU and, if applicable, “USE LNAV MINIMUM” will display on the CDU and MFD (Figure 16A-11). Additionally, the PFD will display a flashing amber “MSG” indicating the CDU has an active message.

“USE LNAV MINIMUM” will appear only if the approach has an LNAV minimum published. For approaches that do not have LNAV minimums published, an “APPR NOT AVAILABLE” message will appear and a missed approach must be flown. If the label “LPV APPR” was already present on the PFD, this label will remain even though the integrity is degraded. The amber messages must be acknowledged and the appropriate changes made to the approach briefing. With SBAS integrity degraded, the vertical deviation indicator will be removed when inside the FACF and a red “VNV” label will appear indicating the loss of vertical integrity. Active VNAV modes will be removed (will change to VPTCH) and armed VNAV modes will be lined out as seen in the figure (Figure 16A-12). Further descent can only be accomplished using nonVNAV modes (e.g., VS, FLC).

Figure 16A-11. SBAS Failure Messages

“LPV NOT AVAILABLE” indicates SBAS integrity is not sufficient for the LPV approach. Similar to an ILS with glideslope failure, a decision can be made to continue the approach but descending only to the published LNAV minimum, or executing a missed approach.

Figure 16A-12. VNAV Flag


16A-9





Prior to the FAF Prior to the FAF, baro-VNAV can be manually selected to recover vertical guidance after the LPV VNAV has failed. VNAV will then be available to continue to LNAV/VNAV minimums or LNAV minimums, as appropriate. This is accomplished on the Active Arrival Data page by pressing DEP/ARR and choosing ARR DATA (Figure 16A-13). Pressing the APPR VNAV GP will select between GPS altitude VNAV (LPV) and baro-VNAV (BARO). Once BARO is selected the change in VNAV must be executed. VNAV will return and the approach can continue to LNAV/VNAV minimums or LNAV minimums. It is critical to understand that LPV minimums are not to be flown during this operation. PFD annunciations will display “TERM” and “GPS APPR” instead of “LPV TERM” and “LPV APPR” (Figure 16A-14) Additionally, “LPV NOT AVAILABLE” and “USE LNAV MINIMUM” messages will be removed from the displays and the CDU message page.

After the FAF If SBAS guidance fails after the FAF, the descent may be continued to the LNAV minimum or a missed approach can be flown. If a descent is continued it can only be done using VS, FLC, or PTCH mode since baro-VNAV is not selectable at this point and VNAV deviation will be flagged inoperative.

MISSED APPROACH Pressing the go-around button will allow the FMS to sequence to missed approach fixes after reaching the missed approach point. Lateral guidance will remain in approach mode while on final and then sequence to terminal mode, as appropriate, when past the missed approach point. PFD annunciations will change to “TERM” to indicate when the CDI scale has changed.

16A-10

Figure 16A-13. Changing VNAV Guidance




Figure 16A-14. PFD Annunciations Non-SBAS

LATERAL GUIDANCE SBAS corrections for lateral guidance will be used on all GPS approaches. If SBAS lateral integrity fails or the aircraft is outside SBAS coverage, the FMS will automatically begin using RAIM as discussed earlier. Should RAIM fail “NO NPA RAIM” will annunciate on the CDU when inside the 31nm terminal area with an approach loaded (NPA =Nonprecision Approach). The FMS must not be used as primary navigation with this message active (Figure 16A-15). Additionally, if a “LOSS OF INTEGRITY” message posts at any time before or during an approach the approach must be abandoned and the FMS must no longer be used as primary navigation.

Figure 16A-15. Loss of Nonprecision Approach RAIM


16A-11



QUICK REFERENCE ROCKWELL COLLINS WAAS FMS (VERSION 4.0)

16A-12


SELECT SBAS PROVIDER Choose the appropriate SBAS provider for world region (Figure 16A-16):

If appropriate provider is not chosen, a “CHK SBAS SVC PRVDR” message will appear on the CDU message line when loading an LPV approach. If no SBAS providers are chosen, the FMS will not use augmented signals.

LOAD LPV APPROACH Procedures for loading an LPV approach are the same as loading a non-LPV approach (Figure 16A-17, Sheet 1 of 2). 1. Conf irm desired airport is in ORIGIN or DESTination on the active flight plan page 2. Choose an APPRoach, and the desired transition (VECTOR is always default)

Figure 16A-16. Select SBAS Provider

Figure 16A-17. LPV Approach (Sheet 1 of 2)

WAAS = North America EGNOS = Europe MSAS = Japan 1. Press IDX

GNSS Control

2. Choose SELECT SBAS (R5) 3. Press left line select key to Enable the desired provider


16A-13





Figure 16A-17. LPV Approach (Sheet 2 of 2)

3. “WAAS LPV” is displayed at R5 a. In Europe, “EGNOS LPV” b. In Japan, “MSAS LPV” c. This label only indicates the selected approach has an LPV minimum published. It is NOT real-time display of system capability. 4. Verify LEGS page or MFD MAP to ensure proper information 5. EXECute after conf irmation The PFD will display “LPV TERM” in white when within 31nm of the desired air por t (Figure 16A-17, sheet 2 of 2). The PFD will display “LPV APPR” in green after passing the Final Approach Course Fix (FACF) if the SBAS system is operational. Baro-VNAV is used up until LPV APPR is annunciated at which time GPS corrected VNAV (LPV VNAV) will be used for the remainder of the approach. A slight jump in the vertical deviation indicator may be noticeable during this transition. Baro-VNAV temperature restrictions do NOT apply to LPV VNAV.

16A-14


NOTES

FAILURE OF SBAS DURING LPV APPROACH

Prior to FAF 1. These messages will appear on the CDU:

The following procedures assume only the SBAS system has failed. The GPS system is still operating normally. RAIM prediction and RAIM checking will automatically be used by the FMS as in nonSBAS units.

a. “LPV NOT AVAILABLE” b. Also, if LNAV minimums are published “USE LNAV MINIMUM” 2. If LNAV minimums are published, this message will appear on the MFD: a. “USE LNAV MINIMUM”

If the whole GPS system fails then a non-GPS approach would have to be flown as per AFM or AFM supplement guidance (Figure 16A-18, Sheet 1 of 3). Inside 31nm to airport but prior to FAF:

3. An amber MSG will flash on the PFD 4. The VNAV deviation will have a red VNV flag with the deviation indicator removed

APPR FMS

VALT GP

185 180 160

700

2 1

10

600 60 6540 20

14 1 0

1

10

120

4000 4

20 DN

3 000

400

2 4

100

LPV APPR MSG

FMS1

251

30.16IN

1000

W

21

30

DTK 251 RALPE 2.5NM

24

Figure 16A-18. Failure of SBAS During LPV Approach (Sheet 1 of 3)


16A-15





5. Aircraft can be descended with nonVNAV (VS, FLC, etc.) modes to the LNAV minimum OR 5. Aircraft can be descended using VNAV with manual selections (Figure 16A-18, Sheet 2 of 3): a. Press DEP / ARR ARR DATA or Press IDX page 2 ARR DATA b. Choose BARO (L4) as the APPR VNAV GP c. EXECute VNAV change d. Verify VNAV indications have returned on the PFD e. Use baro-VNAV to descend to appropriate minimums (LNAV/VNAV or LNAV)

The PFD will display “GPS APPR” in green when within 2nm of the FAF.

Inside the FAF 1. These messages will appear on the CDU: a. “LPV NOT AVAILABLE” b. Also, if LNAV minimums are published “USE LNAV MINIMUM” 2. If LNAV minimums are published, this message will appear on the MFD: a. “USE LNAV MINIMUM” 3. An amber MSG will flash on the PFD (Figure 16A-18, Sheet 3 of 3)

The PFD will display “TERM” in white when within 31nm of the desired airport.

Figure 16A-18. Failure of SBAS During LPV Approach (Sheet 3 of 3)

4. The VNAV deviation will have a red VNV flag with the deviation indicator removed 5. Depending on aircraft altitude, aircraft may be descended with non-VNAV (VS, FLC, etc.) modes to the LNAV minimum OR 5. Execute published missed approach Figure 16A-18. Failure of SBAS During LPV Approach (Sheet 2 of 3)

16A-16

Selections back to baro-VNAV guidance are NOT allowed inside the FAF.


LOAD LNAV/VNAV OR LNAV APPROACH 1. Conf irm desired airport is in ORIGIN or DESTination on the active flight plan page 2. Choose an APPRoach, and the desired transition (VECTOR is always default) 3. “GNSS BARO” is displayed at R5 (Figure 16A-19) a. This label only indicates the selected approach will be using baroVNAV. It is NOT real-time display of system capability.

4. Verify LEGS page or MFD MAP to ensure proper information 5. EXECute after conf irmation The PFD will display “TERM” in white when within 31nm of the desired airport. The PFD will display “GPS APPR” in green when within 2nm of the FAF. Baro-VNAV is used for the entire procedure. Baro-VNAV temperature restrictions apply to LNAV/VNAV minimums.

MOD KICT ARRIVAL STARS NONE

1/2

APPROACHES

RNV

01L TRANS

VECTORS BACAY CUTIK

- - - - - -
GNSS BARO ARR DATA>

LEGS> [ EXEC

EXEC

Figure 16A-19. Load LNAV/VNAV or LNAV Approach


16A-17





FAILURE OF SBAS DURING LNAV/VNAV APPROACH

NOTES

No messages will appear if the SBAS signal fails during an LNAV/VNAV or LNAV approach provided the navigation integrity from the GPS remains within limits. RAIM prediction and RAIM checking will automatically be used by the FMS as in nonSBAS units. Inside 31nm to airport (Figure 16A-20):

ACT LEGS

1/6 SEQUENCE AUTO/INHIBIT

MSG

KICT / o 309 12NM ---/----ICT / o 9.2NM 307 MUGER ---/----/ o 3.3NM 307 WUKOL ---/----/ o 0.5NM / 307 WUKUS ---/----------------LEG WIND> [ [ NO NPA RAIM

EXEC

Figure 16A-20. RAIM Failure after SBAS Failure

1. If RAIM is insufficient for the approach this message will appear on the CDU a. “NO NPA RAIM” 2. An amber MSG will flash on the PFD 3. Accomplish a non-GPS approach as per AFM or AFM supplement

16A-18


LOAD LNAV/VNAV APPROACH WITH WAAS (RARE) The following images and information are available in the Collins FMS but no procedures have been designed, as of this printing, by the FAA.

4. Verify LEGS page or MFD MAP to ensure proper information 5. EXECute after conf irmation The FMS will use any available SBAS provider for lateral navigation.

1. Confirm desired airport is in ORIGIN or DESTination on the active flight plan page

The PFD will display “L/V TERM” in white when within 31nm of the desired airport.

2. Choose an APPRoach, and the desired transition (VECTOR is always default)

The PFD will display “L/V APPR” in green when within 2nm of the FAF.

3. “SBAS L/V” is displayed at R5 (Figure 16A-21)

The FMS will use baro-VNAV until the FACF and then transition to SBAS VNAV just like LPV approaches.

a. This label only indicates the selected approach will be using SBAS VNAV. It is NOT real-time display of system capability.

Baro-VNAV temperature restrictions do not apply when using SBAS VNAV. For failure of SBAS integrity, see the LPV approach section.

Figure 16A-21. LNAV/VNAV Approach with WAAS FOR TRAINING PURPOSES ONLY

16A-19





LOAD NON-GPS APPROACH

A “NO APPR” label will appear on the PFD.

1. Conf irm desired airport is in ORIGIN or DESTination on the active flight plan page

An “APPR FOR REF ONLY” will appear on the CDU.

2. Choose an APPRoach, and the desired transition (VECTOR is always default)

Verify AFM or AFM supplement limitations for navigation guidance requirements.

3. “BARO” is displayed at R5 (Figure 16A-22)

NOTES

a. This label only indicates the selected approach will be using baroVNAV. It is NOT real-time display of system capability. 4. Verify LEGS page or MFD MAP to ensure proper information 5. EXECute after conf irmation

Figure 16A-22. Load Non-GPS Approach

16A-20




NAVIGATION INTEGRITY

NOTES

If the navigation integrity falls outside of tolerance for the phase of flight (enroute or terminal) a message will be displayed on the CDU and PFD. This message is a total FMS integrity message and will appear whether SBAS is being received or not (Figure 16A-23). 1. A “LOSS OF INTEGRITY” message will appear on the CDU 2. A “LOI” or “LOI TERM” will appear on the PFD depending on the 31nm distance from the airport 3. Use another source of navigation

Figure 16A-23. Navigation Integrity


16A-21



RAIM PREDICTION

These are the possible outcomes of approach RAIM prediction:

RAIM prediction will only be necessar y when outside the coverage of SBAS or during SBAS NOTAM’s indicating an outage of signal integrity. 1. Press IDX

GNSS CONTROL

2. C h o o s e N PA R A I M ( L 5 ) ( F i g u r e 16A-24) 3. Destination airport will automatically be f illed with flight plan destination airport 4. E n t e r s a t e l l i t e s t h a t h av e b e e n NOTAM’d out of service in the deselect option in L3 5. The ETA will automatically be f illed when inflight or it can be manually entered in R2 (i.e., when still on the ground)

Figure 16A-24. RAIM Prediction

16A-22


AVAILABLE UNAVAILABLE REQ PENDING

NOTES



ROCKWELL COLLINS FMS DIFFERENCES Non-WAAS

WAAS (v4.0)

“GPS” label on applicable pages

“GNSS” label on applicable pages

No Space Based Augmentation System (SBAS)

Uses Space Based Augmentation System (SBAS) US = WAAS Europe = EGNOS Japan = MSAS India =GAGAN

VNAV

VNAV Enroute / Terminal Uses Baro-VNAV only ( ± 500 FT) Approaches Uses Baro-VNAV only ( ± 250 FT)

Enroute / Terminal Uses Baro-VNAV only ( ± 500 FT) Approaches LPV minimums WAAS only (Angular) LNAV / VNAV minimums Baro-VNAV ( ± 250 FT) WAAS when FAA certified (Angular) LNAV minimums Baro-VNAV only ( ± 250 FT)

RNAV SID / RNAV STAR ± 1nm CDI within 30nm of ARPT ± 5nm CDI outside of 30nm Must do RAIM prediction

RNAV SID / RNAV STAR ± 1nm CDI for entire procedure (“TERM”) ± 1nm CDI when off procedure within 31nm of ARPT ± 2nm CDI when off procedure outside 31nm of ARPT RAIM prediction only when WAAS fails

Q Routes / T Routes ± 1nm CDI within 30nm of ARPT ± 5nm CDI outside of 30nm Must do RAIM prediction

Q Routes / T Routes ± 1nm CDI within 31nm of ARPT ± 2nm CDI outside 31nm RAIM prediction only when WAAS fails

Approaches Cannot choose multiple label approaches

Approaches Can choose multiple label approaches e.g., RNAV (GPS) Y Rwy 10 / RNAV (GPS) Z Rwy 10

GPS APPR mode ~2nm from FAF

LPV APPR mode after FACF L/V APPR mode after FACF GPS APPR mode ~2nm from FAF

Non-GPS approches can be flown without messages

Non-GPS approaches will have “APPR FOR REF ONLY” CDU message “NO APPR” PFD message

No stepdown fixes inside FAF

All stepdown fixes inside FAF (non-ILS)


16A-23


CHAPTER 17 OXYGEN SYSTEM CONTENTS INTRODUCTION ............................................................................................................ 17-1 GENERAL ......................................................................................................................... 17-1 OXYGEN SYSTEM .......................................................................................................... 17-2 Components................................................................................................................. 17-2 Controls and Indicators............................................................................................... 17-4 Operation ..................................................................................................................... 17-6 OXYGEN DURATION ................................................................................................... 17-8 Time of Useful Consciousness .................................................................................. 17-8 SERVICING ....................................................................................................................... 17-9 Purging ....................................................................................................................... 17-10 Cylinder Retesting .................................................................................................... 17-10 QUESTIONS.................................................................................................................... 17-11


17-i

17 OXYGEN SYSTEM

Page


ILLUSTRATIONS Title

Page

17-1

Oxygen Supply Cylinder .................................................................................... 17-2

17-2

Crew Oxygen Masks ........................................................................................... 17-3

17-3

Passenger Masks.................................................................................................. 17-3

17-4

First-Air Oxygen Mask....................................................................................... 17-4

17-5

Oxygen System Controls .................................................................................... 17-5

17-6

Oxygen System Annunciators ........................................................................... 17-5

17-7

Pressure Gauge.................................................................................................... 17-6

17-8

Oxygen System Schematic ................................................................................. 17-7

17-9

Oxygen Duration with Partially Full Bottle..................................................... 17-8

17-10

Oxygen Duration................................................................................................. 17-9

17-11

Oxygen Fill Valve and Gauge............................................................................ 17-9

TABLE Table 17-1

Title

Page

Time of Useful Consciousness........................................................................... 17-8


17-iii

17 OXYGEN SYSTEM

Figure


17 OXYGEN SYSTEM

CHAPTER 17 OXYGEN SYSTEM

INTRODUCTION Pilot and passenger comfort and safety are of prime importance in operating this aircraft. Understanding the proper use of oxygen is crucial for both. Federal Aviation Regulations (FARs) require that any time an aircraft flies above 25,000 feet, oxygen must be immediately available to the crew and passengers. This chapter presents a discussion of the oxygen system. Operation, controls, and procedures along with oxygen duration charts are included.

GENERAL Th e K i n g A i r 3 5 0 a i r c r a f t p r o v i d e s adequate oxygen flow for crew and passengers for a cabin pressure altitude of up to 35,000 feet.

