Cessna Citation II TM [PWD] Dec 1999

Cessna Citation II Technical Manual

Chapter 1 Aircraft General Table of Contents Overview ............................................... 1-1 Publications .......................................... 1-1 Airplane Equipment ............................... 1-1 Airframe Structure ................................. 1-2 Fuselage ........................................... 1-3 Wings ............................................... 1-4 Empennage....................................... 1-6 Nose Section ........................................ 1-7 Pressurized Center Section .................... 1-9 Flight Compartment ......................... 1-12 Passenger Cabin ............................. 1-16 Cabin Door and Stair Assembly ....... 1-20 Cabin/Cargo Door ........................... 1-26 Emergency Exit Door ........................... 1-29 Hand-Held Fire Extinguishers ............ 1-30 Aft Fuselage Section ........................... 1-31 Limitations .......................................... 1-32 Emergency Procedures......................... 1-32

Cessna Citation II Technical Manual Overview

The Cessna Citation II is a high performance, twin-turbofan, pressurized, eight to twelve place airplane certificated in accordance with FAR Part 25 airworthiness standards. The standard airplane is approved for operation in day, night, VFR, IFR, and known icing conditions as defined by the FAA. Publications Cessna Aircraft Company publishes documentation providing detailed airplane systems information and operating procedures. This Technical Manual is not intended to supersede the Operating Manual, FAA approved Airplane Flight Manual (AFM), the Pilot’s Check List, and/or related publications specific to your airplane. Airplane Equipment Airplane systems and equipment provided by the manufacturer as standard from the factory, as well as manufacturer installed optional systems or equipment will be covered in this publication. Vendor supplied Supplemental Type Certificated (STC) accessories or equipment will not be covered. This chapter provides a general description of the airplane structure, accessories, and equipment.

AIRCRAFT GENERAL 12/99

FOR TRAINING PURPOSES ONLY

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Airframe Structure The Cessna Citation II is a low-wing airplane of primarily all-metal construction with retractable landing gear and aft-mounted engines. The airframe is a conventional semi-monocoque design of aluminum alloy with composite structures used in specific areas. Flush riveting, fairings, and aerodynamic joint sealants are used where appropriate to minimize drag.

RUDDER ELEVATOR TRIM

FORWARD SPAR RUDDER TRIM

ELEVATOR DORSAL FIN ENGINE PYLON

AILERON ENGINE CARRY-THRU BEAMS

SPEED BRAKE FLAP

REAR SPAR

STRINGER

FRAME AFT PRESSURE BULKHEAD

FLOOR PANEL/ RAIL ASSEMBLIES

MAIN WHEEL WELL AFT SPAR

STRINGER

AILERON TRIM

NOSE WHEEL WELL

WINDOW FRAME DOOR FRAME

FRAME LIGHTNING STRIPS

FORWARD PRESSURE BULKHEAD

WING CARRY-THRU SPARS

STATIC WICKS RIB FRONT SPAR STRINGER

RADOME

Protection against electromagnetic interference (EMI) and radio frequency interference (RFI) is accomplished primarily by the incorporation of bonding jumpers throughout the airframe. A conductive finish applied to the inner surface of wing skin panels provides additional protection against EMI and RFI. Protection against lightning strikes and accumulation of static electricity is accomplished by lightning strips installed on the nose cap (radome), and by static wicks installed on the trailing edges of the wings, flight control surfaces, and tailcone stinger.

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CITATION II CHAPTER 1 12/99


Fuselage The fuselage is an aluminum alloy, semi-monocoque structure consisting of transverse frames and bulkheads, longitudinal stringers, and external skin panels. The nose wheel well structure, integral to the fuselage, provides attachment points for the nose gear assembly and related components. Frames and fittings are also provided for the attachment of doors and windows. Composite nose and tailcone fairings provide aerodynamic smoothness and access to avionics components. Carrythru spars pass laterally through the lower fuselage for attachment of the wings. Carry-thru beams pass laterally through the aft fuselage for attachment of the engines. Chordwise ribs, spanwise stringers, and external skin panels are fastened to the outboard front (main) and rear wing carrythru spars to form the stub wings. The interior of each stub wing is sealed for fuel storage forward of the rear spar, between the inboard and outboard ribs. The fuel storage area is chemically treated and finished with an epoxy primer for corrosion resistance. Aluminum alloy fairings provide aerodynamic smoothness between the fuselage and stub wing, as well as access to wire bundles and various air and fluid lines.



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Spanwise bulkheads, angles, and stringers; chordwise ribs, and external skin panels are fastened to the outboard fore and aft engine carry-thru beams to form the engine pylons. The outboard ribs are constructed of stainless steel and sealed to form a firewall and vapor barrier through which control cables, wire bundles, and various air and fluid lines are routed to and from the engines. Access panels on the lower surface of the fuselage, stub wings, and engine pylons facilitate inspection and maintenance. Drain holes are provided on the lower surface of the fuselage, stub wings, and pylons where fluids and/or moisture collect. Drainage from unpressurized areas is continuous. Drainage from pressurized areas is regulated by check valve seals which are open only when the airplane is unpressurized. Note: Moisture drain holes must be clear and free of obstructions for proper operation. The nose section, pressurized center section, and aft section of the fuselage are further described individually in this chapter. Wings The wings are aluminum alloy, semi-monocoque structures consisting of front (main) and rear spars, spanwise stringers, chordwise ribs, and external skin panels. The wings and stub wings are mated and secured by threaded fasteners at attachment points on the upper and lower front and rear spars. Main wheel wells, integral to each wing structure, provide attachment points for the main gear assemblies and related components. The skin panel directly above each main wheel well consists of a honeycomb core material bonded between aluminum skin panels. The interior of each wing, excluding the main wheel well, is sealed for fuel storage forward of the rear spar, between the inboard and outboard ribs. The fuel storage area is chemically treated and finished with an epoxy primer for corrosion resistance.

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Cessna Citation II Technical Manual The wing structures also provide attachment points for the ailerons, flaps, speed brakes, and their associated actuators. The ailerons and flaps are of aluminum alloy, semi-monocoque construction incorporating spanwise spars, chordwise ribs, and external skin panels. The speed brakes are of aluminumreinforced, magnesium alloy construction.

The outboard end of each wing is enclosed by a wing tip of aluminum alloy, semi-monocoque construction incorporating chordwise ribs, spanwise stringers, and external skin panels. Wing tip fairings may be of aluminum alloy or composite construction. The inboard leading edge of each wing is formed by an electrically-heated anti-ice panel. A Kevlar insulation shield provides a thermal barrier between the heated leading edge panel and the wing structure. Pneumatic deice boots are installed on the outboard leading edge of each wing. Access panels on the lower surface of the wings facilitate inspection and maintenance of control surface actuators and fuel system components.



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Empennage The empennage is an aluminum alloy, semi-monocoque structure consisting of the vertical stabilizer, horizontal stabilizer, and associated control surfaces. The vertical stabilizer is constructed primarily of fore and aft spars, chordwise ribs, and external skin panels. The fore and aft spars are secured by threaded fasteners to the fuselage structure. The horizontal stabilizer is constructed primarily of fore and aft spars, spanwise auxiliary spars, chordwise ribs, and external skin panels. The fore and aft spars are secured by threaded fasteners to the vertical stabilizer. Attachment points are provided for the rudder, elevators, and their associated actuators. The rudder and elevators are also constructed of spars, ribs, and external skin panels.

A composite dorsal fin and saddle fairing provide aerodynamic smoothness between the upper fuselage and the vertical stabilizer. The emergency locator transmitter is housed within these fairings. Vertical and horizontal stabilizer tip fairings (caps) may be of aluminum alloy or composite construction. Pneumatic deice boots are installed on the leading edge of the vertical and horizontal stabilizers. Access panels facilitate inspection and maintenance of control surface actuators and some navigational components. The access panels located on the vertical stabilizer directly below the horizontal stabilizer consist of a honeycomb core material bonded between aluminum skin panels. 1-6



Cessna Citation II Technical Manual Nose Section The unpressurized nose section contains the avionics bay and the nose baggage compartment.

The avionics bay is located just aft of the fiberglass nose cap (radome) which houses the weather radar antenna. An avionics access panel, constructed of honeycomb core material bonded between aluminum skin panels, encloses the width of the upper surface of the nose section between the nose cap and the nose baggage compartment. ▲ AVIONICS BAY The access panel and nose cap form a single unit attached by quick-disconnect “Tridair” fasteners and secured by two key locks, one installed on each side of the panel. Note: The avionics access panel and nose cap are removed and installed as a single unit. The double row of fasteners should not be disturbed unless separation of the nose cap from the access panel is required for maintenance.



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The nose baggage compartment is located between the avionics bay and forward pressure bulkhead. Access is through left and right side doors attached to the fuselage structure by two hinges each, and secured by two latches and one key lock each. Each door is opened by disengaging the key lock and latches and lifting the door to the open ▲ NOSE BAGGAGE COMPARTMENT position. When fully open, a spring-loaded stop assembly holds the door in position. Closing the doors is accomplished by releasing the stop, lowering the door, and reengaging the latches and key lock. Fore and aft dividers separate the nose baggage compartment from the avionics bay and forward pressure bulkhead respectively. A hinged access panel on the aft divider, accessible through the right baggage door, facilitates fluid and pneumatic servicing. Five inspection windows on the access panel are positioned to permit viewing the sight gages on the brake fluid and windshield anti-ice fluid reservoirs, as well as pneumatic pressure gages for the emergency braking and gear extension storage cylinder and anti-skid accumulator(s). On airplanes 550-0254 and earlier (not incorporating SB550-35-2) the oxygen storage cylinder is also serviced and accessed through the right baggage door. In some installations, a hinged access panel may be located below this door to facilitate oxygen servicing. Though limits vary with equipment installation, the maximum volume and load capacity of the nose baggage compartment are 17 cubic feet and 350 pounds respectively. Indication of nose baggage door security is provided by the amber [DOOR NOT LOCKED] annunciator. The annunciator is controlled by a microswitch integral to each forward latch assembly. With electrical power applied to the airplane and either door unsecured, the door warning circuit is complete and the annunciator illuminates. When both doors are properly secured, the door warning circuit is interrupted and the annunciator is extinguished.

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Cessna Citation II Technical Manual Pressurized Center Section

The airplane center section is reinforced and sealed for pressurization to the skin between the forward and aft pressure bulkheads. Included in the center section are the flight compartment, passenger cabin, standard cabin door, optional cabin/cargo door (if installed), and the emergency exit. A two-piece windshield, two side windows, and a foul weather window provide flight compartment visibility. The windshield is a Plexiglas laminate of stretched acrylic outer and inner layers with a vinyl core.

▲ PILOT'S WINDSHIELD



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On airplanes 550-0681 and earlier, both side windows consist of prestressed Plexiglas outer and inner panes separated by the window frame and spacers to form a thermal air barrier. On airplanes 550-0682 and after, both side windows consist of prestressed Plexiglas outer and middle panes with an acrylic inner “frost” pane. A thermal air barrier is formed between the outer pane and middle pane, separated by the window frame and spacers; and between the middle pane and frost pane, separated by spacers and seals. Six windows are located on each side of the passenger cabin including those located in the cabin door and the emergency exit. Each includes a Plexiglas laminate outer pane of stretched acrylic outer and inner layers with a vinyl core, and an acrylic inner “frost” pane. The outer pane and frost pane are separated by the window retainer and a spacer to form a thermal air barrier. All passenger cabin windows, excluding those located in the passenger door and the emergency exit, incorporate integral sliding shades. All windows, forming part of the pressure vessel, are fixed except for the foul weather window. The foul weather window, located forward of the pilot’s side window, is also a Plexiglas laminate of stretched acrylic outer and inner layers with a vinyl core. The window is hinged at the bottom, latched at the top, and incorporates a peripheral seal which makes the window airtight when properly secured. The window is opened ▲ FOUL WEATHER WINDOW by disengaging the latch and pulling the window inward. Closing and securing the window is accomplished by repositioning the window in the frame and reengaging the latch. During ground operation, the foul weather window may be opened to supplement flight compartment ventilation. During cold weather operation, the flight compartment may be warmed before flight by routing a preheater hose through the foul weather window. For pressurized operation, the foul weather window must be properly secured.

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Cessna Citation II Technical Manual Windshield and window condition is critical to flight compartment visibility as well as pressure vessel structural integrity and should be inspected before flight. Obvious damage or deformities should be evaluated against repair/replacement criteria by maintenance personnel. Generally, light surface damage and discoloration or deformities outside of critical vision areas do not compromise structural integrity or limit flight operations. Conversely, moderate to heavy surface damage, cracks and delamination, and discoloration or deformities inside of critical vision areas require immediate repair or replacement before resuming flight operations. Note: Unrepaired damage may reduce windshield/window service life; therefore, the manufacturer recommends that any windshield/window damage be repaired as soon as practical. Plexiglas is easily damaged by improper handling and cleaning techniques and should be kept clean and waxed using the methods and materials prescribed by the manufacturer. When prescribed materials are unavailable, the windshield and windows may be cleaned using a common, mild liquid soap as long as appropriate precautions are taken: 1. Surfaces should be permitted to cool before cleaning. 2. All hand and wrist jewelry should be removed. 3. Dirt and abrasives should be dislodged with running water. 4. Cloth or sponge may only be used to transfer soap to surface. 5. Only light rubbing force with a bare hand should be applied. 6. After cleaning, surface should be rinsed with running water. 7. Surface may be dried with clean, damp chamois or cloth. 8. After drying, contact with the surface should be avoided.



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Flight Compartment The flight compartment provides for a crew of two, with full instrumentation for the pilot and copilot. Layout is conventional in that all controls, switches, and instruments are accessible to the pilot for single pilot operation.

The pilot’s and copilot’s instrument panels contain primary and navigational flight instruments and controls. Flood and panel light controls, light switches, and the standby gyro switch/ light(s) are located on the pilot’s lower instrument panel. Controls for the parking brake, control surface lock, emergency braking, and auxiliary/emergency gear extension are located below the pilot’s instrument panel. Windshield bleed air controls, fan switches, and gyro switches are located on the copilot’s lower instrument panel. The windshield rain removal augmenter control is located below the copilot’s instrument panel.

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Cessna Citation II Technical Manual The upper left instrument panel contains the pilot’s audio control panel, outside air temperature (OAT) indicator, clock, and optional angle-of-attack (AOA) indicator. The lower left instrument panel contains electrical system switches and indicators, fuel system switches and controls, engine switches, ice protection switches, exterior light switches, and the test selector switch. The right instrument panel contains the copilot’s audio control panel, battery temperature indicator (if installed), gyro pressure gage (550-0626 and earlier), oxygen pressure gage, and flight hours meter. The center instrument panel contains engine instrumentation, the annunciator panel, avionics control panels, weather radar, supplemental navigation equipment, and the landing gear controls and position indicators. Circuit breaker panels are located on the left and right sidewalls. Crew oxygen outlets and audio jacks are located on the left and right side consoles. Oxygen system controls are located on the left side console. Controls and indicators for the engine fire protection system and thrust reversers are located on a panel directly below the glareshield. The center pedestal contains the engine control levers as well as controls for the flaps, speed brakes, manual trim, autopilot, and navigation equipment. The environmental panel contains cabin pressurization switches, controls, and indicators. Refer to corresponding systems chapters for specific detail.



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1

2

3 1. Height Adjustment Handle 2. Fore and Aft Position Handle 3. Seat Back Tilt Handle

The pilot’s and copilot’s seats are installed on independent rail assemblies attached to longitudinal floor beams integral to the fuselage structure. Fore and aft position, height, and tilt angle are manually adjustable. Each seat moves fore and aft along the rails on roller and guide assemblies attached to its frame. Stop pins engage the rails to secure the seat in position. Fore and aft adjustment is accomplished by lifting the handle located below the forward center of the seat to disengage the stop pins from the rails, and sliding the seat to the desired position. Height adjustment is accomplished by lifting the handle located below the inboard forward corner of the seat and weighting or unweighting the seat to the desired position. A shock cord (bungee) and pulley arrangement provides approximately 100 pounds of lift assistance when the seat is unweighted. Tilt angle adjustment is provided by a pneumatic actuator attached to each seat frame, and a handle located on the aft inboard corner of the seat. Pushing down on the forward end of the handle releases air pressure within the actuator allowing the seat to be tilted to the desired position. When the handle is released, air pressure trapped within the actuator holds the seat in the selected position. Armrests are installed on each inboard seat back. When not in use, the armrests are stowed in an upright position behind and flush with the seat backs. For use, each armrest is pulled inboard from its stowed position and lowered by pushing down on its forward end. On airplanes 550-0222 and after, the lowered position of the armrest may be selected by means of an adjustable stop. This adjustment, however, cannot be made during flight. 1-14



Cessna Citation II Technical Manual Each crew seat is equipped with a restraint system incorporating an inertia reel type shoulder harness. Airplanes 550-0431 and after are equipped with a five-point restraint system consisting of two lap belts, a dual-strap shoulder harness, forward restraint strap, and padded rotary buckle. Airplanes 5500356 through 550-0430 are ▲ CREW RESTRAINT SYSTEM equipped with a four-point restraint system consisting of two laps belts, a dual-strap shoulder harness, and padded rotary buckle. In four-point and fivepoint installations, the inertia reel is attached to the seat frame and all restraints engage the rotary buckle. The quick-release design of the rotary buckle permits simultaneous disengagement of all restraints except the outboard lap belt, to which the buckle is attached, by rotating the release mechanism counterclockwise. Airplanes 550-0355 and earlier are equipped with a three-point restraint system consisting of two laps belts, a single-strap shoulder harness, and conventional buckle. In this installation, the inertia reel is attached to the overhead airframe structure aft and outboard of each crew seat. The shoulder harness engages a link on the outboard lap belt which engages the buckle on the inboard lap belt. Other flight compartment equipment and furnishings include a navigation chart case located behind the copilot’s seat, a relief tube storage case located behind the pilot’s seat, sun visors, overhead directional air vents, and overhead flight compartment lighting. The relief tube incorporates an electricallyheated drain/vent which is ▲ DIRECTIONAL AIR VENT AND LIGHT supplied with 28 VDC right main bus power through the 7.5-amp TOILET circuit breaker (not accessible from the flight compartment). On airplanes 5500627 and after, optional tinted sun visors positionable along a monorail track may be installed in place of the standard vinylcovered “pivoting” sun visors. AIRCRAFT GENERAL 12/99


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Passenger Cabin The passenger cabin measures approximately 194.7" from the forward cabin dividers to the aft pressure bulkhead, 59.2" from sidewall to sidewall, and 56.0" from the lowest point of the floor to the ceiling. Passenger cabin configurations vary according to seating arrangement and installation of standard or optional furnishings and equipment.

▲ AFT PASSENGER CABIN

Standard configurations typically include seating for six passengers, a refreshment center, and a non-flushing toilet. Optional configurations include seating arrangements for up to ten passengers, a flush toilet, an executive writing table, and various storage cabinets and/or refreshment centers. Forward and aft dividers separate the passenger area from the flight compartment and aft baggage compartment respectively.

▲ REFRESHMENT CENTER

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▲ EXECUTIVE WRITING TABLE

The standard passenger seats, numbered 5 through 10 for payload computation purposes, are installed on independent rail assemblies attached to longitudinal floor beams integral to the fuselage structure. These seats may be high back, or low back with an adjustable headrest. Seats 5 and 6, located midcabin, may be installed facing forward or aft. Seats 7 and 8, located immediately forward of the aft cabin dividers, are installed forward facing only. Seats 9 and 10, located immediately aft of the forward cabin dividers, are installed aft facing only. Each seat moves fore and aft along the rails on roller and guide assemblies attached to its frame. Stop pins engage the rails to secure the seat in position. Fore and aft adjustment is accomplished by lifting the handle located below the forward center of the seat to disengage the stop pins from the rails, and sliding the seat to the desired position. When located “over spar,” seats 7 and 8 are secured directly to the rails and are not adjustable fore and aft once installed. Seat back angle adjustment from vertical to a reclining position is provided by spring-loaded “hydrolock” actuators attached to each seat frame. Pressing the button on the inboard side of the armrest releases air pressure within the actuators allowing the seat back to be tilted to the desired position. When the button is released, air pressure trapped within the actuators holds the seat back in the selected position. AIRCRAFT GENERAL 12/99


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Passenger seats equipped for lateral adjustment may be positioned inboard away from the sidewalls to provide greater headroom. This is accomplished by lifting the control handle located below the inboard side of the seat and moving the seat to the desired position. Note: Lateral adjustment seats should be locked in the outboard position during takeoff and landing. Armrests are installed on each inboard seat frame. When not in use, armrests are stowed flush with the bottom seat cushions. For use, each armrest is pulled upward until a spring-loaded latching mechanism is engaged. Lifting the lever on the forward end of each armrest disengages the latching mechanism permitting the armrest to be stowed. Each passenger seat is equipped with a restraint system consisting of an adjustable lap belt and inertia reel shoulder harness (550-0550 and after) or an adjustable lap belt only (550-0505 and earlier). Optional passenger seating arrangements may include two additional standard type seats located at midcabin and numbered 3 and 4 for payload computation purposes, an aft portable seat, forward lounge seats, and a two or threeplace forward facing divan (couch). The optional flush toilet may be certified for use as a passenger seat when ▲ AFT FLUSH TOILET located in the aft baggage compartment (550-0550 and after) or when located in the forward passenger cabin (550-0505 and earlier). Non-flushing toilets are generally not certified for use as a passenger seat. Aft located toilets are also equipped with a relief tube incorporating an electrically-heated drain/vent which is supplied with 28 VDC right main bus power through the 7.5-amp TOILET circuit breaker (not accessible from the flight compartment). Flush toilets are also powered by this circuit. Privacy for the aft toilet area may be provided by curtains, sliding doors (5500550 and after), or a folding door (550-0505 and earlier) which extend(s) between the left and right aft cabin dividers.

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1

1. 2. 3. 4.

Passenger Oxygen Masks Reading Light Reading Light Switch Ventilation Air Outlet

Individually controlled reading lights, ventilation air outlets, and oxygen outlets are located on the ceiling above each passenger station. Airflow for heating is provided through registers located just above floor level on each side of the cabin.



2

3

4

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The aft baggage compartment is located between the aft cabin dividers and the aft pressure bulkhead, and is accessible during flight. A cargo net and tie-down straps are provided to secure the contents of the baggage compartment. The net is held in place by attachment fittings that engage anchor plates located on the floor, aft pressure bulkhead, and sidewalls. Though limits vary with passenger cabin configuration, the maximum volume and load capacity of the aft baggage compartment are from 34 to 43 cubic feet and 220 to 600 pounds respectively. Note: Refer to the appropriate AFM and airframe placards for weight and balance limitations specific to your airplane.

Cabin Door and Stair Assembly The cabin door and stair assembly are located in the forward left side of the center section. The door is a single-section assembly of aluminum alloy construction which houses the forward left passenger cabin window and, when secured with all locking pins engaged, forms an integral part of the pressure vessel. The door swings forward to the open position on a vertical hinge attached to the forward door frame structure and provides an opening 50.7” high, 23.5” wide at the bottom, and 19.7” wide at the top. The stair assembly is attached at two hinge points on the lower door frame structure and incorporates two fold-out steps.

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Cessna Citation II Technical Manual The cabin door is held securely closed by twelve locking pins which engage sockets in the door frame structure when the interior or exterior handle is rotated to the “LOCK” position. When either handle is rotated to the “OPEN” position, the locking pins are disengaged. The overcenter locking design of the door handle linkage combined with an interior handle latching mechanism function to prevent inadvertent opening of the door, particularly from inside the airplane.

▲ UPPER LOCKING PINS

▲ LOWER LOCKING PINS

The overcenter locking design requires that either handle be rotated fully to the “LOCK” or “OPEN” position before being stowed. The latching mechanism secures the interior handle in the “STOW” position. Before rotating the interior handle to the “LOCK” or “OPEN” position, the latch must be disengaged by simultaneously pressing a release ▲ INTERIOR HANDLE RELEASE button and squeezing a trigger located on the top and backside of the handle respectively.



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The cabin door is opened from outside the airplane by unstowing the flush-mounted exterior handle, rotating it clockwise to the “OPEN” position, and pulling the door outward and forward. A security lock prevents the exterior handle from being unstowed without the appropriate key. From inside the airplane, the door is opened by disengaging the interior handle latching mechanism as previously described, rotating the handle counterclockwise to the “OPEN” position, and pushing the door outward and forward. When fully open, a springloaded door stop/catch assembly, integral to the hinge, holds the door in position. The catch is disengaged by pushing a release lever located immediately forward of the door frame on the passenger cabin sidewall.

▲ EXTERIOR CABIN DOOR HANDLE

▲ CATCH RELEASE LEVER

▲ INTERIOR CABIN DOOR HANDLE

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Cessna Citation II Technical Manual With the catch engaged, the stair assembly may be unstowed and extended. When fully extended, the stair assembly is supported by two chain assemblies attached to fittings on the fore and aft door frame structure. Chain assembly tension should be inspected regularly to confirm that each carries an equal load with weight on the steps.

▲ CABIN STAIR ASSEMBLY

Note: Uneven chain assembly tension or step deflection should be corrected to prevent damaging the stair assembly or its attachment points on the door frame structure. Before closing the cabin door, the stair assembly must first be stowed. This is accomplished from outside the airplane by lifting the stair assembly until a retraction lever at its forward hinge point engages a spring clip in the lower door frame structure. From inside the airplane, stepping on the retraction lever causes the stair assembly to pivot on its hinges into the stowed position. After stowing the stair assembly and disengaging the door catch as previously described, the cabin door is closed from outside the airplane by pushing the door aft and inward to position it in the frame, and rotating the exterior handle counterclockwise to the “LOCK” position. From inside the airplane, the cabin door is closed by pulling the door aft and inward to position it in the frame, using the assist chain provided for this purpose, and rotating the interior handle clockwise to the “LOCK” position. A rubber bumper is installed on the forward stair frame to prevent damage if the cabin door is inadvertently closed while the stair assembly is extended. A recessed area of the interior door trim panel accommodates the stowed stair assembly when the door is closed. Note: After closing the cabin door from outside or inside the airplane, the handle used must be returned to the stowed position.



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2

1

1. Upper Locking Pin Indicator Windows 2. Lower Locking Pin Indicator Windows 3. Overcenter Linkage Indicator Window

3

Five position indicator windows are provided on the interior door trim panel to visually confirm cabin door security. The upper two and lower two indicator windows permit inspection of locking pin engagement. The center indicator window permits inspection of the overcenter locking position of the door handle linkage. To facilitate inspection, the visible portion of the upper two and lower two “square” locking pins are green with white dots. With the cabin door properly secured, the white dot on each of these locking pins should be visible in its respective window. Through the center window, the alignment of two horizontal markings above the words DOOR CLOSED should be visible.

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2

1

1. Cabin Door Microswitch 2. Inflatable Rubber Door Seal Additional indication of cabin door security is provided by the amber [DOOR NOT LOCKED] annunciator. The annunciator is controlled by a microswitch, installed on the door frame structure, which is actuated by the lower forward locking pin. With electrical power applied to the airplane and the door unsecured, the door warning circuit is complete and the annunciator illuminates. When the door is properly secured, the door warning circuit is interrupted and the annunciator is extinguished. An inflatable rubber seal is installed in a retainer around the periphery of the door frame to enhance fuselage-to-door sealing. The seal is inflated by 23 ± 1 PSIG regulated engine bleed air by way of a valve which is actuated open or closed by the lower forward locking pin. With either or both engines operating and the door properly secured, the valve is actuated open and the seal is inflated. A check valve prevents the seal from deflating should loss of bleed air pressure occur. When the locking pin is disengaged, the valve is actuated closed and the seal is deflated to facilitate opening and closing of the cabin door. Additional protection against cabin door leakage is provided by a weather seal affixed to the inboard periphery of the door except in the area of the hinge which is protected by a separate rain seal.



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Cabin/Cargo Door The optional cabin/cargo door may be installed in place of the standard cabin door in the forward left side of the center section. The door is a two-section assembly of aluminum alloy construction which, when secured with all locking mechanisms engaged, forms an integral part of the pressure vessel. The upper section of the door swings upward to the open position on a full length, piano-type hinge attached to the upper door frame structure and houses the forward left passenger cabin window. The lower (airstair) section of the door swings downward to the open position on a full length, piano-type hinge attached to the lower door frame structure and incorporates three fold-out steps. When fully open, the cabin/cargo door provides an opening 50.7” high and 35.12” wide. The upper and lower sections of the door are each held securely closed by six cable-operated latch fittings which engage eccentric latch posts on the door frame structure. Though each door section is operated independently of the other, two locking pins integral to the upper door handle linkage secure both sections together. The overcenter locking design of the upper door handle linkage functions to prevent inadvertent opening of the door, particularly from inside the airplane by requiring that the interior or exterior handle be rotated fully to the “LOCK” or “OPEN” position before being stowed. The upper section is opened from outside the airplane by unstowing the flush-mounted exterior handle, rotating it clockwise to the “OPEN” position to disengage the latch fittings and locking pins, and pulling the door outward. A security lock prevents the exterior handle from being unstowed without the appropriate key. From inside the airplane, the upper section is opened by unstowing the interior handle, rotating it counterclockwise to the “OPEN” position, and pushing the door outward. A pair of gas-operated extenders assist in opening the upper section and hold it in position when fully open. A locking mechanism on the forward extender stabilizes the upper section when fully open in windy conditions. The upper section must be opened before the lower section is extended.

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Cessna Citation II Technical Manual With the upper section open, lifting the lower handle disengages the latch fittings allowing the lower section to be extended. The steps are cable-operated and fold out automatically during extension. When fully extended, the lower section is supported by two cable assemblies attached to fittings on the fore and aft door frame structure. A spring-loaded reel assembly automatically retracts the support cables when the lower section is closed. A gas-operated snubber dampens the extension rate of the lower section. Cable assembly tension should be inspected regularly to confirm that each carries an equal load with weight on the steps and that no load is carried by the snubber. Note: Uneven cable assembly tension or load on the snubber should be corrected to prevent damaging the lower section attachment or support points on the door frame structure. The cabin/cargo door is closed by lifting the lower section to position it in the frame and returning the lower handle to the closed position to engage the latch fittings. A T-handle on the aft support cable facilitates closing the lower section from inside the airplane. As the lower section is closed, the cableoperated stairs are automatically stowed against the interior trim panel. With the lower section secured, the upper section is pulled down and positioned in the frame, and the upper handle is rotated (exterior-counterclockwise, interior-clockwise) to the “CLOSE” position to engage the latch fittings as well as the locking pins which secure both sections together. Note: After closing the upper section from outside or inside the airplane, the handle used must be returned to the stowed position.



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Position indicator windows are provided on the upper and lower section interior trim panels to visually confirm positive engagement of each latch fitting. To facilitate inspection, the visible portion of a stop arm on each latch fitting shaft is green in color. With the cabin/cargo door properly secured, each stop arm should appear as a green vertical bar within its respective window. An additional indicator window on the upper section interior trim panel permits inspection of the overcenter locking position of the upper door handle linkage. Through this window, the alignment of two vertical markings between the words DOOR CLOSED should be visible. Additional indication of cabin door security is provided by the amber [DOOR NOT LOCKED] annunciator. The annunciator is controlled by five microswitches: two actuated by the upper section latch fittings linkage, one actuated by a lower section latch fitting linkage, and two actuated by the locking pins. With electrical power applied to the airplane and the upper or lower section of the door unsecured, the door warning circuit is complete and the annunciator illuminates. When both sections are properly secured, the door warning circuit is interrupted and the annunciator is extinguished. An inflatable rubber seal is installed in a retainer around the periphery of the door frame to enhance fuselage-to-door sealing. The seal is inflated by 23 ± 1 PSIG regulated engine bleed air by way of a valve which is actuated open or closed by the upper door handle linkage. With either or both engines operating and the upper section of the door properly secured, the valve is actuated open and the seal is inflated. A check valve prevents the seal from deflating should loss of bleed air pressure occur. When the upper section is unsecured, the valve is actuated closed and the seal is deflated to facilitate opening and closing of the cabin/cargo door. A “cross seal,” installed on the lower door section, seals the gap between the upper and lower section.

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Cessna Citation II Technical Manual Emergency Exit Door

▲ EMERGENCY EXIT EXTERIOR

▲ EMERGENCY EXIT INTERIOR

The emergency exit door is located in the forward right side of the center section. The door is a single-section assembly of aluminum alloy construction which houses the forward right passenger cabin window and, when properly secured, forms an integral part of the pressure vessel. The plug-type door is installed from inside the airplane and secured in place by two fixed retainers and a latch pin which engage the door frame structure. The latch pin is operated by an interior or exterior handle. To prevent inadvertent opening of the door, particularly from inside the airplane, the latching mechanism is springloaded to the closed position, and the interior handle is guarded by a plastic cover. Ground security is provided by a locking pin which prevents the latching mechanism from being operated when inserted. The locking pin incorporates a REMOVE BEFORE FLIGHT streamer.