The system consists of an oxygen bottle mounted in the aircraft tail section, crew masks, passenger oxygen masks, and a firstaid mask.


17-1


17 OXYGEN SYSTEM

The crew has two push-pull controls, a pressure gauge, and appropiate annunciators in the cockpit.

for unpressurized, high-altitude flight. The cylinder can come in three sizes: 50, 77, or 115 cubic foot.

The normal mask flow rate is 3.9 liters per minute (liters per minute-normal temperature/pressure differential). The cockpit diluter-demand masks used by the flight crew use twice the normal amount in 100% or EMERG selection.

A shutoff valve and regulator is attached to the end of the oxygen cylinder, The regulator is constant flow and supplies low pressure oxygen through plumbing to the outlets in the aircraft.

Before each flight, check that the crew masks are in the 100% mode.

A push/pull lever in the cockpit controls the shutoff valve and regulator.

OXYGEN SYSTEM

Fill the cylinder through a valve accessible through an access door on the right side of the aft fuselage.

COMPONENTS

Crew Oxygen Masks

Oxygen Supply Cylinder

Th e c re w h a s d i l u t e r- d e m a n d , q u i c k donning oxygen masks stowed in the cockpit overhead between the pilot and copilot seats (Figure 17-2).

A cylinder mounted behind the aft pressure bulkhead (Figure 17-1) supplies the oxygen

Figure 17-1. Oxygen Supply Cylinder

17-2



Donning Masks

Testing Masks Figure 17-2. Crew Oxygen Masks

Even while stowed, these masks are always plugged into the oxygen system. They can be donned easily and quickly with one hand. The diluter-demand crew masks deliver oxygen to the user only upon inhalation. A small switch on each mask permits two modes of operation: NORMAL and 100%. In the NORMAL position, air from the cockpit is mixed with the oxygen supplied through the mask. This reduces the rate of depletion of the oxygen supply. It is also more comfortable to use than 100% aviator breathing oxygen. However, if smoke or fumes fill the cockpit, use the 100% position to prevent breathing contaminated air. For this reason, the selector levers should be left in the 10 0% position when the masks are not in use so they are always ready for emergency use.

Push the red pushbutton/knob on the crew mask to test the mask. When the pushbutton is depressed, oxygen flows.

Passenger Masks The pressure of oxygen in the passenger oxygen system supply line automatically extends a plunger against each of the passenger mask dispenser doors. This forces the doors open. The masks then drop to hang about nine inches below the dispensers (Figure 17-3). The lanyard valve pin at the top of the mask hose must be pulled out so oxygen can flow to the masks. The pin is connected to the oxygen mask via a flexible cord. When the oxygen mask is pulled down for use, the cord pulls the pin out of the lanyard valve. The lanyard valve pin must be manually reinserted into the valve in order to stop the flow of oxygen when the mask is no longer needed.

A red pushbutton/knob on the mask is for testing and selecting the EMERG position. For certain emergencies, the crew masks must be selected to EMERG. When the red knob is turned to the EMERG position, a small constant, positive flow of oxygen flows to the mask. In EMERG, the crew mask is considered a constant-flow type until a small amount of pressure builds in the face mask. Figure 17-3. Passenger Masks


17-3

17 OXYGEN SYSTEM

Squeeze the red finger grip control switch on the face of the mask to inflate the elastic straps with oxygen pressure. Fit the inflated straps over the head. Hold the mask over the nose and mouth as the grip is released. The elastic straps then deflate and hold the mask tightly against the face.


17 OXYGEN SYSTEM

First-Aid Oxygen Mask

CONTROLS AND INDICATORS

Whenever the primary oxygen supply line is charged (PULL ON SYSTEM READY control knob pulled), oxygen is available to the first-aid mask in the toilet compartment in the aft fuselage (Figure 17-4). A manual ON/OFF valve in the overhead container box releases oxygen to the mask.

The crew has two controls for the oxygen system. One arms the system. The second one is an override handle to ensure passenger masks deploy.

PULL-ON Handle The PULL ON- SYSTEM READY handle on the left side of the center pedestal arms t h e o x y g e n s y s t e m ( Fi g u re 17- 5 ) . Th e handle operates a cable that opens and closes the shutoff valve at the oxygen supply cylinder. When the handle is pushed in, oxygen is unavailable in the aircraft. When the handle is pulled out, the primary oxygen supply line is charged with oxygen as long as the cylinder is not empty. Pull the handle out prior to engine starting to ensure oxygen is immediately available any time it is needed.

Passenger Handle A PASSENGER MANUAL DROP OUT push-pull control handle is on the right side of the center pedestal. This is an override handle that manually opens the passenger oxygen shutoff valve. This action pressurizes the passenger oxygen system whenever the primary oxygen supply line is armed. Electrical power is not required for the manual system to operate. Figure 17-4. First-Air Oxygen Mask

17-4


17 OXYGEN SYSTEM


Figure 17-5. Oxygen System Controls

Figure 17-6. Oxygen System Annunciators

Annunciators An oxygen pressure switch at the passenger system shutoff valve senses pressure in the primary oxygen supply line. When the system is not armed and there is no pressure in the primary oxygen supply line, the switch illuminates the amber OXY NOT ARMED annunciator (Figure 17-6). If sufficient pressure is in the line, the OXY NOT ARMED annunciator extinguishes.

A second oxygen pressure switch is in the ceiling in the aft passenger compartment. Oxygen pressure activates the switch, which then illuminates the PASS OXYGEN ON annunciator. This advises the crew that the masks are deployed and oxygen is available to the passengers.


17-5


Pressure Gauges The oxygen system has two pressure gauges: one on the copilot right subpanel and one adjacent to the filler valve to check system pressure while servicing (Figure 17-7). Refer to Servicing section.

The PULL ON SYSTEM READY control in the crew compartment must be on to supply oxygen to the crew and passengers. Oxygen then flows from the cylinder bottle through the plumbing to the various outlets (Figure 17-8).

17 OXYGEN SYSTEM

Passenger Autodeployment The barometric pressure switch (sensor) in the upper sidewall forward of the right emergency exit automatically deploys the passenger oxygen masks when cabin altitude reaches 12,50 0 feet. When activated, oxygen pressure is released to the container boxes. A plunger extends to open the doors and drop the masks. The oxygen valve lanyard pin must be pulled for oxygen to flow to each mask.

Figure 17-7. Pressure Gauge

OPERATION If decompression at altitude occurs, the b o d y ’s p r i m a r y n e e d i s f o r o x y g e n t o prevent hypoxia. Hypoxia is a lack of the oxygen needed to keep the brain and other body tissues functioning properly. Early symptoms such as an increased sense of well-being quickly give way to slow reactions, impaired thinking ability, unusual fatigue, and a dull headache. The crew, therefore, must act quickly to don oxygen masks and provide oxygen to the passengers before the onset of hypoxia.

17-6

When oxygen flows into the passenger oxygen system supply line, a pressure switch illuminates the green PASS OXYGEN ON annunciator. This switch also causes the cabin indirect lights to illuminate bright regardless of the CABIN LIGHTS BRIGHT-DIM-OFF switch position. In addition, the NO SMOKING and FASTEN S E AT B E LT s i g n s i l l u m i n a t e a n d a chime sounds. When the cabin descends below 12,500 feet or electrical power is removed from the circuit, the shutoff valve closes to terminate oxygen flow to the passenger masks. When the masks are no longer required, reinserting the lanyard pin stops the flow of oxygen. Th e OX Y C O N T RO L c i rc u i t b re a ke r on the copilot CB panel powers the automatic sensor.



FORWARD PRESSURE BULKHEAD

DILUTER DEMAND CREW MASK

TO ANNUNCIATOR PASS OXYGEN ON

PASSENGER MANUAL OVERRIDE SHUTOFF VALVE

17 OXYGEN SYSTEM

CREW MASK

TO COCKPIT OXYGEN PRESSURE GAGE

SOLENOID OFF

ON BAROMETRIC PRESSURE SWITCH

CONTROL CABLE

PASSENGER 2 MASK OUTLET (TYPICAL 5 PLACES)

OXYGEN PRESSURE SENSE SWITCH PASSENGER SINGLE MASK OUTLET FIRST AID OXYGEN MASK STOWED IN MANUALLY OPERATED BOX OXYGEN MASK CONTAINER, LINES AND OUTLET FOR FOLD-UP SEATS AFT PRESSURE BULKHEAD

CONTROL CABLE

HIGH PRESSURE OVERBOARD RELIEF

LEGEND HIGH PRESSURE LINE

PRESSURE REGULATOR AND SHUTOFF VALVE COMPOSITE OXYGEN CYLINDER

LOW PRESSURE LINE OXYGEN CYLINDER

Figure 17-8. Oxygen System Schematic


17-7


Manual Deployment

Preflight Requirement

If the barometric pressure switch malfunctions, the pilot may manually arm the system by pulling out on the PASSENGER MANUAL DROP OUT handle.

It is a preflight requirement to check the oxygen available for the crew and passengers to ensure it is sufficient for descent to 12,50 0 feet or until loss of pressure in the aircraft can be corrected and cabin altitude pressure restored. Full oxygen system pressure is 1,80 0 ±50 psi at 21°C.

Crew Only Usage 17 OXYGEN SYSTEM

To shut off passenger oxygen and isolate the remaining oxygen to the crew and first-aid o u t l e t s, p u s h i n t h e PAS S E N G E R MANUAL DROP OUT handle. Then pull the OXY CONTROL circuit breaker in the ENVIRONMENTAL group on the copilot CB panel.

OXYGEN DURATION TIME OF USEFUL CONSCIOUSNESS

For all oxygen duration computations, count each diluter-demand crew mask in use in the 10 0% mode as two people. For example, to compute oxygen duration for four passengers and two crewmembers, consider it as eight people using oxygen. In order to determine if sufficient oxygen is available: 1. Note pressure on oxygen pressure gauge

The Time of Useful Consciousness (Table 17-1) shows the average time of useful consciousness available at various altitudes. This is the time from the onset of hypoxia u n t i l l o s s o f e f f e c t i v e p e r f o r m a n c e. Individual reactions may differ from those shown in the table.

2. Use Oxygen Available with a Partially Full Bottle graph (Figure 17-9) to find the percent of usable capacity

Table 17-1. TIME OF USEFUL CONSCIOUSNESS TIME OF USEFUL CONSCIOUSNESS ALTITUDE

TIME

35,000 FEET 30,000 FEET 28,000 FEET 25,000 FEET 22,000 FEET 12 TO 18,000 FEET

1/2 TO 1 MINUTE 1 TO 2 MINUTES 2 1/2 TO 3 MINUTES 3 TO 5 MINUTES 5 TO 10 MINUTES 30 MINUTES OR MORE

Using the Emergency Descent procedure i n t h e P O H c a n m i n i m i z e t h e e ff e c t s of hypoxia.

Figure 17-9. Oxygen Duration with Partially Full Bottle

3. To obtain duration in minutes, find d u ra t i o n f o r a f u l l b o tt l e f o r t h e number of persons aboard from the Oxygen Duration chart (Figure 17-10)

17-8


17 OXYGEN SYSTEM


Figure 17-10. Oxygen Duration

Figure 17-11. Oxygen Fill Valve and Gauge

4. Multiply full bottle duration by the percent of usable capacity Th i s i s t h e a va i l a b l e o x y g e n d u ra t i o n in minutes.

SERVICING Refer to the manufactuter’s Pilot Operating Handbook and Maintenance Manual prior to purging or servicing the oxygen system.

Service the system through the filler valve. Remove the access plate on the right side of the aft fuselage. A pressure gauge is adjacent to the filler valve to check system pressure during filling (Figure 17-11). Remove the protective cap from the filler valve. Attach the hose from an oxygen recharging unit. Make sure that both the aircraft oxygen system and the servicing equipment are properly grounded before servicing the system.


17-9


17 OXYGEN SYSTEM

WARNING

CAUTION

To prevent overheating, fill the oxygen system slowly by adjusting the recharging rate with the pressure regulating valve on the recharging unit. All oxygen cylinders should be filled to 1,800 psi at a temperature of 21°C. This p re s s u re m a y b e i n c re a s e d a n additional 3.5 psi for each degree above 21°C; similarly, for each degree below 21°C, reduce the pressure for the cylinder by 3.5 psi. When the oxygen system is properly charged, disconnect the filler hose from the filler valve, and replace the protective cap. When servicing, purging, or replacing the oxygen cylinder, if it becomes necessary to disconnect a fitting, the threads of the fitting should be wrapped with MIL-T27730 anti-seize tape prior to being reconnected to the system.

U s e o n l y Av i a t o r s B re a t h i n g Oxygen (MIL-0-27210) for servicing the oxygen system. Do not use oxygen intended for medical purposes or such industrial uses as welding. Such oxygen may contain e x c e s s i v e m o i s t u re t h a t c o uld freeze in the valves and lines of the oxygen system.

WARNING The following precautions should b e o b s e r v e d w h e n p u rg i n g o r servicing the oxygen system:

PURGING Purge the oxygen system to remove offensive odors. The system should also be purged any time system pressure drops below 50 psi or lines are left open. To accomplish purging, connect a recharging unit into the system. Permit oxygen to flow through the lines and outlets until any offensive odors have been carried away. If any offensive odor lingers, continue purging the system for an additional hour. If such odors still remain, replace the supply cylinder. After the system has been adequately purged, return it to its normal operation position, and service the system.

CYLINDER RETESTING • Avoid any operation that could create sparks. Keep all burning cigarettes or fire away from the vicinity of the aircraft when the outlets are in use. • Inspect the filler connection for cleanliness before attaching it to the filler valve. • Make sure that your hands, tools, and clothing are clean, particularly of grease or oil stains, for these contaminants are extremely dangerous in the vicinity of oxygen. • As a further precaution against fire, open and close all oxygen valves slowly during filling.

17-10

Oxygen cylinders are of two types: 3HT or 3A/3AA. Lightweight cylinders marked 3HT on the side plate must be hydrostatically tested every three years. The test date is on the cylinder. The bottle has a service life of 4,380 pressurizations or 24 years, whichever occurs first. It then must be discarded. Regular weight cylinders are stamped 3A or 3AA. They must be hydrostatically tested every five years. Service life on these cy l i n d e r s i s n o t l i m i t e d . Re f e r t o t h e manufacturer’s maintenance manual for more details.



QUESTIONS

17 OXYGEN SYSTEM

1. Deployment of the passenger oxygen masks is indicated by illumination of the _________ annunciator. A. Red [PASS OXYGEN ON] warning B. Red [CABIN ALT HI] warning C. Amber [CABIN OXYGEN ON] caution D. W h i t e [ PAS S O X Y G E N O N ] system status 2. Manual deployment of the passenger oxygen masks is available by _______ the control handle on the _______ side of the center console. A. Pulling out; left B. Pulling out; right C. Pushing in; left D. Pushing in; right 3. Th e a m b e r [ OX Y N OT A R M E D ] caution annunciator illuminates when the: A. Main oxygen system is not armed. B. Crew oxygen system is not armed. C. Passenger oxygen system is being used. D. Crew oxygen system is being used. 4. Crew oxygen is provided by a _________ oxygen mask. A. Constant-flow B. Pressure-demand C. Constant-flow quick-donning D. Diluter-demand quick-donning


17-11

18 MISCELLANEOUS




18-i


CHAPTER 19 MANEUVERS AND PROCEDURES CONTENTS INTRODUCTION............................................................................................................. 19-1 GENERAL ......................................................................................................................... 19-1 MANEUVERS ................................................................................................................... 19-2 One Engine Inoperative............................................................................................. 19-2 Stalls.............................................................................................................................. 19-3 Flutter ........................................................................................................................... 19-4 Turbulent Weather ...................................................................................................... 19-4 Windshear .................................................................................................................... 19-5 Flight in Icing Conditions .......................................................................................... 19-5 Wake Turbulence......................................................................................................... 19-6 Takeoff and Landing Conditions .............................................................................. 19-6

19 MANEUVERS AND PROCEDURES

FLIGHT PROFILES......................................................................................................... 19-6


19-i



Title

Page

Normal Takeoff and Departure ......................................................................... 19-7

19-2

Engine Loss at or above V1 ............................................................................... 19-8

19-3

Rejected Takeoff.................................................................................................. 19-9

19-4

Approach to Stall—Landing Configuration Model 350............................... 19-10

19-5

Approach to Stall—En Route Configuration ............................................... 19-11

19-6

Approach to Stall—Takeoff Configuration ................................................... 19-12

19-7

Approach to Stall—Approach Configuration ............................................... 19-13

19-8

Visual Approach and Landing......................................................................... 19-14

19-9

Visual Approach—No Flaps............................................................................ 19-15

19-10

One Engine Inoperative—Visual Approach and Landing .......................... 19-16

19-11

ILS Approach .................................................................................................... 19-17

19-12

Nonprecision Approach ................................................................................... 19-18

19-13

Circling ............................................................................................................... 19-19


19-1


19-iii


INTRODUCTION The crew must be thoroughly familiar with all information published by the manufacturer about the aircraft. In additional to maintenance inspections and preflight information required by federal regulations, a complete, careful preflight inspection is imperative before each flight.

GENERAL Maintain center of gravity (CG) within the approved envelope throughout the planned flight. Ensure the aircraft is loaded so it does not exceed weight and CG limitations. Refer to the manufacturer’s Pilot Operating Handbook (POH) and Airplane Flight Manual (AFM).

In addition to maintaining the altitude appropriate for the direction of flight, pilots flying VFR at night should maintain a safe minimum altitude as dictated by terrain, obstacles such as TV towers, or communities in the area. This applies especially in mountainous terrain where there is usually very little ground reference.


19-1


CHAPTER 19 MANEUVERS AND PROCEDURES


Minimum clearance is 2,000 feet above the highest obstacle enroute. Do not depend on the ability to see obstacles in time to miss them. Flight on dark nights over sparsely populated country can be the same as IFR.

Asymmetric Thrust

During normal two-engine operations, always fly the published takeoff speeds on initial climb out. Then accelerate to your cruise climb airspeed after you have obtained a safe altitude. Use cruise climb airspeed to give you increased in-flight visibility and better fuel economy.

With only a single engine, airflow over the wing is also reduced. Yaw also affects the lift distribution over the wing. This causes a ro l l t o w a rd t h e i n o p e ra t i v e e n g i n e. Balance these forces by banking slightly (up to 5º) into the operating engine.

Loss of one engine creates yaw because of the asymmetric thrust. To compensate, balance yaw forces with the rudder.

Airspeed

MANEUVERS ONE ENGINE INOPERATIVE Safe flight with one engine out requires an understanding of basic aerodynamics and proficiency in engine-out procedures.

Climb Performance


Loss of power from one engine affects both climb performance and controllability. Climb performance depends on an excess of power normally required for level flight. With twin engine aircraft and one engine inoperative, power loss is even more than 50%. Climb performance is reduced by at least 80%. Consult charts in the manufacturer’s POH and the FAA-approved AFM for confirmation. Single-engine climb performance depends on four factors: • Airspeed—Too little or too much decreases climb performance • Drag—Gear, flaps, and windmilling prop • Power—Amount available in excess of that needed for level flight • Weight—Passengers, baggage, and fuel load

Airspeed is the key to safe single-engine o p e r a t i o n s. A t f i r s t i n d i c a t i o n o f a n engine failure during climb out or while on approach, establish V YSE or V XSE , whichever is appropriate.

VMCA VMCA, the airspeed below which directional control cannot be maintained, is designated by the red radial on the airspeed indicator. Adhering to the practice of never flying at or below the published VMCA virtually eliminates loss of directional control issues.

VYSE VYSE , the airspeed that gives the best single engine rate of climb with an engine out, is designated by the blue radial on the airspeed indicator. V YS E delivers the greatest gain in altitude in the shortest possible time. It is based on the following criteria: • Critical engine inoperative; its propeller in minimum drag position • Operating engine at not more than maximum continuous power • Landing gear retracted • Wing flaps in most favorable (i.e., best lift-drag ratio) position • Aircraft flown at recommended bank angle

19-2



VXSE V XSE , the airspeed that gives the steepest angle of climb with an engine out, is only for clearing obstructions during initial climbout. It gives the greatest altitude gain per unit of horizontal distance. It requires more rudder control input than V YSE .

Single Engine Service Ceiling The single engine service ceiling is the maximum altitude at which the aircraft climbs at a rate of at least 50 feet per minute in smooth air. Use the single-engine service ceiling graph during flight planning to determine whether the aircraft, as loaded, can maintain the minimum enroute altitude (MEA) if IFR, or terrain clearance if VFR, following an engine failure.

3. Reduce drag to an absolute minimum. 4. Secure failed engine and related subsystems. The first three steps are completed from memory without consulting the POH. Consult the checklist for step 4 and beyond. The aircraft must be banked about 5º into the operating engine with the slipskid ball slightly out of center toward the operating engine to achieve rated performance. Be sure to identify the inoperative engine positively before securing it. Verify with cautious throttle movement.

STALLS Th e s t a l l w a r n i n g s y s t e m m u s t b e operational at all times. This is especially important in all high performance multiengine aircraft during engineout practice or stall demonstrations. The single-engine stall speed of a twin-engine aircraft is generally slightly below the power off (engines idle) stall speed for a given weight condition. Single-engine stalls in multi-engine aircraft are not recommended.

Know and follow to the letter the singleengine emergency procedures specified in your POH and AFM. All the procedures have the same basic fundamental steps:

Do not attempt V M CA demonstrations when altitude and temperature are such that the engine-out minimum control speed is known or discovered to be close to stalling speed. Loss of directional or lateral control just as a stall occurs is potentially hazardous.

1. M a i n t a i n a i r c r a f t c o n t r o l a n d airspeed at all times.

Low altitude stalls have not been approved by the Department of Transportation.

Single Engine Procedures

2. Normally apply 100% torque to the operating engine. If the engine failure occurs at a speed below VMCA, however, or during cruise or a steep turn, the crew may elect to use only enough power to maintain a safe speed and altitude. If the failure occurs on final approach, use power only as necessary to complete the landing.

Spins A major cause of fatal accidents in general aviation aircraft is a spin. Stall demonstrations and practice are a means for a pilot to acquire the skills to recognize when a stall is about to occur and to recover as soon as the first signs are evident. If a stall does not occur, a spin cannot occur.


19-3


Drag caused by a windmilling propeller, extended landing gear, or flaps in the landing position severely degrades or destroys single-engine climb performance. Since climb performance varies widely with weight, temperature, altitude, and aircraft configuration, the climb gradient (altitude gain or loss per mile) may be marginal or even negative under some conditions. Study the POH and AFM to know what performance to expect.


The King Air 350 has not been tested for spin recovery characteristics. Intentional spins are prohibited. If application of stall recovery controls is delayed, a rapid rolling and yawing motion may develop even against full aileron and r u d d e r . Th i s r e s u l t s i n t h e a i r c r a f t becoming inverted during the onset of a spinning motion. The longer the pilot delays before taking corrective action, the more difficult recovery becomes. Always remember that extra alertness and good pilot techniques are required for slow flight maneuvers, including the practice or demonstration of stalls or V MCA . Ensure that the CG is as far forward as possible. Fo r w a rd C G a i d s s t a l l re c o v e r y, s p i n avoidance, and spin recovery. An aft CG can create a tendency for a spin to flatten out. This delays recovery.

FLUTTER

If excessive vibration, particularly in the c o n t ro l c o l u m n a n d r u d d e r p e d a l s, i s encountered in flight, this could be the onset of flutter. • Immediately reduce airspeed; lower landing gear if necessary • Restrain controls of the aircraft until vibration ceases • Fly at reduced airspeed; land at the nearest suitable airport. • Have aircraft thoroughly inspected


Flutter is a phenomenon that can occur w h e n a n a e ro d y n a m i c s u r f a c e b e g i n s v i b r a t i n g . Th e e n e r g y t o s u s t a i n t h e vibration is derived from airflow over the surface. The amplitude of the vibration can decrease if airspeed is reduced, remain constant if airspeed is held constant with no failures, or increase to the point of selfdestruction, especially if airspeed is high and/or is allowed to increase. Flutter can lead to an in-flight break up of the aircraft. Aircraft are designed so flutter does not occur in the normal operating envelope as long as the aircraft is properly maintained. Decreasing the damping and stiffness of the structure or increasing the trailing edge weight of control surfaces tends to cause flutter. If a combination of those factors is sufficient, flutter can occur within the normal operating envelope. A thorough preflight inspection can eliminate things that might cause flutter. Improper tension on control cables or any

19-4

other loose condition in the flight control system can cause or contribute to flutter. Pay particular attention to control surface attachment hardware including tab pushrod attachment. Rectify any looseness of fixed surfaces or movement of control surfaces other than in the normal direction of travel. Control surface drain holes must be open to prevent freezing of accumulated moisture. This also could create an increased trailing edge heavy control surface and flutter.

TURBULENT WEATHER A complete and current weather briefing is a requirement for a safe trip. Updating of weather information enroute is also essential. Plan the flight to avoid areas of reported severe turbulence. Regard thunderstorms, squall lines, and violent turbulence as extremely dangerous and avoid. Thunderstorms also pose the possibility of a lightning strike. A roll cloud ahead of a squall line or thunderstorm is visible evidence of violent turbulence. The absence of a roll cloud, however, should not be interpreted as turbulence free. It is not always possible to detect individual storm areas or find the in-between clear a r e a s. I f t u r b u l e n c e i s u n e x p e c t e d l y encountered, note the following. Proper airspeed answers two basic problems when flying through turbulent air.



Beware of over controlling in an attempt to correct for changes in attitude. Applying control pressure abruptly builds up G-forces rapidly. This could cause structural damage or even failure. Watch angle of bank. Make turns as wide and shallow as possible. Be equally cautious in applying forward or back pressure to keep the aircraft level. Maintain straight and level attitude in either up or down drafts. Use trim sparingly to avoid being grossly out of trim as the vertical air columns change velocity and d i re c t i o n . L o w e r t h e l a n d i n g g e a r, i f necessary, to avoid excessive airspeeds.

WINDSHEAR Windshears are rapid, localized changes in wind direction. They can occur vertically or horizontally. Windshear is very dangerous, particularly on approach to landing when airspeeds are slow. A horizontal wind shear is a sudden change in wind direction or speed that can transform a headwind into a tailwind. This produces a sudden decrease in airspeed because of the inertia of the aircraft. A vertical wind shear i s a s u d d e n u p d ra f t o r d o w n d r a f t . Microbursts are intense, highly localized severe downdrafts. The prediction of windshear is far from an exact science. Monitor airspeed carefully when flying in storms, particularly on approach. Be mentally prepared to add power and go around at the first indication of encountering a windshear.

FLIGHT IN ICING CONDITIONS Carefully review the POH and AFM to ascertain the required operable equipment for flight in icing conditions. In addition, understand the limits of approval or certification for the aircraft for flight in icing conditions. Observe procedures in the POH and AFM for operation in icing conditions. Activate deice and anti-icing systems before entering an area of moisture where the aircraft is likely to go through a freezing level. When icing is detected, the pilot must immediately divert by flying out of the area of visible moisture or go to an altitude where icing is not encountered. Maintain the minimum speed for operation in icing conditions. If airspeed deteriorates below this minimum, the angle of attack increases to the point where ice may build up on the under side of the wings aft of areas protected by deicing/anti-icing equipment. Ice build up and its extent in unprotected areas may not be directly observable from the cockpit. As ice accumulates on the aircraft, distortion of the wing airfoil increases drag and reduces lift. Stalling speeds increase. In addition, stall warning devices are not accurate and cannot be relied upon in icing conditions. The pilot must remain sensitive to any indication such as observed ice accumulation, loss of airspeed, need for increased power, reduced rate of climb, or sluggish re s p o n s e t h a t i c e i s a c c u m u l a t i n g o n unprotected surfaces. In icing conditions, disengage the autopilot at an altitude sufficient to permit the pilot to gain the feel of the aircraft prior to landing. The most important ingredients to safe flight in icing conditions are a complete and current weather briefing, sound pilot judgment, close attention to rate and type of ice accumulations, and knowledge that severe icing is beyond the capability of modern aircraft. React promptly.


19-5


If the crew maintains an excessive airspeed, structural damage or failure may occur. If the airspeed is too low, a stall could occur. Reduce speed to the turbulent air penetration speed or to maneuvering speed in the limitations of the POH and AFM. These s p e e d s p re s e n t t h e b e s t a s s u ra n c e o f avoiding excessive stress loads and providing the proper margin against inadvertent stalls due to gusts.


FLIGHT PROFILES

WAKE TURBULENCE Every aircraft generates wake turbulence while in flight. Part of this is from the propeller or jet engine and part is from the wing tip vortices. The larger and heavier the aircraft, the more pronounced and turbulent the wake is. In tests, vortex velocities of 133 knots have been recorded.

Specific flight profiles are graphically depicted on the following pages (Figure 19-1 through 19-13).

Encountering the rolling effect of wing tip vortices within two minutes after passage of large aircraft is extremely hazardous to small aircraft. The roll effect can exceed the maximum counter roll obtainable in a small aircraft. Turbulent areas may remain for three minutes or more. Plan to fly slightly above and to the windward side of other aircraft. There is no set rule to follow because of the wide variety of conditions that can exist. For a thorough discussion of this, consult the Airman’s Information Manual and Advisory Circular 90-23, Aircraft Wake Turbulences.

TAKEOFF AND LANDING CONDITIONS


The landing gear should remain extended for approximately 10 seconds longer than normal when taking off on runways covered with water or freezing slush. This allows the wheels to spin and dissipate the freezing moisture. Cycle the landing gear up, then down. Wait approximately five seconds and then retract again. Ensure the entire operation occurs below the maximum landing gear operating airspeed. Exercise caution when taking off or landing during gusty wind conditions. Be aware of the special wind conditions caused by buildings or other obstructions near the runway in a crosswind pattern.

19-6



AREA DEPARTURE / CLIMB PROFILE 1. 170 KIAS TO 10,000' 2. 160 KIAS 10,000' - 15,000' 3. 150 KIAS 15,000' - 20,000' 4. 140 KIAS 20,000' - 25,000' 5. 130 KIAS 25,000' - 30,000' 6. 120 KIAS 30,000' - 35,000'

CRUISE 1. ACCELERATE TO CRUISE SPEED 2. SET CRUISE POWER 3. COMPLETE CRUISE CHECKLIST

CLIMB-OUT 1. ACCELERATE TO 170 KIAS 2. COMPLETE CLIMB CHECKLIST

TAKEOFF

TAKEOFF ROLL 1. OBSERVE TORQUE AND ITT LIMITS

1. ROTATE AT VR TO APPROX. 10˚ NOSE UP 2. ESTABLISH POSITIVE RATE OF CLIMB 3. LANDING GEAR—UP 4. LANDING/TAXI LIGHTS—OFF 5. AIRSPEED—V35 UNTIL CLEAR OF OBSTACLES

IN POSITION

VYSE OR ABOVE

1. BRAKES—HOLD 2. SET STATIC TAKEOFF POWER 3. PROP RPM—1,700 4. AUTOFEATHER ANNUNCIATORS—ON 5. BRAKES—RELEASE


1. FLAPS—UP 2. YAW DAMP—ON 3. CLIMB POWER—SET

BEFORE TAKEOFF 1. CHECKLIST—COMPLETED 2. TAKEOFF BRIEFING—COMPLETED 3. CONFIRM V1, VR, AND V2

Figure 19-1. Normal Takeoff and Departure


19-7


NOTE:

1,500' AGL

DO NOT RETARD FAILED ENGINE POWER LEVER UNTIL THE AUTOFEATHER SYSTEM HAS COMPLETELY STOPPED PROPELLER ROTATION.

1. COMPLETE ENGINE FAILURE CHECKLIST CLEAN-UP ITEMS 2. LAND AS SOON AS PRACTICAL

CLIMB 1. VYSE (BLUE LINE) 2. APPROX. 9 - 10˚ PITCH

V2 1. CHECK MAX POWER (100% / 820˚) 2. AIRSPEED AT V2 3. VERIFY PROP FEATHERED

ENGINE LOSS 1. MAINTAIN RUNWAY HEADING

TAKEOFF

400' AGL (CLEAR OF OBSTACLES)

1. ROTATE AT VR TO APPROX. 10˚ NOSE UP 2. ESTABLISH POSITIVE RATE OF CLIMB 3. LANDING GEAR—UP

1. REDUCE PITCH TO 5˚ PITCH TO ACCELERATE TO VYSE (BLUE LINE) 2. V2 + 9 KT—FLAPS UP


NOTE: BEFORE TAKEOFF 1. FOLLOW NORMAL TAKEOFF PROCEDURES UNTIL AT OR ABOVE V1

IT MAY BE NECESSARY TO BANK AS MUCH AS 5˚ INTO THE GOOD ENGINE TO MAINTAIN RUNWAY HEADING. IT MAY REQUIRE ALMOST FULL RUDDER ON THE SIDE OF THE GOOD ENGINE TO KEEP THE BALL SLIGHTLY OFF CENTER.

Figure 19-2. Engine Loss at or above V1

19-8



EMERGENCY OR MALFUNCTION AT OR BELOW V1 1. RECOGNIZE REASON FOR REJECTING TAKEOFF 2. POWER LEVERS—GROUND FINE 3. BRAKING—AS NECESSARY 4. MAINTAIN RUNWAY HEADING

CLEAR OF RUNWAY 1. COMPLETE AFTER LANDING CHECKLIST

BEFORE TAKEOFF


1. FOLLOW NORMAL TAKEOFF PROCEDURES UNTIL INITIATING ABORT AT OR BELOW V1

NOTE:

WARNING DO NOT USE REVERSE THRUST WITH ONE ENGINE INOPERATIVE. CARE MUST BE EXERCISED WHEN USING SINGLE-ENGINE GROUND FINE ON SURFACES WITH REDUCED TRACTION.