▲ EMERGENCY EXIT INTERIOR HANDLE

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The emergency exit door is opened from inside the airplane by removing the locking pin and the plastic cover (if installed), then unstowing the interior handle and rotating it clockwise to the “OPEN” position. From outside the airplane, the flush-mounted exterior handle must be unstowed and rotated counterclockwise to the “OPEN” ▲ EMERGENCY EXIT EXTERIOR HANDLE position. With the latch pin disengaged, the top of the door is lowered until clear of the upper frame and the bottom of the door is lifted until the retainers are clear of the lower frame, thereby allowing the door to be removed. Closing the door is accomplished from inside the airplane by carefully positioning the retainers on the bottom of the door in the lower frame, pushing the top of the door into position in the upper frame, and rotating the interior handle fully-counterclockwise to engage the latch pin. A one-piece, self-inflating rubber seal is affixed to the inboard periphery of the door to enhance fuselage-to-door sealing. The seal is inflated by cabin pressure entering the seal through a series of holes. Note: After closing the cabin door from outside or inside the airplane, the door handle used must be returned to the stowed position. Hand-Held Fire Extinguishers Typically, two hand-held fire extinguishers are provided: one is secured to a bracket on the copilot’s seat frame, the other is located in the passenger cabin. Refer to the AFM emergency procedures and the instructions printed on the cylinder for proper operation. ▲ HAND-HELD FIRE EXTINGUISHER

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Cessna Citation II Technical Manual Aft Fuselage Section

The unpressurized aft fuselage section houses control cables and linkage for the tail control surfaces and the engines, and such major components as the autopilot servos, the battery, electrical power junction “J” boxes, the external power receptacle, the environmental air cycle machine (ACM), and the oxygen storage cylinder (550-0255 and after, or earlier airplanes in compliance with SB550-35-2). When optionally installed, the drag chute, components of the Freon air conditioning system, and the tailcone baggage compartment are also housed within the aft fuselage. Access to these components and to the tailcone baggage compartment is through a door attached to the lower left side of the aft fuselage structure by a piano-type hinge. The door swings downward to the open position and is secured by two hook-type latches and a key lock. ▲ TAILCONE ACCESS DOOR



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The tailcone baggage compartment consists of floor, side, and ceiling panels attached to a support structure. A hinged access panel facilitates inspection and servicing of components located forward of the baggage compartment. To reduce the risk of a baggage compartment fire, the structural elements of the baggage compartment and the inner surface of the door are treated with a flame control coating. A cargo net and tie-down straps are provided to secure the contents of the baggage compartment. The net is held in place by attachment fittings which engage anchor plates located on the floor and aft panel of the baggage compartment. Though limits vary with equipment installation, the maximum volume and load capacity of the tailcone baggage compartment range from 13 to 25 cubic feet and 200 to 500 pounds respectively. Note: Refer to the appropriate AFM and airframe placards for weight and balance limitations specific to your airplane. Indication of tailcone access door security is provided by the amber [DOOR NOT LOCKED] annunciator. The annunciator is controlled by a microswitch integral to the forward latch assembly. With electrical power applied to the airplane and the door unsecured, the door warning circuit is complete and the annunciator illuminates. When the door is properly secured, the door warning circuit is interrupted and the annunciator is extinguished. Limitations Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.

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Cessna Citation II Technical Manual Citation II Airplane Dimensions

14.80’

18.35’ 47.25’ 18.83’

8.0’

51.7’

17.59’



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Chapter 2 Engines Table of Contents Overview .................................................... 2-1 Engine Installation ...................................... 2-1 Operational Theory and Power Ratings ............................................ 2-3 Engine Description ...................................... 2-4 Low Compressor Section ....................... 2-4 High Compressor Section ...................... 2-5 Combustion Section .............................. 2-7 High Turbine Section ............................. 2-8 Low Turbine Section .............................. 2-8 Accessory Gearbox ................................ 2-9 Engine Fuel System .................................. 2-10 Engine Driven Fuel Pump .................... 2-11 Fuel Control Unit ................................. 2-12 Oil-to-Fuel Heat Exchanger ................... 2-15 Flow Divider ........................................ 2-16 Fuel Manifold and Nozzles .................. 2-18 Fuel Drains ......................................... 2-18 Ignition System ........................................ 2-19 Engine Oil System .................................... 2-22 Oil Tank ............................................. 2-22 Oil Pressure System ........................... 2-23 Scavenge System ............................... 2-25 Breather System ................................. 2-25 Engine Controls ........................................ 2-26 Throttle Levers .................................... 2-26 Thrust Reverser Levers ........................ 2-27 Engine Synchronizer Selector Switch ................................... 2-28 Automatic Fuel Shutoff Control ............ 2-28


Engines, continued Table of Contents Engine Indicating System .......................... 2-29 Fan Tachometer .................................. 2-29 Inter-Turbine Temperature Gage ........... 2-30 Fuel Flow Gage ................................... 2-32 Turbine Tachometer............................. 2-32 Engine Oil Temperature Gage .............. 2-33 Engine Oil Pressure Gage .................... 2-33 Engine Starting System ............................. 2-35 Engine Synchronizer System ...................... 2-38 Thrust Reverser System ............................ 2-40 Thrust Reverser Hydraulics ........................ 2-40 Thrust Reverser Control Valves ............ 2-41 Thrust Reverser Isolation Valves .......... 2-41 Check Valves and Restrictors .............. 2-42 Thrust Reverser Levers ........................ 2-42 Thrust Reverser Deployment ................ 2-43 Thrust Reverser Stowage ..................... 2-45 Thrust Reverser Emergency Stowage ............................ 2-46 Thrust Reverser Emergency Stowage Test ..................... 2-46 Engine Fire Detection and Extinguishing System ................................ 2-47 Thermal Detectors ............................... 2-47 Detector Control Unit .......................... 2-47 Explosive Cartridges ............................ 2-47 Extinguisher Bottles ............................ 2-48 Annunciator Switches .......................... 2-48 System Operation ............................... 2-48 System Testing ................................... 2-51 Limitations ............................................... 2-51 Emergency Procedures .............................. 2-51


This chapter will discuss the engines of the Cessna Citation II, their major components, controls, operation and indication. The airplane is powered by two JT15D-4 lightweight, mediumbypass, axial-flow, turbofan engines manufactured by Pratt & Whitney Canada, Inc. The engines generate thrust which propels the airplane and rotational torque which powers all enginedriven accessories. Indication of engine operating parameters is displayed by gages on the upper center instrument panel. Other indications are displayed on the annunciator panel. Engine Installation The engines are attached to left and right pylons formed by carry-thru beams which pass laterally through the aft fuselage. Each engine is secured to its associated pylon by two forward mounts and one aft “steady” mount. The mounts incorporate isolators which function to reduce the transmission of engine vibration through the fuselage structure. Each forward mount is secured by four bolts to the engine and one bolt to the forward carry-thru beam. Each steady mount is secured by four bolts to the aft carry-thru beam and one bolt to the engine.

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The outboard rib of each pylon is constructed of stainless steel and sealed to form a firewall and vapor barrier through which control cables, wire bundles, and various air and fluid lines are routed to and from the engines.

2

3

1

4

5 6 1. Engine Air Inlet Lip 2. Forward Cowling 3. Upper Cowling

4. Aft Cowling 5. Aerodynamic Tangs 6. Lower Cowling

The engines are enclosed by cowlings which provide aerodynamic smoothness. The forward cowling houses the engine air inlet lip and starter/generator cooling air inlet scoop and duct. The aft cowling surrounds the outer exhaust nozzle and incorporates aerodynamic “tangs” that enclose the thrust reverser actuators when installed. The upper and lower cowlings are attached to the forward and aft cowlings by quick-disconnect fasteners to facilitate removal for inspection or servicing. The upper cowling incorporates the oil filler access door. The lower cowling incorporates the starter/generator cooling air outlet and an aerodynamic fairing which encloses various engine drain lines that extend through its lower surface. To reduce the transmission of engine noise to the cabin, the forward and aft cowlings are attached directly to the engine such that they make no contact with the pylon.

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Cessna Citation II Technical Manual Operational Theory and Engine Power Ratings The power output of a jet engine, expressed in pounds of thrust, is determined by the velocity to which it is capable of accelerating a given mass of air. As a rule, a smaller volume of air must be accelerated to a higher velocity than that required to generate the same amount of thrust from a larger volume of air. In a turbojet engine, thrust is generated by compressing and combusting the entire volume of inlet air, and discharging it from the exhaust nozzle at high velocity. In a turbofan engine, a portion of the thrust generated by compressing and combusting a smaller volume of air (in the same manner as a turbojet engine) is utilized to drive a fan which generates thrust from a larger volume of air at a lower velocity. This is accomplished by dividing inlet air drawn into the engine by the fan into primary (inner) and secondary (outer) paths. The primary path directs a smaller volume of air through the engine core where it is compressed, combusted, and accelerated to a higher velocity. In the JT15D-4, approximately 42% of this “core thrust” is used drive the compressor, fan, and accessory gearbox while the remaining 58% is discharged from the inner exhaust nozzle. The secondary path directs a larger volume of non-combusted air through a concentric bypass duct surrounding the engine core and discharges it from the outer exhaust nozzle. Of the 2500 LBS thrust produced by each engine at sea level, approximately 66% is “bypass thrust” generated by the fan while 34% is generated by the core. However, because fan efficiency decreases as altitude increases, the ratio of bypass thrust to core thrust (thrust ratio) progressively reverses between sea level and approximately 40,000’ MSL. The relationship between mass airflow through the bypass duct and mass airflow through the engine core determines the engine’s “bypass ratio.” Of the total 77.8 LBS/SEC mass air flow (Wa) developed by the fan, 56.6 LBS/SEC is directed through the bypass duct while 21.2 LBS/SEC is directed through the core thereby producing the engine’s 2.7:1 bypass ratio. Unlike thrust ratio, bypass ratio remains constant regardless of factors which influence air density such as altitude, temperature, and airspeed.

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Power ratings are also based on the engine’s sea level static output in ISA conditions. The power ratings and leading particulars given for the JT15D-4 in the table below are for training purposes only and are not DOT approved. Operating Condition Takeoff Max. Continuous Max. Climb Max. Cruise

Net Thrust LBS 2500** 2375* 2375 2345

Specific Fuel LBS/HR/LBS THRUST 0.562 0.556 0.556 0.555

* Available to 15.0°C (59.0°F) ** Available to 15.0°C (59.0°F) time limited to 5-minutes

Engine Description This section is intended to provide a basic description of the JT15D-4 engine. Each engine is functionally divided into independent high and low pressure sections which rotate on concentric, bearing supported shafts. The outer, high pressure (N2) shaft supports the high turbine and high compressor, associated primarily with the generation of core thrust. The inner, low pressure (N1) shaft supports the low turbines, low compressor (fan), and booster stage, associated primarily with the generation of bypass thrust. Both shafts rotate in a clockwise direction at different speeds. In that no mechanical link exists between the shafts, the engine is classified as a free-turbine. The engine is further divided into six principle sections as follows: Low Compressor Section The low compressor section consists of the fan, booster stage, and associated stator assemblies. The fan is installed on the forward end of the N1 shaft and incorporates 28 blades that function to induct and compress all inlet air entering the engine. The induction of inlet air is optimized by the aerodynamic shape of the nose cone attached to the front of the fan. Immediately aft of the fan, compressed inlet air is divided into concentric primary (inner) and secondary (outer) paths.

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Cessna Citation II Technical Manual The primary path directs a smaller volume of air through a row of 38 core inlet guide vanes (stators) to the booster stage. The booster stage, installed aft of the fan on the N1 shaft, is a single-stage, 43-blade, axial compressor that functions to increase core inlet pressure. After passing through the booster stage, core inlet air is directed through a row of 37 stators to the high compressor. The secondary path directs a larger volume of air to the bypass duct through two staggered rows of 66 stators each. This air passes through the bypass duct without undergoing combustion, after which is it discharged from the outer exhaust nozzle as bypass thrust. A T2 thermocouple probe senses inlet air temperature prior to induction by the fan. A T2.6 thermocouple probe senses bypass air temperature within the outer exhaust nozzle. In combination, these probes constitute the T1 signal producing elements of the inter-turbine temperature (ITT) sensing system. Protection against engine air inlet icing is provided by bleed air heating of the nose cone, core inlet stators, and T2 thermocouple. Refer to Chapter 10 for a complete description of engine ice protection systems. High Compressor Section The high compressor section consists of a single-stage centrifugal impeller, stator assembly, impeller shroud, and an array of 24 diffuser pipes and deflector vanes. The centrifugal impeller is installed on the forward end of the N2 shaft and incorporates 32 blades (16 full/16 splitter) that function to accelerate core inlet air. This air is directed to the impeller through a row of 23 stator vanes. The impeller shroud, diffuser pipes, and deflector vanes are installed within the gas generator case. The impeller shroud contains the accelerated core inlet air and directs it radially to the diffuser pipes. The diffuser pipes function primarily to decelerate this air thereby maximizing its pressure. Their secondary function is to restore axial flow from radial flow by redirecting the air through 90°. Exiting the diffuser pipes, compressor discharge air (P3) passes through the deflector vanes. The angular orientation of the deflector vanes optimizes P3 air flow for delivery to the combustion chamber.

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JT15D-4 General Arrangement N1 TACH GENERATOR PAD

Left Front View

OIL FILLER NECK AND DIPSTICK

ANTI-ICING VALVE

OIL SCAVENGE LINE RETURN TO TANK

STARTER GENERATOR PAD N2 TACH GENERATOR PAD

OIL PUMP ASSEMBLY

BEARING SCAVENGE LINE

P3 TO ANTI-ICING VALVE

ACCESSORY GEARBOX

Right Front View

T2 PROBE

NOSE CONE LOW COMPRESSOR FAN

OIL PRESSURE LINE TO NO 3 1/2 AND 4 BEARINGS P3 LINES FOR FCU AND T1 PROBE ANTI-ICING

MAIN OIL PRESSURE BOSS P3 CABIN BLEED OIL PRESSURE LINE FUEL FILTER HOUSING

2-6

OIL FILTER BYPASS VALVE

OIL-TO-FUEL HEAT EXCHANGER OIL PRESSURE REGULATING VALVE

OIL FILTER HOUSING


ACCESSORY GEARBOX DRAIN


Cessna Citation II Technical Manual Combustion Section The combustion section consists of the combustion chamber, fuel nozzles, and igniters. The annular, reverse-flow combustion chamber, installed within the gas generator case, provides an area for the mixture and ignition of air and fuel, and directs combustion gas flow to the turbines. Concentric outer and inner liners extend forward from the aft (domed) end of the combustion chamber. Twelve adapters are positioned around the domed end for insertion of the fuel nozzles. Two adapters are positioned at 5 and 7 o’clock on the outer liner for insertion of the spark igniters. At the forward end of the combustion chamber, a large exit duct and small exit duct join the outer and inner liners respectively. P3 air enters the interior of the combustion chamber through a series of perforations in the inner and outer liners. Metered fuel from the fuel control unit enters the interior of the combustion chamber through the fuel nozzles. The air/fuel mixture is initially ignited by the spark igniters during engine start, after which combustion is self sustaining under normal engine operating conditions. The shape, size, and location of the perforations, as well as the location of the fuel nozzles and spark igniters, provide the best air/fuel ratio for engine starting and sustained combustion. The exit ducts redirect the combustion gas flow inward then aft through 180° to the turbine inlet, thus the term “reverse-flow.” Air not used in the combustion process, referred to as secondary air, is used by the engine for ice protection, hot section cooling, bearing compartment sealing, and fuel control. To ensure even temperature distribution and prevent flame contact with the interior walls of the combustion chamber, cooling rings direct a layer of P3 air over these surfaces during combustion. To prevent exposing the high compressor to excessive combustion gas flow temperatures, the large and small exit ducts incorporate heat shields through which P3 air is also directed for cooling purposes.

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High Turbine Section The high turbine section consists of the single-stage high turbine, segmented shroud assembly, and turbine inlet guide vane assembly. The high turbine is installed on the aft end of the N2 shaft and incorporates 71 blades. The segmented shroud assembly provides an ideal blade tip running clearance to maximize turbine efficiency. The turbine inlet guide vane assembly incorporates 14 stator vanes that direct combustion gas flow against the blades at an optimal angle and speed. The stator vanes feature cored passages through which P3 air is directed for cooling purposes. Air enters through inlet ports at each vane’s root and is evacuated into the gas flow path through outlet ports at each vane’s trailing edge. The high turbine extracts energy from the combustion gas flow to generate the rotational torque that drives the N2 shaft, high compressor, and engine-driven accessories. The energy not absorbed by the high turbine is directed to the low turbine section. Note: Engines in compliance with P&WC SB7293 feature “D.S.” high turbine blades produced using a directionally-solidified casting process. Low Turbine Section The low turbine section consists of two low turbines and their associated turbine inlet stator vane assemblies. The low turbines (referred to as 2nd and 3rd stage turbines) are installed in tandem on the aft end of the N1 shaft. The 2nd stage turbine incorporates 61 blades. The 3rd stage turbine incorporates 55 blades. Both feature shrouded blade tips to maximize turbine efficiency. Each turbine inlet stator vane assembly incorporates 43 stator vanes that direct combustion gas flow against the turbine blades at an optimal angle and speed. The low turbines extract energy from the combustion gas flow to generate the rotational torque that drives the N1 shaft, fan, and booster stage. After driving the turbines, the combustion gas flow is discharged from the inner exhaust nozzle as core thrust.

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Cessna Citation II Technical Manual Six thermocouple probes sense combustion gas flow temperature within the inner exhaust nozzle. In combination, these probes constitute the T6 signal producing elements of the ITT sensing system. Accessory Gearbox The accessory gearbox (AGB), located on the lower intermediate case of the engine, houses the gearing and provides support for all engine-driven accessories except the N1 tachometer generator. The AGB consists of a cast magnesium alloy housing and rear cover. The housing provides bearing supported gear shafts which drive the centrifugal air/oil separator and all externally-mounted engine accessories. External mounting pads are provided for the starter/generator, engine-driven fuel pump, hydraulic pump, oil pump assembly, and N2 tachometer generator. The AGB main shaft engages the starter/generator directly, and is linked to the N2 shaft by a vertical tower shaft. The tower shaft is splined at each end to engage upper and lower bevel gears. The upper bevel gear meshes with a bevel gear fitted to the N2 shaft. The lower bevel gear meshes with a bevel gear fitted to the AGB main shaft. During engine start, the tower shaft transmits starter/generator rotation to the N2 shaft to turn the engine. During engine operation, the tower shaft transmits N2 shaft rotation to the AGB main shaft to turn the starter/generator and other engine-driven accessories. A second vertical tower shaft engages the N1 tachometer generator at its upper end. A bevel gear fitted to the lower end of this shaft meshes with a bevel gear fitted to the N1 shaft.

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Engine Fuel System The engine fuel system functions to regulate fuel flow to the engines. Major components for each engine include an enginedriven fuel pump, fuel control unit, oil-to-fuel heat exchanger, flow divider, fuel manifold, and 12 dual-orifice type nozzles. Engine-Driven Fuel Pump One single-stage, positive3 displacement, gear-type, engine-driven fuel pump is 1 mounted on and driven by the accessory gearbox of each engine. Each pump functions to supply clean fuel under high pressure to the fuel 2 control unit (FCU) of its associated engine, and motive flow pressure to its associated 1. Fuel Pump Body primary ejector pump. Typical 2. Fuel Filter Housing (inlet) pump capacity is 3935 PPH 3. Fuel Filter Housing (outlet) and 580 PSI at 100% N2. Fuel entering the pump passes through a 74 micron, wire mesh inlet filter before entering the pump chamber. The inlet filter is self relieving at 9 to 12 PSID should it become obstructed. Exiting the pump chamber, high pressure fuel passes through a 10 micron, nonmetallic, disposable outlet filter en route to the FCU. A spring-loaded, ball-type, bypass valve, preset to open at 40 to 60 PSID, allows high pressure fuel to bypass the outlet filter element should it become obstructed. Pump chamber inlet pressure is maintained by a jet pump nozzle located upstream of the inlet filter. When pump pressure exceeds metered pressure, a portion of FCU bypass fuel is returned to the pump chamber inlet through this nozzle. Fuel Control Unit (FCU) The fuel control unit (FCU) is mounted on and driven by the engine-driven fuel pump through an integral splined coupling shaft. The FCU determines the correct fuel schedule to produce desired engine power in response to THROTTLE lever movement. Major components include the fuel metering section, computing section, and N2 governing section. Other associated components include a T2 temperature compensator and step modulator.

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Fuel Metering Section The fuel metering section includes a pressure relief valve, metering valve, and bypass valve. Unmetered fuel pressure (P1) is supplied to the FCU from the engine-driven fuel pump. The pressure relief valve functions to prevent system overpressurization by opening when P1 exceeds 1260 PSID, thereby returning excess fuel pressure (P0) to the pump inlet. The metering valve is essentially a tapered needle valve positioned by the FCU computing section to regulate fuel flow to the engine. The bypass valve functions to maintain a constant 15 to 24 PSI differential (delta P) between P1 and metered fuel pressure (P2) across the metering valve orifice. Because bypass valve position and fuel flow are functions of metering valve position, the bypass valve responds to increased fuel flow through the metering valve by decreasing the amount of P0 returned to the pump inlet. When fuel flow through the metering valve is decreased, the bypass valve responds by increasing the amount of P0 returned to the pump inlet.

3

1

4

2

1. Fuel Control Unit 2. Fuel Outlet (P2)

3. Throttle Lever Input 4. Fuel Bypass Line

Minimum flow through the metering valve is factory adjusted to approximately 155 to 160 pounds per hour (LBS/HR) to provide the correct amount of fuel flow for engine starting and to prevent engine failure during rapid deceleration.

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Computing Section The computing section includes a pneumatic governor bellows and an acceleration bellows connected to a common torque tube linkage that positions the metering valve. In combination, these components control acceleration, deceleration, and steady engine operating states in response to pneumatic signals supplied by the N2 governing section. The body of the computing section is divided into two chambers separated by the governor bellows. The upper chamber is supplied with Py (governing) pressure; the lower chamber is supplied with Px (enrichment) pressure. As such, the exterior of the governor bellows is exposed to Py pressure and the interior of the governor bellows is exposed to Px pressure. The acceleration bellows is sealed at absolute pressure and located within the lower chamber. The force of Px pressure acting on the acceleration bellows is cancelled by the force of Px pressure acting on the same area of the interior of the governor bellows. This “area of cancellation” ensures that any change in Py pressure will have a greater effect on metering valve position than an equal change in Px pressure. Because Px and Py are derived from P3 (compressor discharge air), and because P3 is proportional to N2 and air density, the acceleration bellows provides an absolute pressure reference to compensate for reduced air density at higher altitudes. An air filter is installed in-line to prevent foreign material present in the P3 air from entering the FCU.

N2 Governing Section The N2 governing section is associated primarily with engine speed setting and speed control. Major components include a speed scheduling cam, governor flyweights, feedback springs, a governor lever, enrichment lever, enrichment valve, and backup valve. The speed scheduling cam sets feedback spring resistance in response to THROTTLE lever position. The smaller of the two feedback springs resists enrichment lever movement; the larger spring resists governor lever movement. The enrichment valve is spring-loaded open and regulates the introduction of P3 air into the body of the N2 governing section in response to enrichment lever movement. Movement of the governor lever regulates the bleeding of Py air to the atmosphere through vents in the body of the N2 governing section.

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Cessna Citation II Technical Manual The flyweights are driven by the FCU coupling shaft such that their lifting force varies with engine speed. Above approximately 85% N2, when flyweight lifting force is sufficient to overcome the resistance of the smaller feedback spring, the enrichment lever moves the enrichment valve toward closed and less P3 air enters the FCU. As flyweight lifting force increases, the enrichment lever makes contact with the governor lever pivoting about the same point. When flyweight lifting force is sufficient to overcome the resistance of the larger feedback spring, the enrichment lever moves the governor lever toward open and more Py pressure is bled to the atmosphere. The corresponding decrease in Py pressure at the computing section moves the metering valve toward closed and reduces fuel flow to the engine. Conversely, an increase in Py pressure at the computing section moves the metering valve toward open and increases fuel flow to the engine. When Px and Py are simultaneously decreased during decceleration or increased during acceleration, Py pressure will have a greater effect on metering valve position than Px pressure. Following a change in power setting, the corresponding “lag” in Px and Py pressure change at the computing section regulates the transition from the previous fuel flow rate through the metering valve to that which will produce the desired engine power. As such, an excessively lean condition or excessively rich condition are prevented. When the sum of all forces acting on the position of the metering valve are in equilibrium, fuel flow and N2 remain essentially constant. The backup (overspeed) valve is spring-loaded closed and normally seals a secondary governing air pressure (Py) orifice plumbed to the computing section. When N2 exceeds the selected speed by approximately 9%, the valve is forced open by the governor lever and more Py pressure is bled to the atmosphere. The corresponding rapid decrease in Py pressure at the computing section moves the metering valve toward closed and reduces fuel flow to the engine, thereby preventing N2 overspeed.

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1. T2 Temperature Compensator The T2 temperature compensator functions to maintain consistent acceleration response by modifying the engine’s fuel flow schedule for variations in outside air temperature. To accomplish this, the unit incorporates a bimetallic disk temperature probe that senses ambient air temperature, and a metering pin that regulates the bleeding of Px pressure to the atmosphere through vents in the compensator housing. The metering pin is spring-loaded open and moves toward closed when the bimetallic disks expand under increasing temperature. As such, Py pressure at the computing section is increased and decreased inversely proportional to bypass air temperature.

2 1

2. Step Modulator The step modulator functions to maintain consistent acceleration response by maintaining a constant P3 supply pressure to the FCU. To accomplish this, the unit incorporates an electrically-actuated restrictor orifice that is energized open primarily when the engine ice protection system is activated. When open, more P3 air than normal is supplied to the FCU to compensate for that consumed in the bleed air heating of induction air inlet components. The restrictor orifice is also energized open when the engine ignition system is activated. When closed, the supply of P3 air to the FCU returns to normal.

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Cessna Citation II Technical Manual Oil-to-Fuel Heat Exchanger The oil-to-fuel heat exchanger functions to reduce engine oil temperature and prevent ice formation in the fuel. 4 1

2

3

1. Fuel Inlet 2. Fuel Outlet

3. Oil Inlet Manifold 4. Oil Outlet Manifold

The unit consists primarily of a cylindrical outer “shell” and an inner “core.” The core is formed by 85 transfer tubes which run axially through the interior of the shell. Fuel from the FCU enters the core at the forward end of the unit and exits at the aft end of the unit en route to the flow divider. Oil from the pressure pump enters the shell through an external inlet port and manifold that directs the oil to the aft end of the unit. As the oil flows forward through the interior of the shell, its heat is transferred to the lower temperature fuel flowing in the opposite direction through the core. An array of baffles, alternately positioned on opposite sides of the shell interior, repeatedly diverts the oil laterally over the core to maximize heat transfer. Upon reaching the forward end of the unit, the oil exits the shell through an external manifold and outlet port en route to the flow divider.

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Flow Divider The flow divider integrates a minimum pressurizing and flow divider valve, fuel cut-off valve, spill valve, and dump valve in a single unit. The minimum pressurizing and flow divider valve maintains a sufficient minimum system pressure and divides metered fuel flow between the primary and secondary fuel manifolds. This plunger-type valve is spring-loaded closed and opened by fuel pressure. When fuel pressure reaches approximately 75 PSID, the valve opens to an internal passage that supplies the primary manifold. When engine speed increases above approximately 60% N2, increasing fuel pressure further opens the valve to an internal passage that supplies the secondary manifold also.

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8

6

2 5 4

1. 2. 3. 4.

Flow Divider Fuel Inlet (P1) Bypass Outlet P3 Line

3

5. 6. 7. 8.

Primary Outlet Drain Outlet Secondary Outlet Throttle Linkage

The rotary-type fuel cut-off valve is positioned by the THROTTLE lever and functions to control fuel flow between the FCU and the flow divider. When the THROTTLE lever is positioned to “OFF,” the valve is fully-closed and restricts fuel flow. When positioned to “IDLE” the valve is fully-open.

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Cessna Citation II Technical Manual Engine Fuel System Schematic

© PCW

During engine start, when the THROTTLE lever is initially positioned to “IDLE,” the fuel pressure supplied to the primary manifold is regulated by the spill valve as a function of P3 pressure. When P3 pressure is low, the spill valve is fully-open and fuel in excess of that required for engine starting is returned to the pump inlet. As engine speed increases to idle, the corresponding increase in P3 pressure moves the spill valve toward closed and progressively less fuel is returned to the pump inlet. When P3 pressure reaches approximately 30 PSI, the spill valve is fully-closed. During engine shutdown, when the THROTTLE lever is positioned to “OFF” and the cut-off valve is fully-closed, the minimum pressurizing and flow divider valve returns to its springloaded closed position. As fuel pressure falls below approximately 5 PSI, the dump valve opens and residual fuel is drained from the primary and secondary manifolds to an EPA canister mounted below the engine. During engine operation, the EPA canister is charged with air tapped from the bypass duct and the residual fuel is returned to the tank. Each return line incorporates a check valve to prevent backflow from the fuel tank to the canister. Residual fuel in excess of the canister’s 15-ounce capacity is dumped overboard through a drain tube that extends through the lower engine cowling.

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Fuel Manifold and Nozzles The fuel manifold consists of primary and secondary transfer tubes, and twelve manifold adaptors which carry metered fuel from the flow divider to the fuel nozzles. Each manifold adapter independently supplies primary and secondary fuel flow to its associated nozzle. The dual-orifice type nozzles are enclosed in sheaths which extend into the combustion chamber. Atomized fuel is introduced to the combustion chamber through the primary orifice when the flow divider is supplying the primary manifold, or through both the primary and secondary orifices when the flow divider valve is supplying the primary and secondary manifolds. Each sheath is slotted to permit the passage of secondary air for nozzle cooling and improved atomization. Fuel Drains 1 Two fuel drains, installed in the 6 o’clock position of the gas generator case below the combustion chamber, ensure that all residual fuel which accumulates in this area is drained overboard after engine shut down. Each drain is 3 fitted with a transfer tube that 2 carries residual fuel to a common drain valve installed on 1. Combustion Drain Line the outer bypass duct. The 2. Overboard Breather Tube drain valve is spring-loaded to 3. EPA Canister the open position and held closed during engine operation by P3 air pressure. When open, residual fuel is routed to an overboard breather tube that extends through the lower engine cowling. Residual fuel is also drained from the inner exhaust nozzle through an overboard drain line that extends through the lower engine cowling.

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Cessna Citation II Technical Manual Ignition System

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1. Spark Igniter 2. Ignition Exciter Box 3. High Tension Wire The ignition system functions primarily to supply the electrical spark necessary to initiate combustion during engine start. Each engine’s ignition system operates independently of the other and consists of two engine mounted ignition exciters, twohigh tension cables, and two spark igniters. Nominal 28 VDC power is supplied to the system; however, the system is operable between 9 and 30 VDC. Note: On airplanes 550-0470 and earlier not incorporating SB550-74-01 (P&WC SB 7178), each engine is equipped with a single ignition exciter.

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The ignition exciters are sealed boxes which house the components that transform DC voltage into high-energy electrical pulses that are supplied to the spark igniters. The pulses are supplied to the spark igniters through high tension cables which are encased in flexible metal braiding for protection from heat. On engines with dual exciters, each is a single-output type that supplies its associated igniter only. On engines with a single exciter, each is a dual-output type that supplies both ignitors. In either installation, the system is designed such that one igniter will remain operable should failure of the opposite igniter occur. The spark igniters are located at 5 and 7 o’clock positions on the gas generator case and extend inward to the interior of the combustion chamber through its outer liner. Each igniter is comprised of a threaded outer casing and central electrode separated by a semiconducting material. When the ignition exciter is energized, a capacitor is progressively charged until sufficient voltage is produced to ionize the gap between the positive electrode and the negative casing. When this occurs, capacitor voltage is discharged across the gap in the form of a high-energy spark. To extend service life and reduce the risk of failure, the igniters are cooled by secondary air. Each engine’s ignition system is independently controlled by a corresponding two-position (ON/NORM) LH or RH IGNITION switch on the lower left instrument panel. Under most operating conditions, these switches should remain in the “NORM” (off) position. In this position, ignition system activation occurs automatically during engine start and engine ice protection system operation.

▲ IGNITION SWITCHES

Independent indication of left (LH) or right (RH) ignition system operation is provided by green lights located above each IGNITION switch. When power is being supplied to either system, the corresponding light will be illuminated regardless of operating condition.