IF REJECTED TAKEOFF IS DUE TO REASONS OTHER THAN ONE ENGINE POWER LOSS, REVERSE IS MOST EFFECTIVE AT HIGH SPEEDS; BRAKING IS MOST EFFECTIVE AT LOW SPEEDS

Figure 19-3. Rejected Takeoff


19-9


VMC CONDITIONS—500' AGL—BASE LEG (SIMULATOR TRAINING ONLY)

BEGINNING OF MANEUVER

1. TORQUE—30% 2. MAINTAIN 500' AGL

STALL AND RECOVERY

COMPLETION OF MANEUVER

AT FIRST SIGN OF IMPENDING STALL:

COMPLETION:

1. SIMULTANEOUSLY SET MAXIMUM POWER, ADJUST PITCH ATTITUDE TO 7° (GA), AND ROLL WINGS LEVEL

1. LEVEL AT 1,500' AGL 2. ACCELERATE TO 140 KIAS

2. ACCELERATE TO VREF + 10 KT— FLAPS APPROACH

3. FLAPS—APPROACH 4. GEAR—DOWN

3. POSITIVE RATE—GEAR UP 5. PROPELLERS—1,700 RPM 4. ACCELERATE TO 125 KIAS—FLAPS UP 6. FLAPS—DOWN 7. TORQUE—REDUCE TO 15%

5. MAINTAIN HEADING AND CLIMB TO PATTERN ALTITUDE OF 1,500' AGL

8. MAINTAIN 500' AGL, LEVEL TURN TO FINAL 9. MAINTAIN 15% TORQUE

19 MANEUVERS AND PROCEDURES HORN OR BUFFET NOTES: 1. DECREASE SPEED APPROXIMATELY 1 KT PER SECOND 2. PITCH ATTITUDE PRIOR TO STALL MAY BE APPROXIMATELY 12° NOSE UP 3. HORN SHOULD SOUND APPROXIMATELY 10 KTS ABOVE STALL SPEED

Figure 19-4. Approach to Stall—Landing Configuration Model 350

19-10




APPROACH AND RECOVERY


INITIAL CONDITION: 1. YAW DAMP—AS REQUIRED

1. SIMULTANEOUSLY ADVANCE POWER TO MAXIMUM AND ADJUST PITCH ATTITUDE TO MAINTAIN ALTITUDE

2. PROPELLERS—1,500 RPM


COMPLETION: 1. AT 140 KIAS—RESET CRUISE POWER 2. MAINTAIN ASSIGNED ALTITUDE

3. TORQUE—10% 2. AS AIRSPEED INCREASES, REDUCE PITCH ATTITUDE TO MAINTAIN ALTITUDE

4. MAINTAIN INITIAL HEADING


5. MAINTAIN INITIAL ALTITUDE

HORN OR BUFFET NOTES: 1. DECREASE SPEED APPROXIMATELY 1 KT PER SECOND 2. PITCH ATTITUDE PRIOR TO STALL MAY BE APPROXIMATELY 12° NOSE UP 3. HORN SHOULD SOUND APPROXIMATELY 10 KTS ABOVE STALL SPEED

Figure 19-5. Approach to Stall—En Route Configuration


19-11



STALL AND RECOVERY


HORN OR BUFFET INITIAL CONDITION:

RECOVERY AT HORN OR BUFFET:

1. YAW DAMPER REQUIRED

1. SIMULTANEOUSLY ROLL THE WINGS LEVEL AND REDUCE THE PITCH ATTITUDE, AS NECESSARY, TO STOP THE STALL WARNING (APPROX. 10˚)

2. PROPELLERS—1,700 RPM

COMPLETION: 1. RESET CRUISE POWER 2. MAINTAIN ALTITUDE/HEADING, OR AS DIRECTED

3. TORQUE—10% 2. MAINTAIN ALTITUDE AND ALLOW AIRSPEED TO INCREASE

4. MAINTAIN INITIAL HEADING 5. MAINTAIN INITIAL ALTITUDE

3. FLAPS—UP AT OR ABOVE VYSE (BLUE LINE)

6. FLAPS—APPROACH (BELOW TRIANGLE) 7. AT 110 KIAS OR LESS, SIMULTANEOUSLY SET THE TORQUE TO 50% (SIMULATED 100% TORQUE), ESTABLISH A BANK ANGLE OF 20˚ (NO MORE THAN 30˚), RAISE THE NOSE, AND CLIMB


8. THIS MANEUVER MAY BE PERFORMED WHILE MAINTAINING A 15 - 30˚ ANGLE OF BANK OR WHILE MAINTAINING A HEADING

NOTES: 1. DECREASE SPEED APPROXIMATELY 1 KT PER SECOND 2. PITCH ATTITUDE PRIOR TO STALL MAY BE APPROXIMATELY 20˚ NOSE UP 3. HORN SHOULD SOUND APPROXIMATELY 10 KTS ABOVE STALL SPEED

Figure 19-6. Approach to Stall—Takeoff Configuration

19-12



IMC CONDITIONS—1,500' AGL (SIMULATOR TRAINING ONLY)


STALL AND RECOVERY



COMPLETION: 1. LEVEL AT 1,500' AGL

2. MAINTAIN 1,500' AGL

1. SIMULTANEOUSLY SET MAXIMUM POWER AND ADJUST PITCH ATTITUDE TO 7° (GA)

3. FLAPS—APPROACH

2. POSITIVE RATE—GEAR UP

4. GEAR—DOWN

3. ACCELERATE TO 125 KIAS (BE-350) / 122 KIAS (BE-300)—FLAPS UP

1. TORQUE—30% - 40%

2. ACCELERATE TO 140 KIAS

5. PROPELLERS—1,700 RPM 6. TORQUE—REDUCE TO 15%

4. MAINTAIN HEADING AND CLIMB TO PATTERN ALTITUDE OF 1,500' AGL

7. DESCEND TO 1,000' AGL 8. LEVEL AT 1,000' AGL


9. MAINTAIN 15% TORQUE

HORN OR BUFFET NOTES: 1. DECREASE SPEED APPROXIMATELY 1 KT PER SECOND 2. PITCH ATTITUDE PRIOR TO STALL MAY BE APPROXIMATELY 20° NOSE UP 3. HORN SHOULD SOUND APPROXIMATELY 10 KTS ABOVE STALL SPEED

Figure 19-7. Approach to Stall—Approach Configuration


19-13


REJECTED/BALKED LANDING 1. POWER—MAXIMUM ALLOWABLE 2. PITCH—10˚ NOSE UP 3. AIRSPEED—VREF 4. ESTABLISH NORMAL CLIMB WHEN CLEAR OF OBSTACLES: 5. FLAPS—APPROACH AT OR ABOVE VREF +10 KT 6. GEAR—UP (WHEN POSITIVE CLIMB IS ESTABLISHED) 7. FLAPS—UP AT OR ABOVE VYSE (BLUE LINE)

INITIAL 1. OBTAIN ATIS 2. DESCENT CHECKLIST— COMPLETE

THRESHOLD 1. GEAR—RECHECK DOWN 2. AIRSPEED—VREF 3. POWER—IDLE

ARRIVAL 1. TORQUE—APPROX. 30% 2. 150 - 175 KIAS (TYPICAL) 3. START BEFORE LANDING CHECKLIST

LANDING 1. GROUND FINE OR REVERSE 2. BRAKES—AS NECESSARY

DOWNWIND 1. FLAPS—APPROACH 2. 130 - 140 KIAS

FINAL

ABEAM TOUCHDOWN POINT

1. 130 - 140 KIAS (VYSE MIN) WHEN LANDING ASSURED: 2. FLAPS—DOWN 3. TRANSISTION TO VREF 4. YAW DAMP—OFF 5. PROPS—FULL FORWARD

1. GEAR—DOWN 2. BEFORE LANDING CHECKLIST—COMPLETE


BASE 1. 130 KIAS (MIN REC)

CAUTION

CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICS, THE PROPELLER CONTROL MUST BE IN FULL INCREASE RPM POSITION.

NOTE: REVERSE IS MOST EFFECTIVE AT HIGHER SPEEDS; BRAKING IS MOST EFFECTIVE AT LOWER SPEEDS.

IF POSSIBLE, PROPELLERS SHOULD BE MOVED OUT OF REVERSE AT APPROXIMATELY 40 KNOTS TO MINIMIZE BLADE EROSION. CARE MUST BE EXERCISED WHEN REVERSING ON RUNWAYS WITH LOOSE SAND, DUST, OR SNOW ON THE SURFACE. FLYING GRAVEL WILL DAMAGE PROPELLER BLADES, AND DUST OR SNOW MAY IMPAIR THE PILOT'S VISIBILITY.

Figure 19-8. Visual Approach and Landing

19-14



INITIAL 1. OBTAIN ATIS 2. DESCENT CHECKLIST— COMPLETE

THRESHOLD 1. GEAR—RECHECK DOWN 2. AIRSPEED—FLAPS UP APPROACH SPEED—VREF + 20 KT 3. POWER—IDLE

ARRIVAL 1. TORQUE—APPROX. 30% 2. 150 - 175 KIAS (TYPICAL)


DOWNWIND 1. FLAPS—UP (INOPERATIVE) 2. AIRSPEED—140 KIAS 3. START FLAPS UP LANDING CHECKLIST 4. COMPUTE FLAPS UP APPROACH SPEED AND LANDING DISTANCE

1. AIRSPEED—140 KIAS WHEN LANDING ASSURED: 2. FLAPS—UP 3. TRANSITION TO FLAPS UP APPROACH SPEED— VREF + 20 KT 4. YAW DAMP—OFF 5. PROPS—CONFIRM FULL FORWARD

1. GEAR—DOWN 2. PROPS—FULL FORWARD 3. FLAPS UP LANDING CHECKLIST— COMPLETE UP TO LANDING ASSURED

BASE 1. AIRSPEED—140 KIAS

NOTE: FLAPS UP APPROACH SPEED IS VREF + 20 KT

CAUTION

CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICS, THE PROPELLER CONTROL MUST BE IN FULL INCREASE RPM POSITION.

NOTE: REVERSE IS MOST EFFECTIVE AT HIGHER SPEEDS; BRAKING IS MOST EFFECTIVE AT LOWER SPEEDS.


Figure 19-9. Visual Approach—No Flaps


19-15


FINAL ABEAM TOUCHDOWN POINT


INITIAL 1. OBTAIN ATIS 2. DESCENT CHECKLIST—COMPLETE

GO-AROUND 1. POWER—MAXIMUM ALLOWABLE 2. LANDING GEAR—UP 3. AIRSPEED—INCREASE TO 125 KIAS 4. FLAPS—UP


ARRIVAL 1. TORQUE—APPROXIMATELY 80% 2. AIRSPEED—150 - 175 KIAS (TYPICAL) 3. START ONE-ENGINE-INOPERATIVE APPROACH AND LANDING CHECKLIST

LANDING 1. GROUND FINE—AS NECESSARY 2. BRAKES—AS NECESSARY

DOWNWIND 1. FLAPS—APPROACH 2. AIRSPEED—130 - 140 KIAS


FINAL

ABEAM TOUCHDOWN POINT 1. GEAR—DOWN 2. PROP—FULL FORWARD

BASE 1. AIRSPEED—VREF + 15 KT

1. AIRSPEED—VREF + 15 KT WHEN IT IS CERTAIN THERE IS NO POSSIBILITY OF GO-AROUND: 2. FLAPS—DOWN 3. AIRSPEED—VREF 4. YAW DAMP—OFF 5. ONE-ENGINE-INOPERATIVE APPROACH AND LANDING CHECKLIST—COMPLETE

WARNING: DO NOT USE REVERSE THRUST WITH ONE ENGINE INOPERATIVE. CARE MUST BE EXERCISED WHEN USING SINGLE-ENGINE GROUND FINE ON SURFACES WITH REDUCED TRACTION.

Figure 19-10. One Engine Inoperative—Visual Approach and Landing

19-16



OM INITIAL 1. OBTAIN ATIS 2. BRIEF APPROACH AND MISSED APPROACH 3. FMS/NAV AIDS—SET UP/IDENT 4. DESCENT CHECKLIST— COMPLETE

GLIDE SLOPE INTERCEPT 1. TORQUE—20% - 30% 2. 130 - 140 KIAS (VYSE MIN)

MM DH-MISSED APPROACH 1. POWER—MAXIMUM ALLOWABLE 2. PITCH—7˚ NOSE UP (FD-GA) 3. GEAR—UP 4. FLAPS—UP 5. COMPLETE MISSED-APPROACH PROCEDURE (SNAP)

ARRIVAL 1. TORQUE—30% - 40% 2. 150 - 175 KIAS (TYPICAL) 3. FD—AS DESIRED 4. START BEFORE LANDING CHECKLIST

DH

APPROACH INBOUND 1. FLAPS—APPROACH 2. AIRSPEED—130 - 140 KIAS 3. RESET ALTITUDE PRESELECT TO M.A.P. ALTITUDE

APPROACHING GLIDE SLOPE 1. GEAR—DOWN 2. PROPS—FULL FORWARD 3. COMPLETE BEFORE LANDING CHECKLIST

DH-VISUAL AND LANDING ASSURED


1. FLAPS—DOWN 2. TRANSITION TO VREF 3. YAW DAMP—OFF 4. PROPS—CONFIRM FULL FORWARD

THRESHOLD LANDING 1. GROUND FINE OR REVERSE 2. BRAKES—AS NECESSARY

1. GEAR—RECHECK DOWN 2. AIRSPEED—VREF 3. POWER—IDLE

CAUTION

CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICS, THE PROPELLER CONTROL MUST BE IN FULL INCREASE RPM POSITION. NOTE: REVERSE IS MOST EFFECTIVE AT HIGHER SPEEDS; BRAKING IS MOST EFFECTIVE AT LOWER SPEEDS


Figure 19-11. ILS Approach


19-17


PROCEDURE TURN OUTBOUND

INITIAL 1. OBTAIN ATIS 2. BRIEF APPROACH AND MISSED APPROACH 3. FMS/NAV AIDS—SETUP/IDENT 4. DESCENT CHECKLIST— COMPLETE

1. START TIMING 2. FLAPS—APPROACH 3. AIRSPEED—130 - 140 KIAS FAF

PROCEDURE TURN INBOUND 1. FD—AS DESIRED 2. RESET ALTITUDE PRESELECT

ARRIVAL 1. TORQUE—30% - 40% 2. AIRSPEED— 150 - 175 KIAS (TYPICAL) 3. FD—AS DESIRED 4. START BEFORE LANDING CHECKLIST

MAP-MISSED APPROACH

STATION PASSAGE 1. START TIMING 2. SET ALTITUDE PRESELECT

1. POWER—MAX (100% / 820˚) 2. PITCH—7˚ NOSE UP (FD-GA) 3. GEAR—UP 4. FLAPS—UP 5. COMPLETE MISSED-APPROACH PROCEDURE (SNAP)

INTERCEPT FINAL APPROACH 1. COURSE INBOUND MAP

APPROACH INBOUND 1. RESET ALTITUDE PRESELECT TO APPROACH MINIMUMS 2. GEAR—DOWN 3. PROPS—FULL FORWARD

FAF

MDA

FINAL APPROACH FIX


1. START TIMING 2. GEAR—CONFIRM DOWN 3. TORQUE—APPROX. 15% 4. COMPLETE BEFORE LANDING CHECKLIST 5. AIRSPEED—130 - 140 KIAS


MINIMUM DESCENT ALTITUDE (MDA)


MAP-LANDING ASSURED

1. LEVEL OFF AT MDA, AT LEAST 1 MILE PRIOR TO MAP, IF POSSIBLE 2. TORQUE—40% 3. AIRSPEED—130 - 140 KIAS (VYSE MIN)

1. FLAPS—DOWN 2. TRANSITION TO VREF 3. YAW DAMP—OFF 4. PROPS—CONFIRM FULL FORWARD

CAUTION

CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICS, THE PROPELLER CONTROL MUST BE IN FULL INCREASE RPM POSITION. NOTE: REVERSE IS MOST EFFECTIVE AT HIGHER SPEEDS; BRAKING IS MOST EFFECTIVE AT LOWER SPEEDS


Figure 19-12. Nonprecision Approach

19-18



NOTE:

ARRIVAL 1. PLAN CIRCLING MANEUVER 2. FOLLOW NORMAL APPROACH PROCEDURES TO MDA

THIS IS A CATEGORY B AIRCRAFT, BUT AIRSPEEDS OF 121 THROUGH 140 KIAS REQUIRE CATEGORY C MINIMUMS

MDA MAP FINAL


MINIMUM DESCENT ALTITUDE (MDA) 1. LEVEL OFF AT MDA AT LEAST 1 MILE PRIOR TO MAP, IF POSSIBLE 2. TORQUE—40% 3. 130 - 140 KIAS (VYSE MIN) 4. MANEUVER WITHIN VISIBILITY CRITERIA 5. MAINTAIN MDA

1 NM

1. 130 - 140 KIAS (VYSE MIN) WHEN LANDING ASSURED: 2. FLAPS—DOWN 3. TRANSITION TO VREF 4. YAW DAMP—OFF 5. PROPS—FULL FORWARD

MAP AND DURING CIRCLING MANEUVER

BASE 1. COMMENCE DESCENT FROM A POINT WHERE A NORMAL LANDING CAN BE MADE

CAUTION TO ENSURE CONSTANT REVERSING CHARACTERISTICS, THE PROPELLER CONTROL MUST BE IN FULL INCREASE RPM POSITION. NOTE: REVERSE IS MOST EFFECTIVE AT HIGHER SPEEDS; BRAKING IS MOST EFFECTIVE AT LOWER SPEEDS.

CAUTION IF POSSIBLE, PROPELLERS SHOULD BE MOVED OUT OF REVERSE AT APPROXIMATELY 40 KNOTS TO MINIMIZE BLADE EROSION. CARE MUST BE EXERCISED WHEN REVERSING ON RUNWAYS WITH LOOSE SAND, DUST, OR SNOW ON THE SURFACE. FLYING GRAVEL WILL DAMAGE PROPELLER BLADES, AND DUST OR SNOW MAY IMPAIR THE PILOT'S VISIBILITY.