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Cessna Citation II Technical Manual Ignition System Schematic

© PCW

During engine start, ignition system activation is carried out by a series of relays in the start control circuit which supply 28 VDC hot battery bus power to the appropriate exciter(s) through a 7.5-amp IGNITION circuit breaker located in the aft fuselage electrical power junction “J” box. When the ENGINE ANTI-ICE switches or IGNITION switches are set to the upper (on) position, each exciter receives 28 VDC power through its associated 15-amp LH IGN or RH IGN circuit breaker on the left CB panel, correspondingly supplied by the left or right main bus. Note: The position of the LH IGN and RH IGN circuit breakers has no effect on ignition system operation during the engine starting sequence. Note: The IGNITION switches should be positioned to “ON” during takeoff, approach and landing, and turbulent air penetration.

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Engine Oil System The engine oil system consists of pressure, scavenge, and breather sub-systems. The pressure system pumps oil from the tank to lubricate and cool all internal moving engine components by way of cored passages and transfer tubes. The scavenge system returns oil to the tank by way of gravity drains, transfer tubes, cored passages, and scavenge pumps. The breather system functions essentially to vent air pressure from the scavenge oil system. An oil-to-fuel heat exchanger is provided to regulate engine oil temperature. Engine oil temperature and pressure indication is provided by the independent OIL °C and OIL PSI gages on the upper center instrument panel. Total oil capacity is 2.08 US gallons, of which 1.25 US gallons are usable. Oil Tank The oil tank is integral to the intermediate case. The tank is serviced through a filler neck located on the outboard intermediate case, and is accessible through a hinged door on the upper cowling. A dipstick, integral to the filler cap, is provided for oil level inspection. The dipstick is calibrated to indicate the approximate ▲ OIL FILLER NECK tank quantity in US quarts when the engine is hot, therefore, oil level should be inspected within ten minutes of engine shutdown. Note: The engine should be serviced with approved synthetic oils listed in the most current revision of P&WC SB 7001.

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Cessna Citation II Technical Manual Engine Oil System Schematic

© PCW

Oil Pressure System Engine oil pressure is developed by a gear-type pressure pump that is mounted on and shaft-driven by the accessory gear box (AGB). The pressure pump is integral to an oil pump assembly that also houses the scavenge pumps. From the tank, oil is drawn through a filter screen to the inlet side of the pump. Engine ▲ OIL PUMP ASSEMBLY oil pressure developed by the pump is routed through an external transfer tube to an assembly that houses a pressure regulating valve, check valves, oil filter, and filter bypass valve. The spring-loaded, piston-type pressure regulating valve prevents operating pressure from exceeding 73 ± 6 PSI by diverting excess oil pressure back to the pump inlet. Regulated oil pressure is directed through the oil-to-fuel heat exchanger en route to the oil filter.

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The oil filter utilizes a 10 micron, cartridge-type element which prevents particles of foreign matter present in the oil from circulating through the engine. Should the filter element become partially restricted or blocked, an integral spring-loaded bypass valve permits continued though unfiltered oil flow to the engine. The bypass valve is opened when the differential between filter inlet and outlet pressure reaches approximately 15 to 24 PSID.

1 4 3 2

1. 2. 3. 4.

Oil Filter Housing Fuel Filter Housing Oil Pressure Regulating Valve Oil-to-Fuel Heat Exchanger

The check valves open at 5 PSI when the engine is operating to permit oil flow from the pressure pump to the engine, and from the pressure regulating valve to the pump inlet. When the engine is not operating, the check valves function to prevent gravity draining of oil from the tank to the bearing cavities (primarily after engine shutdown), and allow the oil filter to be removed for inspection without draining the tank. From the filter, pressure oil flow is divided into three paths. The first path directly supplies the AGB through an internal transfer tube. The second path supplies the N1 shaft front (#1) bearing, the N2 shaft front and rear (#2 and #3) bearings, and the tower shaft bevel gears through an external transfer tube and internal passages. The third path supplies the N1 shaft intermediate and rear (#3½ and #4) bearings through an external transfer tube and internal passages. Labyrinth air seals are used to confine pressure oil to the bearing compartments.

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Cessna Citation II Technical Manual Scavenge System The scavenge system incorporates a pair of gear-type pumps that are integral to the oil pump assembly and shaft-driven by the AGB. The gearshaft that drives the oil pressure and scavenge pumps also drives the N2 tachometer generator. Scavenge oil gravity drains from the #1, 2, 3, and #3½ bearings to the AGB assisted by secondary air pressure from the bearing compartment labyrinth seals. Scavenge oil from the #4 bearing is drawn through an external transfer tube by the smaller of the two scavenge pumps. Scavenge oil from the AGB is drawn through a filter screen by the larger scavenge pump. From these pumps, scavenge oil is returned to the tank through an external transfer tube and internal passage. Breather System The breather system incorporates a centrifugal impeller driven by the AGB main shaft which also drives the starter/generator. The impeller, also referred to as the centrifugal breather, separates air from the AGB scavenge oil by centrifugal force. In operation, air is drawn radially inward while oil is thrown radially outward. Once separated, the relatively oil-free air is vented to the atmosphere through an overboard breather tube that extends through the lower engine cowling.

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Engine Controls Primary engine control is provided by left and right THROTTLE levers located on the center pedestal between the pilot’s and copilot’s seats. When thrust reversers are optionally installed, thrust reverser control levers are mounted “piggyback” on the THROTTLE levers. When the engine synchronizer system is optionally installed, an ENGINE SYNC selector switch is also located on the center pedestal.

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1

1. Throttle Levers 2. Release Triggers

3. Thrust Reverser Lever 4. Engine Sync Switch

Throttle Levers The THROTTLE levers function to set engine speed and shut down the engines. To accomplish this, THROTTLE lever movement is transmitted to its associated FCU by “controlex” cables and bellcrank assemblies. The FCU, in turn, is mechanicallylinked to the flow divider cut-off valve by an interconnect rod. When positioned forward of IDLE, the THROTTLE lever sets the FCU to maintain the selected engine speed. When positioned to “IDLE,” the FCU is set to provide approximately 49% N2 minimum. The IDLE position also provides the correct fuel schedule during engine start. When positioned to “OFF,” the flow divider cut-off valve restricts fuel flow to the engine.

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Cessna Citation II Technical Manual A detent prevents inadvertent THROTTLE lever movement from “IDLE” to “OFF.” To clear this detent, spring-loaded triggers on the outboard sides of the levers must be lifted. Throttle lever friction is adjusted using an INCREASE FRICTION knob located on the right side of the control pedestal. Rotating this knob clockwise increases throttle lever friction; counterclockwise rotation decreases friction. Thrust Reverser Levers The thrust reverser (TR) levers are mounted on the THROTTLE levers and function to control thrust reverser operation. To accomplish this, independent left and right, deploy and stow microswitches, located within the center pedestal, are actuated by TR lever movement. The deploy switches are actuated when the TR levers are pulled aft; the stow switches are actuated when the TR levers are pushed forward. When these switches are actuated, corresponding deploy or stow solenoids integral to the thrust reverser control valves are energized, thereby positioning these valves to permit thrust reverser operation as selected. To prevent inadvertent thrust reverser operation, the THROTTLE levers must be set to “IDLE” before the TR levers can be moved. During thrust reverser transition, the throttle feedback system holds the throttle linkage at idle until the reverser doors are fully-deployed or fully-stowed. A pair of locking solenoids prevents aft movement of the TR levers to increase reverse thrust until the reverser doors are fully-deployed. TR lever stops limit reverse thrust N1 to approximately 92% in sea level ISA conditions. Refer to the thrust reverser section of this chapter for functional detail. Warning: The THROTTLE levers should not be restrained manually or by lever friction during takeoff as this will prevent the throttle feedback system from automatically returning the throttle linkage and THROTTLE levers to “IDLE” should inadvertent thrust reverser deployment occur.

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Engine Synchronizer Selector Switch The ENGINE SYNC selector switch is located aft of the right THROTTLE lever on the center pedestal and functions to control engine synchronizer system operation. When positioned to “FAN,” the system synchronizes left (master) and right (slave) fan RPM (N1). When positioned to “TURB,” the system synchronizes turbine RPM (N2). When positioned to “OFF,” the system is deactivated. Refer to the engine synchronizer section of this chapter for functional detail. Automatic Fuel Shutoff Control The automatic fuel shutoff control is a mechanically-operated system that functions to shutdown the engine in the event of #4 bearing failure or N2 shaft decoupling. To accomplish this, the system incorporates an actuator rod that is linked by bellcrank to a piston-type valve. Should 0.070” rearward, axial displacement of the N2 shaft occur, N2 shaft contact with the rod will actuate the valve thereby cutting off the supply of fuel to the nozzles.

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Cessna Citation II Technical Manual Engine Indicating System The engine indicating system provides visual indication of critical engine operating parameters and provides annunciation where these parameters are exceeded. Each engine’s indicating system operates independently of the other and includes an N1 (fan) tachometer generator, inter-turbine temperature (ITT) sensing system, fuel flow transmitter, N2 (turbine) tachometer generator, oil pressure transmitter, and oil temperature sensor which supply input signals to the corresponding left or right channel of their associated gages. Other engine indications are displayed on the annunciator panel. The primary engine gages are arranged in a horizontal row on the upper center instrument panel. From left to right these are FAN % RPM, ITT, FUEL FLOW, TURBINE % RPM, OIL TEMP, and OIL PRESS. With the exception of the TURBINE % RPM digital readout, indication is provided by the position of independent left (L) and right (R) vertical tape bars relative to a graduated, vertical instrument scale. The FAN % RPM gage also features a digital readout. Where appropriate, colored markings on the instrument scale denote operating ranges and limitations corresponding to the indicated parameter. Fan Tachometer The FAN tachometer provides indication of N1 shaft rotational speed in % RPM, with 100% equaling approximately 15,904 RPM. The instrument scale is graduated in 10% increments between 20 and 90%, and 2% increments between 90 and 110%. Numerical values are marked at each 10% increment between 20 and 110%. Green markings between 25 and 104% denote the normal indicating range. Red warning lines at 104% denote the maximum N1 limitation (16,540 RPM). Input signals are supplied to each channel of the gage by its associated left or right N1 tachometer generator located on the upper intermediate case.

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The digital readout on the lower instrument face provides supplemental indication of FAN % RPM. The readout displays in tenths between 00.0 and 99.9%. A mechanical counter, mounted above the gage, may be set to reference a target power setting. Each channel of the gage receives 28 VDC power through its associated 2-amp LH FAN SPEED or RH FAN SPEED circuit breaker, correspondingly supplied by the left or right main bus. When power is removed from either channel, the affected digital readout will be extinguished. The vertical tape indicating portion of the gage is essentially “self-generating” and will remain operable above 50 N1 even when power is removed from the instrument. No OFF flag is provided. Inter-Turbine Temperature Gage The ITT gage provides indication of inter-turbine temperature in °C. The instrument scale is graduated in 100° increments between 200 and 800°C, and 10° increments between 500 and 800°C. Numerical values are marked at each 100° increment between 200 and 800°C. Green markings between 150 and 680°C denote the normal indicating range. Yellow markings between 680 and 700°C denote the caution range. Red warning lines at 700°C denote the maximum ITT limitation during engine start, which is time limited to 2-seconds. Note: During normal engine start, ITT indication should not exceed 500°C. Note: ITT indication exceeding 500°C during normal engine start should be investigated for cause. Each channel of the gage receives 28 VDC power through its associated 2-amp LH ITT SPEED or RH ITT circuit breaker, correspondingly supplied by the left or right main bus. A red OFF warning flag will appear at the top of the vertical scale whenever power is removed from the corresponding channel of the gage.

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Cessna Citation II Technical Manual Input signals are supplied to each channel of the gage by its associated left or right ITT sensing system which computes a “simulated” inter-turbine temperature (T5) from signals produced by eight thermocouple probes. One thermocouple probe senses inlet air temperature (T2) prior to induction by the fan. Another senses bypass air temperature (T2.6) within the outer exhaust nozzle. In combination, the T2 and T2.6 probes constitute the T1 signal producing elements of the system. Six probes sensing combustion gas flow temperature within the inner exhaust nozzle constitute the T6 signal producing elements of the system. The bimetallic (alumel/chromel) thermocouples generate a mild electrical current which varies in response to the temperature sensed. The T2 and T2.6 probes contain three thermocouples each, wired in series. The T6 probes contain one thermocouple each, wired in parallel to compensate for uneven heat distribution and to obtain an average temperature reading. Each probe is connected to its associated T1 or T6 wiring harness. The wiring harnesses are also bimetallic, having an alumel negative circuit and chromel positive circuit. In operation, the difference between T2.6 and T2 represents the temperature rise in non-combusted airflow as it passes through the bypass duct. Because the three thermocouples in each of these probes are wired in series, the temperature rise is multiplied by three to produce the T1 signal. The T1 signal and composite T6 signal are corrected for sampling errors by a variable resistor. The variable resistor is set to a required trim value during final engine acceptance checks and is sealed at that setting. The sum of the corrected signals, computed from the formula T1 + T6 = T5, is displayed by the ITT gage.

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Fuel Flow Gage The FUEL FLOW gage provides indication of engine fuel flow in LBS/ HR. The instrument scale is graduated in 100 pound per hour increments between 0 and 2000 LBS/HR. Numerical values are marked at each 200 pound per hour increment between 0 and 2000 LBS/HR. Input signals are supplied to each channel of the gage by its associated left or right fuel flow transmitter. Each channel of the gage receives 28 VDC power through its associated 2-amp LH FUEL FLOW or RH FUEL FLOW circuit breaker, correspondingly supplied by the left or right main bus. A red OFF warning flag will appear at the top of the vertical scale whenever power is removed from the corresponding channel of the gage. Refer to Chapter 3 for a complete description of the fuel flow indicating system. Turbine Tachometer The TURBINE tachometer provides digital indication of high pressure (N2) shaft rotational speed in % RPM, with 100% equaling approximately 32,760 RPM. The readout displays in tenths between 00.0 and 99.9%, and single-digits above 100%. A red light is located below each digital readout. When N2 reaches the maximum 96% limitation (31,450 RPM), the corresponding light will illuminate and readout will flash. Input signals are supplied to each channel of the gage by its associated left or right N2 tachometer generator installed on the forward AGB. Each channel of the gage receives 28 VDC power through its associated 2-amp LH TURB SPEED or RH TURB SPEED circuit breaker, correspondingly supplied by the left or right main bus. When power is removed from either channel, the affected digital readout will be extinguished.

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Engine Oil Temperature Gage The OIL TEMP gage provides indication of oil temperature in °C. The instrument scale is graduated in 10° increments between 0 and 130°C. Numerical values are marked at each 20° increment between 0 and 120°C. Green markings denote the normal indicating range. Red warning lines at 121°C denote the maximum oil temperature limitation. Input signals are supplied to each channel of the gage by its associated left or right oil temperature sensor installed in the oil filter housing. Each channel of the gage receives 28 VDC power through its associated 2-amp LH OIL TEMP or RH OIL TEMP circuit breaker, correspondingly supplied by the left or right main bus. A red OFF warning flag will appear at the top of the vertical scale whenever power is removed from the corresponding channel of the gage. Engine Oil Pressure Gage The OIL PRESS gage provides indication of oil pressure in PSI. The instrument scale is graduated in 10 PSI increments between 0 and 100 PSI. Numerical values are marked at each 20 PSI increment between 0 and 100 PSI. Green markings denote the normal indicating range. Yellow markings denote the low pressure caution range. Red warning lines at 35 PSI denote the minimum oil pressure limitation. Input signals are supplied to each channel of the gage by its associated left or right oil pressure transmitter installed in the oil filter housing. Each channel of the gage receives 28 VDC power through its associated 2-amp LH OIL PRESS or RH OIL PRESS circuit breaker, correspondingly supplied by the left or right main bus. A red OFF warning flag will appear at the top of the vertical scale whenever power is removed from the corresponding channel of the gage. ENGINES 12/99


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Should oil pressure fall below approximately 35 PSI, a red [LH] or [RH] [OIL PRESS WARN] annunciator (550-0550 an after) or [L OIL PRESS LO] or [R OIL PRESS LO] annunciator (550-0505 and earlier) will illuminate. Input signals are supplied by an oil pressure transmitter and low pressure switch installed on the AGB. Note: The normal oil pressure indicating range applies to engine speeds above 60% N2. Caution: Engine operation at oil pressures below 70 PSI is undesirable. Should oil pressure fall below 70 PSI in-flight, flight may be completed at oil pressures between 35 and 70 PSI when engine power is reduced. Warning: Engine operation at oil pressures below 35 PSI is unsafe. Should oil pressure fall below 35 PSI in-flight, the airplane should be landed as soon as possible using the minimum engine power required to sustain flight.

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Cessna Citation II Technical Manual Engine Starting System Engine starting is accomplished semi-automatically by independent left and right start control circuits which activate and deactivate the starter/generator, ignition system, and boost pump through a series of relays. The engines may be started using battery power or external power. With one engine operating, the opposite engine may also be started using generator power. LH and RH momentary-on, push-button ENGINE START switches, located on the lower left instrument panel, control start sequence activation. A STARTER DISENGAGE switch, located between the ENGINE START switches, permits manual termination of the start sequence. Either engine may be started first, however, the left engine is typically started first due to its proximity to the battery.

Of the numerous relays in each start control circuit, only three are associated primarily with engine starting, these are the auxiliary (aux) start relay, start relay, and start control relay. A left engine start sequence is activated by pressing and releasing the LH ENGINE START switch. The resulting momentary flow of current energizes the aux start relay, after which it remains energized through a latching circuit. Current flowing across the closed contacts of the aux start relay simultaneously energizes the start relay and start control relay and illuminates the engine instrument floodlights.

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With the start relay energized, the LH ENGINE START switch illuminates and the starter-generator begins turning the engine. Current flowing across the closed contacts of the start control relay energizes the left boost pump and illuminates the amber [LH] [FUEL BOOST ON] annunciator (550-0550 and after) or [L FUEL BOOST ON] annunciator. When N2 reaches 8 to 10%, lifting the spring-loaded trigger and advancing the left THROTTLE lever to “IDLE” opens the flow divider cut-off valve to supply fuel to the engine. Simultaneously, a throttle position switch permits current flow across the closed contacts of the start control relay to the ignition system and illuminates the green light above the LH IGNITION switch. From this point on, normal initiation of combustion should be indicated by a steady rise in N1, N2, and ITT. If no indication of N1 exists when N2 reaches 20 to 25%, or if ITT approaches 700°C or fails to rise within 10-seconds, the start sequence should be manually terminated by pressing the STARTER DISENGAGE switch. If indications are normal, the start control circuit will automatically terminate the start sequence by deenergizing the aux start relay, thereby deenergizing the start relay and start control relay, under the following conditions: On airplanes 550-0406 and after, and earlier airplanes incorporating SB550-28-1, the starting sequence is terminated by the engine speed sensing circuit of the generator control unit (GCU) when N2 reaches approximately 40%. On airplanes 5500405 and earlier not incorporating SB550-28-1, the starting sequence is terminated by the motive flow pressure switch when the engine-driven fuel pump is supplying at least 180 ± 5 PSI motive flow to the primary ejector pump. This pressure is typically achieved when N2 reaches approximately 30%. Should the motive flow switch fail to open, the engine speed sensing circuit of the GCU will terminate the starting sequence when N2 reaches approximately 40%. Note: During normal engine start, ITT indication should not exceed 500°C. Note: ITT indication exceeding 500°C during normal engine start should be investigated for cause.

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Cessna Citation II Technical Manual The following precautions apply primarily during engine starting and ramp area ground operations. Warning: An engine running at higher power is capable of generating sufficient suction to pull a person into the intake duct; even the suction generated by an idling engine is capable of pulling small objects through the intake duct sufficient to cause injury to persons and/or engine damage. As a rule, personnel and equipment should not be positioned within an arc extending a minimum of 25 feet forward of and 90° either side of the intake duct when starting or running the engine. Warning: An engine running at higher power is capable of generating an exhaust wake sufficient in temperature and velocity to cause injury to persons and/or damage to property. During engine start, it is possible for accumulated fuel within the exhaust duct to be ejected from the engine as long streams of flame. Exhaust gases can cause respiratory and/or eye irritation. At high engine speeds, the jet wake may propel loose dirt, sand, stones, and/or other debris for considerable distances. As a rule, personnel, equipment, structures, and flammable material should be clear of an area extending aft from the exhaust nozzle to a minimum width of 30 feet and a minimum distance of 160 feet when starting or running an engine.

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Engine Synchronizer System The engine synchronizer system automatically matches the fan or turbine speed of the right (slave) engine to the left (master) engine. Major components of the system include a control box, motorized speed setting actuator, trimmer assembly, flexible drive shaft, selector switch, and indicator light. The control box monitors RPM signals supplied by the N1 (fan speed) and N2 (turbine speed) tachometer generators of each engine and provides operating commands to the speed setting actuator. The speed setting actuator, located within the right engine pylon, operates the trimmer assembly via the flexible drive shaft. The trimmer assembly ▲ SPEED SETTING ACTUATOR adjusts the right (slave) engine’s fuel control unit to maintain N1 or N2 within 1.5% of the left (master) engine. Based on the RPM signals received and the mode selected, the control box commands the speed setting actuator to adjust slave N1 or N2 The system is controlled by the ENGINE SYNC selector switch located aft of the right THROTTLE lever on the center pedestal. When positioned to “FAN,” the system synchronizes N1. When positioned to “TURB,” the system synchronizes N2. When positioned to “OFF,” the system is deactivated. The amber indicator light, located adjacent to the ▲ ENGINE SYNCHRONIZER SWITCH selector switch, illuminates when the engine synchronizer system is operating.

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Cessna Citation II Technical Manual The “FAN” position should be selected when passenger comfort is a primary consideration, because fan speed mismatch is more noticeable in the passenger cabin. The “TURB” position should be selected when crew comfort is a primary consideration, because turbine speed mismatch is more noticeable in the flight compartment. On airplanes 550-0165 and earlier, indication of synchronizer system performance may be monitored using an optional synchroscope mounted to the right of the engine instruments on the upper center instrument panel. The engine synchronizer system is powered by 28 VDC from the left main bus (550-0550 and after) or right main bus (5500505 and earlier). Circuit protection is provided by a 5-amp (550-0626 and earlier) or 2-amp (550-0627 and after) ENGINE SYNC circuit breaker located on the left CB panel.

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Thrust Reverser System

Optional external target-type thrust reversers may be installed allowing engine thrust to be used to help decelerate the airplane during landing rollout. The thrust reversers consist of two hydraulically-actuated doors per engine. Each door is attached to the outer exhaust nozzle by means of an aluminum support casting and four interconnecting links. The interconnecting links are attached to sliding carriage mechanisms that are driven by the hyrdraulic actuators. When deployed, the reverser doors are positioned behind the exhaust nozzles. The upper doors direct thrust upward and forward; the lower doors direct thrust downward and forward. When stowed, the reverser doors are flush contoured to form the aft portion of the engine nacelle. The reverser doors are held in the stowed position by the overcenter locking design of the linkage. Thrust Reverser Hydraulics Hydraulic pressure for thrust reverser operation is supplied by the same system that supplies the and landing gear and speed brakes. This section will describe the various valves and switches that control thrust reverser operation. Refer to Chapter 8 for a complete description of the hydraulic system.

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Cessna Citation II Technical Manual Thrust Reverser Control Valves The solenoid-operated thrust reverser control valves function to direct hydraulic pressure to, and return flow from, the deploy or stow ports of their corresponding actuators. To accomplish this, each control valve contains an internal selector spool that is spring-loaded to a neutral position and operated by independent deploy and stow solenoids. In the neutral position, when both solenoids are deenergized, the deploy and stow ports are both connected to the return line so that hydraulic pressure will not be trapped in the lines between each control valve and its associated actuators. When the deploy solenoid is energized, the selector spool is positioned to direct hydraulic pressure to the deploy ports of the actuators, and direct return flow from the stow ports of the actuators to the reservoir. Conversely, when the stow solenoid is energized, the selector spool is positioned to direct hydraulic pressure to the stow ports, and direct return flow from the deploy ports to the reservoir. The solenoids are energized and deenergized primarily by the thrust reverser (TR) levers through corresponding deploy and stow microswitches. The deploy solenoid is energized through the deploy microswitch when its associated TR lever is pulled aft to the “idle deploy” position. The stow solenoid is energized through the stow microswitch when its associated TR lever is pushed forward to the “stow” position. To prevent overheating the deploy solenoid when energized, holding voltage is gradually reduced from 28 VDC to 6 VDC by a transistorized circuit while its associated thrust reverser doors are deployed. To prevent overheating the stow solenoid, limit switches deenergize the solenoid when its associated thrust reverser doors are fully-stowed. Thrust Reverser Isolation Valves The thrust reverser isolation valves, installed upstream of each control valve pressure inlet, function to isolate the hydraulic components of the corresponding thrust reverser system from the main hydraulic system when the thrust reversers are not in use. The isolation valves are normally spring-loaded closed, and energized open primarily by the TR levers through the deploy and stow microswitches. Each isolation valve remains open when its associated TR lever is in the “idle deploy” position, and closes when its associated thrust reverser doors are fully-stowed.

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Check Valves and Restrictors Check valves are installed in the return line to prevent reverse fluid flow through the control valve and isolation valve. A restrictor is installed between the control valve and the actuator stow ports to prevent the rate of return flow from the actuators from exceeding the rate of pressure flow to the actuators when the thrust reversers are deployed at ground speeds above 120 knots. Thrust Reverser Levers Thrust reverser operation is controlled by the TR levers mounted “piggyback” on the THROTTLE levers. To prevent inadvertent thrust reverser operation, the THROTTLE levers must be set to “IDLE” before the TR levers can be moved. Additionally, thrust reverser deployment is inhibited by either main gear safety switch when the airplane is in flight. During thrust reverser transition, the throttle feedback system holds the throttle linkage at idle until the reverser doors are fully-deployed or fully-stowed. A pair of locking solenoids prevents aft movement of the TR levers to increase reverse thrust until the reverser doors are fully-deployed. Warning: The THROTTLE levers should not be restrained manually or by lever friction during takeoff as this will prevent the throttle feedback system from automatically returning the throttle linkage and THROTTLE levers to “IDLE” should inadvertent thrust reverser deployment occur. Note: The thrust reversers should be deployed only during landing rollout following touchdown and only after all three wheels are on the ground. Thrust reverser deployment prior to nose wheel touchdown can generate sufficient pitch up movement to cause aft fuselage contact with the ground. Note: The thrust reversers should not be deployed during touch-and-go landings due to increased takeoff distance resulting from the time required to restow the reverser doors prior to takeoff power being set.

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Cessna Citation II Technical Manual Thrust Reverser Deployment Thrust reverser deployment is initiated by pulling the TR levers aft to the “idle deploy” position. In this position, the control valve deploy solenoid, isolation valve, and hydraulic system center valve are energized to permit thrust reverser deployment.

1

1. Stow Position 2. Idle Deploy Position 3. Full Reverse Position

2

3

When the center valve is energized and hydraulic pressure is being supplied to the thrust reversers, the amber [HYD PRESS ON] annunciator will be illuminated. When pressure downstream of each isolation valve reaches approximately 200 PSI, a pressure switch causes the corresponding amber [ARM] annunciator on the glareshield panel to illuminate. As the actuators begin driving the sliding carriages forward along guides rods to extend the interconnecting links and deploy the doors, a stow limit switch is actuated causing the corresponding amber [UNLOCK] annunciator on the glareshield panel to illuminate. When the doors have fully-deployed, a deploy limit switch is actuated causing the corresponding white [DEPLOY] annunciator on the glareshield panel to illuminate. Following thrust reverser deployment, engine thrust is increased by pulling the TR levers further aft. Full reverse thrust is obtained when the TR levers are positioned fully-aft against their stops. The TR lever stops limit reverse thrust N1 to approximately 92% in sea level ISA conditions. ENGINES 12/99


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Thrust Reverser Deployment Schematic

© PCW

Thrust Reverser Stowed Schematic

© PCW

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Cessna Citation II Technical Manual Thrust Reverser Stowage Thrust reverser stowage is initiated by pushing the TR levers forward to the “stow” position. In this position, the control valve deploy solenoid, isolation valve, and hydraulic system center valve are energized to permit thrust reverser stowage. As the actuators begin driving the sliding carriages aft, the corresponding [DEPLOY] annunciator extinguishes. When the doors are fully-stowed, the corresponding [UNLOCKED] annunciator extinguishes and the control valve deploy solenoid, isolation valve, and hydraulic system center valve will are deenergized. In this condition, the [HYD PRESS ON] and [ARM] annunciators will also be extinguished.

2

1

4 6 5 3

1. Hydraulic Actuator 2. Sliding Carriage 3. Interconnect Links

4. Stow Limit Switch 5. Deploy Limit Switch 6. Throttle Feedback Linkage

Electrical components of the left and right thrust reverser systems are supplied with 28 VDC power from the left main bus and right main bus respectively, through corresponding 7.5amp LH and RH THRUST REVERSER (550-0550 and after) or LH and RH THRU REV (550-0505 and earlier) circuit breakers on the left CB panel. The control valve and isolation valve function with an input power of 18 to 30 VDC. When electrical power is removed from the system, the thrust reversers will fail to the stowed position providing the linkage is not in its overcenter position. ENGINES 12/99


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Thrust Reverser Emergency Stowage 2 Should either reverser fail to stow in response to TR lever movement due to failure of the deploy or stow switches or due to loss of electrical power to the control circuit, the affected reverser(s) can be 1 stowed by positioning the corresponding emergency stow switch (STOW SW) to the “EMER” position. The STOW 1. Emergency Stow Switch SW for each thrust reverser is 2. Thrust Reverser correspondingly located on Annunciators the left and right side of the glareshield panel. Each emergency stow circuit receives electrical power through the opposite thrust reverser circuit breaker. Thrust Reverser Emergency Stowage Test Emergency stow switch function can be verified on the ground by deploying the reversers normally and then positioning each switch to “EMER.” In this condition, the reversers should stow normally, sequentially extinguishing the [DEPLOY] and [UNLOCK] annunciators, while the [ARM] and [HYD PRESS ON] annunciators remain illuminated. After testing, each TR lever should be returned to the “stow” position, and the corresponding emergency stow switch should be positioned to “NORMAL.”

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Cessna Citation II Technical Manual Engine Fire Detection and Extinguishing System This system is designed to give immediate warning and provides extinguishing capabilities should a fire occur in either engine nacelle. The system includes thermal detectors, a detector control unit, two explosive cartridges, two extinguisher bottles, discharge tubing, and four annunciator/switches. Thermal Detectors Each engine is equipped with a thermal detector installed in high-heat areas within each engine nacelle. The detector consists of a 188-inch flexible steel tube containing a single wire centered within a highly compacted semiconductor material. The steel tube is hermetically sealed with fireproof connectors at each end. The sensing elements form a closed-loop circuit that detects fire within its associated engine compartment by decreasing electrical resistance in response to heat. Detector Control Unit The detector control unit, located within the aft fuselage, evaluates the resistance of the sensing elements to minimize false alarms, and illuminates the [ENG FIRE PUSH] annunciator/ switches and sounds a fire warning bell (if installed) when sensor cable resistance indicates that fire is present within the nacelle. If resistance gradually falls below the preset fire alarm point of 200 ohms, the detector control unit illuminates the appropriate [ENG FIRE PUSH] annunciator/switch and sounds the fire warning bell. If resistance instantaneously falls below the preset short discriminating value, indicating a short circuit rather than a fire, an electronic lockout is actuated that prevents illumination of the [ENG FIRE PUSH] annunciator/switch and sounding of the fire warning bell. Explosive Cartridges The explosive cartridges are pyrotechnic devices installed within the discharge valve assembly on each extinguisher bottle. When the corresponding [BOTTLE ARMED PUSH] annunciator/switch is depressed, its respective explosive cartridge detonates, rupturing the cartridge housing assembly, thereby removing the mechanical locking force against the valve plug and allowing the valve plug to be expelled by extinguishing agent pressure. The valve plug and housing assembly parts are captured by a strainer assembly.

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Extinguisher Bottles The extinguisher bottles, located within the aft fuselage, contain Halon-1301 fire extinguishing agent pressurized with dry nitrogen. Each bottle is equipped with two discharge valves and outlets, a combination safety outlet/fill port, and a pressure gage. On airplanes equipped with thrust reversers, the bottles are cylindrical with a capacity of 125 cubic inches. On airplanes without thrust reversers, the bottles are spherical with a capacity of 86 cubic inches. Annunciator/Switches The fire extinguishing system is controlled using the left and right red [ENG FIRE PUSH] and white [BOTTLE ARMED PUSH] annunciator/switches on the glareshield panel. The To prevent inadvertant actuation, the [ENG FIRE PUSH] annunciator switches are guarded by hinged, springloaded, transparent covers. System Operation If an overheat condition is detected, the appropriate [ENG FIRE PUSH] annunciator/switch will illuminate. Depressing the [ENG FIRE PUSH] annunciator/switch closes the corresponding fuel and hydraulic firewall shutoff valves, takes the corresponding starter/generator off-line, and arms both extinguisher bottles. When the appropriate [BOTTLE ARMED PUSH] annunciator/ switch is depressed, the corresponding explosive cartridge is detonated and extinguishing agent is routed through tubing and discharged within the forward nacelle. Airplanes equipped with thrust reversers also discharge extinguishing agent into the area between the upper thrust reverser door and the engine exhaust duct assembly through holes in the front flange of the upper thrust reverser door.

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Cessna Citation II Technical Manual Engine Fire Detection System Schematic

© PCW

Engine Fire Extinguishing System Schematic

© PCW

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The fire extinguishing system is plumbed such that either or both extinguishing bottles can be discharged into either engine by pressing the appropriate [ENG FIRE PUSH] and [BOTTLE ARMED PUSH] annunciator/switches. When each extinguisher bottle has been discharged, the corresponding red [ENG FIRE PUSH] annunciator/switch should no longer be illuminated, indicating that the fire in the associated engine has been extinguished. If the [ENG FIRE PUSH] annunciator/switch remains illuminated, the remaining extinguishing agent can be discharged into the affected engine by pressing the appropriate [BOTTLE ARMED PUSH] annunciator/switch. A gage indicating extinguishing agent supply pressure is located on each extinguisher bottle within the aft fuselage compartment. Each gage provides the only indication that extinguishing agent may be leaking from its respective bottle. A table of acceptable bottle pressures per ambient temperature is placarded adjacent to each pressure gage and/or printed in the Operating Limitations section of the Operating Manual. On airplanes 550-0550 and after, 28 VDC electrical power is supplied to the left fire detection system and the left firewall shutoff valves/extinguishing system by the right main bus, and the right fire detection and the right firewall shutoff valves/ extinguishing system by the left main bus, through corresponding 2-amp LH and RH FIRE DET and 7.5-amp LH and RH F/W SHUTOFF circuit breakers on the left CB panel. On airplanes 550-0505 and earlier, the left and right fire detection system and the left and right firewall shutoff valves/extinguishing systems are supplied with 28 VDC power from the left main bus and right main bus respectively, through corresponding 2-amp LH and RH FIRE DETECT circuit breakers on the left CB panel.

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Cessna Citation II Technical Manual System Testing A test function is provided to test the fire detection system annunciator/actuators and to check continuity of their associated sensors and detector control units. When the rotary TEST selector switch on the lower left instrument panel is positioned to “FIRE WARN”, both red [ENG FIRE PUSH] annunciator/actuators should ▲ TEST SELECTOR SWITCH illuminate. Pressing either [ENG FIRE PUSH] annunciator/actuator will then illuminate both [BOTTLE ARMED PUSH] annunciator/actuators. The fire extinguishing system may be tested on the ground or in flight. Limitations Refer to the applicable aircraft manufacturers FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable aircraft manufacturers FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.

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Chapter 3 Fuel System Table of Contents Overview ............................................... 3-1 Fuel Storage System ............................. 3-1 Fuel Tanks ....................................... 3-1 Servicing........................................... 3-2 Fuel Venting System ............................. 3-5 Fuel Drains ........................................... 3-6 Distribution System ............................... 3-8 Motive Flow Ejector Pumps ................ 3-8 Motive Flow Pressure Switches .......... 3-9 Motive Flow Shutoff Valves ................ 3-9 Boost Pumps .................................... 3-9 Fuel Filters ..................................... 3-11 Maintenance Shutoff Valves ............. 3-11 Fuel Firewall Shutoff Valves ............. 3-12 Crossfeed Valves ............................ 3-13


Fuel System, continued Table of Contents Fuel System Indication ........................ 3-14 Fuel Quantity Indication ................... 3-14 Low Fuel Level Warning ................... 3-15 Low Fuel Pressure Warning .............. 3-15 Fuel Flow Indication ......................... 3-16 Fuel Remaining/ Consumed Indication ..................... 3-17 Operational Summaries ....................... 3-18 Engine Starting ............................... 3-18 Normal Operation ............................ 3-19 Low Fuel Pressure .......................... 3-19 Crossfeeding ................................... 3-20 Limitations .......................................... 3-26 Emergency Procedures......................... 3-26

Cessna Citation II Technical Manual Overview This chapter will describe the Cessna Citation II fuel system, its major components, controls, operation, and indication. The airplane fuel supply is contained in two independent wing systems connected by crossfeed lines. Fuel transfer and delivery of fuel to the engines is carried out by a distribution system of motive flow ejector pumps, electrically-powered boost pumps, engine-driven fuel pumps, and control valves. The distribution system supplies fuel to the engines in excess of that required for all operations. Indication of fuel quantity and flow are displayed by gages on the upper center instrument panel. Other system indications are displayed on the annunciator panel. Emergency fuel shutoff systems are provided. Fuel Storage System The fuel storage system includes the fuel tanks, filler caps, venting system, and fuel drains. Fuel Tanks The airplane fuel supply is stored in one integral tank per wing. Each tank occupies roughly the entire inner area of the wing forward of the rear spar, excluding the main wheel well, and is formed by sealing all structural joints between the extreme inboard and outboard ribs. The front spars and all interior ribs incorporate holes that permit fuel migration within the tank. Outboard interior ribs incorporate baffle plates to prevent rapid fuel load shift when the airplane transitions to and from a winglow attitude. The fuel storage area is chemically-treated and finished with an epoxy primer for corrosion resistance. During normal operation, each engine is supplied by its associated tank. During crossfeeding, fuel may be supplied from one tank to both engines or, in the event of engine failure, from either tank to the operative engine. Refer to the fuel system operational summaries in this chapter for a complete description of these conditions.

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Each fuel tank is divided into a main area and sump area. The sump area is located within the main area, inboard of the main wheel well, and functions as a reservoir to which fuel is either gravity-fed or transferred, and from which fuel is supplied to the engines. Each sump area is enclosed by a cover assembly that houses the boost pump, the primary motive flow ejector pump, and a pair of flapper-type check valves. The flapper valves permit the gravity feeding of fuel from the main area to the sump area while preventing back flow from the sump area to the main area. This arrangement ensures that sufficient fuel is contained within each sump area to supply the engines during all normal maneuvering, and a minimum of five seconds fuel supply during negative gravity maneuvers not exceeding -0.5 G. Aside from gravity feeding, all other transferring of fuel is carried out by the distribution system. Servicing Each tank is serviced through a single, flush-mounted filler cap located on the outboard upper surface of the wing. The location of the filler caps ensures that sufficient fuel expansion space will exist within the tanks when topped off. Filler cap security should be checked during preflight. A fuel nozzle grounding point is located on the lower surface of each wing tip.

The fuel tanks have an approximate usable capacity of 371 US gallons each, and 742 US gallons (5000 LBS) total. The fuel tanks should be kept full between flights (providing weight and balance considerations permit) to reduce explosive vapors and condensation. Maintaining fuel load symmetry during servicing is unnecessary; however, the maximum permissible asymmetry is 200 LBS during normal flight operations and 600 LBS in an emergency.

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Cessna Citation II Technical Manual Approved fuels are JET A, JET A-1, JET B, JP-4, JP-5, or JP-8, all with 0.15% PFA55MB anti-icing additive in solution. When preblended fuel is not available, anti-icing additives conforming to MIL-I-27686E (Ethylene Glycol Monomethyl Ether (EGME)) or MIL-I-85470 (Diethylene Glycol Monomethyl Ether (DIEGME)) specifications such as “Prist” may be introduced directly into the nozzle fuel stream during servicing. Concentrations of less than 0.06% (20 fluid ounces of additive per 260 gallons of fuel or more) may be insufficient to prevent fuel system icing or microbiological contamination. Conversely, concentrations of more than 0.15% (20 fluid ounces of additive per 104 gallons of fuel or less) could cause damage to internal components of the fuel system or erroneous fuel quantity indications. Caution: EGME and DIEGME are aggressive chemicals and should not exceed 0.15% of fuel volume. Improperly handled, these materials will damage the epoxy primer and sealants used in the fuel tanks, O-ring seals, and any part of the airplane’s exterior finish with which it comes in contact. Warning: Anti-icing additives containing EGME or DIEGME are harmful if inhaled, swallowed, or absorbed through the skin, and will cause eye irritation. Refer to all instructions and warnings regarding toxicity and flammability before using these materials. All grades of aviation gasoline (AVGAS) conforming to MIL-G5572 specifications are approved for use under emergency circumstances only, providing the airplane is operated in accordance with related procedures and limitations specified in the AFM. Use of AVGAS is limited to no more than 3500 US gallons or 50 hours of engine operation during any period between engine overhaul. For record keeping purposes, 1 hour of engine operation may be considered equivalent to 70 US gallons. Note: For a complete listing of approved fuels and additives, refer to the latest revision of Pratt & Whitney Service Bulletin 7144. Warning: Do not allow open flame or smoking in the vicinity of the airplane during fuel system servicing. Warning: Do not operate avionics, communications, or other electrical equipment on the airplane during fuel system servicing. FUEL SYSTEM 12/99


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Warning: Appropriate fire fighting equipment must be available during fuel system servicing. Access panels on the lower surface of each wing are provided to facilitate inspection and maintenance of the fuel tanks as well as components of the distribution and indicating systems. These panels incorporate liquid-tight perimeter seals. Prior to flight, the lower surface of each wing should be inspected for evidence of fuel leakage. If observed, the source and cause of leakage should be determined by maintenance personnel and evaluated against classification criteria and repair action requirements. Generally, light to moderate seepage does not limit flight operations, unless observed in proximity to the heated leading edge panels. Conversely, heavy seepage, running leaks, or any fuel leakage observed in proximity to the heated leading edge panels require immediate repair before resuming flight operations. Caution: Any fuel leakage caused by structural failure such as cracks, or failure of components such as fuel lines, must be repaired before resuming flight operations.

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Cessna Citation II Technical Manual Fuel Venting System An independent venting system is installed within each wing to equalize external and internal tank pressures. Each tank is vented to the atmosphere through a flushmounted, NACA-type, antiicing vent scoop located on the outboard lower surface of the wing. The design of the vent scoop prevents icing without the need for supplemental heat. Additional venting system components include a vent float valve, two check valves, a vent line, and a canistertype vent surge tank (550-0006 and after) located within each tank. On airplanes 550-0005 and earlier, the vent surge tank is integral to the wing structure. Pressure caused by expanding fuel is automatically vented from the tank through the vent line which extends from the inboard end of the tank to the vent scoop near the outboard end of the tank. Air pressure building within the tank normally enters this line through the vent float valve and escapes to the atmosphere through the vent scoop. When the float valve is closed or when the airplane is in a climb attitude with a wing low, air pressure building within the affected tank enters the vent line through an opening at its inboard end. When the float valve is closed during level flight or descent, the check valves allow ambient air to enter the tank. When the airplane is parked on an inclined surface such that the float valve is closed, internal tank pressure caused by expansion will force fuel through the open end of the vent line to the surge tank inlet. Because the outlet is positioned higher in the tank than the inlet, fuel accumulations in the surge tank must reach the level of the outlet before being forced overboard through the vent scoop. A secondary venting system is incorporated in the fuel filler caps to ensure fuel flow in the event of normal venting system obstruction. Should internal tank pressure drop below external pressure, the filler cap vent allows ambient air to be drawn into the tank thereby equalizing these pressures. Note: All vents should be free of obstructions prior to flight.

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Fuel Drains

Fuel drain valves, used to check for contamination and to remove water and sediment from the tanks, are located on the lower surface of each wing. Four (550-0060 and after) or six (550-0059 and earlier) “quick-drain” type valves are provided for each fuel tank. One drain valve is also provided for each fuel filter assembly. All drain valves are actuated by pushing up the inner portion (poppet) of the valve assembly. Note: Rotating the poppet of the quick-drain valve assemblies while draining fuel will lock the valve in its open position. This is accomplished using a phillips head screwdriver (550-0060 and after) or flat-bladed screwdriver (550-0059 and earlier).

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Cessna Citation II Technical Manual Airplanes 550-0409 and after incorporate semi-flush mounted fuel tank drain valves which permit poppet O-ring replacement without defueling the airplane or valve assembly removal. Normal preflight fuel draining procedures will generally remove most excess water from the fuel tanks. However, small amounts of water will remain in solution within the fuel. This residual water will facilitate microbe and bacterial growth in settlement areas of the fuel system which can lead to fuel flow obstruction and/or corrosion. To minimize the effects of fuel contamination, the pilot should ensure that the fuel does not contain unapproved additives and has been properly handled by the source. Before every flight, a sample should be taken from each fuel tank drain and inspected for contamination. If contamination is detected, draining from that point should continue until contamination is no longer present. If after continued draining contamination still exists, the airplane should not be flown. Note: Operators not acquainted with a particular airfield should confirm that the fuel supply there is routinely checked for contamination, contains approved additives, and is properly filtered before allowing the fuel system to be serviced. Note: At least 30 minutes should elapse between fuel system servicing and sample taking. Warning: Under no circumstances should the airplane be flown with contaminated or unapproved fuel.

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Distribution System The distribution system functions to maintain a continuous supply of fuel to the engines in excess of that required for all operations. Major components of the system include primary and transfer motive flow ejector pumps, electrically-powered boost pumps, and engine-driven fuel pumps. Other components, located in the “dry bay” area between the front and rear spars within each lower fuselage to stub wing fairing, include maintenance shutoff valves, firewall shutoff valves, crossfeed valves, motive flow check valves, fuel filters, motive flow pressure switches (550-0405 and earlier not incorporating SB55028-1), and motive flow valves. Motive Flow Ejector Pumps Motive flow ejector pumps are operated by fuel pressure supplied primarily by the engine-driven fuel pumps. During engine start and crossfeeding, motive flow pressure is supplied by the boost pumps. One primary ejector pump and two transfer ejector pumps are located in each fuel tank. The transfer pumps transfer fuel from the main area of the fuel tank to the sump area. The primary pump supplies fuel from the sump area to the engine-driven fuel pump. Each primary ejector pump operates on bypass fuel from its associated enginedriven fuel pump when the engine is operating and the enginedriven fuel pump is developing sufficient pressure to maintain motive flow. Each pair of transfer ejector pumps operates on bypass fuel from its associated primary ejector pump. As motive flow fuel enters each ejector pump through its pressure inlet, pressure is reduced by venturi effect thereby drawing fuel through its suction inlet. The suction inlet incorporates a wire mesh fuel strainer to prevent solid particles from entering the pump. A spring-loaded, ball-type check valve is installed in each primary pump outlet to prevent reverse flow through the pump when its associated boost pump is in operation. On airplanes 550-0405 and earlier not incorporating SB550-28-1, a motive flow pressure switch is located in each primary ejector pump supply line. Refer to the motive flow pressure switch description and engine starting operational summary in this chapter for functional details.

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Cessna Citation II Technical Manual Motive Flow Pressure Switches On airplanes 550-0405 and earlier not incorporating SB550-28-1, a motive flow pressure switch is installed in each primary ejector pump supply line. Each switch is normallyclosed permitting current flow through the start control circuit to its respective starter motor, ignition system, and boost pump during the engine starting sequence. After light-off has occurred and the engine-driven fuel pump is developing at least 180 ± 5 PSI motive flow, the switch is opened interrupting current flow to the starter motor, ignition system, and boost pump, thereby terminating the starting sequence. During engine shutdown, as motive flow pressure decreases below approximately 120 PSI, each switch returns to its normallyclosed position. The motive flow pressure switches should not be confused with the fuel pressure switches also described in this chapter. On airplanes 550-0406 and after, and earlier airplanes incorporating SB550-28-1, termination of the starting sequence occurs as a function of engine speed. Refer to the engine starting operational summary in this chapter for functional details. Motive Flow Shutoff Valves An electrically-operated motive flow shutoff valve is installed in each primary ejector pump supply line. Each shutoff valve is normally-open permitting motive flow to its corresponding primary ejector pump. During crossfeeding, the shutoff valve for the system not supplying fuel is energized closed by 28 VDC power supplied through its associated 15-amp LH BOOST or RH BOOST circuit breaker on the left CB panel. Refer to the crossfeeding operational summary in this chapter for functional details. Boost Pumps One electrically-powered, centrifugal-type boost pump is submerged in the sump area of each fuel tank. Each boost pump provides fuel pressure to its respective engine during engine start and opposite engine during crossfeeding. During conditions of low fuel pressure or low fuel level, the boost pumps may be activated to ensure uninterrupted fuel flow to the engines. A check valve is installed in each pump outlet to prevent reverse flow through the pump when it is not in operation.

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Each boost pump is independently controlled by a threeposition (ON/OFF/NORM) switch labeled FUEL BOOST, located on the lower left instrument panel. Under most operating conditions, these switches should remain in the “NORM” (normal) position. In this position, boost pump activation occurs automatically during engine start, crossfeeding, and conditions of low fuel pressure.

▲ BOOST PUMP SWITCHES

During engine start, automatic boost pump activation is carried out by a series of relays in the start control circuit that supply 28 VDC power to the appropriate pump. During crossfeeding and conditions of low fuel pressure, automatic boost pump activation is carried out by various relays and switches that supply 28 VDC power to the appropriate pump(s). Refer to the corresponding operational summaries in this chapter for complete descriptions of these conditions. With the FUEL BOOST switches set to “ON,” 28 VDC power is supplied directly to each boost pump through its associated 15-amp LH BOOST or RH BOOST circuit breaker on the left CB panel. In this condition, on airplanes 550-0550 and after, the left boost pump is supplied by the right main bus and the right boost pump is supplied by the left main bus. On airplanes 5500505 and earlier, the left boost pump and right boost pump are correspondingly supplied by the left main bus and right main bus. When set to “OFF,” automatic activation of the boost pumps will occur only during engine start or crossfeeding as previously described. Note: The position of the LH BOOST and RH BOOST circuit breakers on the left CB panel has no effect on boost pump operation during the engine starting sequence.

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Cessna Citation II Technical Manual Indication of boost pump operation is provided by an amber, three lens [LH] [RH] [FUEL BOOST ON] annunciator (550-0550 and after) or by independent, amber [L FUEL BOOST ON] and [R FUEL BOOST ON] annunciators (550-0505 and earlier). When power is being supplied to either boost pump, the corresponding annunciator will be illuminated regardless of operating condition. Fuel Filters Each fuel filter incorporates a disposable paper element that functions to trap solid particle contaminants present in the fuel. A differential pressure sensing switch and bypass valve are integral to each filter head assembly. Should the differential between filter inlet and outlet pressure exceed approximately 3.75 PSID, the corresponding amber [LH] or [RH] [FUEL FLTR BYPASS] annunciator (550-0550 and after), or the amber [FUEL FILT BYPASS] annunciator (550-0505 and earlier) will illuminate indicating filter element obstruction and an impending bypass condition. Should this differential exceed approximately 4.75 PSID, the bypass valve will open providing continued though unfiltered fuel flow to the engine. A spring-loaded drain valve is installed in the base of each filter bowl and extends through the lower surface of the stub wing. Before flight, a sample should be drained from each of these valves and inspected for contamination. A manually-operated shutoff valve in the outlet port of each filter head permits replacement of the paper element without fuel drainage from the engine supply line. Maintenance Shutoff Valves The maintenance shutoff valves function to isolate components of the distribution system downstream of the fuel tanks during maintenance operations. Each ball-type valve is spring-loaded to the open position and manually closed by rotating its handle to a detent in the valve body.

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Fuel Firewall Shutoff Valves The fuel firewall shutoff valves function to terminate fuel supply in the event of an engine fire. Each ball-type valve is motoroperated and electrically-controlled by independent [LH ENG FIRE] and [RH ENG FIRE] annunciator/switches on the glare shield panel fire tray. A transparent, spring-loaded guard is installed over each switch to protect against inadvertent actuation.

When either switch is pressed, the corresponding valve is closed by 28 VDC power supplied through its associated 7.5amp LH F/W SHUTOFF or RH F/W SHUTOFF circuit breaker on the left CB panel. When fully closed, the corresponding amber [LH] or [RH] [F/W SHUT OFF] annunciator (550-0550 and after) or amber [L F/W SHUTOFF] or [R F/W SHUTOFF] annunciator (550-0505 and earlier) should be illuminated. Additionally, both [BOTTLE ARMED PUSH] annunciator/switches on the glare shield panel fire tray should be illuminated. A closed valve may be reopened by pressing the corresponding annunciator/ switch to release it from its locked position. Note: The hydraulic firewall shutoff valve is also closed when the fuel firewall shutoff valve is closed. Refer to Chapter 8 for a complete description of the hydraulic firewall shutoff valve. Refer to Chapter 2 for a complete description of the engine fire protection system.

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Cessna Citation II Technical Manual Crossfeed Valves The crossfeed valves function to permit supplying fuel from one tank to both engines or, in the event of engine failure, from either tank to the operative engine. Each ball-type valve is motor-operated and electrically-controlled by a three-position LH TANK/OFF/RH TANK CROSSFEED selector switch on the lower left instrument panel. When positioned to “LH TANK” or “RH TANK,” both valves are simultaneously energized open by 28 VDC power supplied through corresponding 15-amp LH BOOST and RH BOOST circuit breakers on the left CB panel. When open, either or both engines are supplied with fuel from the selected tank. When positioned to “OFF,” both valves are simultaneously deenergized closed and each engine is supplied by its associated tank. Refer to the crossfeeding operational summary in this chapter for functional details. Note: During crossfeeding, the motive flow shutoff valve for the system not supplying fuel is energized closed by 28 VDC power supplied through the same circuit that opens the crossfeed valves.

FUEL SYSTEM 12/99


3-13

Fuel System Indication Indication of fuel quantity and flow are displayed by independent gages located on the upper center instrument panel. Fuel used and fuel remaining are displayed on optional gages, typically located on the lower copilot’s instrument panel. Fuel Quantity Indication The fuel quantity indicating system consists of capacitancetype probes, compensator modules, and a dual-channel, vertical-scale FUEL QTY gage. Five probes measure the level of fuel in each tank. The combined capacitance signal generated by the probes is transmitted to the corresponding channel of the gage where they are displayed as a measurement of fuel quantity in pounds (LBS). The compensator module, located in the sump area of each tank, modifies these signals to correct for changes in fuel temperature. The instrument scale is graduated in 100 pound increments between 0 to 3000 LBS. The position of independent left (L) and right (R) white vertical tape bars against the instrument scale, indicates the fuel quantity of the corresponding tank. 28 VDC left main bus power is supplied to the left fuel quantity indicating system through the 2-amp LH FUEL QTY circuit breaker. 28 VDC right main bus power is supplied to the right fuel quantity indicating system from the right main bus through the 2-amp RH FUEL QTY circuit breaker. On airplanes 5500550 and after, the LH FUEL QTY and RH FUEL QTY circuit breakers are located on the left and right CB panels respectively. On airplanes 550-0505 and earlier, the LH FUEL QTY and RH FUEL QTY circuit breakers are both located on the right CB panel.

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Cessna Citation II Technical Manual Low Fuel Level Warning A normally-open, float-type switch monitors the level of fuel in the sump area of each tank. When the fuel quantity in either tank decreases to between 169 and 219 pounds, the affected float switch closes thereby illuminating the corresponding amber [LH] or [RH] [FUEL LOW LEVEL] annunciator (550-0550 and after) or amber [L FUEL LEVEL LO] or [R FUEL LEVEL LO] annunciator (airplanes 550-0505 and earlier). The low fuel level warning system operates independently of the fuel quantity and optional fuel remaining indicating systems. Low Fuel Pressure Warning A normally-open, pressure switch monitors motive flow fuel pressure between each primary ejector pump and its associated engine-driven fuel pump. When this pressure falls below approximately 5 PSI, the affected fuel pressure switch closes thereby illuminating the corresponding amber [LH] or [RH] [FUEL LOW PRESS] annunciator (550-0550 and after) or amber [L FUEL PRESS LO] and [R FUEL PRESS LO] annunciator (5500505 and earlier), and activating the corresponding boost pump. The low fuel pressure warning system operates independently of all other fuel indicating systems. Refer to the fuel system operational summaries in this chapter for a complete description of low fuel pressure conditions.

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3-15

Fuel Flow Indication The fuel flow indicating system consists of fuel flow transmitters and a dual-channel, vertical-scale FUEL FLOW gage. The fuel flow transmitters are positioned in-line between the FCU and the oil-to-fuel heat exchanger of each engine. Each transmitter generates an electrical signal proportional to the rate of fuel flow to its associated engine. These signals are supplied to the corresponding channel of the gage where they are displayed as a measurement of fuel flow in pounds per hour (LBS/HR). The instrument scale is graduated in 100 pound per hour increments between 0 to 2000 LBS/HR. The position of independent left (L) and right (R) white vertical tape bars against the instrument scale, indicates the rate of fuel flow to the corresponding engine. Although the indicating range of the gage is 0 to 2000 LBS/HR, the operating range of each transmitter is approximately 145 to 1800 LBS/HR. Typical fuel flow rates at cruise power settings vary between approximately 450 and 650 LBS/ HR per engine depending on operating conditions. 28 VDC left main bus power is supplied to the left fuel flow indicating system through the 2-amp LH FUEL FLOW circuit breaker. 28 VDC right main bus power is supplied to the right fuel flow indicating system through the 2-amp RH FUEL FLOW circuit breaker. On airplanes 550-0550 and after, the LH FUEL FLOW and RH FUEL FLOW circuit breakers are located on the left and right CB panels respectively. On airplanes 550-0505 and earlier, the LH FUEL FLOW and RH FUEL FLOW circuit breakers are both located on the right CB panel. To prevent erratic indication at low engine power settings, each fuel flow indicating system channel is disabled by a corresponding throttle cutoff switch when its associated THROTTLE lever is positioned below approximately 10% N2.

3-16



Cessna Citation II Technical Manual Fuel Remaining/Consumed Indication Optional fuel remaining or fuel consumed indicators may be located on the lower copilot’s instrument panel if installed. On airplanes 550-0550 and after, a four-digit LCD readout displays the total FUEL LBS remaining from signals supplied by the left and right fuel quantity indicating channels. The indicator has a range of 0 to 6000 LBS. On airplanes 550-0505 and earlier, a four-digit electromechanical counter displays the total FUEL REMAINING in LBS. On airplanes 550-0062 through -0505, an additional four-digit electromechanical counter displays the total FUEL CONSUMED in LBS. Each of these indicators receives signals from the left and right fuel flow indicating channels and has a range of 0 to 9999 LBS. Before takeoff, the FUEL REMAINING indicator must be set to the known pounds of fuel on board the airplane. The FUEL CONSUMED indicator must be set to 0000. This is accomplished using the PRESET knob on the indicator face plate. Rotating this knob clockwise or counterclockwise correspondingly increases or decreases the displayed value. The normal rate of change is approximately 4 pounds per second in both directions. When the knob is pressed inward, the rate of change increases to approximately 20 pounds per second. The knob may be locked in this “fast” position when pressed fullyinward, and returned to the “slow” position when pulled outward. When released, the knob returns to its spring-loaded center position.

FUEL SYSTEM 12/99


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Operational Summaries The following operational summaries describe the function and interaction of fuel system components under particular operating conditions. Refer to the fuel distribution section of this chapter for specific component descriptions. Engine Starting During the engine starting sequence, activation and deactivation of the starter motor, ignition system, and boost pump is carried out by a series of relays in the start control circuit. When activated, the boost pump supplies fuel pressure directly to the engine. After light-off has occurred and the engine is operating at a speed sufficient to sustain combustion, the starter motor, ignition system, and boost pump are deactivated, thereby terminating the starting sequence. On airplanes 550-0406 and after, and earlier airplanes incorporating SB550-28-1, the starting sequence is terminated by the engine speed sensing circuit of the generator control unit (GCU) when N2 reaches approximately 40%. On airplanes 5500405 and earlier not incorporating SB550-28-1, the starting sequence is terminated by the motive flow pressure switch when the engine-driven fuel pump is supplying at least 180 ± 5 PSI motive flow to the primary ejector pump. This pressure is typically achieved when N2 reaches approximately 30%. Should the motive flow switch fail to open, the engine speed sensing circuit of the GCU will terminate the starting sequence when N2 reaches approximately 40%. Automatic boost pump activation will occur during engine start with the FUEL BOOST switches in the “NORM” or “OFF” position. Under most operating conditions, however, these switches should remain in the “NORM” position. When activated during engine start, each boost pump receives 28 VDC power through its associated 15-amp LH BOOST or RH BOOST circuit breaker, located in the aft fuselage electrical power “J” box, correspondingly supplied by the left or right main bus. Boost pump operation is indicated by illumination of the corresponding amber [LH] or [RH] [FUEL BOOST ON] annunciator (5500550 and after) or amber [L FUEL BOOST ON] and [R FUEL BOOST ON] annunciator (550-0505 and earlier). Note: The position of the LH BOOST and RH BOOST circuit breakers on the left CB panel in the flight compartment has no effect on boost pump operation during the engine starting sequence. 3-18



Cessna Citation II Technical Manual Normal Operation With the engine operating and the engine-driven fuel pump developing sufficient pressure to maintain motive flow, the primary ejector pump supplies the engine-driven fuel pump and transfer ejector pumps. En route to the engine-driven pump, fuel passes through the maintenance shutoff valve, fuel filter, and firewall shutoff valve. From the engine-driven pump, fuel is directed through the FCU, fuel flow transmitter, oil-to-fuel heat exchanger, and flow divider, to the fuel nozzles. Low Fuel Pressure To ensure uninterrupted fuel flow to the engine-driven fuel pumps, boost pump activation occurs automatically as a function of fuel pressure. When the output of either primary ejector pump falls below approximately 5 PSI, the corresponding fuel pressure switch closes thereby energizing its associated pressure switch relay and boost pump relay. With the boost pump relay closed, 28 VDC is supplied to the boost pump which then supplies fuel pressure to the engine-driven fuel pump. Closure of the fuel pressure switch is indicated by illumination of the corresponding amber [LH] or [RH] [FUEL LOW PRESS] annunciator (550-0550 and after) or amber [L FUEL PRESS LO] and [R FUEL PRESS LO] annunciator (550-0505 and earlier). Boost pump operation is indicated by illumination of the corresponding amber [LH] or [RH] [FUEL BOOST ON] annunciator (550-0550 and after) or amber [L FUEL BOOST ON] and [R FUEL BOOST ON] annunciator (550-0505 and earlier). The boost pump relay is initially energized through the pressure switch relay, but remains energized through an integral latching circuit. In this condition, the relay will remain energized closed and the boost pump will continue operating as long as the associated FUEL BOOST switch remains in the “NORM” position, regardless of fuel pressure. Should indication of boost pump operation exist without corresponding indication of low fuel pressure, the boost pump relay circuit should be reset by moving the switch to “ON” and returning it to “NORM.” Should indication of low fuel pressure accompany indication of boost pump operation, the associated FUEL BOOST switch should remain in the “NORM” position with the pump operating. Should indication of low fuel pressure exist without indication of boost pump operation, the associated FUEL BOOST switch should be set to the “ON” position.

FUEL SYSTEM 12/99


3-19

Automatic boost pump activation will occur during conditions of low pressure only when the FUEL BOOST switches are in the “NORM” position. When activated during conditions of low pressure, each boost pump receives 28 VDC power through its associated 15-amp LH BOOST or RH BOOST circuit breaker, located in the aft fuselage electrical power “J” box, correspondingly supplied by the left or right main bus. Each pressure switch relay and boost pump relay is energized closed by 28 VDC power through its associated 15-amp LH BOOST or RH BOOST circuit breaker, located on the left CB panel in the flight compartment. To prevent low fuel pressure boost pump activation at low engine power settings, each pressure switch relay is disabled by a corresponding throttle cutoff switch when its associated THROTTLE lever is positioned below approximately 10% N2. Crossfeeding Crossfeeding permits fuel to be supplied from one tank to both engines or, in the event of engine failure, from either tank to the operative engine. Under normal operating conditions, crossfeeding for the purpose of maintaining fuel load symmetry is seldom necessary unless asymmetry exceeds 200 LBS. When crossfeeding is necessary, the crossfeed selector switch should be positioned to the tank indicating the higher fuel quantity until fuel load symmetry is achieved. When either tank is selected, the corresponding boost pump is activated and both crossfeed valves are simultaneously energized open. Approximately one second is required for the crossfeed valves to fully open. During this time, a green [INTRANSIT] annunciator, located above the crossfeed selector switch, should be illuminated. Approximately three seconds after the crossfeed valves have fully opened, the non-selected tank motive flow shutoff valve is energized closed through a time delay relay. This time delay provides sufficient time for the crossfeed valves to fully open before motive flow is interrupted, and prevents non-selected tank boost pump activation due to low motive flow pressure. Should activation of both boost pumps occur, crossfeeding would be prevented due to equal pressure at each crossfeed valve. Therefore, indication of boost pump operation should be closely monitored when crossfeeding is initiated.