Figure 19-13. Circling


19-19


1. DETERMINE THAT VISUAL CONTACT WITH THE RUNWAY CAN BE MAINTAINED AND A NORMAL LANDING CAN BE MADE FROM A CIRCLING APPROACH, OR INITIATE A MISSED APPROACH 2. MAINTAIN MDA DURING CIRCLING MANEUVER


CHAPTER 20 WEIGHT AND BALANCE CONTENTS Page INTRODUCTION............................................................................................................. 20-1 GENERAL ......................................................................................................................... 20-1 Weighing....................................................................................................................... 20-2 LOADING .......................................................................................................................... 20-2 Cargo Loading............................................................................................................. 20-4 COMPUTING .................................................................................................................... 20-5

20 WEIGHT AND BALANCE

Procedure ..................................................................................................................... 20-5


20-i



Title

Page

Dimensional and Loading Data........................................................................ 20-2

20-2

Passenger Seating Configurations Payload Locations.................................... 20-3

20-3

Loading Data (Cargo Configuration) .............................................................. 20-3

20-4

Useful Load Weights and Moments Cargo...................................................... 20-4

20-5

Weight and Balance Loading Form .................................................................. 20-6

20-6

Cabnet Contents and Baggage.......................................................................... 20-7

20-7

Useful Load Weights and Moments—Useable Fuel ...................................... 20-8

20-8

Moment Limits Vs. Weight ................................................................................ 20-9

20-9

Center of Gravity.............................................................................................. 20-10


20-1


20-iii


CHAPTER 20 WEIGHT AND BALANCE

INTRODUCTION Maintaining center of gravity (CG) within the approved envelope throughout the planned flight is an important safety consideration. The aircraft must be loaded so it does not exceed the weight and CG limitations. This chapter presents an overview on how this is accomplished.

GENERAL

If the aircraft is loaded above maximum takeoff weight, takeoff distance and landing distance are longer. Stalling speeds are

higher. Rate of climb, cruising speed, and range are lower. If an aircraft is loaded so that the CG is forward of the forward limit, additional control movements for maneuvering the aircraft are required as well as higher control forces. The crew may experience d i ff i c u l t y d u r i n g t a ke o ff a n d l a n d i n g because of the elevator control limits.


20-1


Aircraft loaded above the maximum takeoff or landing weight limitations have an overall lower performance level. Refer to the Performance section of the Pilot's Operating Handbook and FAA-approved Airplane Flight Manual.


If an aircraft is loaded aft of the aft CG, the crew experiences a lower level of stability. These are characterized by lower control forces, difficulty in trimming, lower control forces for maneuvering with the danger of structural overload, decayed stall charact e r i s t i c s, a n d l o w e r l e v e l o f l a t e r a l directional damping.

LOADING

WEIGHING

It is the responsibility of the operator to ensure that the aircraft is properly loaded. At the time of delivery, Raytheon Aircraft Company provides the necessary weight and balance to compute individual loadings. All subsequent changes in weight and balance are the responsibility of the owner and/or operator.

Periodic weighing of the aircraft may be required to keep the basic empty weight current. The operator determin e s t h e frequency of weighing.

Different charts are available for the 350, 350C, 350ER, and 350CER. Samples of these for the 350 are depicted in Figures 201 and 20-2. Figure 20-3 is for cargo aircraft.


Figure 20-1. Dimensional and Loading Data

20-2




Figure 20-2. Passenger Seating Configurations Payload Locations

Figure 20-3. Loading Data (Cargo Configuration)


20-3


CARGO LOADING Determine the method of loading cargo, its placement in the aircraft, and method of restraint before starting the actual loading. For loads that are evenly distributed in a given section, use the Useful Load Weights and Moments—Cargo tables (Figure 204). These are available in both pounds and kilograms.

For any load that cannot be located at the centroid of a section or that extends over more than one section, determine the CG and its location in the aircraft. • Determine CG arm (fuselage station) by measuring in inches from a known location in the cabin to CG of the load • Next, multiply weight by CG arm to determine moment for the load • Divide the moment by 10 0 to be compatible with other loading data


Figure 20-4. Useful Load Weights and Moments Cargo

20-4



COMPUTING Th e f o l l o w i n g t a b l e s a n d c h a r t s a r e necessary for loading and computing weight and balance: • Basic Empty Weight and Balance form—Contains the basic empty weight and moment of the aircraft at the time of delivery • Useful Load Weights and Moments tables—Contains load items that may be loaded into the aircraft; these include tables for occupants, cabinet contents and baggage, cargo, and fuel • Moment Limits vs. Weight graph or table—Displays the minimum and maximum moments approved by the FAA; these moments correspond to the forward and aft CG flight limits (landing gear down) for a particular weight

4. Enter fuel weight and moment/100 from t h e U s e f u l L o a d We i g h t s a n d Moments—Usable Fuel table (Figure 20-7). 5. Subtotal weight column and moment column. Check to ensure weight and moment are within ramp weight limits (Figure 20-8 and 20-9). 6. Subtract start, taxi, and takeoff fuel weight and moment/10 0, which is normally 10 0 pounds at an average moment/10 0 of 277 pound per inches, to determine takeoff weight and moment. Check to ensure weight and moment does not exceed maximum takeoff limits. 7. C o m p l e t e l i n e s 2 3 t h r o u g h 2 7 t o determine landing weight.

Divide all moments by 10 0 to simplify computations. Ensure that all cargo and baggage is properly secured before takeoff. A sudden shift in balance at rotation can cause controllability problems.

PROCEDURE 1. Record basic empty weight and moment f r o m t h e B a s i c E m p t y We i g h t a n d Balance form (Figure 20-5). Divide the moment by 10 0 to correspond to the U s e f u l L o a d We i g h t a n d M o m e n t s tables. 2. Re c o r d w e i g h t a n d c o r r e s p o n d i n g moment/100 of each item. These values are in the Useful Load Weights and Moments tables (Figure 20-6).



3. Subtotal weight column and moment column. The weight without usable fuel must not exceed the maximum zero fuel weight limitation of 12,500 pounds. All weight in excess of this limitation must be fuel. The auxiliary tanks may be used only when main tanks are completely filled.

20-5


WEIGHT AND BALANCE LOADING FORM SERIAL NO.________________REG NO.________________DATE_________________ LINE

ITEM

1

Basic Empty Weight

2

Pilot

3

Copilot

4

Passenger 1 or Cargo in Section A

5

Passenger 2 or Cargo in Section B

6

Passenger 3 or Cargo in Section C

7

Passenger 4 or Cargo in Section D

8

Passenger 5 or Cargo in Section E

9

Passenger 6

10

Passenger 7

11

Passenger 8

12

Passenger 9

13

Passenger 10

14

Total Cabinet Contents

15

Baggage

16

Baggage

17

Baggage

18

Subtotal - Zero Fuel Weight. DO NOT EXCEED 12,500 LBS. (5670 KG)

19

Fuel Loading

20

Subtotal - Ramp Weight DO NOT EXCEED 15,100 LBS. (6349 KG)

21

Less Fuel for Start, Taxi and Take-off**

22

Total - Take-Off Weight. DO NOT EXCEED 15,000 LBS. (6804 KG)

WEIGHT *( )

F.S. (IN)

MOM/100 *( )

* Enter units used. Lbs and Lb-In. or Kg and Kg-In. ** Fuel for start, taxi and take-off is normally 100 lbs (45 kg) at an average moment/100 of 227 LbIn. (103 Kg-In.). LANDING WEIGHT DETERMINATION


23

Fuel Loading from Line 19

24

Less Fuel used to Destination (including fuel for start, taxi and take-off)

25

Total Fuel Remaining. Moment/100 from Usable Fuel Weights and Moments Table

26

Zero Fuel Weight from Line 18

27

Total Landing Weight (line 25 + 26)

NOTE: Shaded areas in the above tables indicate values that are not required to arrive at a final weight and balance.

Figure 20-5. Weight and Balance Loading Form

20-6




Figure 20-6. Cabinet Contents and Baggage


20-7


20-8

USEFUL LOAD WEIGHTS AND MOMENTS - USABLE FUEL (GALLONS) 6.3 LB/GAL

GAL

WEIGHT

MOMENT/100

WEIGHT

63 126 189 252 315 378 441 504 567 630 693 756 819 882 945 1008 1071 1134 1197 1260 1323 1386 1449 1512 1575 1638 1701 1764 1827 1890 1953 2016 2079 2142 2205 2268 2331 2394 2457 2520 2583 2646

115 230 347 464 581 699 825 952 1079 1208 1335 1463 1591 1721 1849 1973 2098 2222 2346 2472 2606 2732 2860 2985 3107 3232 3357 3482 3612 3740 3869 3995 4124 4251 4382 4512 4643 4779 4912 5048 5187 5324

65 130 195 260 325 390 455 520 585 650 715 780 845 910 975 1040 1105 1170 1235 1300 1365 1430 1495 1560 1625 1690 1755 1820 1885 1950 2015 2080 2145 2210 2275 2340 2405 2470 2535 2600 2665 2730

6.7 LB/GAL

MOMENT/100

118 238 358 478 600 721 851 982 1114 1247 1378 1510 1641 1775 1907 2035 2164 2292 2421 2551 2688 2818 2951 3079 3205 3334 3463 3592 3726 3858 3992 4122 4255 4386 4521 4655 4791 4930 5068 5208 5351 5493

6.9 LB/GAL

WEIGHT

MOMENT/100

WEIGHT

67 134 201 268 335 402 469 536 603 670 737 804 871 938 1005 1072 1139 1206 1273 1340 1407 1474 1541 1608 1675 1742 1809 1876 1943 2010 2077 2144 2211 2278 2345 2412 2479 2546 2613 2680 2747 2814

122 245 369 493 618 743 877 1012 1148 1285 1420 1556 1692 1830 1966 2098 2231 2363 2495 2629 2771 2905 3042 3174 3304 3437 3570 3703 3841 3977 4115 4249 4386 4521 4660 4798 4938 5082 5224 5368 5516 5662

69 138 207 276 345 414 483 552 621 690 759 828 897 966 1035 1104 1173 1242 1311 1380 1449 1518 1587 1656 1725 1794 1863 1932 2001 2070 2139 2208 2277 2346 2415 2484 2553 2622 2691 2760 2829 2898

Figure 20-7. Useful Load Weights and Moments—Useable Fuel

MOMENT/100

126 252 380 508 636 765 903 1042 1182 1323 1462 1602 1743 1885 2025 2161 2298 2434 2569 2707 2854 2992 3133 3269 3403 3540 3677 3814 3956 4096 4238 4376 4517 4656 4799 4941 5085 5234 5380 5528 5681 5831



10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 170 180 190 200 210 220 230 240 250 260 270 280 290 300 310 320 330 340 350 360 370 380 390 400 410 420

6.5 LB/GAL



Figure 20-8. Moment Limits Vs. Weight


20-9



Figure 20-9. Center of Gravity

20-10


21 FLIGHT PLANNING AND PERFORMANCE


CHAPTER 21 FLIGHT PLANNING AND PERFORMANCE CONTENTS Page INTRODUCTION............................................................................................................. 21-1 GENERAL FLIGHT PLANNING................................................................................. 21-1 PERFORMANCE ............................................................................................................. 21-2 Limitations ................................................................................................................... 21-2 Factors Affecting Performance.................................................................................. 21-2 Using Graphs............................................................................................................... 21-3 Calculations ................................................................................................................. 21-8


21-i


Title

Page

21-1

Airspeed Calibration .......................................................................................... 21-4

21-2

Maximum Takeoff Weight to Achieve Takeoff Climb Requirements .......... 21-6

21-3

Maximum Landing Weight ................................................................................ 21-7

21-4

Takeoff Path Profile ............................................................................................ 21-8

21-5

ISA Conversions ............................................................................................... 21-10

21-6

Fahrenheit–Celsius Temperature Conversion ............................................... 21-11

21-7

Feet–Meters Conversion.................................................................................. 21-12

21-8

U.S. Gallons–Liters Conversion ...................................................................... 21-13

21-9

Pounds–Kilograms Conversion ....................................................................... 21-14

21-10

Inches–Millimeters Conversion ...................................................................... 21-15


21-iii



CHAPTER 21 FLIGHT PLANNING AND PERFORMANCE

INTRODUCTION The pilot must be completely familiar with the performance of the aircraft and performance data in the Pilot's Operating Handbook (POH) and the FAA-approved Airplane Flight Manual ( AFM ). Aircraft performance depends on the effects of temperature and pressure altitude. The AFM must be aboard the aircraft at all times.

GENERAL FLIGHT PLANNING The pilot-in-command should be familiar with all available information concerning a flight. Obtain a current and complete preflight briefing. This should consist of the following: • L o c a l , e n r o u t e, a n d d e s t i n a t i o n weather

• Enroute terrain and obstructions • Alternate airports • Airport runways active • Length of runways • Takeoff and landing distances for conditions expected

• Enroute navaid information


21-1





The pilot reviews his planned enroute track and stations and makes a list for quick reference. A flight plan filed with Flight Service Stations is also recommended even if it is VFR. Also, advise Flight Service Stations of changes or delays of one hour or more.

FACTORS AFFECTING PERFORMANCE Tables in the AFM detail variables affecting performance. Assumptions that relate to all performance calculations, unless otherwise stated, are the following:

Re m e m b e r t o c l o s e t h e f l i g h t p l a n at destination.

• Cabin pressurized

PERFORMANCE

• Winds taken as tower winds (30 feet above runway surface); factors applied as prescribed in applicable regulations. In the tables, negative re p re s e n t s t a i l w i n d a n d p o s i t i v e represents headwind.

LIMITATIONS Th e m a x i m u m o p e r a t i n g w e i g h t s a r e limited by the manufacturer’s performance criteria. Compliance is mandatory. Fo r Pa r t 9 1 a n d Pa r t 1 3 5 o p e r a t i o n s, these include: • Maximum takeoff weight to achieve takeoff climb requirements • Maximum takeoff weight as limited by tire speed • Takeoff field length • Maximum landing weight to achieve climb requirements • Normal landing distance-flaps down Additional limitations for 14 Part 135 operations include: • Service ceiling for one engine inoperative • Takeoff flight path requirements to 1,500 ft AGL

• Humidity corrections to power according to applicable regulations

Icing Flight Degradations in performance are also determined for selected conditions. The performance information for icing flight is for reference only; it was obtained in controlled conditions with simulated ice shapes attached to the aircraft. These accumulations, referred to as "normal ice accumulations," include the following: • Leading edges of boots (simulated ice that may accumulate by the time the pilot activates the boots (i.e., one inch) • Fo l l o w i n g u n p r o t e c t e d s u r f a c e s (simulate ice that may accumulate during a 45-minute holding condition):

° Nose radome ° Wing center sections not protected by boots

° Wing outer panels not protected by boots

° Vertical stabilizer and bullet ° Horizontal stabilizer not protected by boots

21-2


Actual performance degradations may be more or less than the values quoted in the graphs, depending on type and duration of icing encounter. Icing notes are on the following graphs and tables: • Stall Speeds - Power Idle • Maximum Cruise Power table • Normal Cruise Power table • Maximum Range Power table • One-Engine-Inoperative Maximum Cruise Power table • Maximum Landing Weight

USING GRAPHS When using the graphs in the manufact u r e r ’ s m a n u a l s, k e e p t h e f o l l o w i n g information in mind. All power settings and performance are predicated on OAT from the pilot side console display. Do not use the temperature displayed on the pilot PFD and MFD. In addition to presenting the result for a particular set of conditions, the example on a graph also presents the order in which the various scales on that graph should be used. Fo r i n s t a n c e, i f t h e f i r s t i t e m i n t h e example is OAT, then enter the graph at the existing OAT.

• Approach Climb Gradient • Climb - Balked Landing • Landing Distance-Flaps Down The notes approximate performance with engine anti-ice on power. The effect varies depending on airspeed and ambient c o n d i t i o n s. A t l o w e r a l t i t u d e s w h e re operation at or near the torque limit is possible, the effect of using engine antiice is less (depending on how much power can be recovered after engine anti-ice has been turned on). If the power set before engine anti-ice activation during flight can b e a c h i e v e d a f t e r a c t i va t i o n w i t h o u t exceeding engine limitations, anti-ice effects are negated. For all takeoff charts, during operation requiring engine anti-ice on, the results read from the graphs remain the same if the power per the Static Takeoff Power-1,70 0 RPM With Engine Anti-Ice Off graph can be set without exceeding engine limitations. If this value cannot be achieved, then power set per the Static Takeoff Power1,700 RPM with Engine Anti-Ice On graph requires the results read from the takeoff graph be altered by the amount specified.

Reference lines indicate where to begin following the guidelines. Always project to the reference line first, and then follow the guidelines to the next item. Maintain the same proportional distance between the guideline above and the guideline below the projected line. For instance, if the projected line intersects the reference line in the ratio of 30% down/70% up between the guidelines, then maintain this same 30%/70% relationship between the guidelines all the way to the next item. Th e a s s o c i a t e d c o n d i t i o n s d e f i n e t h e specified conditions from which performance parameters have been determined. They are not intended as instructions. Pe r f o r m a n c e va l u e s d e t e r m i n e d f ro m charts can only be achieved if the specified conditions exist. The full amount of usable fuel is available for all approved flight conditions. All airspeeds are indicated airspeeds (IAS) u n l e s s o t h e r w i s e n o t e d . A s s u m e z e ro instrument error. With the exception of stall speeds presented, all were derived from calibrated airspeeds corrected per either the Airspeed Calibration-Normal System-Takeoff Ground Roll graph or the A i r s p e e d C a l i b ra t i o n - N o r m a l Sy s t e m graph (Figure 21-1).


21-3





Figure 21-1. Airspeed Calibration

21-4


Maximum Takeoff Weight Graph

Maximum Landing Weight Graph

The Maximum Takeoff Weight to Achieve Ta k e o f f C l i m b Re q u i r e m e n t s g r a p h presents the most restrictive maximum takeoff weight (Figure 21-2):

Th e M a x i m u m La n d i n g We i g h t g ra p h presents the most restrictive maximum landing weight (Figure 21-3):

• Aircraft in takeoff configuration with m o s t c r i t i c a l C G, c r i t i c a l e n g i n e inoperative and propeller feathered, and remaining engine at maximum takeoff power

° With landing gear extended results

in a steady gradient of climb between liftoff and point where landing gear is retracted that is measurably positive

° With landing gear retracted results

• With aircraft in the discontinued approach configuration (i.e., gear up and flaps approach) at normal approach speed, critical engine inoperative, and the other engine at available takeoff power, results in a steady gradient of climb of 2.1% • With landing gear extended, flaps extended, and both engines operating at takeoff power, results in a steady g ra d i e n t o f c l i m b o f 3 . 3 % a t the most critical CG and normal approach speed

in a steady gradient of climb of 2%

• Aircraft in enroute configuration at an altitude 1,500 feet above takeoff surface with critical engine inoperative, remaining engine at maximum continuous power, and most critical CG results in a steady gradient of climb of 1.2%


21-5





Figure 21-2. Maximum Takeoff Weight to Achieve Takeoff Climb Requirements

21-6




Figure 21-3. Maximum Landing Weight


21-7



To determine the approximate pressure altitude at departure and destination airports:

CALCULATIONS Takeoff Figure 21-4 is a takeoff path profile to aid the pilot in the various calculations that need to be accomplished. Conversion charts for ISA conversion, feet to meters, pounds to kilograms, etc., are shown in Figures 21-5 through 21-10 at the end of this chapter. The following steps summarize how to calculate the takeoff performance. 1. Determine the following at departure: OAT, field elevation, altimeter setting, wind, runway 35L length, and gradient 2. Th e n d e t e r m i n e t h e f o l l o w i n g a t d e s t i n a t i o n : OAT, f i e l d e l e va t i o n , a l t i m e t e r s e tt i n g, w i n d , r u n w ay 2 5 length, and gradient

• Add 1,0 0 0 ft to field elevation for each 1.0 0 in Hg that the reported altimeter setting value is below 29.92 in Hg or subtract 1000 ft for each 1.00 in Hg above 29.92 in Hg • Find the difference between 29.92 in Hg and the reported altimeter setting • Multiply the answer by 10 0 0 to find the difference in feet between field elevation and pressure altitude 3. Next determine the maximum takeoff weight with graphs such as Maximum Takeoff Weight to Achieve Takeoff C l i m b R e q u i r e m e n t s — F l a p s U p, Maximum Takeoff Weight as Limited by Tire Speed-Flaps Up; 135 Operators also need to determine maximum enroute weight at minimum enroute altitude.

TAKE - OFF PATH PROFILE

Figure 21-4. Takeoff Path Profile

21-8


4. Then determine the minimum field length for takeoff with Takeoff Field Length—Flaps Up and Takeoff SpeedsFlaps Up graphs 5. Finally, determine takeoff path with one engine inoperative

Enroute Graphs With appropriate graphs, the pilot can then determine the following: • Time, fuel, and distance to cruise climb • Time, fuel, and distance to descend

One-Engine Inoperative Computations

• Cruise true airspeed

Graphs estimate the horizontal distance required to reach a height of 1,500 feet, or the minimum climb gradient required to clear an obstacle along the takeoff flight path. If clearance of obstacles beyond the runway is required, these may restrict takeoff weight accordingly.

• Cruise fuel flow

The takeoff distance extends from brake release to reference zero, which is the horizontal point along the runway at which the aircraft is 35 feet above the runway. Th e n e t t a ke o ff f l i g h t p a t h b e g i n s a t reference zero and consists of the following segments: • First segment climb extends from reference zero to the point where the landing gear completes the retraction cycle; airspeed is maintained at V 2 • Second segment climb begins at the end of the first segment and extends to 40 0 feet above the runway; airspeed during the second segment is V 2

• Cruise power settings • Reserve fuel • Total fuel requirements • Zero-fuel weight limitation

Landing Weight To determine the landing weight, subtract fuel required for the trip from the ramp weight. Then: 1. With Maximum Landing Weight graph, determine maximum landing weight 2. With Approach Climb Gradient graph, determine approach climb gradient and climb speed 3. With Normal Landing Distance—Flaps Down graph, read landing distance and approach speed 4. With Climb-Balked Landing graph, read rate of climb, climb gradient, and climb speed

• The horizontal acceleration and flap retraction segment consists of an acceleration from V 2 to V YSE at a constant height of 40 0 feet; flap retraction is completed during this segment • Third segment begins when oneengine-inoperative climb speed is reached at 40 0 feet and extends to 1,500 feet above the runway; airspeed i s m a i n t a i n e d a t V YS E d u r i n g this segment


21-9





Figure 21-5. ISA Conversions

21-10




Figure 21-6. Fahrenheit–Celsius Temperature Conversion


21-11



Figure 21-7. Feet–Meters Conversion

21-12




Figure 21-8. U.S. Gallons–Liters Conversion


21-13



Figure 21-9. Pounds–Kilograms Conversion

21-14




Figure 21-10. Inches–Millimeters Conversion


21-15


CHAPTER 22 CREW RESOURCE MANAGEMENT

Page CREW CONCEPT BRIEFING GUIDE ....................................................................... 22-3 Introduction ................................................................................................................. 22-3 Common Terms ........................................................................................................... 22-3 Pretakeoff Briefing (IFR/VFR)................................................................................. 22-4 Crew Coordination During the Approach Sequence ............................................. 22-4 ALTITUDE CALLOUTS................................................................................................. 22-5 Enroute......................................................................................................................... 22-5 Approach—Precision.................................................................................................. 22-5 Approach—Nonprecision .......................................................................................... 22-6 Significant Deviation Callouts ................................................................................... 22-7


22-i

22 CREW RESOURCE MANAGEMENT

CONTENTS


ILLUSTRATIONS Title

Page

22-1

Situational Awareness in the Cockpit......................................................................... 22-1

22-2

Command and Leadership ................................................................................. 22-1

22-3

Communication Process ..................................................................................... 22-2

22-4

Decision Making Process.................................................................................... 22-2


22-iii


Figure


CAPTAIN INDIVIDUAL S/A


CHAPTER 22 CREW RESOURCE MANAGEMENT REMEMBER

COPILOT INDIVIDUAL S/A

2+2=2 – or –

GROUP S/A

2+2=5 (Synergy)

IT'S UP TO YOU!

CLUES TO IDENTIFYING:

HUMAN

OPERATIONAL

• Loss of Situational Awareness • Links in the Error Chain 1. FAILURE TO MEET TARGETS 2. UNDOCUMENTED PROCEDURE 3. DEPARTURE FROM SOP 4. VIOLATING MINIMUMS OR LIMITATIONS 5. NO ONE FLYING AIRPLANE 6. NO ONE LOOKING OUT WINDOW 7. COMMUNICATIONS 8. AMBIGUITY 9. UNRESOLVED DISCREPANCIES 10. PREOCCUPATION OR DISTRACTION 11. CONFUSION OR EMPTY FEELING 12.

Figure 22-1. Situational Awareness in the Cockpit

LEADERSHIP STYLES AUTOCRATIC AUTHORITARIAN DEMOCRATIC STYLE LEADERSHIP LEADERSHIP (EXTREME) STYLE STYLE

LAISSEZFAIRE STYLE (EXTREME)

PARTICIPATION LOW

HIGH

Command — Designated by Organization — Cannot be Shared Leadership — Shared among Crewmembers — Focuses on "What's right," not "Who's right"

Figure 22-2. Command and Leadership


22-1


INTERNAL

EXTERNAL

INTERNAL

BARRIERS

BARRIERS

BARRIERS

NEED

SEND

RECEIVE

OPERATIONAL GOAL


THINK: • Solicit and give feedback • Listen carefully • Focus on behavior, not people • Maintain focus on the goal • Verify operational outcome is achieved

FEEDBACK

ADVOCACY: to increase others' S/A

INQUIRY: to increase your own S/A

* State Position

* Decide What, Whom, How to Ask

* Suggest Solutions

* Ask Clear, Concise Questions

* Be Persistent and Focused

* Relate Concerns Accurately

* Listen Carefully

* Draw Conclusions from Valid Information * Keep an Open Mind

— REMEMBER — Questions enhance communication flow Don't give in to the temptation to ask questions when Advocacy is required Use of Advocacy or Inquiry should raise a "red flag".

Figure 22-3. Communication Process

EVALUATE RESULT

RECOGNIZE NEED

IDENTIFY AND DEFINE PROBLEM

IMPLEMENT RESPONSE

COLLECT FACTS SELECT A RESPONSE

IDENTIFY ALTERNATIVES WEIGH IMPACT OF ALTERNATIVES

HINTS: • Identify the problem: — Communicate it — Achieve agreement — Obtain commitment • Consider appropriate SOPs • Think beyond the obvious alternatives • Make decisions as a result of the process • Resist the temptation to make an immediate decision and then support it with facts

Figure 22-4. Decision Making Process

22-2



COMMON TERMS PIC

Designated by the company for flights requiring more than one pilot. Responsible for conduct and safety of the flight. Designates pilot flying and pilot not flying duties.

INTRODUCTION To a large extent the success of any aircrew depends on how effectively crewmembers coordinate their actions using standardized and approved procedures. In other chapters you have been exposed to standardized maneuvers, procedures, and checklists. This chapter illustrates standard aircrew calls and briefing guidelines. When used in a logical sequence with aircrew checklists and flight procedures, these callouts can improve aircrew efficiency and enhance safety. These callouts and briefings are only recommendations to be used in a larger system of standard operating procedures that become the core of an effective crew resource management program. They are not intended to supersede any individual company SOP, but are examples of good operating practices.

Pilot in Command

PF

Pilot Flying Controls the aircraft with respect to assigned airway, course, altitude, airspeed, etc., during normal and emergency conditions. Accomplishes other tasks as directed by the PIC.

PM

Pilot Monitoring Maintains ATC communications, copies clearances, accomplishes checklists, and other tasks as directed by the PIC.

B

Both


22-3


CREW CONCEPT BRIEFING GUIDE


PRETAKEOFF BRIEFING (IFR/VFR)

PF

Requests the pilot monitoring to obtain destination weather. (Transfer of communication duties to the pilot flying may facilitate this task.)

NOTE


The following briefing is to be completed during item 1 of the Pretakeoff checklist. The pilot flying will accomplish the briefing. 1. Review the ATC clearance and departure procedure (route and altitude, type of takeoff, significant terrain features, etc.). 2. Review those items that are not standard procedure to include deferred or MEL items (if applicable).

PM

PF

Advises the pilot of current destination weather, approach in use, and special information pertinent to the destination. Requests the pilot monitoring to perform the approach setup.

PM

PF

Ac c o m p l i s h e s t h e a p p r o a c h setup and advises of frequency tuned, identified and course set. Transfers control of the aircraft to the pilot monitoring, advising, “You have control, heading , altitude ” and special instructions. (Communications duties should be transferred back to the pilot monitoring at this point.)

3. Review required callouts, unless standard calls have been agreed upon, in which case a request for “Standard Callouts” may be used. 4. Review the procedures to be used in case of an emergency on departure.

PM

5. As a final item, ask if there are any questions.

PF

The pilot who will fly the approach will review, then brief the approach procedure.

PF

Advises, “I have control, heading , altitude .”

CREW COORDINATION DURING THE APPROACH SEQUENCE

PM

Responds, “I have control, heading , altitude .”

Confirms “You have control, heading , altitude .”

NOTE The following crew coordination approach sequence should be completed as early as possible, prior to initiating an IFR approach.

22-4

NOTE The above sequence should be completed prior to the FAF.



ALTITUDE CALLOUTS ENROUTE PM

PF

State altitude leaving and assigned level off altitude

“CHECKED”

“200 above/below”

“LEVELING”


1,000 ft prior to level off

APPROACH—PRECISION PM

PF At 1,0 0 0 ft above minimums

“1,000 feet above”

“DH _________”

At 50 0 ft above minimums “500 feet above minimums”

“NO FLAGS”

At 10 0 ft above minimums “100 feet above” At decision height (DH) “Decision Height, approach lights at (clock position)”

“CONTINUING” OR “LANDING”

OR “Decision Height, runway at (clock position)” OR

“CONTINUING OR “LANDING”

“Decision Height, runway not in sight”


“MISSED APPROACH”

22-5


APPROACH—NONPRECISION PM

PF At 1,0 0 0 ft above MDA

“1,000 feet above”

“MDA _________”


At 50 0 ft above MDA “500 feet above.”

“NO FLAGS” At 10 0 ft above MDA

“100 feet above.” At minimum descent altitude (MDA) “MDA”

“MAINTAINING MDA”

At or prior to the missed approach point (MAP) “Approach lights at (clock position)”


“Runway at (clock position)”


“Runway not in sight”

“MISSED APPROACH”

22-6



SIGNIFICANT DEVIATION CALLOUTS PM

PF IAS ±10 KIAS “CORRECTING TO _________” 22 CREW RESOURCE MANAGEMENT

“VREF ± __________”

Heading ±10° enroute, 5° on approach “Heading __________ degrees left/right”

“CORRECTING TO _________”

Altitude ±100 ft enroute, +50/-0 ft on final approach “Altitude __________ high/low”

“CORRECTING TO _________”

CDI left or right one dot “Left/right of course__________ dot”

“CORRECTING”

RMI course left or right ±5° “Left/right of course__________ degrees”

“CORRECTING”

Vertical descent speed greater than 1,0 0 0 fpm on final approach “Sink rate__________”

“CORRECTING” Bank in excess of 30°

“Bank__________ degrees”

“CORRECTING”


22-7



WALKAROUND

NOTES

The following section is a pictorial walkaround. Each item listed in the exterior power-off preflight inspection is displayed. The foldout pages contain photographs that depict the specific area to be inspected. The general photographs contain circled numbers that correspond to specific steps displayed on the subsequent pages.



WA-1

106

107 108



LEFT WING AND NACELLE


72 73 74 75

105

102 103 104

93 94 95 96

101

115

90 91 92

76 88 89 97 98

83 84 85 86

79 80 82 87

71 77 78

81

99 63 100 64 65 66 67 68 69 70

54

62 61

50 51

60

55

57 58 59 1.

WA-2

FLAPS (CONDITION, ASYMMETRY PROTECTION, AND FLAP TRACKS)—CHECK

3B.

FLAPS (CONDITION, ASYMMETRY PROTECTION, AND FLAP TRACKS)—CHECK

CABIN DOOR SEAL, STEP EXTENSION CABLE, LIGHT WIRE, DAMPER, AND HANDRAILS—CHECK

111

2.

120

3A.

113 114 119

117 112 118 130 131

110


LEFT SIDE WINDOWS—CHECK

109




AILERON AND TAB—CHECK

5.

STATIC WICKS (AILERON AND WINGLET)—CHECK

6.

SIPHON BREAK VENT—CLEAR

LIGHTS—CHECK

8.

MAIN FUEL TANK CAP—SECURE

9.

STALL WARNING VANE—CHECK

WALKAROUND

4.

7.


WA-3



10.

TIE-DOWN—REMOVE

WALKAROUND 11.

12.

OUTBOARD DEICE BOOT AND STALL STRIP— CHECK

13.

WING PANELS—SECURE

FLUSH OUTBOARD FUEL DRAIN—CHECK

WA-4




17A. LANDING GEAR (DOORS, TIRES, STRUT, WHEEL WELL)—CHECK (CONTINUED)

LEADING-EDGE FUEL TANK AND GRAVITY-LINE DRAINS—DRAIN

WALKAROUND

14.

17B. LANDING GEAR (DOORS, TIRES, STRUT, WHEEL WELL)—CHECK (CONTINUED)

15.

INVERTER COOLING LOUVERS—CLEAR

17C. LANDING GEAR (DOORS, TIRES, STRUT, WHEEL WELL)—CHECK (CONTINUED)

16.

FLUSH FUEL VENT AND HEATED FUEL VENT— CLEAR


WA-5



WALKAROUND

17D. LANDING GEAR (DOORS, TIRES, STRUT, WHEEL WELL)—CHECK

18.

19.

20.

TORQUE KNEE ASSEMBLY AND SAFETY SWITCH— CHECK

21.

BRAKES—CHECK

22

FIRE EXTINGUISHER PRESSURE—CHECK

ENGINE OIL VENT—CLEAR

BRAKE LINE AND BRAKE DEICE PLUMBING (IF INSTALLED)—CHECK

WA-6




CHOCKS—REMOVE 26.

ICE VANE AND OIL RADIATOR EXHAUST—CLEAR

27.

ICE LIGHT—CHECK

28.

ENGINE OIL—CHECK QUANTITY/CAP SECURE

WALKAROUND

23.

24.

FUEL STRAINER AND FUEL FILTER DRAINS—DRAIN

25.

COLLECTOR DRAIN—CLEAR


WA-7



COWLING AIR EXHAUST—CLEAR

32.

TOP COWLING CAMLOCKS (LEFT SIDE)—SECURE

30.

ENGINE COWLING, DOORS, AND PANELS (LEFT SIDE)—SECURE

33.

PROPELLER—CHECK AND ROTATE

31.

EXHAUST STACK AND FAIRING (IF INSTALLED) (LEFT SIDE)—CHECK FOR CRACKS

34.