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Cessna Citation II Technical Manual If both boost pumps are activated, the non-selected tank boost pump should be deenergized by moving the associated FUEL BOOST switch to “ON” and returning it to “NORM.” When crossfeeding, indication of boost pump operation should occur immediately, correspond to the selected tank only, and be continuous until crossfeeding is terminated. When both crossfeed valves are open, either or both engines are supplied with fuel under boost pump pressure from the selected tank. Regardless of which tank is supplying the engine(s), a portion of crossfed fuel is returned to the nonselected tank through the transfer ejector pumps at a rate of approximately 600 LBS/HR. As a result, the indicated quantity of the non-selected tank will increase as the indicated quantity of the selected decreases. When fuel load symmetry is achieved, the crossfeed selector switch should be positioned to “OFF.” In this position, the nonselected tank motive flow shutoff valve is deenergized open. Approximately three seconds later, both crossfeed valves and the corresponding boost pump are deenergized. Approximately one second is required for the crossfeed valves to fully close. During this time, the green [INTRANSIT] annunciator should be illuminated. When both crossfeed valves are fullyclosed, each engine is supplied by its associated tank. Automatic boost pump activation will occur during crossfeeding with the FUEL BOOST switches in the “NORM” or “OFF” position. When activated during crossfeeding, the left boost pump, crossfeed valve, and motive flow valve are supplied with 28 VDC power from the left main bus (550-0550 and after) or right main bus (550-0505 and earlier) through the 15amp LH BOOST circuit breaker located on the left CB panel. The right boost pump, crossfeed valve, and motive flow valve are supplied with 28 VDC power from the right main bus (5500550 and after) or left main bus (550-0505 and earlier) through the 15-amp RH BOOST circuit breaker also located on the left CB panel.

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Fuel Distribution System Schematic (normal operation)

3-22

© PCW


© PCW

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Fuel Distribution System Schematic (left tank feeds both engines)

FUEL SYSTEM 12/99


Fuel Distribution System Schematic (right engine inoperative)

3-24

© PCW


© PCW

3-25



Fuel Distribution System Schematic (left fuel firewall valve shutoff closed)

FUEL SYSTEM 12/99

Limitations Refer to the applicable aircraft manufacturers FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable aircraft manufacturers FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.

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Chapter 5 Flight Controls Table of Contents Overview ............................................... 5-1 Control Wheels ..................................... 5-2 Ailerons ................................................ 5-3 Aileron Trim ...................................... 5-4 Elevators .............................................. 5-6 Elevator Trim .................................... 5-7 Electric Elevator Trim ......................... 5-9 Rudder Pedals .................................... 5-10 Rudder ............................................... 5-11 Rudder Trim .................................... 5-12 Control Lock ....................................... 5-14 Nosewheel Steering............................. 5-15 Wing Flaps ......................................... 5-18 Flap Actuation System ..................... 5-19 Flap Control .................................... 5-19 Flap Position Indication ................... 5-20


Flight Controls, continued Table of Contents Speed Brakes ..................................... 5-22 Speed Brake Hydraulics ................... 5-23 Speed Brake Control Valve .............. 5-23 Speed Brake Safety Valve ............... 5-23 Speed Brake Thermal Relief Valve ... 5-24 Speed Brake Switch ........................ 5-24 Stall Warning ...................................... 5-28 Stick Shaker ................................... 5-28 Stick Shaker Self Test .................... 5-28 Limitations .......................................... 5-29 Emergency Procedures......................... 5-29

Cessna Citation II Technical Manual Overview The flight controls of the Cessna Citation II consist of statically mass-balanced ailerons, elevators, rudder, and associated trim systems for each. The control surfaces are bearing supported and operated through conventional cable systems and mechanical linkage. Trim tabs are attached to the trailing edge of the left aileron, right elevator, and rudder. The elevator trim tab is positioned manually or by an electrically operated servoactuator. The rudder trim tab and aileron trim tab are positioned manually only. The manual trim controls and their associated mechanical position indicators are located on the center pedestal. This chapter also includes coverage of the wing flaps, speed brakes, nosewheel steering system, stick shaker, and stall warning system.

FLIGHT CONTROLS 12/99


5-1

Control Wheels

Dual control wheels for aileron and elevator operation are positioned on columns in front of the pilot’s and copilot’s seats. Control wheel rotation is transmitted to a drum around which the aileron control cables are attached. The control wheel and drum are installed on opposite ends of a common shaft that is bearing supported within a cover assembly attached to the top of each column. The control cables are guided by pulleys through the interior of the column to the aileron operating linkage. The control wheels are interconnected by crossover and synchronizing cables such that both rotate simultaneously. Each control column is bearing supported at its base and pivots about this point in response to fore and aft control wheel movement. The control columns are interconnected by a torque tube such that both move simultaneously. Control column movement is transmitted to the elevator operating linkage by a push-pull rod attached to the torque tube.

5-2



Cessna Citation II Technical Manual Ailerons

The ailerons are of aluminum alloy, semi-monocoque construction and are attached at two hinge points on the rear spar of each wing, outboard of the flaps. On airplanes 550-0049 and after, and earlier airplanes incorporating SB550-57-2, a small fence installed on the inboard edge of each aileron functions to maintain aileron effectiveness by reducing airflow spillage when the flaps are extended. Control wheel rotation is transmitted through cables and pulleys to an aileron sector assembly located below the passenger cabin floor. Rotation of the sector is transmitted through cables and pulleys to an aileron actuator assembly located within each wing. Each actuator assembly consists of a quadrant, yoke, and pivot/stop plate. The quadrant rotates on a sealed bearing at its center axis and provides attachment points for the aileron control cables. An additional sealed bearing pressed into the arm of the quadrant provides an off-center pivoting attachment point for the yoke which serves as the mechanical link between the quadrant and the aileron. As the quadrant rotates, the eccentric motion of the yoke positions the aileron up or down accordingly.



5-3

To prevent overtravel, aileron deflection between approximately 19° up and 15° down is limited by the quadrant arm’s range of motion between up and down travel limit stop bolts fitted to the pivot/stop plate. During autopilot operation, the aileron autopilot servo-actuator rotates the sector assembly through separate cables and pulleys, and sector rotation is transmitted back to the control wheels. In the event of malfunction, the servo-actuator may be overridden with control wheel pressure. Aileron Trim Aileron trim is provided by a trim tab attached to the inboard trailing edge of the left aileron by a full-length, pianotype hinge. The trim tab is actuated by a pair of push-pull rods attached to a dual jack screw type actuator installed within the left wing, forward of the aileron. The actuator is driven through a chain/cable assembly that is operated by the aileron trim control knob. Due to the positioning of the actuator, the trim tab moves in the opposite direction of aileron movement, thereby functioning as a servo-type trim tab and reducing the control forces required to position the aileron during flight. To prevent overtravel, aileron trim tab deflection between approximately 20° up and 20° down is limited by travel stop blocks fitted to the chain/cable assembly.

5-4



Cessna Citation II Technical Manual Ailerons and Aileron Trim

AILERON SECTOR ASSEMBLY

WING CABLES

FUSELAGE CABLES

COPILOT’S CONTROL WHEEL

PILOT’S CONTROL WHEEL

RIGHT CONTROL COLUMN CABLE ASSEMBLY

TURNBUCKLES

SYNCHRONIZING CABLES PULLEYS TRIM TAB HORN

INTERCONNECT CABLES AND TURNBUCKLE

AILERON CABLES

CABLE AND PULLEY ASSEMBLY

INBOARD ADJUSTABLE PUSHROD AILERON QUADRANT TRIM KNOB

AILERON CABLES

ACTUATOR CHAINS

AILERON TAB ACTUATOR CABLE AND PULLEY ASSEMBLY

CHAIN GUARD



5-5

Elevators

The elevators are of aluminum alloy, semi-monocoque construction and are attached at two hinge points each on the rear spar of the horizontal stabilizer. The left and right elevator are interconnected by torque tubes and an elevator horn thereby forming a single moving elevator assembly. Fore and aft control wheel movement is mechanically transmitted by a push-pull rod to an elevator sector assembly located below the flight compartment floor. Rotation of the sector is transmitted through cables and pulleys to an elevator bellcrank assembly located within the aft fuselage. Bellcrank rotation is mechanically transmitted to the elevator assembly by a pair of push-pull rods. An elevator bob weight (attached to the control column interconnect tube) and an elevator down spring (attached to the bellcrank) function to improve stability and elevator balance during flight. To prevent overtravel, elevator deflection between approximately 20° up and 15° down is limited by the sector’s range of motion between up and down travel limit stop bolts, and by the bellcrank’s range of motion between up and down limit stop blocks fitted to their corresponding support brackets.

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Cessna Citation II Technical Manual During autopilot operation, the elevator autopilot servo-actuator rotates the sector assembly and bellcrank assembly through separate cables and pulleys attached by clevis fittings to the elevator control cables within the aft fuselage. Sector rotation is transmitted back to the control column and in the event of malfunction, the servo-actuator may be overridden with control wheel pressure. Elevator Trim Elevator trim is provided by a trim tab attached to the inboard trailing edge of the right elevator by a full-length, piano-type hinge. The trim tab is actuated by a pair of push-pull rods attached to a dual jack screw type actuator installed within the right horizontal stabilizer, forward of the elevator. The actuator ▲ ELEVATOR TRIM CONTROL WHEEL is driven through chain/cable assemblies that are operated by the manual or electric elevator trim controls. An elevator trim sector assembly, located within the aft fuselage, provides a mechanical link between the actuator and the trim controls. The sector assembly rotates on a sealed bearing at its center axis and consists of two sectors which provide attachment points for the trim control chain/cable assemblies. One sector is round and provides attachment points for the chain/cable assembly operated by the trim controls. The other sector is elliptical and provides attachment points for the chain/ cable assembly that drives the actuator. The elliptical sector functions primarily to increase the speed of actuator response to nose up trim control input to expedite the reduction of elevator control forces during go-around. Actuator response to trim control input progressively increases in speed between the NOSE DN and NOSE UP ranges, and vice versa. In the neutral (T.O.) range, actuator response to trim control input is approximately equal. To prevent overtravel, elevator trim tab deflection between approximately 15° up and 17° down (550-0576 and earlier) or 17° up and 15° down (550-0577 and after) is limited by travel stop blocks fitted to the chain/cable assembly between the sector and the actuator. FLIGHT CONTROLS 12/99


5-7

Elevators and Elevator Trim

SPROCKET

ELEVATOR TRIM SECTOR CABLE/CHAIN ASSEMBLY

TURNBUCKLES

TRIM CONTROL WHEEL

PULLEYS

TAB HORN

ELEVATOR TRIM TAB AND PUSHROD ASSEMBLY

ELEVATOR TAB ACTUATOR TAB PUSHRODS

ELEVATOR HORN

ELEVATOR PUSHRODS

TORQUE TUBE

ELEVATOR BELLCRANK ASSEMBLY

ELEVATOR SECTOR ASSEMBLY BEARING

STOP BLOCK

ELEVATOR BELLCRANK

CABLE

5-8



Cessna Citation II Technical Manual Electric Elevator Trim 2 Electric elevator trim is provided by a servo-actuator 1 located within the aft fuselage. The servo-actuator consists of a motor that drives a clutch/ drum assembly around which the elevator nose down trim control cable is wrapped. Controls consist of a twopiece, momentary-on type trim switch labeled DOWN/UP, and a push-button type disconnect 1. Trim Switch switch labeled AP/TRIM DISC, 2. Trim Disconnect Switch located on the outboard horn of the pilot’s control wheel and copilot’s control wheel (if installed). When both halves of the trim switch are held in the “UP” or “DOWN” position, the servo-actuator motor and clutch are energized by 28 VDC power supplied from the left main bus (550-0550 and after) or right main bus (550-0505 and earlier) through the 5-amp PITCH TRIM circuit breaker. When energized, the clutch engages the drum and the motor rotates the assembly in the desired direction. When released, the motor and clutch are deenergized, and both halves of the trim switch return to their spring-loaded center (off) positions. On airplanes 550-0162 and after, when simultaneous or conflicting elevator trim commands are given by the pilot and copilot, the pilot’s commands override those of the copilot. In the event of electric elevator trim malfunction, the servoactuator may be overridden by pressing and holding the AP/ TRIM DISC switch until the PITCH TRIM circuit breaker, located on the left CB panel, is pulled. Until the fault is identified and corrected, the PITCH TRIM circuit breaker should remain pulled, the elevator should be trimmed manually, and the autopilot should not be used. Proper operation of the electric trim system can be checked prior to flight by confirming correct rotation of the elevator trim control wheel and position indicator in response to “UP” and “DOWN” trim switch input. While the trim switch is being held in the “UP” position and “DOWN” position, proper operation of the AP/TRIM DISC switch can be checked by pressing and holding the switch and confirming that the elevator trim control wheel and position indicator stop rotating. FLIGHT CONTROLS 12/99


5-9

Rudder Pedals

Pilot and copilot rudder pedals are provided for rudder operation, nosewheel steering, and airplane braking. Each rudder pedal is fitted to an arm assembly that is suspended from a pair of concentric, bearing supported torque tubes. The pilot’s and copilot’s left rudder pedals are connected to the inner torque tube; the pilot’s and copilot’s right rudder pedals are connected to the outer torque tube. The outer torque tube is comprised of two sections interconnected by a bridge assembly which provides travel clearance for the copilot’s left rudder pedal. Each torque tube incorporates link arms for the attachment of its associated rudder and nosewheel steering control cables. The inner and outer torque tubes are linked by an interconnect cable such that each pair of pedals moves simultaneously. Each rudder pedal is independently adjustable fore and aft by means of a lever located on the pedal arm. Pressing on the lower end of this lever disengages a spring-loaded locking pin from a hub that is fitted to the torque tube. With the locking pin disengaged, the pedal may be moved to one of three positions provided by the hub. Reengaging the locking pin secures the pedal arm in the selected position.

5-10



Cessna Citation II Technical Manual Rudder

The rudder is of aluminum alloy, semi-monocoque construction and is attached at three hinge points on the rear spar of the vertical stabilizer. Rudder pedal operation is transmitted through cables and pulleys to a bellcrank in the aft fuselage. The bellcrank is attached directly to a torque tube extending from the base of rudder. In addition to the three hinge points, the bellcrank pivots on a bearing/stop plate. To prevent overtravel, rudder deflection up to approximately 22° either side of center is limited by the bellcrank’s range of motion between left and right travel limit stop bolts fitted to the bearing/stop plate. During autopilot operation, the rudder autopilot servo-actuator rotates the bellcrank through separate cables and pulleys attached by clevis fittings to the rudder control cables within the aft fuselage. Bellcrank rotation is transmitted back to the rudder pedals and in the event of malfunction, the servo-actuator may be overridden with rudder pedal pressure.



5-11

Rudder Trim Rudder trim is provided by a trim tab attached to the trailing edge of the rudder by a full-length, piano-type hinge. The trim tab is actuated by a pair of push-pull rods attached to a dual jack screw type actuator installed within the vertical stabilizer, forward of the rudder. The actuator is driven through a chain/ cable assembly that is operated by the rudder trim control knob. Due to the positioning of the actuator, the trim tab moves in the opposite direction of rudder movement, thereby functioning as a servo-type trim tab and reducing the control forces required to position the rudder during flight. To prevent overtravel, rudder trim tab deflection up to approximately 10° either side of center is limited by travel stop blocks fitted to the chain/cable assembly.

5-12



Cessna Citation II Technical Manual Rudder and Rudder Trim TRIM TAB HORN

RUDDER TORQUE TUBE

RUDDER BELLCRANK

ADJUSTABLE PUSHROD

AUTOPILOT SERVO CABLE

LEFT CABLE

PULLEY AUTOPILOT SERVO CABLE

RUDDER TRIM ACTUATOR TURNBUCKLE

TRIM CONTROL WHEEL

SPROCKET

INTERCONNECT CABLE

CABLE/CHAIN ASSEMBLY

OUTER TUBE ASSEMBLY

RUDDER PEDAL ASSEMBLY

LEFT FORWARD CABLE


BEARING SUPPORT BRACKET

RIGHT FORWARD CABLE

PEDAL ASSEMBLY


5-13

Control Lock The control lock (also called the gust lock) prevents movement of and possible damage to the ailerons, elevator, and rudder by locking these surfaces in a neutral position and locking the THROTTLE levers in the "OFF" position. With the control lock engaged, nosewheel deflection is limited to 60°. The control lock system consists of a control handle, cable assemblies, push rods, slide plate, and bellcranks. Operationally, pulling the control lock handle moves the slide plate, which pulls the aileron, elevator, and rudder cables by capturing balls swaged onto the cables. The cables also move the bellcrank, which rotates the throttle locking cams into the locked position. When control surfaces reach the neutral position and the throttle cams reach the locked position, the control lock handle should reach its locking detent. Before pulling the control lock handle, the nosewheel should be centered, the aileron control wheels should be level, and the THROTTLE levers should be in the "OFF" position. The control lock is released by rotating the control lock handle 45° and lowering it into the release position. Warning: The control lock should be released before starting the engine. Caution: The control lock should be released before towing the airplane.

5-14



Cessna Citation II Technical Manual Nosewheel Steering The nosewheel steering system is operated by the rudder pedals and 1 allows airplane directional control during ground operations. Rudder pedal operation is 3 transmitted by cables to a bellcrank located within the nose 2 wheel well. Bellcrank movement is transmitted by a spring-loaded 1. Steering Bellcrank steering rod to a 2. Steering Rod steering arm that 3. Steering Arm operates a steering gear mechanism mounted atop the trunnion. The steering arm and gear mechanism are interconnected by a universal joint that automatically centers the nosewheel during retraction. The spring-loaded steering rod (bungee) allows the nosewheel to be positioned beyond rudder pedal travel limits when using differential braking or power, or when the airplane is being towed. The nosewheel steering gear mechanism is attached using shear bolts that are designed to protect nose gear components by breaking at a torque load exceeding 15,000 inch pounds, which corresponds to a left or right tow bar excursion of more than 60° left or right with the control lock engaged or more than 95° under any condition. Should the shear bolts break, the nose gear strut becomes free wheeling. Differential braking must then be used for steering control.



5-15

Steering is accomplished by allowing the aircraft to roll while displacing the appropriate rudder pedal. The nosewheel can be steered up to 20° left or right of center during taxi. When taxiing, the minimum wing tip turning radius using differential braking and partial power is 69.3 feet. This method causes excessive wear on the tires, and should be employed sparingly. If the airplane is parked with the nose wheel cocked to one side, initial taxiing should proceed with caution. The nosewheel steering gear mechanism is attached using shear bolts that are designed to protect nose gear components by breaking at a torque load exceeding 15,000 inch pounds, which corresponds to a left or right tow bar excursion of more than 95°. The airplane may be towed if the parking brake is not engaged. If the control locks are engaged, turning angle during tow is limited to 60° to avoid control lock damage. If the control locks are not engaged, turning angle during tow is limited to 95° to avoid breaking the steering gear shear bolts. Should the shear bolts break, the nose gear strut becomes free wheeling. Differential braking must then be used for steering control. Caution: Nose gear forced beyond the towing stop (95° limit) will shear bolts attaching steering gear assembly to cylinder.

5-16



Cessna Citation II Technical Manual Minimum Turning Radius (wing tip) AIRPLANES 550-0627 AND AFTER = 69’4" AIRPLANES 550-0626 AND EARLIER = 69’2 ‰ "

PIVOT POINT

17.6’

20.2’

AIRPLANES 550-0627 AND AFTER = 35’ AIRPLANES 550-0626 AND EARLIER = 35’2"

TURNING RADIUS: AIRPLANES 550-0627 AND AFTER = 34’8" AIRPLANES 550-0626 AND EARLIER = 34’7 ‰ "



5-17

Wing Flaps

Each wing is fitted with a single-piece, Fowler-type flap positioned by an electromechanical drive system and “track-androller” type operating linkage. When extended, the flaps move rearward and downward, effectively modifying wing camber and increasing wing area to reduce the stalling speed of the airplane. When retracted, the flaps form the trailing edge of the wing. Each flap is of aluminum alloy, semi-monocoque construction and is attached to the wing structure at three positions by its operating linkage. A pair of rollers is attached to the inboard and outboard ends, and lower middle surface of each flap. Each pair of rollers engages a corresponding inboard, outboard, and center track which extend aft from their attachment points on the wing structure. Three bracket assemblies form attachment points for inboard, center, and outboard bellcrank assemblies, each of which is linked to the flap by a push-pull rod.

5-18



Cessna Citation II Technical Manual Flap Actuation System 1 The flap operating linkage is electromechanically driven by 3 two 28 VDC motors, reduction gearing, and chain/cable assemblies that actuate the bellcranks in each wing. The chain/cable assemblies mate 2 with sprockets fitted to drive shafts that each engages its associated reduction gearbox. Each reduction gearbox is driven by its associated motor. 1. Bellcrank 2. Flap Track The drive shafts are linked by 3. Roller Pivot Point an interconnect chain such that they operate simultaneously. Normally, both motors and drive shafts operate together to position the flaps. Should one motor or drive shaft fail, the functioning motor or drive shaft should permit continued flap operation. Flap Control Flap position is controlled using the FLAP lever located on the center pedestal to the right of the THROTTLE levers. The FLAP lever can be set to any flap position between “FLAPS UP” (0°) and “LAND” (40°); the FLAP lever incorporates a mechanical detent at the “T.O. & APPR” (15°) position. Full flap extension is selected by pushing the FLAP lever fully down past the “T.O. & APPR” detent to the “LAND” position.



5-19

Flap Position Indication A flap position indicator is located on the center pedestal to the left of the FLAP lever. The indicator is mechanically connected to the flaps by a cable-operated sector assembly; the indicator therefore moves with the flaps and provides verification that the flaps have assumed the selected position. Down (extend) and up (retract) position switches are attached to the position indicator sector assembly. A cam on the FLAP lever actuates these switches. When the FLAP lever is positioned to extend the flaps, the cam contacts the down position switch, thereby energizing the flap actuator motors such that they extend the flaps. When the FLAP lever is positioned to retract the flaps, the cam contacts the up position switch, thereby energizing the flap actuator motors such that they retract the flaps. As the flap indicator moves to correspond with FLAP lever position, the sector assembly carries the respective position switch out of contact with the cam, thereby deenergizing the flap motors. Up and down flap limit switches function as backups to the position switch, and deenergize the flap motors when the flaps reach the fully-retracted or fully-extended position. The landing gear warning horn sounds if the FLAP lever is set below the T.O. & APPR position and the gear is not down and locked, regardless of airspeed or THROTTLE lever position. The horn is energized by one or more downlock switches within the landing gear actuators in conjunction with a flap approach switch incorporated within the FLAP lever assembly. The flap approach switch is actuated closed by the FLAP lever when set below the T.O. & APPR position (approximately 15°). The flap control circuit and actuator motors receive 28 VDC left main bus power through the 5-amp FLAP CONTROL and 15amp FLAP MOTOR circuit breakers respectively, each located on the left CB panel.

5-20



Cessna Citation II Technical Manual Flap Motor and Actuator Assembly

BRACKET AND PULLEY ASSEMBLY

RIGHT FLAP INTERCONNECT CABLES

RIGHT FLAP ACTUATION CABLES

TO FLAP INTERCONNECT ASSEMBLY

FLAP DRIVE MOTOR

LH RETURN CABLE ASSEMBLY

TO INBD BELLCRANK ACTUATOR

FLAP DRIVE GEARBOX

ACTUATION CABLE

OUTBOARD FLAP BELLCRANK

ACTUATION PULLEY ASSEMBLY



5-21

Speed Brakes

1

1. Operating Linkage 2. Limit Switch 3. Hydraulic Actuator

2

3 Hydraulically-operated, electrically-controlled speed brakes are located on the upper and lower surfaces of the wings, forward of the flaps. When extended, the speed brakes increase drag sufficiently to allow increased airplane rateof-descent without exceeding VMO/MMO. The speed brakes may also be extended during landing rollout, to spoil lift and provide aerodynamic braking. The speed brakes are of aluminum-reinforced magnesium alloy construction and are attached to the rear wing spar at five hinge points each. Operating linkage for each pair of speed brakes consists of a bellcrank, push-pull rods, and a hydraulic actuator. The push-pull rods link the upper and lower speed brakes to the bellcrank such that they operate simultaneously. The cylinder end of the hydraulic actuator is attached to the wing structure; the rod end is attached to the bellcrank. Normal operation is initiated by the SPEED BRAKE switch on the center pedestal. 5-22



Cessna Citation II Technical Manual Speed Brake Hydraulics Hydraulic pressure for speed brake operation is supplied by the same system that supplies the landing gear and thrust reversers. This section will describe the various valves and switches that control speed brake operation. Refer to Chapter 8 for a complete description of the hydraulic system. Speed Brake Control Valve The solenoid-operated speed brake control valve functions to direct hydraulic pressure to, and return flow from, the extend or retract ports of the actuators. To accomplish this, the control valve contains an internal selector spool that is spring-loaded to a neutral position and operated by independent extend and retract solenoids. In the neutral position, when both solenoids are deenergized, the extend and retract ports are blocked, trapping hydraulic pressure in the lines between the control valve and the actuators. When the extend solenoid is energized, the selector spool is positioned to direct hydraulic pressure to the extend ports of the actuators, and direct return flow from the retract ports of the actuators to the reservoir. Conversely, when the retract solenoid is energized, the selector spool is positioned to direct hydraulic pressure to the retract ports, and direct return flow from the extend ports to the reservoir. The solenoids are energized and deenergized primarily by the SPEED BRAKE switch through up (extend) and down (retract) limit switches. The limit switches are mechanically-actuated during speed brake extension and retraction. On airplanes 5500015 and after, and earlier airplanes incorporating SB550-27-2, the limit switches are actuated by the speed brake bellcranks and lower speed brakes. On airplanes 550-0014 and earlier not incorporating SB550-27-2, the limit switches are integral to the speed brake actuators. Speed Brake Safety Valve The speed brake safety valve, installed between the control valve extend port and the return line, functions to inhibit speed brake extension by relieving hydraulic pressure from the extend lines when engine speed is set above approximately 85% N2. The safety valve is normally energized closed, spring-loaded open, and controlled by throttle position switches. A check valve, installed downstream of the safety valve, prevents backflow from the return line to the extend lines.



5-23

Speed Brake Thermal Relief Valve The speed brake thermal relief valve, installed between the control valve retract port and the return line, functions to relieve hydraulic pressure in excess of approximately 1500 PSI from the retract lines to prevent system overpressure caused primarily by thermal expansion. A check valve, installed downstream of the relief valve, prevents backflow from the return line to the retract lines. Speed Brake Switch Speed brake operation is initiated by the SPEED BRAKE switch on the center pedestal. The switch is solenoid-operated, spring-loaded to the upper (RETRACT) position and requires electrical power to remain in the lower (EXTEND) position. The circuit that supplies electrical power to the solenoid is completed through the throttle position switches when the speed of both engines is set below approximately 85% N2. When the speed of either or both engines is set above approximately 85% N2, the circuit is interrupted, the switch returns to the "RETRACT" position, and speed brake extension is inhibited. Note: On airplanes 550-0231 and earlier not incorporating SB550-27-4, the throttle position switch circuit may be bypassed and the speed brakes extended by holding the SPEED BRAKE switch in the "EXTEND" position. Electrical components of the speed brake system are supplied with 28 VDC power from the left main bus (550-0550 and after) or right main bus (550-0505 and earlier) through the 5-amp SPEED BRAKE circuit breaker on the left CB panel. The control valve and safety valve function with an input power of 18 to 30 VDC. When electrical power is removed from the system, the speed brakes fail to the retracted position.

5-24



Cessna Citation II Technical Manual Speed Brakes

UPPER SPEEDBRAKE

PUSHROD

TRUNNION

HINGE PINS ACTUATOR RETRACT HOSE ATTACH 90 PORT EXTEND HOSE ATTACH 45 PORT

BELLCRANK ACTUATOR PUSHROD LUG

ACTUATOR HOUSING



5-25

Speed Brake Extension Schematic

© PCW

Speed brake extension is initiated by setting the SPEED BRAKE switch to "EXTEND." In this position, the control valve extend solenoid, safety valve, and hydraulic system center valve are energized to permit speed brake extension. During extension, the extend solenoid and center valve are energized through the up limit switches. When the speed brakes are fully-extended and the up limit switches are open, the extend solenoid and center valve are deenergized and the white [SPEED BRAKE EXTEND] annunciator (550-0550 and after) or [SPD BRAKE EXTENDED] annunciator (550-0505 and earlier) is illuminated. In this condition, the hydraulic system returns to "open center" mode and the control valve selector spool returns to its neutral position, thereby trapping hydraulic pressure in the extend lines and holding the speed brakes in the selected position. Note: Setting the speed of either or both engines above approximately 85% N2 with the speed brakes extended will cause them to retract. Note: When the center valve is energized and hydraulic pressure is being supplied to the speed brakes, the amber [HYD PRESS ON] annunciator will be illuminated.

5-26



Cessna Citation II Technical Manual Speed Brake Retraction Schematic

© PCW

Speed brake retraction is normally initiated by setting the SPEED BRAKE switch to "RETRACT." In this position, the control valve retract solenoid, safety valve, and hydraulic system center valve are energized to permit speed brake retraction. During retraction, the retract solenoid and center valve are energized through the down limit switches. When the speed brakes are fully-retracted and the down limit switches are open, the retract solenoid and center valve are deenergized and the white [SPEED BRAKE EXTEND] or [SPD BRAKE EXTENDED] annunciator is extinguished. In this condition, the hydraulic system returns to "open center" mode and the control valve selector spool returns to its neutral position, thereby trapping hydraulic pressure in the retract lines and holding the speed brakes in the selected position. On airplanes 550-0014 and earlier not incorporating SB550-272, a self-locking device integral to each actuator secures its associated speed brake in the retracted position. On airplanes 550-0015 and after and earlier airplanes incorporating SB55027-2, each lower speed brake incorporates two spring-loaded retainers that function to prevent droop when the actuators are depressurized following retraction.



5-27

Stall Warning Stall warning is provided by stall strips installed on the inboard section of each wing deice boot. These strips disrupt airflow, thereby providing aerodynamic warning of an impending stall by inducing prestall buffet. Aerodynamic prestall warning buffet commences at an airspeed of approximately VS1 +10 in the clean configuration and VSO +5 in the landing configuration. Stick Shaker The optional stick shaker functions to provide warning of an impending stall by imparting a low-frequency vibration to the pilot’s control column when the AOA system senses an impending stall. The stick shaker is mounted on the forward side of the pilot’s control column and consists of an electric motor, rotating weights, stick shaker relay, resistor, and test switch. In-flight activation of the stick shaker is initiated by a signal from the AOA system that closes an angle-of-attack stall warning switch in the AOA indicator and energizes the stick shaker relay. With the relay energized, 28 VDC left main bus power is supplied to the motor through the 5-amp ANG OF ATTACK circuit breaker. The resistor regulates power to maintain a stick shaker frequency of between 23 and 24 Hz. Activation of the stick shaker during ground operations is inhibited by the left main gear safety switch except during system testing. Stick Shaker Self Test Operation of the stick shaker may be verified by rotating the TEST selector switch on the lower left instrument panel to the “STICK SHAKER” position. If the system is functioning normally, the AOA indicator should flag and drive to zero. The indicator flag should then pull from view and the indicator needle should drive to 1.0. As the needle passes 0.75, the stick shaker should activate for several seconds. The cycle should repeat until the TEST selector switch is repositioned.

5-28



Cessna Citation II Technical Manual Limitations Refer to the applicable aircraft manufacturer’s FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable aircraft manufacturer’s FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.



5-29


Chapter 6 Electrical System Table of Contents Overview ............................................... 6-1 DC Power Sources ................................ 6-2 Battery ............................................. 6-2 Battery Overheat Warning .................. 6-3 Starter/Generators ............................ 6-4 Generator Control Units ..................... 6-5 Starter/Generator Ground Cooling ...... 6-7 External Power System .......................... 6-8 External Power Requirements ............. 6-8 Overvoltage/ Overcurrent Protection ....................... 6-9 DC Power Distribution ......................... 6-10 Hot Battery Bus .............................. 6-10 Battery Bus .................................... 6-10 Left and Right Main Busses ............ 6-11 Emergency Bus ............................... 6-11 Battery Switch................................. 6-12 Starter/Generator Switch ................. 6-13 DC System Indication .......................... 6-14 Voltmeter ........................................ 6-14 Ammeters ....................................... 6-14 DC Circuit Protection ........................... 6-15 Circuit Breakers .............................. 6-15 Current Limiters .............................. 6-15


Electrical System, continued Table of Contents Bus Distribution Tables ....................... 6-34 Alternating Current (AC) Power System.............................. 6-44 Dual Split Bus AC System ................... 6-44 Inverters ......................................... 6-44 AC Control ...................................... 6-44 Inverter Testing ............................... 6-45 “Tied” Split Bus AC System ................ 6-46 Inverters ......................................... 6-46 AC Control ...................................... 6-46 Inverter Testing ............................... 6-47 Single Bus System .............................. 6-48 Inverters ......................................... 6-48 AC Control ...................................... 6-48 Inverter Testing ............................... 6-48 AC Circuit Protection ........................... 6-51 Lighting .............................................. 6-52 Exterior Lighting .................................. 6-52 Taxi/Landing Lights ......................... 6-54 Wing Recognition Lights................... 6-54 Rotating/Flashing Beacon ................ 6-55 Wing Inspection Light ...................... 6-55 Tail Floodlights ................................ 6-55


Electrical System, continued Table of Contents Interior Lighting ................................... 6-56 Control and Instrument Lighting ....... 6-56 Instrument Panel Floodlights ............ 6-56 Vertical Scale Engine Instrument Light .............................. 6-56 Counter Light .................................. 6-57 Cathode Tube Lights ....................... 6-57 Map Lights ..................................... 6-57 Indirect Cabin Lights ....................... 6-58 Overhead Console Sign ................... 6-58 Entrance Lights ............................... 6-58 Emergency Exit Sign ........................ 6-59 PSU Light ....................................... 6-59 Passenger Reading Lights................ 6-59 Aft Fuselage Interior Light ................ 6-60 Baggage Compartment Lights .......... 6-60 Limitations .......................................... 6-61 Emergency Procedures......................... 6-61

Cessna Citation II Technical Manual Overview This chapter describes systems that supply and control airplane electrical power. Interior and exterior airplane illumination will also be discussed. The Cessna Citation II electrical system is powered by two 28 Volt Direct Current (VDC), 400-ampere (amp), negativeground, engine-driven starter/generators. A 24 VDC nickelcadmium battery provides current for engine starting and serves as an emergency source of power. An external power receptacle located below the left engine pylon allows an external power source to energize the airplane for ground operations or engine start. Electrical power from these sources is distributed to the airplane’s systems through a multiple bus arrangement designed to provide continued operation in the event of an electrical source failure. Items in the system requiring alternating current (AC) for their operation are powered by two inverters rated at 300 to 600 volt-amp (VA).

▲ PRIMARY ELECTRICAL SYSTEM CONTROLS AND INDICATORS

ELECTRICAL SYSTEM 12/99


6-1

DC Power Sources The section describes the sources of airplane DC power, including the battery, starter/generators, and external power system. Battery The 20-cell 44 amp2 hour (550-0550 and after) or 19-cell 40 amp-hour (550-0505 and earlier) nickelcadmium (NiCad) battery is secured by a hold-down clamp to a battery tray mounted within the 1 aft fuselage. The amp-hour rating 3 indicates that when new and fully charged, the battery 1. Battery Vent Tube is capable of deliver2. Battery Hold Down ing one amp of 3. Battery Tray current for 44 hours (in the case of 44 amp-hr battery), before reaching a fully discharged condition. The following formula: amp-hr rating ÷ amp load = hours available, may be used to approximate hours of battery power available in the event of a dual generator failure. In the case of a 44 amp-hr battery, this is accomplished as shown in the following example: 44 amp-hr ÷ 88-amp load = 0.5 hours. Note: A nickel-cadmium battery will maintain a constant output voltage during approximately 90 percent of its discharge cycle, after which available power will rapidly deteriorate. The electrolyte in a NiCad battery serves only as a conductor and does not react with the battery plates. Because of these characteristics, the condition of a NiCad battery cannot be reliably determined by voltage checks or specific gravity readings. However, the plates within a NiCad battery absorb electrolyte as the battery discharges, so battery condition can be approximated by observing electrolyte level.

6-2



Cessna Citation II Technical Manual Electrolyte level in the NiCad battery should be 1/8 inch above the visible insert or plates two to four hours following a full charge, or 1/4 inch above the visible insert or plates immediately after full charge. To prevent the accumulation of fluid and vapors, the battery box is equipped with independent drain and vent tubes that extend through the lower surface of the aft fuselage. It is recommended that the battery electrolyte level be checked every 100 flight hours or every 14 days, whichever occurs first. Distilled water should be used when servicing is required. Battery Overheat Warning Indication of excessive battery temperature is provided by the battery overheat warning system. It consists of a battery temperature sensor, temperature module, and battery overheat annunciator. The temperature sensor is installed between the cells near the center of the battery and provides input to the temperature module. Bat▲ BATTERY TEMPERATURE GAGE tery temperature should remain below 145°F. Should battery temperature exceed 145 to 160°F, the temperature module will cause illumination of the red [BATT O’TEMP] (airplanes 550-0550 and after) or [BATT O’HEAT] (550-0505 and earlier) annunciator. If battery temperature exceeds 160°F, the [BATT O’TEMP] or [BATT O’HEAT] annunciator will flash. A battery temperature gage may be optionally installed on airplanes 550-0626 and earlier and is standard on airplanes 550-0627 and after. The battery overheat warning system receives 28 VDC left main bus power through the 2-amp BATT TEMP circuit breaker on the left CB panel. Note: The battery should be serviced if battery temperatures exceeding 145°F are indicated. Battery overheat warning system operation may be verified by rotating the TEST selector switch, located on the lower left instrument panel, to the “BATT TEMP” position. In this position, a test mode is activated that simulates a temperature exceeding 160°F, thereby causing the [BATT O’HEAT] or [BATT O’TEMP] annunciator to flash.



6-3

Starter/Generators The starter/generators supply 28 VDC electrical power to all distribution busses, provide charging current to the battery, and are also used as engine starting motors. The starter/generators are each capable of producing a continuous current of 400 amps.

The starter/generators can each sustain an overload of up to 600 amps for five minutes. On airplanes 550-0550 and after, and earlier airplanes incorporating SB550-54-4, sustained generator load is limited to 325 amps above 35,000 feet. On airplanes 550-0505 and earlier not incorporating SB550-54-4, sustained generator load is limited to 250 amps above 25,000 feet. Each generator is capable of powering the entire airplane electrical system with the exception of the optional Freon air conditioning system, which should be turned off in the event of generator failure.

6-4



Cessna Citation II Technical Manual Generator Control Units Once on-line and supplying power, each generator’s output is controlled by an independent generator control unit (GCU). The GCU provides voltage regulation, load sharing, ground fault protection, overvoltage and overexcitation protection, as well as differential voltage and reverse current protection. The GCU also regulates field excitation and starter cut-off functions during engine start. Voltage Regulation Voltage regulation is accomplished by an integrated circuit which essentially compares actual generator output to a regulated reference voltage (28.5 VDC ± 1 VDC). When sensing a differential between these two voltages, the integrated circuit regulates field excitation thereby regulating generator output. The circuit also incorporates a field relay which operates in conjunction with its respective GCU’s control relay. Input to each GCU relative to the output of its respective generator is routed through the 10-amp LH GEN SENSE and RH GEN SENSE circuit breakers (not accessible from the flight compartment.) Load Sharing Load sharing is accomplished by an equalizer connection between the left and right GCUs and an integrated equalizer circuit resident to each GCU. This circuit essentially “regulates” voltage regulator output to maintain load sharing within ± 40 amps (under normal operating conditions). The circuit incorporates an equalizer relay which operates in conjunction with its respective GCU control relay. Essentially, whenever the control relay is deenergized, the equalizer relay is deenergized thereby interrupting the equalizer circuit and isolating a “tripped” generator should a ground fault occur.



6-5

Reverse Current Protection Differential voltage and reverse current protection are provided by each GCU through independent sensing circuits. Each differential voltage sensing circuit enables closure of its associated power relay (permitting current flow from the generator to its associated main bus) when the output voltage of the generator is within .30 volts of its associated main bus. Input to each GCU relative to the voltage of its corresponding main bus is through the 2-amp LH BUS SENSE and RH BUS SENSE circuit breakers (not accessible from the flight compartment). Once the power relay is closed, reverse current protection is enabled thereby preventing a failing generator from imposing a load on the other. In this condition, when generator output falls 10% or more below its rated output, the generator is taken off-line until output is restored to a level which will ensure forward current flow to its corresponding main bus. Overvoltage/Overexcitation Protection Overvoltage and overexcitation protection is provided by each GCU in conjunction with its voltage regulation and load sharing equalizer circuits. Should either GCU’s voltage regulation circuit fail, generator output will increase to 35 VDC and an overvoltage integrator will trip the associated field relay after a predetermined period of time thereby taking the generator off-line. When the generators are paralleled and sufficiently loaded, overvoltage may not occur; however, a malfunctioning voltage regulation circuit can result in its associated generator assuming a greater percentage of the load. When this occurs, a “deexcitation” signal is provided by the equalizer circuit to the voltage regulation circuit and the overvoltage integrator thereby taking the generator off-line. Field Weakening The GCU field weakening feature regulates field excitation to assure that the starter/generator operates as a starter and does not generate power during engine start. This function is accomplished by controlling field excitation so as to maintain starter/generator interpole winding current below the level needed to initiate power generation. The field weakening circuitry initiates current regulation when starter/generator interpole winding current drops below the field weakening threshold value, and continues until starter speed reaches the cut-off value (approximately 40% N2). To avoid nuisance trips, all other GCU protection functions are disabled during engine start. 6-6



Cessna Citation II Technical Manual Ground Fault Protection Ground fault protection is provided by a “feeder to ground short” sensing circuit integral to each GCU. After engine start and during initial generator “build-up” (between opening of the start relay and closing of the control relay), if the generator begins to assume a load equal to or greater than its rated output before the control relay has closed, a ground short will be sensed. In this condition, closure of the control relay and continuous excitation of the generator are prevented until “tripping” (opening) of the start relay occurs. After the control relay has closed and the generator is on-line, occurrence of a ground fault condition will cause its associated field relay to trip open thereby taking the generator off-line. Additional protection against generator build-up with an open field relay is a function of the voltage regulation reset circuit. This circuit operates in conjunction with the RESET position of the generator switches. Essentially, generator build-up with an open field relay cannot occur until the switch is momentarily positioned to “RESET.” With the field relay reset, the reset circuit is isolated such that build-up cannot occur if the generator is reset into a ground fault condition. Starter Cut-Off A starter cut-off circuit, integral to each GCU, functions to terminate the start sequence as a function of engine speed (approximately 40% N2). Speed sensing is provided by a tach drive integral to the starter/generator. The field weakening circuitry of each GCU, functions to regulate field excitation during the engine start sequence until starter cut-off occurs. Should the starter cut-off circuit malfunction below cut-off speed, the power relay will be closed, the equalizer relay will be open, and the starter will continue motoring the engine until manually disengaged. Starter/Generator Ground Cooling When operating on the ground, each starter/generator is cooled by an internal fan connected to the generator shaft. The fan draws cooling air through an inlet scoop and duct, located in the lower forward engine cowling, after which it is exhausted through an outlet in the lower cowling. In flight, starter/generator cooling is primarily accomplished by ambient air entering through the inlet scoop.



6-7

External Power System The external power system consists of an external power receptacle located below the left engine pylon and an external power relay located in the aft fuselage electrical power junction “J” box. The receptacle is a three-pin type design with permanent markings identifying the positive and negative pins. The relay is energized closed upon application of external power, thereby permitting electrical flow to the hot battery bus.

▲ EXTERNAL POWER RECEPTACLE

With the battery switch set to “BATT” and the battery relay closed, external power is made available to the distribution system and charging current is supplied to the battery. During engine start using external power, a battery disconnect relay opens, thereby isolating the battery from the distribution system. With 28 VDC external power applied, battery power is conserved during engine starting, or when testing electrical equipment on the ground. External Power Requirements Maximum external power source output should be at least 28 VDC and no more than 1,000 amps. Caution: If an external power source without reverse current protection is turned off while connected to the airplane, rapid battery discharge and battery damage can result. If the external power source is turned off, it should be disconnected from the airplane.

6-8



Cessna Citation II Technical Manual Note: External power engine starts may be accomplished with the generator switches positioned to either “ON” or “OFF;” however, it is recommended that the generator switches be positioned to “OFF” during external power engine start. If the generator switch is in the “ON” position, the generator will come on-line automatically upon completion of the start sequence. If the generator switch is in the “OFF” position, the generator must be brought on-line by positioning the generator switch to “ON.” When the generator begins to supply power to the DC bus, an external power disable relay automatically disconnects external power. Therefore, the generator switch for the operating engine must be positioned to “OFF” to start the second engine using external power. Overvoltage/Overcurrent Protection The overvoltage/overcurrent protection system prevents damage to the starter/generators, avionics equipment, lights, and/or other electrical equipment if external power source voltage and/ or current exceeds limits. The system also prevents battery current from augmenting external power source current, which would cause an electrical overload. The system includes an overvoltage/overcurrent monitor and a current sensor. The monitor disconnects external power from the airplane electrical system if external power voltage remains between 32 and 33 VDC for more than 200 milliseconds. The monitor also operates in conjunction with the current sensor to protect against overcurrent. The current sensor compares external input voltage to a reference voltage supplied by the monitor. When the current sensor detects external power current exceeding 1100 to 1300 amps for 1.7 to 2.3 seconds, the monitor disconnects the airplane electrical system from external power.



6-9

DC Power Distribution Electrical power is distributed to the airplane’s systems through interconnected busses normally supplied by the battery or the starter/generators. The busses may also be supplied by an external power source. Hot Battery Bus The hot battery bus is connected directly to the battery, and is connected to the battery bus and the emergency bus through their associated relays. External power is also supplied directly to the hot battery bus. Battery Bus Battery power is routed from the hot battery bus to the battery bus through the battery relay, which is controlled by the battery switch. When the battery switch is in the “BATT” position and battery voltage is at least 17 volts, the battery relay is energized closed and battery power is supplied to the battery bus. When set to “OFF” the battery is isolated from all but the hot battery bus. When set to “EMER” the battery is isolated from all but the hot battery bus and the emergency battery bus. With both engines operating and both generators on-line, the battery bus is supplied with 28 VDC power from the left and right main busses. With the starter/generators or an external power source supplying 28 volts and the battery supplying 24 volts or less, current flow reverses, thereby charging the battery. Note: If there are no indications of battery power availability to the system with the battery switch in the “BATT” or “EMER” positions, battery service may be required. Note: Charging current from the starter/generator will be available to the battery only when the battery switch is set to “BATT.”

6-10



Cessna Citation II Technical Manual Left and Right Main Busses Electrical power may be supplied to the left and right main busses by the battery, the starter/generators, or an external power source. Battery power or external power are supplied to the main busses through the battery bus. When the starter/ generators are on-line, 28 VDC power is routed directly to the main busses through their corresponding power relays. The left and right main busses are tied together by the battery bus through corresponding 225-amp current limiters. Each main bus supplies its associated circuit breaker panel through three 80-amp current limiters and three 75-amp circuit breakers. The majority of the airplane’s electrical components receive their power from these busses. Emergency Bus The emergency bus is powered by the battery when the battery switch is set to “BATT” or “EMER”. When in the “EMER” position, the battery relay is opened, disconnecting the main DC busses and the battery bus from the emergency bus. Use of the emergency bus enables critical airplane components to be powered by the battery but electrically isolated from malfunctioning starter/generator(s) or other components. The emergency bus circuit is protected by a 20-amp EMER POWER circuit breaker (not accessible from the flight compartment).



6-11

Battery Switch The battery switch primarily controls the battery relay through which battery power is supplied to the battery bus or charging current is supplied to the battery. The battery switch also controls the emergency relay through which power is supplied to the emergency bus. When set to the upper “BATT” position, the battery relay and emergency relay are simultaneously energized closed thereby connecting the battery to the battery bus and the emergency bus to the hot battery bus (550-0626 and earlier) or to the battery bus (550-0627 and after). When set to the center “OFF” position, the battery relay and emergency relay are deenergized open thereby isolating the battery from all but the hot battery bus. When set to the lower “EMER” position, the battery relay is deenergized open while the emergency relay remains energized closed thereby isolating the battery from all but the hot battery bus and the emergency bus. In this condition, no more than 30 minutes of battery power is available to the emergency bus.

6-12



Cessna Citation II Technical Manual Starter/Generator Switch Each starter/generator switch, located on the lower left instrument panel, controls the operating state of its associated starter/generator. When set to the upper (on) position each generator switch supplies an “on-line” signal to its associated GCU thereby activating its generator mode functions. When either switch is set to the center “OFF” position, the online signal is interrupted. The momentary “RESET” position of each generator switch functions to restore the signal supplied to the GCU. It is sometimes necessary to set the generator switch to “RESET” following a windmilling airstart of an engine.



6-13

DC System Indication Indication of electrical system operation is provided by a voltmeter and two ammeters located on the lower left instrument panel, and by annunciators located on the upper center instrument panel.

Voltmeter The voltmeter provides a means of monitoring starter/generator output voltage or battery bus voltage as determined by the position of a selector switch adjacent to the meter. When set to “LH GEN” (left starter/generator), or “RH GEN” (right starter/ generator), the voltage of the selected source is indicated. When set to “BATT,” electrical system voltage monitored at the battery bus is indicated. The voltmeter scale is graduated in 1-volt increments between 10 and 40 D.C. Volts, with numerical values marked at each 10volt increment. Ammeters Independent left and right ammeters indicate the load carried by each generator. The ammeter scale is graduated in 50-amp increments from 0 to 400 amps, with numerical values marked at each 100-amp increment.

6-14



Cessna Citation II Technical Manual DC Circuit Protection Individual electrical system circuits and components are protected against overload by circuit breakers and current limiters. Circuit Breakers Most circuit breakers are located on left and right CB panels on the flight compartment sidewalls so as to be readily accessible to the flight crew. The applicable amperage is marked on each circuit breaker. Labels above each breaker identifies the circuit protected. These push-to-reset type circuit breakers will pop out, or “trip,” when heat is generated by an electrical overload. Should an overload occur, a tripped circuit breaker may be reset after a cooling period of approximately three minutes by pushing it back in. If the circuit breaker trips a second time, a short circuit is indicated and it should not be reset, as this could cause system damage. Additional circuit breakers are installed in various electrical power junction “J” boxes located within the aft fuselage. Current Limiters Primary bus tie circuit protection is provided by a 225-amp current limiter in-line between the battery bus and the left and right main DC busses. Three 80-amp bus feeder current limiters are wired in parallel between each main bus and its associated circuit breaker panel bus. Note: The flight crew should ensure that all circuit breakers are engaged and serviceable fuses are installed before all flights. The airplane should never be operated with any disengaged circuit breakers or open current limiters without a thorough knowledge of the consequences.



6-15


Electrical Schematic (hot battery bus energized, 550-0626 and earlier)

6-16

© PCW


© PCW

6-17



Electrical Schematic (battery powering system, 550-0626 and earlier)



Electrical Schematic (left engine starting, 550-0626 and earlier)

6-18

© PCW


© PCW

6-19



Electrical Schematic (generator assisted right engine start, 550-0626 and earlier)




© PCW

Electrical Schematic (both engines operating, both generators on-line, 550-0626 and earlier)

6-20

© PCW

6-21



Electrical Schematic (external power applied, 550-0626 and earlier)



Electrical Schematic (left generator failed, 550-0626 and earlier)

6-22

© PCW


© PCW

6-23



Electrical Schematic (right main extension bus overload, 550-0626 and earlier)




© PCW

Electrical Schematic (battery switch in emergency position, 550-0626 and earlier)

6-24

© PCW

6-25



Electrical Schematic (hot battery bus energized, 550-0627 and after)



Electrical Schematic (battery powering system, 550-0627 and after)

6-26

© PCW


© PCW

6-27



Electrical Schematic (left engine starting, 550-0627 and after)



© PCW


Electrical Schematic (generator assisted right engine start, 550-0627 and after)

6-28



© PCW

6-29

Electrical Schematic (both engines operating, both generators on-line, 550-0627 and after)



Electrical Schematic (external power applied, 550-0627 and after)

6-30

© PCW


© PCW

6-31



Electrical Schematic (left generator failed, 550-0627 and after)



© PCW


Electrical Schematic (right main extension bus overload, 550-0627 and after)

6-32

© PCW

6-33



Electrical Schematic (battery switch in emergency position, 550-0627 and after)


DC Bus Distribution Table (550-0550 and after) Electrical Power Junction Box

6-34

Power Source and Equipment

Circuit Breaker

Hot Battery Bus Aft & Nose Baggage Compartment Lights Emergency Lights Engine Ignition (during start) Emergency Power

AFT/FWD COMP LT EMER LT IGNITION EMERGENCY POWER

Battery Bus Battery Voltmeter

BATT VOLTAGE

2

Left Isolation Bus Left Generator Ammeter Left Generator Sense Left Engine Start Light Left Generator Voltmeter

LH AMMETER (2) LH GEN SENSE LH START LT LH VOLTMETER

2 10 2 2

Right Isolation Bus Right Generator Ammeter Right Generator Sense Right Engine Start Light Right Generator Voltmeter

RH AMMETER (2) RH GEN SENSE RH START LT RH VOLTMETER

2 10 2 2

Left Main Bus Left Bus Sense Left Fuel Boost Pump Left Generator Off Light Left Landing Light Passenger Advisory Lights Left Recognition Light Indirect Cabin Lights Entertainment Center Tail Flood Lights

LH BUS SENSE LH BOOST LH GEN OFF LH LDG LT OXY/SEAT BELT LH RECOG LT INDIRECT LT ENT CTR TAIL LIGHTS

2 15 2 15 5 5 7.5 5 5

Right Main Bus Right Bus Sense Right Fuel Boost Pump Right Generator Off Light Right Landing Light Right Recognition Light Cabin Lights Toilet/Shaver Outlet

RH BUS SENSE RH BOOST RH GEN OFF RH LDG LT RH RECOG LT CABIN LT TOILET

2 15 2 15 5 7.5 7.5


Amperage

3 5 7.5 20


Cessna Citation II Technical Manual DC Bus Distribution Table (550-0550 and after) Left Circuit Breaker Panel



Circuit Breaker

Left CB Panel Left Engine Fan Speed Digital Indicator Left Engine Turbine Speed Digital Indicator Left 5V Panel Lights AC Inverter 1 Anti-Collision Light AOA System Control AOA System Heater Battery Overtemp System Cabin Fan Cockpit Voice Recorder Left Digital Clock Electroluminescent Panel Lights Left Engine Control Engine Synchronization System Right Engine Fire Detection Right Engine Firewall Shutoff Flap Control Flight Data Recorder Right Fuel Boost Pump Left Engine Fuel Flow Indicator Left Fuel Quantity Indicator Right Engine Ignition System Left Engine ITT Indicator Landing Gear Control Landing Gear Warning Left Engine Start Control Wing Inspection Light Nose Wheel Spinup System Outside Air Temperature Left Engine Oil Pressure Indicator Left Engine Oil Temp Indicator Pitch Trim Control Left Pitot/Static Heater Power Brakes and Anti-Skid Control Normal Pressurization System Right Circuit Breaker Panel Rotating Beacon Speed Brake Control Standby Gyro True Airspeed Probe Heater (Sperry) Cabin Temperature Control Left Engine Thrust Reverser Control Warning Lights 1 Windshield Bleed Air Control Windshield Bleed Air Power Wing Navigation Lights Cockpit Voice Recorder Flap Motors

LH FAN SPEED LH TURB SPEED LH PANEL AC INVERTER NO 1 ANTICOLL ANG OF ATTACK AOA HTR BATT TEMP CABIN FAN VOICE RECORDER LH CLOCK EL PANEL LH ENG ENGINE SYNC RH FIRE DET RH FW SHUTOFF FLAP CONTROL FLIGHT RECORDER RH BOOST LH FUEL FLOW LH FUEL QTY RH IGN LH ITT GEAR CONTROL LDG GEAR LH START WING INSP NOSE WHL RPM OAT LH OIL PRESS LH OIL TEMP PITCH TRIM LH PITOT STATIC SKID CONTROL NORM PRESS RH CB PANEL ROTATING BEACON SPEED BRAKE STBY GYRO TAS HTR TEMP LH THRUST REVERSER WARN LTS 1 W/S BLEED AIR TEMP W/S BLEED AIR NAV VOICE RECORDER FLAP MOTOR


Amperage

2 2 5 15 7.5 5 5 2 20 5 2 1 7.5 5 2 7.5 5 5 15 2 2 7.5 2 5 2 7.5 5 2 2 2 2 5 7.5 20 5 35 5 5 5 5 5 7.5 2 15 5 5 5 15

6-35

DC Bus Distribution Table (550-0550 and after) Left Circuit Breaker Panel

6-36


Circuit Breaker

Left Main Bus Left CB Panel Bus Left CB Panel Bus Left CB Panel Bus

LH BUS NO 1 LH BUS NO 2 LH BUS NO 3

75 75 75

Right Crossover Bus Center 5V Panel Lights Right 5V Panel Lights Windshield Alcohol Pump Right Altimeter Vibrator Right Digital Clock Emergency Pressurization System Right Engine Control Left Engine Fire Detection Left Engine Firewall Shutoff Flight Hour Meter Left Engine Ignition Overspeed Warning Right Pitot/Static Heater Surface Deice Boots Right Thrust Reverser Control Warning Lights 2 Left Fuel Boost Pump

CENTER PANEL RH PANEL W/S ALCOHOL RH ALT RH CLOCK EMER PRESS RH ENG LH FIRE DET LH FW SHUTOFF FLT/HR LH IGN OVERSPEED RH PITOT STATIC SURFACE DEICE RH THRUST REVERSER WARN LTS 2 LH BOOST

5 5 5 2 2 5 7.5 2 7.5 2 7.5 2 7.5 5 7.5 5 15


Amperage


Cessna Citation II Technical Manual DC Bus Distribution Table (550-0550 and after) Right Circuit Breaker Panel



Circuit Breaker

Avionics Bus 1 Automatic Direction Finder 1 Audio Control 1 Autopilot Warning DME 1 (550-0627 and after) EADI 1 (550-0627 and after) EFIS 1 (550-0627 and after) EFIS 1 Control (550-0627 and after) EHSI 1 (550-0627 and after) Flight Director 1 Navigation 1 Radio Altimeter Radio Magnetic Indicator 1 Transponder 1 B&D True Airspeed System Communication 2 Directional Gyro 1

ADF 1 AUDIO 1 WARN DME 1 EADI 1 EFIS 1 EFIS 1 CONT EHSI 1 FD 1 NAV 1 RAD ALT RMI 1 XPDR 1 TAS COMM 2 DG 1

3 3 3 3 5 5 1 5 3 3 2 2 3 2 7.5 5

Avionics Bus 2 EFIS 2 (550-0627 and after) EFIS 2 Control (550-0627 and after) EHSI 2 (550-0627 and after) Flight Director 2 Transponder 2 Communication 2 DME 2 (550-0627 and after) EADI 2 (550-0627 and after) True Air Speed Heater Audio Control 2 Automatic Direction Finder 2

EFIS 2 EFIS 2 CONT EHSI 2 FD 2 XPDR 2 COMM 2 DME 2 EADI 2 TAS HTR AUDIO 2 ADF2

5 1 5 3 3 7.5 3 5 2 3 2

Avionics Bus 3 AFIS Autopilot Servo Communication 3 Flight Management System VLF Navigation

AFIS AP COMM 3 FMS VLF

7.5 7.5 5 5 5

Avionics Bus 4 Flitefone Multifunction Display Multifunction Symbol Generator Weather Radar Radio Magnetic Indicator 2

PHONE MFD DISP MFD SYM GEN RADAR RMI 2

5 5 7.5 7.5 2

Emergency Bus Directional Gyro 2 Cockpit Flood Lights Radio Magnetic Indicator 2 Navigation 2 Attitude Director Indicator 2 Audio Control 1 Communication 1

DG 2 FLOOD RMI 2 NAV 2 ADI 2 AUDIO 1 COMM 1

3 7.5 2 3 2 3 7.5


Amperage

6-37

DC Bus Distribution Table (550-0550 and after) Right Circuit Breaker Panel Power Source and Equipment

Circuit Breaker

Right CB Panel Right Engine Fan Speed Digital Indicator Right Engine Turbine Speed Digital Indicator Right Engine ITT Indicator Right Fuel Flow Indicator Right Fuel Quantity Indicator Right Oil Pressure Indicator Right Oil Temperature Indicator Right Engine Start Control AC Inverter 2 Left CB Panel

RH FAN SPEED RH TURB SPEED RH ITT RH FUEL FLOW RH FUEL QTY RH OIL PRESS RH OIL TEMP RH START AC INVERTER 2 LH CB PANEL

2 2 2 2 2 2 2 7.5 25 35

Right Main Bus Right CB Panel Bus Right CB Panel Bus Right CB Panel Bus

RH BUS NO 1 RH BUS NO 2 RH BUS NO 3

75 75 75

Amperage

Dual Split AC Bus Distribution Table (550-0550 and after) Right Circuit Breaker Panel

6-38


Circuit Breaker

115 VAC BUS 1 Autopilot Control Flight Director 1 Sperry True Airspeed System Vertical Gyro 1 Weather Radar

AP FD 1 AIR DATA VG 1 RADAR

1 1 2 1 1

115 VAC BUS 2 Flight Director 2 Vertical Gyro 2

FD 2 VG 2

1 1

26 VAC BUS 1 EFIS 1 (550-0627 & after) Navigation 1 Radio Magnetic Indicator 1

EFIS 1 NAV 1 RMI 1

2 3 2

26 VAC BUS 2 Navigation 2 Radio Magnetic Indicator 2 Horizontal Situation Indicator 2 EFIS

NAV 2 RMI 2 HSI 2 EFIS

3 2 2 5

Right Sub CB Panel 115VAC 1 115VAC 2 26VAC 1 26VAC 2

115VAC 1 115VAC 2 26VAC 1 26VAC 2


Amperage

5 5 10 10


Cessna Citation II Technical Manual DC Bus Distribution Table (550-0505 and earlier) Electrical Power Junction Box



Circuit Breaker

Hot Battery Bus Aft/Forward Cargo Compartment Lights Emergency Flood Lights Engine Ignition Emergency Power

AFT/FWD COMP LT EMER LT IGNITION EMERGENCY POWER

Battery Bus Battery Voltmeter

BATT VOLTAGE

2

Left Isolation Bus Left Generator Ammeter Left Generator Sense Left Engine Start Light Left Generator Voltmeter

LH AMMETER (2) LH GEN SENSE LH START LT LH VOLTMETER

2 10 2 2

Right Isolation Bus Right Generator Ammeter Right Generator Sense Right Engine Start Light Right Generator Voltmeter

RH AMMETER (2) RH GEN SENSE RH START LT RH VOLTMETER

2 10 2 2

Left Main Bus Left Bus Sense Left Fuel Boost Pump Left Generator Off Light Left Landing Light Passenger Advisory Lights Left Recognition Light Indirect Lighting System Entertainment Center

LH BUS SENSE LH BOOST LH GEN OFF LH LDG LT OXY/SEAT BELT LH RECOG LT INDIRECT LT ENT CTR

2 15 2 15 5 5 7.5 5

Right Main Bus Right Bus Sense Right Fuel Boost Pump Right Generator Off Light Right Landing Light Right Recognition Light Cabin Lights Toilet/Shaver Outlet

RH BUS SENSE RH BOOST RH GEN OFF RH LDG LT RH RECOG LT CABIN LT TOILET

2 15 2 15 5 7.5 7.5


Amperage

3 5 7.5 20

6-39

DC Bus Distribution Table (550-0505 and earlier) Left Circuit Breaker Panel

6-40


Circuit Breaker

Left CB Panel Left Engine Fan Speed Digital Indicator Left Engine Turbine Speed Digital Indicator Left 5V Panel Lights AC Inverter 1 Anti-Collision Light AOA System Control AOA System Heater Battery Overtemp System Cabin Fan Cockpit Voice Recorder Left Digital Clock Electroluminescent Panel Lights Left Engine Control Engine Synchronization System Left Engine Fire Detection Left Engine Firewall Shutoff Flap Control Flight Data Recorder Right Fuel Boost Pump Left Engine Fuel Flow Indicator Left Fuel Quantity Indicator Right Engine Ignition System Left ITT Indicator Landing Gear Control Landing Gear Warning Left Engine Start Control Left Wing Inspection Light Nose Wheel Spinup System Outside Air Temperature Left Engine Oil Pressure Indicator Left Engine Oil Temp Indicator Pitch Trim Control Left Pitot/Static Heater Power Brakes and Anti-Skid Control Normal Pressurization System Right Circuit Breaker Panel Rotating Beacon Speed Brake Control Standby Gyro True Air Speed Probe Heater (Sperry) Cabin Temperature Control Left Engine Thrust Reverser Control Warning Lights 1 Windshield Bleed Air Control Windshield Bleed Air Power Wing Navigation Light Cockpit Voice Recorder Flap Motors