FORWARD INTAKES ON TOP COWLING—CLEAR

WALKAROUND

29.

WA-8




38.

SWING CHECK VALVE EXHAUST—CHECK

39.

GENERATOR AIR INTAKE—CLEAR

WALKAROUND

35A. ENGINE AIR INTAKE—CHECK

35B. OIL RADIATOR AIR INTAKE—CHECK

36. 37.

TOP COWLING CAMLOCKS (RIGHT SIDE)—CHECK FOR CRACKS EXHAUST STACK AND FAIRING (IF INSTALLED) (RIGHT SIDE)—CHECK FOR CRACKS


WA-9



42.

EXTENDED FUEL TANK CAP (350ER)—SECURE

40B. ENGINE COWLING, DOORS, PANELS AND VGS (RIGHT SIDE)—SECURE

43.

HYDRAULIC GEAR SERVICE DOOR—SECURE

44.

INBOARD DEICE BOOT—CHECK

WALKAROUND

40A. ENGINE COWLING, DOORS, PANELS AND VGS (RIGHT SIDE)—SECURE (CONTINUED)

41.

AUXILIARY FUEL TANK CAP—SECURE

WA-10




46.

47.

HEAT EXCHANGER AIR INTAKE AND EXHAUST— CLEAR

48.

EXTENDED FUEL TANK DRAIN (350ER)—DRAIN

49.

LOWER ANTENNAS AND PANELS—SECURE

HYDRAULIC GEAR OVERFILL AND VENT LINES— CLEAR

AUXILIARY FUEL TANK DRAIN—DRAIN


WA-11

WALKAROUND

45.


NOSE

50A. OAT PROBE/RELIEF TUBE VENT—CHECK (CONTINUED)

BRAKE PRESSURE RESERVOIR VENT—CLEAR

52.

LEFT AVIONICS ACCESS PANEL—SECURE

53.

AIR CONDITIONER CONDENSER EXHAUST DUCT— CLEAR

WALKAROUND

51.

50B. OAT PROBE/RELIEF TUBE VENT—CHECK (CONTINUED)

WA-12



NOSE

LANDING GEAR AND TAXI LIGHTS—CHECK

WALKAROUND

57.

54. 55.

WINDSHIELD AND WIPERS—CHECK RADOME CONDITION—CHECK

56.

PITOT MASTS—CLEAR

58A. NOSE GEAR (SHIMMY DAMPER)—CHECK (CONTINUED)

58B. NOSE GEAR (STOP BLOCK)—CHECK (CONTINUED)


WA-13


NOSE

58C. NOSE GEAR (TORQUE KNEE, STRUT)—CHECK (CONTINUED)

60B. NOSE GEAR DOORS AND WHEEL WELL —CHECK

WALKAROUND 58D. NOSE GEAR (TIRE)—CHECK

59. CHOCKS—REMOVE 60A. NOSE GEAR DOORS AND WHEEL WELL—CHECK (CONTINUED)

WA-14

61.

AIR CONDITIONER CONDENSER INTAKE DUCT— CLEAR

62.

RIGHT AVIONICS ACCESS PANEL—SECURE



RIGHT WING AND NACELLE

68.

EJECTOR EXHAUST—CLEAR

69.

HEAT EXCHANGER AIR INTAKE AND EXHAUST— CLEAR INBOARD DEICE BOOT—CHECK

70.

BATTERY ACCESS PANEL—SECURE

71.

AUXILIAR FUEL TANK CAP—SECURE

WALKAROUND

63.

64. 65. 66.

67.

AUXILIARY FUEL TANK DRAIN—DRAIN EXTENDED FUEL TANK DRAIN (350ER)—DRAIN BATTERY BOX DRAIN—CLEAR

LOWER PANELS—SECURE


WA-15



EXTENDED FUEL TANK CAP (350ER)—SECURE

73.

ENGINE OIL—CHECK QUANTITY, CAP SECURE

74. 75.

COWLING AIR EXHAUST—CLEAR ENGINE COWLING, DOORS, AND PANELS (LEFT SIDE)—SECURE

76.

COLLECTOR DRAIN—CLEAR

77.

EXHAUST STACK AND FAIRING (IF INSTALLED) (LEFT SIDE)—CHECK FOR CRACKS TOP COWLING CAMLOCKS (LEFT SIDE)—SECURE

WALKAROUND

72.

78.

WA-16




TOP COWLING CAMLOCKS (RIGHT SIDE)—SECURE EXHAUST STACK AND FAIRING (IF INSTALLED) (RIGHT SIDE)—CHECK FOR CRACKS

84.

SWING CHECK VALVE EXHAUST—CLEAR

85.

GENERATOR AIR INTAKE—CLEAR

WALKAROUND

82. 83.

79. PROPELLER—CHECK AND ROTATE 80. FORWARD AIR INTAKES ON TOP COWLING—CLEAR 81A. ENGINE AIR INTAKE—CHECK

81B. OIL RADIATOR AIR INTAKE—CHECK


WA-17



86.

ENGINE COWLING, DOORS, AND PANELS (RIGHT SIDE)—SECURE

FUEL FILTER AND FUEL STRAINER DRAINS—DRAIN

90.

LANDING GEAR (DOORS, TIRES, STRUT, WHEEL WELL)—CHECK

WALKAROUND

89.

87.

88.

ICE LIGHT—CHECK

OIL RADIATOR AND ICE VANE EXHAUST—CLEAR

WA-18




FIRE EXTINGUISHER PRESSURE—CHECK

94.

92.

93.

95.

TORQUE KNEE ASSEMBLY AND SAFETY SWITCH— CHECK CHOCKS—REMOVE

96.

BRAKES—CHECK

ENGINE OIL VENT—CLEAR

BRAKE LINE AND BRAKE DEICE PLUMBING (IF INSTALLED)—CHECK


WA-19

WALKAROUND

91.



97.

100.

EXTERNAL POWER DOOR—SECURE

101.

OUTBOARD DEICE BOOT AND STALL STRIP— CHECK

102. 103.

TIE-DOWN—REMOVE FLUSH OUTBOARD FUEL DRAIN—DRAIN

HEATED FUEL VENT AND FLUSH FUEL VENT— CLEAR

WALKAROUND 98.

99.

INVERTER COOLING LOUVERS—CLEAR

GRAVITY LINE AND LEADING EDGE FUEL TANK DRAINS—DRAIN

WA-20




105.

106.

STATIC WICKS (WINGLET AND AILERON)—CHECK SIPHON BREAK VENT—CHECK

109.

AILERON—CHECK

110.

FLAPS (CONDITION, ASYMMETRY PROTECTION, FLAP TRACKS, LIMIT SWITCHES, POSITION TRANSMITTER)—CHECK

WALKAROUND

104.

107. 108.

WING PANELS—SECURE

MAIN FUEL TANK CAP—SECURE

LIGHTS—CHECK


WA-21



111.

RIGHT SIDE WINDOWS—CHECK

WALKAROUND

WA-22



RIGHT AFT FUSELAGE

ELT ANTENNA—CHECK

116.

AFT COMPARTMENT BOTTOM ACCESS DOOR— SECURE

117.

TIE-DOWN—REMOVE

WALKAROUND

115.

112.

113. 114.

LOWER ANTENNAS AND BEACON—CHECK

OXYGEN SERVICE ACCESS DOOR—SECURE STATIC PORTS—CLEAR


WA-23


RIGHT AFT FUSELAGE

120. VENTRAL FIN DRAIN HOLES—CLEAR

119.

CABIN AIR EXHAUST—CLEAR

WALKAROUND

118.

ACCESS PANEL—SECURE

WA-24


121.



TAIL

LEFT AFT FUSELAGE

VENTRAL FIN AND STATIC WICK—CHECK

128.

124. 125. 126.

122. 123.

ACCESS PANELS—SECURE

130.

OXYGEN OVERPRESSURE DISCHARGE AND AFT COMPARTMENT DRAIN TUBES—CLEAR

131.

RELIEF TUBE—CLEAR

HORIZONTAL STABILIZER, BOOTS, AND STATIC WICKS (RIGHT AND LEFT)—CHECK ELEVATOR, TAB, AND STATIC WICKS—CHECK TABS IN NEUTRAL POSITION—VERIFY

VOR ANTENNAS (RIGHT AND LEFT)—CHECK RUDDER, RUDDER TAB, STINGER, AND STATIC WICKS—CHECK

129.

127.

STATIC PORTS—CLEAR

POSITION LIGHT, TAIL FLOODLIGHTS (LEFT AND RIGHT)—CHECK



W-25


5 6 7

28 29 30 31 33 34 32

8 9

4

3

27 41

2


1

3

10 11

48

24 25 26

49

129

121 122 128

123

124 125 126 127

2

4 1

36 37 38 39 40

53 52

5

7 42 43 46 12

6

56

WA-26

44 45

35 47

15 17 16 18 19 20 21 22 23


13 14



APPENDIX SYMBOLS, ABBREVIATIONS, AND TERMINOLOGY CONTENTS Page AIRSPEED................................................................................................................... APPA-1 METEOROLOGICAL ............................................................................................... APPA-2 POWER ......................................................................................................................... APPA-3 CONTROL AND INSTRUMENT ........................................................................... APPA-3 GRAPH AND TABULAR ........................................................................................ APPA-4 WEIGHT AND BALANCE ...................................................................................... APPA-5

APPENDIX A

ABBREVIATIONS AND ACRONYMS ................................................................ APPA-6


APPA-i


APPENDIX SYMBOLS, ABBREVIATIONS, AND TERMINOLOGY CAS—Calibrated airspeed is the indicated airspeed of an aircraft corrected for position and instrument error. Calibrated airspeed is equal to true airspeed in standard atmosphere at sea level. G S — G ro u n d s p e e d i s t h e s p e e d o f a n aircraft relative to the ground. IAS—Indicated airspeed is the speed of an aircraft as shown on the airspeed indicator when corrected for instrument error. IAS values published in the manufacturer’s manual assume zero instrument error. KCAS—Calibrated airspeed expressed in knots K I A S — I n d i c a t e d a i r s p e e d e x p re s s e d in knots M — M a c h n u m b e r i s t h e ra t i o o f t r u e airspeed to the speed of sound. TAS—True airspeed is the airspeed of an aircraft relative to undisturbed air, which is the CAS corrected for altitude, temperature, and compressibility. V 1 —Takeoff decision speed V 2 —Takeoff safety speed is the speed at 35 feet AGL, assuming an engine failure at V1. V 35 —Takeoff safety speed at 35 feet AGL with both engines operating

V A—Maneuvering speed is the maximum speed at which application of full available aerodynamic control will not overstress the aircraft. VF—Design flap speed is the highest speed permissible at which wing flaps may be actuated. V FE —Maximum flap extended speed is the highest speed permissible with wing flaps in a prescribed extended position. V LE —Maximum landing gear extended speed is the maximum speed at which an aircraft can be safely flown with the landing gear extended. V LO —Maximum landing gear operating speed is the maximum speed at which the landing gear can be safely extended or retracted. V MCA —Air minimum control speed is the minimum flight speed at which the aircraft is directionally controllable, as determined i n a c c o r d a n c e w i t h Fe d e r a l Av i a t i o n Regulations. The aircraft certification conditions include: one engine becoming inoperative with autofeather armed, a 5° bank toward the operative engine, takeoff power on the operative engine, landing gear up, flaps in the takeoff position, and most rearward CG. For some conditions of weight and altitude, stall can be encountered at speeds above V MCA as established by the certification procedure described above in which event stall speed must be regarded as the limit of effective directional control.


APPA-1

APPENDIX A

AIRSPEED


V MCG —Ground minimum control speed

METEOROLOGICAL

V MO /M MO —Maximum operating limit speed is the speed limit that may not be deliberately exceeded in normal flight operations. V MO is expressed in knots and M MO in Mach number.

Altimeter Setting—Barometric pressure corrected to sea level

V R —Rotation speed V REF —Reference landing approach speed with the landing gear and flaps down V S —Stalling speed or the minimum steady flight speed at which the aircraft is controllable in a specific configuration.

Indicated Pressure Altitude—The number actually read from an altimeter when the barometric subscale has been set to 29.92 inches of mercury (10 13.2 millibars) IOAT—Indicated outside air temperature is the temperature value read from an indicator. ISA—International standard atmosphere in which:

VS1—Stalling speed or the minimum steady flight speed at which the aircraft is controllable.

• Air is a dry, perfect gas.

VSO—Stalling speed or the minimum steady flight speed at which the aircraft is controllable in the landing configuration.

• Pressure at sea level is 29.92 inches of mercury (10 13.2 millibars).

V SSE—Intentional one-engine-inoperative speed is a speed above both V MCA and stall speed, selected to provide a margin of lateral and directional control when one engine is suddenly rendered inoperative.

APPENDIX A

V X —Best angle-of-climb speed is the airspeed that delivers the greatest gain of altitude in the shortest possible horizontal distance. VXSE—One-engine-inoperative best angle-of-climb speed is the airspeed that delivers the greatest gain in altitude in the shortest possible horizontal distance with one engine inoperative. VY—Best rate-of-climb speed is the airspeed that delivers the greatest gain in altitude in the shortest possible time. V YSE —One engine-inoperative best rateof-climb speed is the airspeed that delivers the greatest gain in altitude in the shortest possible time with one engine inoperative.

APPA-2

• Te m p e r a t u r e a t s e a l e v e l i s 5 9 ° Fahrenheit (15°Celsius).

• Temperature gradient from sea level to the altitude at which the temperature is –69.7°F (–56.5°C) is –0.003566°F (–0.00198°C) per foot, and is zero above that altitude. OAT—Outside air temperature is the free air static temperature obtained either from the temperature indicator (IOAT) adjusted for compressibility effects or from ground meteorological sources. Pressure Altitude—Altitude measured from standard sea level pressure (29.92 inches Hg [1013.2 millibars]) by a pressure (barometric) altimeter. It is the indicated pressure altitude corrected for position a n d in s tr u me n t e r ro r. I n th is manual, altimeter instrument errors are assumed to be zero. Position errors may be obtained from the altimeter correction graphs. Station Pressure—Actual atmospheric pressure at field elevation



Temperature Compressibility Effects—An error in the indication of temperature caused by airflow over the temperature probe. The error varies depending on altitude and airspeed. Wind—The wind velocities recorded as variables on the charts of this manual are to be understood as the headwind or tailwind components of the reported winds.

POWER Beta Range—The region of the power lever control, which is aft of the idle gate and forward of the reversing range where blade pitch angle can be changed with o u t a change of gas generator rpm. Cruise Climb—Cruise climb is the maximum power approved for cruise climb. These powers are torque or temperature (ITT) limited. Cycle—A normal or full cycle includes the following: • Engine start • Idle • Takeoff

Maximum Continuous Power—Maximum continuous power is the highest power rating not limited by time. Use of this rating is at the discretion of the pilot. Maximum Cruise Power—Maximum cruise power is the highest power rating for cruise and is not time limited. Propeller Ground Fine—Propeller ground fine operation is used to provide deceleration on the ground during landing and acceleratestop conditions by taking advantage of the maximum available propeller drag without creating negative thrust. Reverse—Reverse thrust is obtained by lifting the power levers and moving them aft of the beta and ground fine range. SHP—Shaft horsepower S t a t i c Ta ke o f f Powe r — S t a t i c t a ke o ff power is the static power that must be available for takeoff without exceeding the engine limitations. Ta ke o f f Powe r — Ta ke o ff p o w e r i s the maximum power rating. Use of this rating should be limited to normal takeoff operations and other operations at the discretion of the pilot.

• Landing

APPENDIX A

• Flight

CONTROL AND INSTRUMENT

• Idle • Shutdown High Idle—High idle is obtained by placing the condition lever in HIGH IDLE position. This limits the power operation to a minimum of 70% of N 1 rpm.

Condition Lever (Fuel Shutoff Lever)— The fuel shutoff lever actuates a valve in the fuel control unit that controls the flow of fuel at the fuel control outlet and regulates the idle range from low to high idle.

Low Idle—Low idle is obtained by placing the condition lever in LOW IDLE position. This limits the power of operation to a minimum of 62% of N 1 rpm.

ITT (Interstage Turbine Temperature)— Eight probes, wired in parallel, indicate the temperature between the compressor and power turbines.


APPA-3


N 1 Tachometer (Gas Generator RPM)— The N 1 tachometer registers the rpm of the gas generator in percent, with 10 0% representing a gas generator speed of 37,500 rpm.

Best Angle-Of-Climb—The best angle-ofclimb speed is the airspeed that delivers the greatest gain of altitude in the shortest possible horizontal distance with gear and flaps up.

Power Lever (Gas Generator N 1 rpm)— The power lever serves to modulate engine power from full reverse thrust to takeoff. Th e p o s i t i o n f o r i d l e r e p r e s e n t s t h e lowest recommended level of power for flight operation.

Best Rate-Of-Climb—The best rate-of-climb speed is the airspeed that delivers the greatest gain of altitude in the shortest possible time with gear and flaps up.

Propeller Control Lever (N P rpm)—The propeller control lever is used to control the rpm setting of the propeller governor. Movement of the lever results in an increase or decrease in propeller rpm. Propeller feathering is the result of lever movement beyond the detents at the low rpm (high pitch) end of the lever travel. P r o p e l l e r G o v e r n o r — Th e p r o p e l l e r governor senses changes in rpm and hydraulically changes propeller blade angle to compensate for the changes in rpm. Constant propeller rpm is thereby maintained at the selected rpm setting. Torquemeter—The torquemeter system i n d i c a t e s t h e s h a f t o u t p u t t o r q u e. Instrument readout is in percent. APPENDIX A

GRAPH AND TABULAR Ac c e l e r a t e - G o — Ac c e l e ra t e - g o i s t h e distance to accelerate to takeoff decision speed (V 1 ), experience an engine failure, c o n t i n u e a c c e l e ra t i n g t o l i f t o ff, t h e n climb and accelerate in order to achieve takeoff safety speed (V 2 ) at 35 feet above the runway. Accelerate-Stop—Accelerate-stop is the distance to accelerate to takeoff decision speed (V 1 ) and then bring the aircraft to a stop.