LH FAN SPEED LH TURB SPEED LH PANEL AC INVERTER NO 1 ANTICOLL ANG OF ATTACK AOA HTR BATT TEMP CABIN FAN VOICE RECORDER LH CLOCK EL PANEL LH ENG ENGINE SYNC LH FIRE DET LH FW SHUTOFF FLAP CONTROL FLIGHT RECORDER RH BOOST LH FUEL FLOW LH FUEL QTY RH IGN LH ITT GEAR CONTROL LDG GEAR LH START LH WING INSP NOSE WHL RPM OAT LH OIL PRESS LH OIL TEMP PITCH TRIM LH PITOT STATIC SKID CONTROL NORM PRESS RH CB PANEL ROTATING BEACON SPEED BRAKE STBY GYRO TAS HTR TEMP LH THRUST REVERSER WARN LTS 1 W/S BLEED AIR TEMP W/S BLEED AIR NAV Voice Recorder FLAP MOTOR

2 2 5 15 7.5 5 5 2 20 5 2 1 7.5 5 2 7.5 5 5 15 2 2 7.5 2 5 2 7.5 5 2 2 2 2 5 7.5 20 5 35 5 5 5 5 5 7.5 2 5 5 5 5 15

Left Main Bus Left CB Panel Bus Left CB Panel Bus Left CB Panel Bus

LH BUS NO 1 LH BUS NO 2 LH BUS NO 3

75 75 75


Amperage


Cessna Citation II Technical Manual DC Bus Distribution Table (550-0505 and earlier) Left Circuit Breaker Panel



Circuit Breaker

Right Crossover Bus Center 5V Panel Lights Right 5V Panel Lights Winshield Alcohol Pump Right Altimeter Vibrator Right Digital Clock Emergency Pressurization System Right Engine Control Right Engine Fire Detection Right Engine Firewall Shutoff Flight Hour Meter Left Engine Ignition Overspeed Warning Right Pitot/Static Heater Surface Deice Boots Right Thrust Reverser Control Warning Lights Left Fuel Boost Pump Equipment Cool

CENTER PANEL RH PANEL W/S ALCOHOL RH ALT RH CLOCK EMER PRESS RH ENG RH FIRE DET RH FW SHUTOFf FL/ HR LH IGN OVERSPEED RH PITOT STATIC SURFACE DEICE RH THRUST REVERSER WARN LTS LH BOOST Equipt Cool

5 5 5 2 2 5 7.5 2 7.5 2 7.5 2 7.5 5 7.5 5 15 7.5

Emergency Bus Directional Gyro 2 Cockpit Flood Lights Navigation 2 Communication 1

DG 2 FLOOD NAV 2 COMM 1

3 7.5 3 7.5


Amperage

6-41

DC Bus Distribution Table (550-0505 and earlier) Right Circuit Breaker Panel

6-42


Circuit Breaker

Right CB Panel Right Engine Fan Speed Digital Indicator Right Engine Turbine Speed Digital Indicator AC Inverter 2 Right Fuel Flow Indicator Right Fuel Quantity Indicator Right Engine ITT Indicator Left Circuit Breaker Panel Right Engine Oil Pressure Indicator Right Oil Temperature Indicator Right Engine Start Control Communication 2 Distance Measuring Equipment 2 Transponder 2 Automatic Direction Finder 2 Communication 3 Audio Control 2 AC Switch Autopilot Warning Area Navigation Ground Proximity Warning True Air Speed Heater VLF Navigation Nav Data Bank Flight Management Systems Weather Radar

RH FAN SPEED RH TURB SPEED AC INVERTER NO 2 RH FUEL FLOW RH FUEL QTY RH ITT LH CB PANEL RH OIL PRESS RH OIL TEMP RH START COMM 2 DME 2 XPDR 2 ADF 2 COMM 3 AUDIO 2 AC SWITCH WARN AREA NAV GROUND PROX HTR TAS VLF NAV NAV DATA BANK FMS RADAR

2 2 15 2 2 2 35 2 2 7.5 7.5 3 3 2 5 3 3 3 2 1 15 7.5 5 7.5 7.5

Right Main Bus Right DC Bus Number 1 Right DC Bus Number 2 Right DC Bus Number 3

RH BUS NO 1 RH BUS NO 2 RH BUS NO 3

75 75 75

Left Crossover Bus Navigation 1 Automatic Direction Finder 1 Audio Control 1 Distance Measuring Equipment 1 Attitude Director Indication 1 EFIS Disp EFIS EHSI 1 Flight Director 1 Radio Altimeter Radio Magnetic Indicator 1 Transponder 1 Communication 2 Directional Gyro 1 Autopilot Servo Flitefone Horizontal Situation Indicator 1 Cockpit Voice Advisory Flight Management System

NAV 1 ADF 1 AUDIO 1 DME 1 ADI 1 EFIS Disp EFIS EHSI 1 FD 1 RAD ALT RMI 1 XPDR 1 COMM 2 DG 1 AP PHONE HSI 1 VOICE ADV FMS

3 3 3 3 5 1 5 5 3 2 2 3 7.5 5 7.5 5 5 5 5

Emergency Bus Navigation 2 Communication 1 Directional Gyro 2 Cockpit Flood Lights

NAV 2 COMM 1 DG 2 FLOOD

2 7.5 3 5


Amperage


Cessna Citation II Technical Manual Split AC Bus Distribution Table (550-0505 and earlier) Right Circuit Breaker Panel Power Source and Equipment

Circuit Breaker

115 VAC BUS 1 Autopilot Control Flight Director 1 Sperry True Airspeed System Vertical Gyro 1 Weather Radar

AP FD 1 AIR DATA VG 1 RADAR

1 1 2 1 1

115 VAC BUS 2 Flight Director 2 Vertical Gyro 2

FD 2 VG 2

1 1

26 VAC BUS 1 Navigation 1 RMI /ADF 1 Attitude Director Indicator 1

NAV 1 RMI /ADF 1 ADI 1

3 2 1

26 VAC BUS 2 Navigation 2 Radio Magnetic Indicator 2 Horizontal Situation Indicator 2 EFIS RMI /ADF 2 Attitude Director Indicator 2

NAV 2 RMI 2 HSI 2 EFIS RMI /ADF 2 ADI 2

3 2 2 2 2 2

Right Sub CB Panel 115VAC 26VAC

115VAC 26VAC

Amperage

5 10

Single AC Bus Distribution Table (550-0505 and earlier) Right Circuit Breaker Panel



Circuit Breaker

115 VAC BUS 1 Flight Director 1 Flight Director 2 Vertical Gyro 1 Vertical Gyro 2 AC Monitor Sperry True Airspeed System Autopilot Control Weather Radar

FD 1 FD 2 VG 1 VG 2 AC MONITOR AIR DATA AP RADAR

1 1 1 1 3 2 1 1

26 VAC BUS 1 Navigation 1 Radio Magnetic Indicator 1 Radio Magnetic Indicator 2 Automatic Direction Finder 1 Automatic Direction Finder 2 Attitude Director Indicator 1 Horizontal Situation Indicator 1

NAV 1 RMI 1 RMI 2 ADF 1 ADF 2 ADI 1 HSI 1

3 2 1 2 2 2 1


Amperage

6-43

Alternating Current (AC) Power System Various airplane components requiring AC power for their operation are supplied by AC inverters through a multiple bus distribution system which varies depending on airplane configuration and unit number. Dual Split Bus AC System (550-0550 and after) Inverters During normal operation, each AC bus is supplied by its associated inverter. Should failure of either inverter occur, the opposite inverter supplies power to all AC busses through an automatic switching circuit. When an inverter has failed, the red [AC FAIL] annunciator and the corresponding [1] or [2] [INV FAIL] annunciator will illuminate, and the red [MASTER WARNING] light/switch will flash. Resetting the [MASTER WARNING] light/ switch will extinguish the [AC FAIL] annunciator but will not reset the failed inverter. Should an AC Bus circuit breaker trip, either or both [INV FAIL] annunciators, the [AC FAIL] annunciator and the red [MASTER WARNING] light/switch will illuminate. The tripped circuit breaker may be reset by pushing it back in, and the annunciators extinguished by pressing the [MASTER WARNING] light/ switch which may return either or both inverters to operation providing the fault has cleared. If the A/C Bus circuit breaker cannot be reset, the bus isolated by the associated circuit breaker is no longer energized and all systems powered by it become inoperative. AC Control Positioning the AC switch to the upper (on) position activates inverter 1 and 2, momentarily illuminating the [AC FAIL] annunciator until both inverters are on-line and in phase. A synchronization circuit between the inverters is used as the reference for phase relationship.

6-44



Cessna Citation II Technical Manual Inverter Testing Each inverter may be tested by positioning the AC switch to the upper (on) position and holding the test switch to the “INV 1” or “INV 2” position and observing illumination of the corresponding [1] or [2] [INV FAIL] annunciator. When the test switch is released the annunciator should extinguish. Each 115 VAC and 26 VAC dual split bus circuit breaker, located on the right flight compartment sidewall, functions to isolate its associated AC bus when overloaded and to illuminate the [AC FAIL] annunciator.

▲ AC SYSTEM CIRCUIT BREAKERS



6-45

“Tied” Split Bus AC System (optional 550-0162~0505 with dual flight director installation) Inverters During normal operation, each AC bus is supplied by its associated inverter. Each 26 VAC and 115 VAC bus are tied through a bus tie circuit breaker. Should failure of either inverter occur, the opposite inverter supplies power to all AC busses through the bus tie circuit breakers. Failure of either inverter will cause the corresponding [1] or [2] [INV FAIL] annunciator, and the red [MASTER WARNING] light/switch to illuminate. Resetting the [MASTER WARNING] light/switch will extinguish the light but will not reset the failed inverter. Should a bus-tie circuit breaker trip, either or both [INV FAIL] annunciator(s), and the red [MASTER WARNING] light/switch will illuminate. The tripped circuit breaker may be reset by pushing it back in and the annunciator extinguished by pressing the [MASTER WARNING] light/switch which may return either or both inverters to operation providing the fault has cleared. If the bus tie circuit breaker(s) cannot be reset, the bus isolated by the associated circuit breaker is no longer energized and all systems powered by it become inoperative. AC Control Positioning the AC switch to the upper (on) position activates inverter 1 and 2, momentarily illuminating the [AC FAIL] annunciator until both inverters are on-line and in phase. A synchronization circuit between the inverters is used as the reference for phase relationship.

6-46



Cessna Citation II Technical Manual Inverter Testing Each inverter may be tested by positioning the AC switch to the upper (on) position and holding the test switch to the “INV 1” or “INV 2” position and observing illumination of the corresponding [1] or [2] [INV FAIL] annunciator. When the test switch is released the annunciator should extinguish. Each 26 VAC bus and 115 VAC bus are tied together through corresponding 26V and 115V AC POWER BUS TIE circuit breakers located on the right flight compartment sidewall.

▲ AC SYSTEM CIRCUIT BREAKERS



6-47

Single Bus System (standard 550-0505 and earlier) Inverters During normal operation, inverter 1 supplies all AC bus requirements. Should failure of inverter 1 occur, the voltage sensor which monitors the 115 VAC power bus will cause illumination of the [AC FAIL] annunciator. Positioning the inverter switch to “INV 2” will provide a ground for the inverter on/off circuit and energize the inverter switching relay, thereby supplying power from inverter 2 to the AC power busses, and extinguish the [AC FAIL] annunciator. AC Control Positioning the inverter switch to “INV 1” provides the ground path for the inverter on/off circuit resulting in power being supplied to the AC power busses from inverter 1.

Inverter Testing Each inverter may be tested by positioning the inverter switch to the “INV 1” or “INV 2” position which should cause the [AC FAIL] annunciator to extinguish. Setting the inverter switch to its “OFF” position should cause the [AC FAIL] annunciator to illuminate.

6-48



Cessna Citation II Technical Manual Dual Split Bus AC Power System (550-0550 and after)

AC Power Indication (550-0550 and after)



6-49

AC Power System (550-0505 and earlier)

Split Bus AC Power System (550-0550 and after)

6-50



Cessna Citation II Technical Manual AC Circuit Protection Circuit breakers for the airplane’s AC powered flight instruments and avionics are located on the right CB panel on the flight compartment sidewall so as to be readily accessible to the flight crew. The applicable amperage is marked on each circuit breaker. Labels above each breaker identify the circuit protected. These push-to-reset type circuit breakers will pop out, or “trip,” when heat is generated by an electrical overload. Should an overload occur, a tripped circuit breaker may be reset after a cooling period of approximately three minutes by pushing it back in. If the circuit breaker trips a second time, a short circuit is indicated and it should not be reset, as this could cause system damage. Note: The flight crew should ensure that all circuit breakers are engaged and serviceable fuses are installed before all flights. The airplane should never be operated with any disengaged circuit breakers without a thorough knowledge of the consequences.



6-51

Lighting The airplane is equipped with a variety of exterior lights to facilitate takeoff, landing, in-flight recognition, and wing inspection, plus interior lights that provide varying degrees of cockpit, cabin, and baggage compartment illumination. Exterior Lighting Exterior lighting consists of navigation lights, anti-collision lights, taxi/landing lights, a rotating or flashing beacon, tail flood lights, a wing inspection light, and recognition lights. Exterior lighting switches are located on the lower left instrument panel and the pilot’s lower instrument panel.

▲ LOWER LEFT INSTRUMENT PANEL

6-52


▲ PILOT’S LOWER INSTRUMENT PANEL


Cessna Citation II Technical Manual 1. Navigation Lights Navigation lights are installed on each wing tip and on the tailcone stinger. The lights are controlled by the NAV switch and supplied with 28 VDC power from the right main bus (550-0550 and after) or left main bus (550-0505 and earlier) through the NAV circuit breaker on the left CB panel.

1

2

When darkness permits, operation of the navigation lights can be confirmed during preflight by observing their reflection on the ground and/or other surrounding objects. During daylight hour preflight, the operation of each light should be confirmed from outside of the airplane. 2. Anti-Collision Strobe Lights High-intensity anti-collision strobe lights are installed on each wing tip adjacent to the navigation lights. The strobe lights are energized by bus voltage boosted through independent power supplies installed within the wing tip. The lights are controlled by the ANTI COLL switch and supplied with 28 VDC left main bus power through the ANTI COLL circuit breaker on the left CB panel. Note: Strobe lights should not be operated in clouds, fog, or haze as their reflection on water droplets in the atmosphere can induce disorientation or vertigo. Note: To avoid interfering with the vision of other pilots, strobe lights should not be operated when taxiing in the vicinity of other aircraft.



6-53

Taxi/Landing Lights A sealed beam taxi/landing light is installed on the forward drag brace of each main gear assembly (airplanes not incorporating SB550-33-5), or on each main gear door (airplanes incorporating SB55033-5). The lights are exposed and operable only when the main gear is extended. Each light is independently controlled by its associated LH or RH LANDING switch. 28 VDC power is supplied to each from its corresponding left or right main bus, through circuit breakers located in the aft fuselage electrical power junction “J” box. Wing Recognition Lights Optional recognition lights, used to provide additional exterior lighting and increase airplane visibility, are installed on the leading edge of each wing tip. Glareshields are installed slightly inboard of the lights to reduce glare within the cockpit. The recognition lights are controlled by the RECOG switch. 28 VDC power is supplied to each from its corresponding left or right main bus, through circuit breakers located in the “J” box. Caution: Recognition lights use a pressurized, halogen cycletype lamp which produces extremely high intensity light which could cause eye damage if viewed directly. The lamp should be protected from abrasions, scratches, impact, and contact with liquids. Handling the lamp should be avoided. Allow the lamp to cool, and wear protective clothing and dark glasses if contact is necessary. Note: Recognition lights should be turned on shortly before takeoff and during descent, and extinguished during climb, cruise, and after landing as soon as the airplane is clear of the runway.

6-54



Cessna Citation II Technical Manual Rotating/Flashing Beacon Rotating or flashing beacons, used to increase airplane visibility, may be installed on the top of the rudder. Any of three different types of beacons may be installed on individual airplanes, two of which use flashing lamps and one of which uses a rotating light assembly that employs a stationary lamp and electrically-driven rotating reflector. The beacon is controlled by the BEACON switch and supplied with 28 VDC left main bus power through the ROTATING BEACON circuit breaker on the left CB panel. Wing Inspection Light The wing inspection light, used to check for ice accumulation on the leading edge of the wing during night operations, is flush-mounted on the left side of the fuselage forward of the wing. The light is controlled by the WING INSP switch and supplied with 28 VDC power from the left main bus (550-0550 and after) or right main bus (550-0505 and earlier) through the WING INSP circuit breaker on the left CB panel. Note: Operation of the wing inspection light is mandatory for flight in icing conditions as defined by the FAA. Tail Floodlights Optional floodlights may be installed on the upper left and right surfaces of the horizontal stabilizer to illuminate the vertical stabilizer. These lights may also be referred to as identification lights, logo lights, or tail lights. The floodlights are controlled by the BEACON switch (550-0038 and after) or the NAV switch (550-0037 and earlier), and supplied with 28 VDC left main bus power through the TAIL LIGHTS circuit breaker located in the “J” box.



6-55

Interior Lighting Airplane interior lighting systems provide illumination, advisory and warning within the flight compartment and passenger cabin, plus baggage compartment illumination. Control and Instrument Lighting Electroluminescent “backlit” panels provide primary illumination of labels on various switch, control, and circuit breaker panels as well as the center pedestal. The electroluminescent panels are powered by 40 to 60 VAC, 400 Hz inverters. Additional panel and instrument illumination is provided by integral 5 VDC powered lights. Panels and instruments that are not internally lighted are illuminated by 5 VDC or 28 VDC powered post lights. Panel and instrument illumination is controlled by ON/OFF switches and rheostats located on the pilot’s lower instrument panel.

▲ PANEL AND INSTRUMENT LIGHT CONTROLS

Instrument Panel Floodlights The instrument panel floodlights, located on the aft overhead console behind a blue tinted lens, are normally used during thunderstorms to provide supplemental instrument panel illumination to compensate for lightning-induced night vision loss. These lights are controlled by the FLOOD LTS rheostat and supplied with 28 VDC emergency bus power through the FLOOD circuit breaker on the right CB panel. Vertical Scale Engine Instrument Light The vertical scale engine instrument light, located on the lower surface of the glareshield panel fire tray, illuminates the vertical scale engine instruments on the upper center instrument panel.

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Cessna Citation II Technical Manual This light is automatically illuminated during the engine start sequence and extinguished upon start sequence completion. During engine start, the light is supplied with 28 VDC emergency bus power through the EMER LT circuit breaker located in the “J” box. During normal operation, the light is supplied with 28 VDC emergency bus power through the FLOOD circuit breaker. In the event of electrical system failure the light is supplied with 28 VDC power from the standby gyro battery pack (if installed). The light is also supplied with 28 VDC power from an emergency lighting battery pack, located above the cabin headliner, through an inertial switch which closes when exposed to an acceleration force of 5Gs or more. Counter Light The post-type counter light illuminates the mechanical counter located above the FAN tachometer on the upper center instrument panel. This light is controlled by the CENTER panel rheostat and supplied with 28 VDC left main bus power through the CENTER PANEL circuit breaker on the left CB panel Cathode Tube Lights Two cathode tube lights are installed under the glareshield to provide supplemental instrument panel lighting. The lights are powered by a high-voltage inverter located within the left side console. This light and the inverter are controlled by the EL panel rheostat and supplied with 28 VDC left main bus power through the EL PANEL circuit breaker on the left CB panel. Map Lights Map lights, providing directional flight compartment illumination, are located on the overhead console. Each light is independently controlled by an associated rheostat, located on the forward end of each side console, and supplied with 28 VDC right main bus power through the 5-amp RH PANEL circuit breaker on the left CB panel.



6-57

Indirect Cabin Lights Primary cabin illumination is provided by an indirect lighting system, consisting of twelve overhead fluorescent lights powered by two inverters located above the cabin headliner. The lights and the inverters are controlled by a three-position (OFF/BRIGHT/ DIM) switch, located on the cabin sidewall forward of the ▲ OFF/BRIGHT/DIM SWITCH entrance door, and supplied with 28 VDC left main bus power through the INDIRECT LT circuit breaker located in the “J” box. The indirect lighting system is optional on airplanes 550-0550 and after, and standard on airplanes 550-0505 and earlier. Overhead Console Sign The lighted overhead console sign, mounted just aft of the forward divider, displays universal “no smoking” and “fasten belt” symbols. This light is controlled by a PASS SAFETY switch and supplied with 28 VDC left main power through the OXY/SEAT BELT SIGN circuit breaker located in the “J” box. Entrance Lights Entrance lights are provided to illuminate the passenger door entrance, emergency exit door, and aft baggage compartment. Each entrance light may be illuminated by an integral switch, or by the PASS SAFETY switch. 28 VDC hot battery bus power is supplied to these lights through the CABIN LIGHTS circuit breaker located in the “J” box. In the event of electrical system failure these lights may be supplied with 28 VDC power from the emergency lighting battery pack and are automatically illuminated through the inertial switch with an acceleration force of 5Gs or more.

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Cessna Citation II Technical Manual Emergency Exit Sign A lighted EMERGENCY EXIT sign is located over the emergency exit door to clearly identify its location. The light is controlled by an illuminated switch located on the cabin sidewall forward of the entrance door, or by the PASS SAFETY switch. 28 VDC left main bus power is supplied to the light through the OXY/SEAT BELT circuit breaker located in the “J” box. In the event of electrical system failure this light may be supplied with 28 VDC power from the emergency lighting battery pack and is automatically illuminated through the inertial switch with an acceleration force of 5Gs or more. PSU Light On airplanes 550-0550 and after, an optional fluorescent light is available to provide passenger service unit (refreshment center or vanity) area illumination. The light is controlled by the threeposition (OFF/BRIGHT/DIM) switch, located on the cabin sidewall forward of the entrance door, and supplied with 28 VDC left main bus power through the INDIRECT LT circuit breaker located in the “J” box. Passenger Reading Lights Reading lights, providing directional illumination, are located above each passenger station. Each light is independently controlled by an integral switch and supplied with 28 VDC right main bus power through the CABIN LIGHTS circuit breaker on the left CB panel.



6-59

Aft Fuselage Interior Light Aft fuselage interior illumination is provided by a detachable light mounted on a swiveling base between the forward and aft engine carry-thru beams. The light is controlled by an ON/OFF switch located on the tailcone access door frame. The ON/OFF switch is wired through a microswitch such that when the switch is ▲ ON/OFF SWITCH set to “ON” the light will illuminate when the door is opened and extinguish when the door is closed. The light is supplied with 28 VDC hot battery bus power through the AFT/FWD COMPARTMENT LIGHT circuit breaker located in the “J” box. Baggage Compartment Lights Nose baggage compartment illumination is provided by a light mounted overhead near the center of the compartment. The light is controlled by an illuminated, rocker-type, on/off switch located adjacent to the light. The on/off switch is wired through a microswitch such that when the switch is set to “on” the light will illuminate when the door is opened and extinguish when the door is closed. Tailcone baggage compartment illumination is provided by a light mounted above the access door. The light is controlled by an ON/OFF switch located on the tailcone access door frame. The ON/OFF switch is wired through a microswitch such that when the switch is set to “ON” the light will illuminate when the door is opened and extinguish when the door is closed. The baggage compartment lights are supplied with 28 VDC hot battery bus power through the AFT/FWD COMPARTMENT LIGHT circuit breaker located in the “J” box.

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Cessna Citation II Technical Manual Limitations Refer to the applicable airplane manufacturers FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable airplane manufacturers FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.



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Chapter 7 Flight Instrumentation Table of Contents Overview ............................................... 7-1 Flight Environment Data System ............. 7-2 Pitot-Static System ................................ 7-2 Pitot Tubes ...................................... 7-2 Static Ports ..................................... 7-3 Alternate Static Air Source ................ 7-3 Airspeed/Mach Indicators ...................... 7-4 Airspeed Pointer............................... 7-4 Mach Sub-Dial ................................. 7-4 Airspeed Mach Indicator Markings ..... 7-5 Index Marker ................................... 7-5 Overspeed Warning System ................... 7-6 True Airspeed System (optional) ............. 7-7 Sperry TAS System .......................... 7-7 B&D TAS System ............................. 7-8 Barometric Altimeters ....................... 7-9 Pilot’s Altimeter ............................. 7-10 Copilot’s Altimeter .......................... 7-11 Altitude Alerting and Reporting System ................................ 7-12 Altitude Alerting .............................. 7-12 Altitude Reporting ........................... 7-13 Vertical Speed Indicators ..................... 7-15


Flight Instrumentation, continued Table of Contents Radio Altimeter System (optional) ........ 7-16 Transceiver and Antennas ............... 7-16 Conventional Indicator (excluding RAD/BAR) ...................... 7-17 Conventional Indicator (RAD/BAR) ... 7-18 Mechanical Flight Director Radio Altitude Indication ................. 7-19 EFIS Radio Altitude Indication ......... 7-20 Attitude and Direction System .............. 7-21 Conventional Attitude Indicator ........ 7-21 Air Driven Gyro............................... 7-22 Gyro Pressure Gage ....................... 7-23 Electrically-Driven Gyro .................... 7-23 Turn-and-Bank Indicator ................... 7-24 Mechanical Flight Directors .................. 7-25 Attitude Directional Indicator ........... 7-25 Horizontal Situation Indicator .......... 7-25 Gyro Slaving .................................. 7-26 ADI Display Features ...................... 7-27 HSI Display Features ...................... 7-31 Electronic Flight Instrumentation System ....................... 7-33 Autopilot/Flight Director System ........... 7-33 Limitations .......................................... 7-33 Emergency Procedures......................... 7-33


▲ PILOT’S FLIGHT INSTRUMENTATION

The primary flight instruments of the Cessna Citation II are positioned on panels directly in front of the pilot and copilot. Pitot-static instrumentation consists of airspeed/mach indicators, barometric altimeters, and vertical speed indicators. Gyroscopic instrumentation consists of attitude directional indicators (ADIs) and horizontal situation indica▲ COPILOT’S FLIGHT INSTRUMENTATION tors (HSIs) of mechanical or electronic (EFIS) type depending on installation. Turn coordination information is provided either by independent mechanical indicators or by a rate-of-turn indicator and conventional inclinometer integral to each ADI/EADI. A standby attitude indicator may be installed as an emergency backup to the ADI(s) or EADI(s). Additional navigational guidance is provided by radio magnetic indicators (RMIs) displaying both VOR and ADF magnetic bearing information, an optional radio altimeter, and a magnetic compass mounted on the windshield center post above the glareshield. An outside air temperature (OAT) indicator, clock, and optional angle-of-attack (AOA) indicator are located on the upper left instrument panel. An optional true airspeed (TAS) system may also be installed. FLIGHT INSTRUMENTATION 12/99


7-1

For organizational purposes, this chapter is divided into separate flight environment data and attitude and direction systems according to the conditions or forces utilized in their operation. The flight environment data system includes equipment and instruments which are sensitive to environmental conditions which influence navigation. The attitude and direction system includes equipment and instruments which are sensitive to gyroscopic, inertial, and magnetic forces which influence navigation. Flight Environment Data System The flight environment data system includes the pitot-static system and associated flight instruments, the overspeed warning system, altitude alerting and reporting system, as well as the optional true airspeed (TAS), radio altimeter, and AOA systems when installed. Pitot-Static System The pilot’s and copilot’s pitot-static flight instruments are supplied by independent pitot-static systems consisting of one pitot tube and two static ports each. Pitot Tubes The pitot tubes are located on the lower left and right surface of the nose section. The left pitot tube supplies ram pressure to the pilot’s airspeed/ mach indicator, the air data computer, and airspeed/mach warning switch. The right pitot tube supplies ram pressure to the copilot’s airspeed indicator, the landing gear warning ▲ PITOT TUBE airspeed switch (550-0627 and after), B&D TAS pressure transducer (if installed), and airspeed/mach warning switch (Canadian (CAA) certified airplanes).

7-2



Cessna Citation II Technical Manual Static Ports The static ports are located on the left and right sides of the fuselage. The upper right and lower left static port are interconnected and supply ambient pressure to the pilot’s airspeed/mach indicator, IVSI, the air data computer, and airspeed/mach warning switch. The upper left and lower right static port are ▲ STATIC PORTS interconnected and supply ambient pressure to the copilot’s airspeed/mach indicator, barometric altimeter, IVSI, the landing gear warning airspeed switch (550-0627 and after), B&D TAS pressure transducer (if installed), and airspeed/mach warning switch (CAA certificated airplanes). Ambient pressure is also provided to the cabin differential pressure indicator through the copilot’s static ports. The interconnection and location of the static ports on opposite sides of the fuselage minimizes system pressure errors caused by uncoordinated flight. The pitot tubes and static ports are protected against icing by integral, electrically-powered heating elements. Refer to Chapter 10 for a complete description of pitot-static ice protection. Note: The pitot tubes and static ports must be clear and free of obstructions for proper operation. Alternate Static Air Source (if installed) Should restriction of the static ports occur, as evidenced by erratic indication of the pitot-static flight instruments, the copilot’s instruments may be provided with ambient pressure from within the nose section through an alternate static air source. A manually-operated control valve, located below the copilot’s instrument panel, is used to select the normal or alternate static air source. Though optional on most Citations, the alternate static air source system is standard on French (DGAC) certificated airplanes.

FLIGHT INSTRUMENTATION 12/99


7-3

Airspeed/Mach Indicators The airspeed/mach indicators provide visual indication of the 1 2 airplane’s speed in relation to the ambient air and the speed of sound. Each unit consists of a stationary indicator dial, airspeed pointer, rotating mach sub-dial, moveable index marker, and an airtight instrument case which houses 3 a pair of airtight diaphragms linked to drive mechanisms 1. Airspeed Pointer that operate the pointer and 2. Mach Sub-Dial sub-dial. The instrument case 3. Index Marker Knob is supplied with static (ambient) air pressure through its associated static ports. The airspeed diaphragm is supplied with pitot (ram) pressure through its associated pitot tube while the mach diaphragm is sealed at standard sea level atmospheric pressure (29.92 inHg/1013.2 mb (reference pressure)). Airspeed Pointer As airspeed increases or decreases, the differential between static pressure and pitot pressure causes the airspeed diaphragm to expand or contract. As it does, its movement is transmitted by the drive mechanism to position the pointer at the corresponding KIAS value on the indicator dial. Mach Sub-Dial The inner (rotating) mach sub-dial is visible through a window on the face of the instrument between 140 and 320 knots on the outer (stationary) KIAS indicator dial. Unlike the airspeed pointer which is positioned relative to speed, the mach sub-dial is positioned relative to altitude. As altitude increases or decreases, the differential between static pressure and reference pressure causes the mach diaphragm to expand or contract. As it does, its movement is transmitted by the drive mechanism to rotate the sub-dial such that the relationship between its position and that of the pointer will correspond to the approximate mach number (M).

7-4



Cessna Citation II Technical Manual Airspeed/Mach Indicator Markings The airspeed indicator dial is calibrated in knots per hour (KIAS) and incorporates two radial slots (9500 LB ZFW) or one radial slot (11,000 LB ZFW) in positions corresponding to the airplane’s Maximum Operating Speed (VMO) limitation(s) listed below. Rotation of the sub-dial, occurring with changes in altitude, causes the color red to become visible through the slot(s) within the appropriate altitude range. Zero Fuel Weight

Altitude Range

VMO

9500 LB 9500 LB 11,000 LB

SL ~ 14,000' 14,000' ~ 28,000' SL ~ 30,500'

262 KIAS 277 KIAS 262 KIAS

A red radial line on the mach sub-dial denotes the airplane’s Maximum Mach Operating (MMO) limitation of 0.705 M at altitudes above 28,000 feet (9500 LB ZFW) or 30,500 feet (11,000 LB ZFW). Essentially, when operating above these altitudes, alignment of the airspeed pointer with the MMO radial line will produce a true airspeed of 0.705 M regardless of indicated speed. To avoid exceeding this limitation, the airspeed pointer must never be permitted to rotate beyond the MMO radial line. Index Marker The index marker is controlled by a knob located on the lower left corner of the instrument face and may be moved to any position around the KIAS scale to reference a desired airspeed.



7-5

Overspeed Warning System The overspeed warning system consists primarily of an airspeed/mach warning switch located on the forward pressure bulkhead, and a warning horn located in the overhead flight compartment outboard of the pilot’s station. The warning switch senses airspeed and altitude via the pilot’s pitot-static system, or copilot’s pitot-static system on Canadian (CAA) certified airplanes, and causes the warning horn to sound when VMO/ MMO is reached or exceeded. On 9500 LB ZFW airplanes, the warning switch integrates two airspeed switches, an altitude switch, and a mach number switch. One airspeed switch (S1) operates in conjunction with the altitude switch (S2) to sound the horn when airspeed reaches 262 KIAS at altitudes below 14,000 feet. The other airspeed switch (S3) and the mach number switch (S4) function to sound the horn at 277 KIAS and 0.705 M respectively, regardless of altitude. On 11,000 LB ZFW airplanes, the warning switch utilizes airspeed switch S1 and the mach number switch S4 to sound the horn at 262 KIAS and 0.705 M respectively, regardless of altitude (switches S2 and S3 are not required to be operational in this installation). 28 VDC power is normally supplied to the warning horn by the right main bus through the 2-amp OVERSPEED circuit breaker. Operation of the overspeed warning system may be verified by rotating the TEST selector switch, located on the lower left instrument panel, to the “OVERSPEED” position. In this position, power supplied through the 2▲ TEST SELECTOR SWITCH amp WARN LTS 1 circuit breaker will cause the horn to sound if the system is functional.