Clearway—A clearway is an area beyond the airport runway not less than 500 feet wide, centrally located about the extended centerline of the runway, and under the control of the airport authorities. The clearway is expressed in terms of a clear plane, extending from the end of the runway with an upward slope not exceeding 1.25%, above which no object nor any terrain protrudes. However, threshold lights may protrude above the plane if their height above the end of the runway is 26 inches or less and if they are located to each side of the runway. Climb Gradient—Climb gradient is the ratio of the change in height during a portion of a climb to the horizontal distance traversed in the same time interval. Demonstrated Crosswind—Demonstrated crosswind is the maximum 90° crosswind component for which adequate control of the aircraft during trakeoff and landing was actually demonstrated during certification. The value shown is not limiting. Landing Distance—Landing distance is the distance from a point 50 ft above the runway surface to the point at which the aircraft would come to a full stop utilizing the technique in the Performance section of the manufacturer’s manual. These distances do not include landing factors that may be required by the operating regulations for destination or alternate airports.

AGL—Above ground level

APPA-4



Net Gradient Of Climb—Net gradient of climb is the gradient of climb with the flaps in the takeoff position and the landing gear retracted. Net indicates that the actual gradients of climb have been reduced by a regulatory increment to allow for turbulence and pilot technique. The net gradient of climb graphs are constructed so that the value(s) obtained using the airport pressure altitude and outside air temperature will be the average gradient from 35 ft above the runway up to 1,500 ft above the runway. Route Segment—Route segment is a part of a route. Each end of that part is identified by a: • Geographic location or • Point at which a definite radio fix can be established Ta k e o f f F i e l d L e n g t h ( TO F L ) — Th e m i n i m u m r u n w a y l e n g t h re q u i re d f o r departure. This distance is the longest of the: • Distance to accelerate and recognize an engine failure at V ! , accelerate to and rotate at V R , then climb and accelerate in order to achieve V 2 at 35 feet above the runway, or • Distance to accelerate and recognize an engine failure at V 1 and bring the aircraft to a stop, or • All-engine-operating distance to accelerate to and rotate at V R , then c l i m b a n d a c c e l e ra t e i n o rd e r t o achieve V 35 at 35 feet above the runway, increased by 15%.

Takeoff Flight Path—Takeoff flight path is the minimum gradient of climb required to clear obstacles in excess of 35 feet, measured horizontally from reference zero and vertically at the altitude above the runway. Reference zero is the point where the aircraft has reached 35 feet above the runway, as determined from the takeoff field length graphs.

WEIGHT AND BALANCE A p p r ove d L o a d i n g E nve l o p e — Th o s e combinations of aircraft weight and center of gravity that define the limits beyond which loading is not approved. Arm—Arm is the distance from the center of gravity of an object to a line about which moments are to be computed. Basic Empty Weight—Basic empty weight is the weight of an empty aircraft, including full engine oil and unusable fuel. This equals empty weight plus the weight of unusable fuel and the weight of all the engine oil required to fill the lines and tanks. Basic empty weight is the basic configuration from which loading data is determined. Center Of Gravity—Center of gravity is the point at which the weight of an object may be considered concentrated for weight and balance purposes. CG Limits—CG limits are the extreme center-of-gravity locations within which the aircraft must be operated at a given weight. Datum—Datum is a vertical plane perpendicular to the aircraft longitudinal axis f ro m w h i c h f o re a n d a f t ( u s u a l l y a f t ) measurements are made for weight and balance purposes.


APPA-5

APPENDIX A

MEA—Minimum enroute altitude


Empty Weight—Empty weight is the weight of an empty aircraft before any oil or fuel h a s b e e n a d d e d . Th i s i n c l u d e s a l l permanently installed equipment, fixed ballast, full hydraulic fluid, full chemical toilet fluid, and all other operating fluids full, except that the engines, tanks, and lines do not contain any engine oil or fuel. Engine Oil—That portion of the engine oil that can be drained from the engine. Jack Point—Jack points are points on the aircraft identified by the manufacturer as suitable for supporting the aircraft for weighing or other purposes. Landing Weight—Landing weight is the weight of the aircraft at landing touchdown. Leveling Points—Leveling points are those points that are used during the weighing process to level the aircraft.

Takeoff Weight—Takeoff weight is the weight of the aircraft at liftoff from the runway. Ta re — Ta re i s t h e w e i g h t t h a t m a y b e indicated by a set of scales before any load is applied. Unusable Fuel—Unusable fuel is the fuel remaining after consumption of usable fuel. Usable Fuel—Usable fuel is that portion of the total fuel that is available for consumption as determined in accordance with applicable regulatory standards. Useful Load—Useful load is the difference between the aircraft ramp weight and the basic empty weight. Zero Fuel Weight—Zero fuel weight is the aircraft ramp weight minus the weight of fuel on board.

Maximum Weight—Maximum weight is the greatest weight allowed by design, structural, performance, or other limitations.

ABBREVIATIONS AND ACRONYMS

Maximum Zero Fuel Weight—Any weight above the value given must be loaded as fuel.

AFD—Adaptive flight display

APPENDIX A

Moment—Moment is a measure of the rotational tendency of a weight, about a specified line, mathematically equal to the product of the weight and the arm. Pa y l o a d — Pa y l o a d i s t h e w e i g h t o f occupants, cargo, and baggage.

ESIS— E le ctro n ic s ta n d by in s trument system EIS—Engine indication system ISS—Impending stall speed LSC—Low speed cue PFD—Primary flight display

PPH—Pounds per hour Ramp Weight—Ramp weight is the aircraft weight at engine start, assuming all loading is completed.

MFD—Multifunction display SAT—Static air temperature

Station—Station is the longitudinal distance from some point to the zero datum or zero fuselage station.

APPA-6



APPENDIX B ANSWERS TO QUESTIONS Chapter 1—Aircraft General 1. Aircraft equipped with dual strakes require yaw damper operation above _________ feet: C. 19,000 2. Lateral-tracking seats must be in the full _________ position for _______ D. Outboard; takeoff and landing. 3. Illumination of the red master warning annunciator [DOOR UNLOCKED] indicates: B. The airstair door is open or not secure. 4. The maximum allowed operating altitude limit is ________ feet. B. 35,000 5. The maximum allowed operating temperature limit above 25,000 feet is ISA + ______°C. C. 31 6. Single pilot operations require: A. The pilot to use a headset with a boom microphone. 7. With appropriate equipment, the kinds of operations allowed: A. Permit flight at night. 8. Passenger briefing cards are required at one per seat for: B. 14 CFR Part 135 operations. 9. V XSE is _______ KIAS. B. 125

APPENDIX B

10. V MCA for Flaps Approach is ______ KIAS. B. 93


APPB-1


Chapter 2—Electrical Power Systems 1. During a battery start, prior to selecting ON with the IGNITION AND ENGINE START switch and before starting the second engine, the DC percent loadmeter should read approximately _______ percent or less. A. 50 2. The minimum battery voltage required for an external power start is _______ volts. C. 20 3. Control switches which are operable during a dual generator failure are indicated by ______________ the switch. A. A white circle around 4. A generator bus tie will open automatically to protect the electrical system from a malfunction when excessive current is sensed on _____________________ bus. C. The same-side generator 5. The external power cart will be set to _______ volts and be capable of generating a minimum of 1000 amps momentarily and 30 0 amps continuously. C. 28.0 – 28.4 6. The maximum sustained generator load at 30,0 0 0 feet is _______ percent. D. 100 7. The first immediate action item for a DUAL GENERATOR FAILURE is: C. Instrument Emergency Lights (if requied)............ON

Chapter 3—Lighting 1. Selecting the landing light switches on will illuminate both landing: C. Lights regardless of gear position. 2. Both wing ice lights are required to be operable during flight during _______ operations. D. Icing conditions 3. The EXIT signs automatically illuminate during normal flight operations when: D. Rapid deceleration is sensed. APPENDIX B

APPB-2



Chapter 4—Master Warning System 1. The MASTER WARNING FLASHERS illuminate when ___________ annunciator illuminate(s). A. A red warning 2. A red warning annunciator will extinguish when: B. The fault is no longer sensed. 3. Faults that illuminate the ______________ annunciators require immediate attention and reaction of the pilot. A. Red warning

Chapter 5—Fuel System 1. If auxiliary fuel is required, the auxiliary tank _______ be filled _______ filling the main fuel tanks. D. Must; after 2. Illumination of the amber [L/R FUEL QTY] annunciator indicates less than 30 minutes of fuel remaining: C. At maximum continuous power. 3. Illumination of the red [L/R FUEL PRESS LO] warning annunciator during normal flight operations indicates: A. Insufficient pressure at the fuel pressure switch. 4. According to the checklist, crossfeed is selected: B. Only during single engine operations. 5. The approved military grade fuels are: A. JP-4, JP-5, and JP-8.

APPENDIX B

6. The maximum allowed lateral fuel imbalance is _______ lbs. B. 300


APPB-3


Chapter 7—Powerplant 1. The minimum N 1 required to select LOW IDLE on the condition lever during engine start is: B. 12% 2. Overfilling the oil may cause: A. Discharge until a satisfactory level is reached. 3. If the compressor bleed valve fails to close as static take-off power is set, torque will indicate ________ than normal and ITT will indicate _______ than normal. D. Lower; higher 4. The shaft horse power rating of 1,050 is a direct function of: C. Torque and Propeller RPM. 5. Ignition operation occurs during engine start and during operations of _____________ or less when engine auto ignition is ____________. B. 17% torque; armed 6. The minimum oil temperature limit allowed for engine start is _____°C. A. –40 7. The maximum allowed continuous ITT for takeoff is _______°C. B. 820 8. The minimum allowed oil pressure for idle is _______ PSI. A. 60 9. Oil temperatures between 99°C and 110°C are limited to _______ minutes. D. Ten 10. The maximum gas generator N 1 RPM limit for takeoff is: B. 104 11. The first immediate action item for an ENGINE FIRE OR FAILURE IN FLIGHT is affected engine: B. Condition Lever ....................................FUEL CUTOFF

APPENDIX B

12. The immediate action items for and ENGINE FAILURE DURING TAKEOFF (AT OR BELOW V 1 ) — TAKEOFF ABORTED are: D. Power Levers ..........................................GROUND FINE, Brakes.......................................................MAXIMUM

APPB-4



13. In order to select ground fine after landing, the pilot: A. Lifts the power lever and moves them aft to the first gate. 14. The propeller governor is scheduled to control RPM between _______ RPM. C. 1450–1700 15. The autofeather system will feather the inoperative engine’s propeller when the opposite engine torque drops below: D. 10% torque. 16. The fuel topping governor limits propeller RPM in flight to ____ percent of selected RPM. D. 106 17. The overspeed governor limits propeller RPM to a maximum of: B. 1768. 18. The maximum allowed continuous RPM for takeoff is _______ RPM. C. 1700

Chapter 8—Fire Protection 1, Engine fire detection and extinguishing is available when the battery bus switch is selected to _________ and the battery switch to __________. D. NORM; ON 2. Engine fire extinguishing is available for the engine: A. Compartment. 3. With the hot battery bus powered, an engine fire extinguisher may be discharged: D. After depressing the on-side firewall fuel valve switch.

APPENDIX B

4. Th e f i r s t i m m e d i a t e a c t i o n f o r E N V I RO N M E N TA L SYST E M S M O K E O R FUMES is: A. Oxygen Mask(s).....................................DON


APPB-5


Chapter 9—Pneumatics 1. Regulated pneumatic air pressure is used to: B. Inflate the deice boots. 2. After selecting the bleed air valve to pneumatic and environmental off after illumination of a single [L or R BLEED FAIL] red master warning annunciator, the annunciator will: B. Remain illuminated. 3. Vacuum air is provided for: D. Wing deice boot hold-down.

Chapter 10—Ice and Rain Protection 1. During flight in visible moisture, or at night when flight from visible moister cannot be assured, engine anti-ice must be on at temperatures below _____°C. B. 5 2. In the event of windshield icing, reduce speed to ______ knots or below. D. 226 3. Operating the propeller deice in the _______ mode provides _______ timer operation. D. AUTO; automatic 4. The surface deice system removes ice build up on the leading edge(s) of the: C. Wing and horizontal stabilizer. 5. The minimum airspeed for sustained flight in icing conditions is _____ knots. A. 140

Chapter 11—Air Conditioning 1. The vapor-cycle refrigeration compressor is located: A. On the right engine accessory section. 2. If the engine speed is too low for the air conditioning compressor to properly engage, the: A. White [AIR COND N1 LOW] status annunciator illuminates. APPENDIX B

3. For more efficient cooling on the ground, place the BLEED AIR VALVES switches to the __________ position. C. ENVIR OFF 4. In the MAN HEAT mode on the ECS, the pilot controls temperature with the: D. MAN TEMP INCR DECR switch

APPB-6



Chapter 12—Pressurization 1. The CABIN ALT gauge indicates cabin _______ and cabin _______ altitude. A. Differential pressure; pressure 2. The cabin _______ pressurize on the ground by selecting the _______ position of the CABIN PRESS switch. B. Will; TEST 3. The white [CABIN ALTITUDE] status annunciator illuminates when cabin pressure altitude indicates _______ feet. A. 10,000 4. The red [CABIN ALT HI] warning annunciator illuminates when cabin pressure altitude exceeds _______ feet. C. 12,000 5. The first immediate item for PRESSURIZATION LOSS is: D. Oxygen Mask(s).....................................DON 6.

The first immediate action item for the EMERGENCY DESCENT is: C. Power Levers ..........................................IDLE

Chapter 14—Landing Gear and Brakes 1. The landing gear handle is designed to work airborne: A. Weight off wheels. 2. The green GEAR DOWN annunciators indicate the gear: C. Is down and locked. 3. The LDG GEAR CONTROL red light illuminates when the gear position may be unsafe and: D. Cannot be dimmed. 4. The alternate landing gear extension system is available for gear: B. Extension.


APPENDIX B

5. The maximum permitted landing gear extended speed is _______ KIAS. B. 184

APPB-7


Chapter 15—Flight Controls 1. Secondary flight controls surfaces are: D. Manually and electrically controlled. 2. Rudder boost aids the pilot in rudder deflection during engine failure operation by sensing: C. Torque differential. 3. Electric pitch trim is available when: B. Both trim switches are activated simultaneously. 4. The mechanical aileron trim is located _______ of the power quadrant on the _______ side. A. Aft; left 5. The maximum speed permissible with flaps in the approach position is _______ KIAS. D. 202 6. Rudder boost: D. Must be on and operational for takeoff, climb, approach and landing.

Chapter 16—Pro Line Avionics 1. The minimum autopilot use height during an approach is _______ feet. A. 79 2. A copilot side heading failure can be corrected by placing the: A. AHRS switch to No. 1. 3. The active No. 2 bearing pointers are: B. Cyan. 4. ISA deviation can be found on the: D. MFD. 5. In order for the BARO MINS to be displayed, the values on the REFS page must be: C. Cyan. APPENDIX B

6. The composite mode is activated by selecting the _______ reversion switch. A. PILOT DISPLAY

APPB-8



7. Pressing the BARO knob on the display control panel (DCP) will: C. Set Flight Level altitudes on the altitude preselector display. 8. The color of the to-waypoint on the CDU is: D. Green. 9. Airspeed trend information is available: C. From a magenta indicator on the airspeed indicator. 10. The minimum autopilot use height during an approach is _______ feet. A. 79 11. An FMS preflight includes: B. Checking the Database. 12. Using the FMS for guidance is not authorized: D. Inside the FAF on a localizer approach. 13. For an FMS preflight procedure, the I in VIPP stands for: A. Initialize. 14. For an FMS preflight procedure, the second P in VIPP stands for: B. Performance. 15. APPR must be pressed: C. When VNAV to a decision altitude is desired. 16. VNAV guidance is: D. Prohibited during a missed approach. 17. If GPS APPR is not displayed inside the final approach fix: A. The GPS must not be used for flight guidance.

APPENDIX B

18. Magenta color text on the CDU LEGs page indicates _______ information. B. Airspeed and altitude


APPB-9


Chapter 17—Oxygen 1. Deployment of the passenger oxygen masks is indicated by illumination of the _________ annunciator. D. White [PASS OXYGEN ON] system status 2. Manual deployment of the passenger oxygen masks is available by _______ the control handle on the _______ side of the center console. B. Pulling out; right 3. The amber [OXY NOT ARMED] caution annunciator illuminates when the: A. Main oxygen system is not armed. 4. Crew oxygen is provided by a _________ oxygen mask. D. Diluter-demand quick-donning

APPENDIX B

APPB-10


ANNUNCIATOR PANEL


ANNUNCIATORS SECTION The Annunciators Section presents a color representation of all the annunciator lights in the aircraft. Please unfold page ANN-3 to the right and leave it open for ready reference as the annunciators are cited in the text.


ANN-1



Figure ANN-1. Annunciators FOR TRAINING PURPOSES ONLY


ANN-3

Beechcraft Be350/350C

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