7-6



Cessna Citation II Technical Manual True Airspeed (TAS) System (optional) True airspeed systems incorporate equipment which senses and measures pitot-static pressures and ambient air temperature to compute a correction for compressibility and ram rise. Once computed, a signal corresponding to true airspeed is supplied to the TAS indicator and/or flight guidance systems which utilize this data in their operation. One of two systems manufactured by Sperry or B&D may be installed. Sperry TAS System Components of the Sperry TAS system include the air data computer (ADC), located in the nose avionics bay; a temperature probe, located on the lower right surface of the nose section; and an indicator, located on the center instrument panel. AC power is supplied to the ADC, temperature probe, and indicator through the 2-amp AIR DATA circuit breaker. The ADC is supplied by the number one 115 VAC bus (split bus AC configuration), or by the 115 VAC bus (single bus AC configuration) and incorporates a transformer which steps 115 VAC down to 26 VAC to supply the temperature probe and indicator. The temperature probe is protected against icing by an integral, electrically-powered heating element controlled by the PITOT & STATIC switch on the lower left instrument panel. The heating element is supplied with 28 VDC power through the 15amp TAS HTR circuit breaker.



7-7

B&D TAS System Components of the B&D system include a true airspeed (TAS) computer, located below the flight compartment floor; a temperature probe, located on the lower right surface of the nose section; and a pressure transducer located on the forward side of the lower right instrument panel. The TAS computer is installed in addition to, and functions independently of, the ADC addressed in the Sperry TAS system description. The TAS computer processes signals supplied by the copilot’s pitot-static system and the temperature probe. Like the Sperry system, the temperature sensed by the probe is corrected for Mach effect to obtain a measurement of static air temperature (SAT) which is computed with pitot-static pressures to produce an electrical signal proportional to true airspeed (TAS). Unlike the Sperry system, however, this signal is supplied only to the flight guidance systems which require TAS data for their operation. The B&D system does not feature a digital indicator or temperature probe ice protection. Should malfunction of this system occur, as evidenced by erratic indication of the copilot’s pitot-static flight instruments, the copilot’s pitotstatic system may be isolated from the TAS computer by closing a pair of valves located on the lower right instrument panel. The valves are labeled TAS COMP - STATIC PITOT and protected by red ▲ TAS COMPUTER PITOT-STATIC VALVES guard covers which must be lifted when actuation is required. In the normal “OPEN” position, TAS signals are supplied to the flight guidance systems requiring this data for their operation. In the “CLOSED” position, TAS data is no longer supplied by the computer, but may still be supplied to flight guidance equipment which provides for manual entry of this data. The system is supplied with 28 VDC power through the 2-amp TAS circuit breaker. When closing the TAS COMP valves is required, this circuit breaker should be pulled to ensure that flight guidance equipment does not receive erroneous airspeed data.

7-8



Cessna Citation II Technical Manual Barometric Altimeters The barometric altimeters provide visual indication of airplane altitude in relation to mean sea level (MSL) atmospheric pressure. Each unit incorporates an indicating pointer, drum-type counter, and barometric scale. The pointer dial is calibrated in increments of 20 and 100 feet, while the counter displays altitude in increments of 20, 100, 1000, and 10,000 feet. A black and white crosshatched area appearing in the left most digit of the counter signifies that indicated altitude is below 10,000 feet. The barometric scale, calibrated in inches of mercury (inHg) and millibars (mb), displays the current correction setting for non-standard atmospheric conditions. At flight altitudes below 18,000 feet, the barometric scale must be set to the current altimeter setting provided by en route reporting stations within 100 nautical miles of the airplane’s position. This setting reflects a computed correction, for nonstandard conditions, of the barometric pressure measured near ground level in the vicinity of the reporting station and provides indication of true altitude above mean sea level (MSL). At flight altitudes 18,000 feet (FL 180) and above, the barometric scale must be set to standard sea level pressure (29.92 inHg/1013.2 mb). This setting provides indication of pressure altitude above the standard datum plane, a theoretical level where atmospheric pressure is equal to standard sea level pressure. Since the assigned altitudes of all aircraft operating at FL 180 and above are referenced to this setting, collision avoidance and vertical separation are assured. To ensure adequate separation from aircraft operating below 18,000 feet and compliance with minimum altitude rules, local reported pressure must be monitored and cruise altitudes in the vicinity of FL 180 proportionally increased by 500 feet for every half inch that current altimeter setting falls below 29.92 inHg/ 1013.2 mb as specified in FAR 91 and the table below: Altimeter Setting 29.92 or higher 29.91 ~ 29.42 29.41 ~ 28.92 28.91 ~ 28.42 28.41 ~ 27.92 27.91 ~ 27.42 27.41 ~ 26.92



Lowest Usable Flight Level FL FL FL FL FL FL FL

180 185 190 195 200 205 210

Adjustment Factor 0' 500' 1000' 1500' 2000' 2500' 3000'

7-9

Pilot’s Altimeter The pilot’s altimeter is a servotype indicator which is electrically-driven by the air data computer (ADC) located in the nose avionics bay. The ADC processes pneumatic signals supplied by the pilot’s pitotstatic system and computes an electrical signal proportional to the airplane’s altitude. This signal is transmitted to the altimeter drive mechanism such that the corresponding altitude is displayed by the pointer and counter. Correction for non-standard atmospheric conditions is accomplished by rotating the barometric setting knob, located on the lower left instrument bezel, until the desired pressure appears on the inHg or mb scale. This setting transmits an electrical signal to the ADC which, in turn, transmits a barometrically-corrected signal to the altimeter drive mechanism. Should an error between these signals occur, a failure warning flag will extend across the counter indicating that altitude readout is not reliable. AC power is supplied to the ADC and the pilot’s altimeter through the 2-amp AIR DATA circuit breaker. The ADC is supplied directly by the 115 VAC bus (single bus AC configuration) or the number one 115 VAC bus (split bus AC configuration) and incorporates a transformer which steps 115 VAC down to 26 VAC to power the altimeter. The failure warning flag will also extend across the counter whenever electrical power is removed from the instrument. An optional radio/barometric (RAD/BAR) altimeter may be installed in place of the standard pilot’s altimeter. Operation and setting of the barometric portion of the instrument are consistent with the standard altimeter. Operation of the radio altimeter portion of the instrument is addressed in the description of radio altimeter systems.

7-10



Cessna Citation II Technical Manual Copilot’s Altimeter The copilot’s conventional altimeter consists of an airtight instrument case which houses an aneroid diaphragm assembly linked to a drive mechanism that operates the pointer and counter. The instrument case is supplied with static (ambient) air pressure through the static ports, while the diaphragm assembly is sealed at standard sea level atmospheric pressure (29.92 inHg/1013.2 mb (reference pressure). As altitude increases or decreases, the differential between static pressure and reference pressure causes the diaphragm assembly to expand or contract. As it does, its movement is mechanically transmitted by the drive mechanism such that the corresponding altitude is displayed by the pointer and counter. Correction for non-standard atmospheric conditions is accomplished by rotating the barometric setting knob, located on the lower left instrument bezel, until the desired pressure appears on the inHg or mb scale. This setting rotates the indicator drive mechanism to produce the necessary altimeter correction. The copilot’s altimeter incorporates a vibrator which functions to optimize indicator response. 28 VDC power is supplied to the vibrator by the right main bus through the 2-amp RH ALT circuit breaker. When the airplane is configured for dual-altitude reporting capability, the standard copilot’s altimeter is replaced by an optional encoding altimeter. Operation and setting of the barometric portion of the instrument are consistent with the standard altimeter. Operation of the encoder portion of the instrument is addressed in the description of altitude alerting and reporting systems.



7-11

Altitude Alerting and Reporting System The altitude alerting and reporting system incorporates those components which provide visual and aural alerting of preselected reference altitudes, and those which supply encoded altitude data to the transponder(s). Altitude Alerting Altitude alerting components include the air data computer (ADC), located in the nose avionics bay; vertical navigation computer/controller (VNCC), located on the center instrument panel; altitude alert light(s) and a warning horn. One altitude alert light is located on the upper right instrument bezel of the standard pilot’s altimeter. Others are located on the upper left bezel of the optional pilot’s RAD/BAR altimeter and optional copilot’s encoding altimeter when these instruments are installed. The altitude alert warning horn is located within the left flight compartment sidewall. The VNCC references altitude signals transmitted to the pilot’s altimeter from the ADC, as previously described, and provides data input and output for altitude alerting and various vertical navigation (VNAV) modes. Mode selection and data input are controlled by a rotary switch and concentric setting (SET) knob, and annunciated by a three-digit ▲ VNAV COMPUTER/CONTROLLER incandescent display. A dimmer (DIM) knob permits adjustment of display intensity. When “ALT” (altitude mode) is selected, rotating the SET knob slews the display in increments of 100 feet between 000(00) and 500(00) feet to the desired altitude preselect. Slew rate is proportional to the speed at which the SET knob is rotated.

7-12



Cessna Citation II Technical Manual As the airplane climbs or descends to within 1000 feet of the preselected altitude, the warning horn will sound for one second and the altitude alert light(s) will illuminate and remain illuminated until the airplane is within 250 feet of the preselected altitude. Should the airplane deviate from the preselected altitude by 250 feet or more, the warning horn will sound for one second and the altitude alert light(s) will illuminate and remain illuminated until the airplane returns to within 250 feet of the preselected altitude or until a new altitude preselect is set. AC power is supplied to the VNCC by the number one 115 VAC bus (split bus AC configuration) or the 115 VAC bus (single bus AC configuration) through the 1-amp FD 1 circuit breaker. 28 VDC power is supplied to the warning horn and altitude alert light(s) by the corresponding DC AVIONICS bus through the 3amp FD 1 circuit breaker. Altitude Reporting Altitude reporting components include the air data computer (ADC), transponder(s), and the optional copilot’s encoding altimeter when installed. These components function to generate a logic code corresponding to the airplane’s pressure altitude which is transmitted to the Air Traffic Control Radar Beacon System (ATCRBS) through the transponder in response to Mode C (altitude reporting) interrogations. Barometric pressure setting of the altimeter(s) has no effect on this function since encoder output is always referenced to standard sea level atmospheric pressure (29.92 inHg/1013.2 mb). In standard configuration, a single transponder (transponder 1) receives encoded altitude signals from the ADC. One optional configuration adds a second transponder (transponder 2) which also receives encoded altitude signals from the ADC. In this installation, either transponder may be assigned to altitude reporting by means of a select switch, co-located with the transponder controls, while the other remains in standby condition. Another option configures the system for dual-altitude reporting. In this installation, the selected transponder may receive encoded altitude signals from the ADC (primary source) or the optional copilot’s encoding altimeter (secondary source) by means of a combination annunciator/switch (550-0258 and after) or crossover switch and relay (550-0257 and earlier).



7-13

The encoder portion of the copilot’s altimeter incorporates a lightweight optical digitizer which imparts no load on the barometric portion of the instrument. 28 VDC power is supplied to the encoder by the right main bus through the 2-amp RH ALT circuit breaker. A failure warning flag will appear in the upper left corner of the instrument face whenever electrical power is removed from the instrument or when the encoder has failed. Should this occur, the instrument will continue to function as a barometric altimeter, but the ADC will provide the only source of encoded altitude.

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Cessna Citation II Technical Manual Vertical Speed Indicators The vertical speed indicators are instantaneous type (IVSI) and provide visual indication of the airplane’s rate of climb or descent in feet per minute (FT/MIN X 1000). Each unit consists of an indicator dial, a pointer, and an airtight instrument case which houses an airtight diaphragm and accelerometer linked to a drive mechanism that operates the pointer. Static (ambient) air pressure from the static ports is supplied to the diaphragm directly, and to the instrument case through a calibrated restrictor orifice. As the airplane climbs or descends, the static air pressure supplied to the instrument decreases or increases accordingly. Due to the restrictor orifice, however, the rate of pressure change within the instrument case occurs more slowly than within the diaphragm. This produces a pressure differential which causes the diaphragm to expand or contract in proportion to the rate of altitude change. As it does, its movement is transmitted by the drive mechanism to position the pointer at the corresponding value on the indicator dial. During a climb, the pressure within the diaphragm decreases at a faster rate than the pressure within the instrument case. The resulting differential causes the diaphragm to contract and the pointer to indicate a rate of climb. During descent, the pressure within the diaphragm increases at a faster rate than the pressure within the instrument case. The resulting differential causes the diaphragm to expand and the pointer to indicate a rate of descent. As the airplane resumes level flight, the pressure within the instrument case and the diaphragm become equalized and pointer indication returns to zero. The time required to stabilize the pressure differential which causes pointer deflection can result in a delay of up to nine seconds before vertical speed indication becomes reliable. To compensate for this, the accelerometer’s sensitivity to vertical G-loading provides instantaneous indication of vertical speed by displacing the pointer prior to the instrument’s response to changes in pressure. During level flight or steady rates of climb or descent, the IVSI function as a conventional VSI. FLIGHT INSTRUMENTATION 12/99


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Radio Altimeter System (optional) The radio altimeter system measures absolute altitude above ground level (AGL) and supplies this data to flight guidance systems, ground proximity warning systems (GPWS), and/or indicators when installed. Though models and manufacturers of these systems differ, all share a basic configuration which includes an indicator, a transceiver, and corresponding transmit and receive antennas. Operation is continuous when electrical power is supplied to the system, however, radio altitude readout is limited by the indicator’s usable range. Transceiver and Antennas The transceiver incorporates solid-state circuitry which makes instantaneous comparisons between the frequency of a frequency-modulated microwave signal that is beamed down from the transmit antenna to a return signal that is reflected back to the receive antenna from the terrain. Because the difference between these frequencies is proportional to the transmit signal’s “round-trip” time to the terrain and back, the frequency difference is processed to generate an electrical signal proportional to absolute altitude. Once generated, this signal is supplied to the indicator(s) and/or flight guidance systems which utilize radio altitude data in their operation. Depending on installation, the transceiver may be located within the right flight compartment sidewall or below the cabin floor. The antennas are located on the lower surface of the fuselage.

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Cessna Citation II Technical Manual Conventional Indicator (excluding RAD/BAR) The indicator may be located on the pilot’s or copilot’s instrument panel and incorporates an altitude scale and pointer, decision height (DH) index marker and setting knob, DH alert light, and a test switch. The altitude scale is calibrated in 10 foot increments below 500 feet, 100 foot increments above 500 feet, and identifies the indicator’s usable range. The pointer is visible at the corresponding AGL value within the usable range only. Outside this range, or if the airplane is steeply banked, the pointer will be positioned behind a mask in the upper left corner of the indicator dial. The DH setting knob, located on the lower right instrument bezel, is rotated to preselect a desired decision height. The DH alert light, located on the upper right instrument bezel, illuminates in conjunction with the sounding of an alert tone when decision height is reached or when the system is tested. The test switch is located on the lower left instrument bezel. The DH alert horn is located in the overhead flight compartment outboard of the pilot’s station. 28 VDC power is supplied to the system by the emergency avionics bus through the 5-amp RAD ALT circuit breaker. A failure warning flag will appear across the upper instrument face whenever electrical power is removed from the instrument or when radio altitude indication becomes invalid. Refer to the appropriate Operating Manual or Airplane Flight Manual for test procedures and information regarding specific capabilities of the system installed in your airplane.



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Conventional Indicator (RAD/BAR) The RAD/BAR indicator is located on the pilot’s instrument panel as previously described and incorporates a four-digit incandescent radio altitude display, decision height (DH) setting knob and alert light, and a test switch. The radio altitude display functions independently of the barometric portion of the instrument. During initial climb-out, radio altitude is displayed in 10 foot increments between 0 and 990 feet AGL. During descent, radio altitude is displayed in 100 foot increments between 2500 and 1000 feet AGL, and 10 foot increments between 1000 and 0 feet AGL. Above 2500 feet AGL the display is blank. The DH setting knob, located on the upper right instrument bezel, is rotated to preselect a desired decision height. The ALT alert light, located on the upper left instrument bezel, illuminates in conjunction with the sounding of an alert tone or GPWS vocal alert when decision height is reached or when the system is tested. The test switch is located on the lower left instrument bezel. The ALT alert light, also functions as a barometric altitude alert light by way of the VNCC as previously described. The RAD/BAR system also incorporates a converter, located in the nose avionics bay, which processes radio altitude signals supplied by the transceiver and glideslope signals supplied by the flight guidance system to support GPWS vocal alert functions. 28 VDC power is supplied to the system by the emergency avionics bus through the 5-amp RAD ALT circuit breaker. Should radio altitude data become invalid, a failure warning flag will appear adjacent to the RAD/ALT display. The failure warning flag in the upper center of the instrument face appears only when power is removed from the barometric portion of the instrument. Refer to the appropriate Operating Manual or Airplane Flight Manual for test procedures and information regarding specific capabilities of the system installed in your airplane.

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Cessna Citation II Technical Manual Mechanical Flight Director Radio Altitude Indication Radio altitude and decision height may also be displayed on the ADI(s) in mechanical flight director installations. Though configurations vary, reference to absolute altitude is generally provided by a bar with chevron markings or a rising runway symbol which appears in the lower center of the attitude sphere as the airplane descends below 200 feet AGL and moves toward the airplane symbol in relation to ground proximity until contact between the two occurs at the point of touchdown. Decision height is generally annunciated by a DH alert light, located on the upper right instrument bezel, which illuminates in conjunction with the sounding of an alert tone when decision height is reached or when the radio altitude system is tested. With the exception of the Sperry AD-650, the DH alert light does not illuminate when the ADI is tested. In addition to the features described above, radio altitude and decision height are also digitally displayed on the Sperry AD-650 (pictured) when installed. The four-digit incandescent RAD ALT dis2 1 play, located on the lower right instrument face, is calibrated in 5 foot increments between 0 3 and 200 feet AGL, and 10 foot 4 increments between 200 and 1. RAD ALT Display 2500 feet AGL. Above 2500 2. DH Display feet AGL the RAD ALT display 3. DH SET/DIM Knob is blank. Reference to ground 4. RA Test Button proximity is provided by a rising runway symbol as previously described. Should radio altitude data become invalid, four dashes will appear in the RAD ALT display. The three-digit decision height (DH) display, located in the lower left corner of the ADI, is calibrated in 10 foot increments between 0 and 990 feet AGL. The inner decision height (DH SET) setting knob, located on the lower right instrument bezel, is rotated to preselect a desired decision height between 0 and 990 feet AGL. The surrounding DIM ring permits adjustment of RAD ALT and DH display intensity. Should decision height data become invalid, three dashes will appear in the DH display.



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In dual AD-650 flight director installations, the alert tone is activated through the DH setting of the pilot’s ADI only. The DH setting of the copilot’s ADI and/or conventional radio altitude indicator, if installed, have no effect on alert tone activation. Refer to the appropriate Operating Manual or Airplane Flight Manual for test procedures and information regarding specific capabilities of the system installed in your airplane. EFIS Radio Altitude Indication Radio altitude and decision height may also be digitally displayed on the EADI(s) in EFIS installations. The four-digit radio altitude (RA) display, located in the lower right corner of the EADI, is calibrated in 5 foot increments between 0 and 200 feet AGL, and 10 foot increments between 200 and 2500 feet AGL. Above 2500 feet AGL the RA display is blank. Additional reference to absolute altitude is provided by a rising runway symbol which appears in the lower center of the attitude sphere as the airplane descends below 200 feet AGL and moves toward the airplane symbol in relation to ground proximity until contact between the two occurs at the point of touchdown. Should radio altitude data become invalid, four amber dashes will appear in the RA display and the rising runway will not be visible in the attitude sphere. The three-digit decision height (DH) display, located in the lower left corner of the EADI, is calibrated in 5 foot increments between 0 and 200 feet AGL, and 10 foot increments between 200 and 990 feet AGL. The decision height (DH/TST) setting knob, located on the EFIS display controller, is rotated to preselect a desired decision height between 0 and 990 feet AGL. Rotating this knob fully counterclockwise removes the DH display from the EADI. As the airplane descends to within 100 feet of decision height, a white box will appear above and left of the radio altitude (RA) display. An amber DH will appear within this box in conjunction with the sounding of an alert tone when decision height is reached or when the EFIS system is tested. Should decision height data become invalid, three amber dashes will appear in the DH display. In dual EFIS installations, the alert tone is activated through the DH setting of the pilot’s EADI only. The DH setting of the copilot’s EADI and/or conventional radio altitude indicator, if installed, have no effect on alert tone activation. Refer to the appropriate Operating Manual or Airplane Flight Manual for test procedures and information regarding specific capabilities of the system installed in your airplane. 7-20



Cessna Citation II Technical Manual Attitude and Direction System The attitude and direction system consists of all gyroscopic flight instruments (including mechanical and EFIS flight directors), turn-and-bank indicators, the RMIs, and the magnetic compass. Conventional Attitude Indicator The attitude indicator provides visual indication of pitch and roll in relation to the actual horizon. The unit consists of a gyro stabilized in the horizontal plane, an attitude sphere, airplane symbol, and roll index pointer. The gyro is air-driven on airplanes 550-0626 and earlier, or electrically-driven on airplanes 550-0627 and after, and responds to pitch and roll movements of the airplane. An electrically driven attitude indicator may also be installed on airplanes 550-0550 through 0626 as optional equipment. The attitude sphere is divided into sky and ground hemispheres by a horizon bar which provides visual reference to the actual horizon. The airplane symbol is secured to the instrument face and provides visual reference of the airplane’s attitude relative to the horizon bar. The roll index pointer is located at the top of the instrument face and provides visual reference of the airplane’s bank angle relative to the actual horizon. An inclinometer is installed on the lower instrument bezel of the electrically-driven attitude indicator to provide visual indication of turn coordination. The inclinometer is comprised of a ball contained in a sealed, silicone liquid filled, glass tube and responds to gravitational and centrifugal forces acting on the airplane. The tube is curved and mounted such that the ball will rest in the center lowest position when the airplane is in coordinated flight. In uncoordinated flight, the ball will move from the center to the outside of a turn (indicating a skid) or the inside of a turn (indicating a slip).



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Air-Driven Gyro On airplanes 550-0626 and earlier, the gyro is air-driven by regulated engine bleed air. Bleed air used by the system is tapped from the supply tubes between the compressor section of each engine and the environmental flow control/shutoff valves within the fuselage tailcone. On airplanes 550-0484, 0483, 0481 and earlier, the bleed air is routed through a precooler within each engine nacelle prior to entering the fuselage tailcone. On airplanes 550-0482, 0485 and after, availability of bleed air to the system is continuous when either or both engines are operating and is not influenced by the position of the PRESS SOURCE selector. On airplanes 550-0484, 0483, 0481 and earlier, however, setting the PRESS SOURCE selector to “LH” or “RH” correspondingly results in bleed air being made available to the system from the left engine or right engine only. The bleed air tapped from each source is routed through independent supply tubes to a common cross fitting within the fuselage tailcone. A check valve in each of these tubes prevents the backflow of bleed air to either engine when the opposite engine has failed or is operating at a sufficiently lower RPM. From the cross fitting, bleed air is routed to the windshield antiice/rain removal system, cabin pressurization control system ejector, and pneumatic distribution pressure regulator. From this regulator, 23.0 +/- 1.0 PSIG bleed air is routed to the pneumatic surface deice system, cabin temperature manual control system, inflatable cabin door seal, and the instrument air system. The instrument air system consists of a water separator/ filter, instrument pressure regulators, the air-driven attitude indicator, and gyro pressure gage. The water separator/filter is installed on the forward pressure bulkhead within the nose section and functions to filter and extract moisture from the bleed air prior to being routed to the instrument pressure regulators. Extracted moisture collects in the lower filter bowl and is eliminated through an orificed drain either to be vented overboard through an instrument air vent/ drain line which extends through the lower surface of the nose section (550-0173 ~ 0626), or evaporated within the nose section (550-0172 and earlier).

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Cessna Citation II Technical Manual The instrument pressure regulators are also installed on the forward pressure bulkhead within the nose section and function in series to reduce and stabilize bleed air pressure prior to supplying the instruments. The “first-stage” regulator reduces the pressure to approximately 6.0 PSIG; the “second-stage” regulator further reduces the pressure to approximately 2.5 PSIG. Regulated bleed air entering the attitude indicator case is directed against “buckets” machined into the rim of the gyro causing it to spin at a high rate of speed. After driving the gyro, bleed air is exhausted from the instrument case through the overboard instrument air vent/drain line previously described. Gyro Pressure Gage On airplanes 550-0626 and earlier, the gyro pressure gage provides visual indication of the bleed air pressure driving the gyro. The normal indicating range is denoted by a green arc between 2.0 and 3.0 PSIG. Pressure indications which are erratic and/or outside the normal range indicate that a malfunction may exist in the pneumatic system. From the gyro pressure gage, bleed air is also exhausted through the overboard instrument air vent/drain line. Electrically-Driven Gyro On airplanes 550-0627 and after, the gyro is electrically driven by 28 VDC power through the 2-amp ADI 2 circuit breaker. Power is supplied to the gyro from the emergency avionics bus whenever the battery switch is in the “BATT” (on) position and the DC avionics power switch is also in the “ON” position. Power is also supplied to the gyro when the battery switch is in the “EMER” (emergency) position, regardless of DC avionics power switch position. A red GYRO failure warning flag will be visible in the upper left instrument face when power is removed from the instrument. A spring-loaded caging knob is located on the lower right instrument bezel.



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Turn-and-Bank Indicator In single mechanical flight director systems, a turn-and-bank indicator is installed on the pilot’s and copilot’s instrument panel. In dual mechanical flight director installations, a turn-and-bank indicator is typically installed on the copilot’s instrument panel only. In dual 5” mechanical flight director installations, an air-driven “standby” attitude gyro replaces the copilot’s turn-and-bank indicator. The turn pointer is attached to a DC electrically-driven gyro which indicates the airplane’s turning rate in degrees per second. If installed, the copilot’s turn indicator is vacuum-driven. Operation of the turn indicator can be checked by initiating a standard rate turn and cross checking the turn rate with the heading indicator. An indicated standard rate turn should show a turning rate of 3° per second on the heading indicator. Visual indication of turn coordination is provided by an inclinometer on the lower instrument face. The inclinometer is comprised of a ball contained in a sealed, silicone liquid filled, glass tube and responds to gravitational and centrifugal forces acting on the airplane. The tube is curved and mounted such that the ball will rest in the center lowest position when the airplane is in coordinated flight. In uncoordinated flight, the ball will move from the center to the outside of a turn (indicating a skid) or the inside of a turn (indicating a slip).

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Cessna Citation II Technical Manual Mechanical Flight Directors Mechanical flight director system configurations are generally classified as 4” or 5”, single or dual flight director. In dual flight director systems, the pilot’s and copilot’s ADI and HSI may be identical or one of various combinations of 4” or 5” instruments depending on installation. The combination of instruments in each installation varies sufficiently between airplanes as to make specific descriptions of every possible system configuration impractical. The following description, therefore, primarily addresses the functional characteristics and operational features which are common among the mechanical ADIs and HSIs, as well as any relevant technical differences that exist between them. Attitude Directional Indicator The ADI or flight director indicator (FDI) functions as a conventional attitude indicator and displays command information provided by the autopilot/flight director computer. Depending on avionics equipment installed and modes available, the ADI may be utilized to intercept and maintain a desired heading, altitude, VOR radial, or localizer course and glideslope. Refer to the appropriate operating manual for specific capabilities of the system installed in your airplane. Horizontal Situation Indicator The HSI functions essentially as a slaved heading indicator and (depending on avionics equipment installed and modes available) provides visual indication of airplane position relative to VOR radials, RNAV courses, localizer courses, and glideslope beams. Refer to the appropriate operating manual for specific capabilities of the system installed in your airplane.



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To function as a slaved heading indicator, the HSI incorporates components which electromechanically produce a controlled precession of the gyro which corrects the compass card to agree with the correct magnetic heading. This installation eliminates the need for periodic resetting of the gyro due to precession drift. Gyro Slaving Gyro slaving is provided by a remotely-mounted magnetic flux detector and slaving accessory. The flux detector senses the direction of the earth’s magnetic field and transmits these signals to the slaving accessory where they are compared with gyro reference signals corresponding to the position of the compass card. The resulting error signal is amplified and transmitted to components which electromechanically produce a controlled precession of the gyro that corrects indicated heading to agree with the magnetic heading. The gyro slaving system is controlled by switches located on the lower left instrument panel and lower copilot’s instrument panel. A slaving meter, used to monitor displacement errors which may exist between indicated heading and magnetic heading, may also be installed. Indicator needle deflection in either direction from center corresponds to the polarity of the error.

1

2

1. Mode Selection Switch 2. Manual Slaving Switch

The gyro slaving switches permit selection of automatic slaving (slaved gyro) when positioned to “AUTO,” or manual slaving (free gyro) modes when positioned to “MAN.” In the slaved gyro mode, heading displacement errors are corrected automatically through the gyro slaving circuit. In the free gyro mode, heading displacement errors are corrected using the corresponding LH/RH switch to rotate the compass card left or right to agree with the magnetic heading, and return the slaving meter indicating needle to center (if installed). The rate of manual compass card rotation is approximately 30° per minute. Refer to the appropriate operating manual for specific capabilities of the system installed in your airplane.

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Cessna Citation II Technical Manual ADI Display Features The following summary describes the functional characteristics and operational features which are common to mechanical ADIs typically installed in the Cessna Citation II. Refer to the appropriate operating manual for specific capabilities of the system installed in your airplane. Attitude Sphere The attitude sphere is divided into sky (blue) and ground (brown) hemispheres by a horizon line which provides visual reference to the actual horizon. Eyelid Display The eyelid display surrounds the attitude sphere and provides visual reference of the relative position of the sky (blue) and ground (brown), independent of attitude sphere position, to facilitate recovery from unusual flight attitudes. Airplane Symbol The airplane symbol is located in the center of the instrument face and provides visual reference of the airplane’s attitude relative to the horizon line. Depending on ADI installation, the symbol may be conventional or reference delta type. Pitch Attitude Indication Pitch attitude is indicated by the relative position of the airplane symbol’s nose above or below the horizon line. A pitch scale on the attitude sphere references deviation above or below the horizon line in increments of 5°. Roll Attitude Indication Roll attitude is indicated by the relative position of the airplane symbol’s wings in relation to the horizon line. A roll attitude pointer and scale on the upper center instrument face references deviation from wings level attitude. The scale is marked at 10, 20, 30, 45, 60, and 90° to the left and right of the triangular 0° index mark at its center. To facilitate roll attitude recognition, the 30 and 60° marks are longer and heavier, while each 45° mark is displayed as a dot or a triangle.



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Depending on ADI installation, the pointer may be a moveable “sky pointer” or fixed “roll pointer.” In sky pointer configuration, the moveable pointer references the relative position of the sky in relation to the fixed scale index mark which references the airplane’s vertical axis. Roll recovery, therefore, is made in the direction of the pointer. In roll pointer configuration, the fixed pointer references the airplane’s vertical axis in relation to the moveable scale index mark which references the relative position of the sky. Roll recovery, therefore, is made in the direction of the scale index mark. Flight Director Command Bar(s) Depending on ADI installation, the command bar(s) may be double or single-cue. In double-cue flight directors, computed commands are displayed by independent pitch (horizontal) and steering (vertical) command bars as a conventional VOR/ILS indicator. Movement of the bars indicates pitch and steering control inputs required to satisfy computed commands of the selected flight director operating mode. In single-cue flight directors, computed pitch and steering commands are displayed by a single (delta) command bar. To satisfy computed pitch and steering commands of the selected flight director operating mode, the airplane symbol is “flown” to align with the command bar. Should loss of pitch or steering command signals from the flight director computer occur, the single-cue command bar or the affected double-cue command bar will retract from view. The non-affected double-cue command bar will continue to display normally. Glideslope Indication Glideslope deviation is indicated by the relationship between a moveable pointer and fixed vertical deviation scale commonly located on the right side of the instrument face. When a valid glideslope signal is being received, deviation above the beam centerline is indicated by displacement of the pointer below scale center. Deviation below the beam centerline is indicated by displacement of the pointer above scale center. Each dot above and below scale center represents approximately 0.4° deviation from the beam centerline. A green area on the scale denotes the category II (CAT II) approach window. Pointer displacement in this area indicates glideslope deviation within approximately 0.2° of the beam centerline.

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Cessna Citation II Technical Manual Expanded Localizer Indication Expanded localizer deviation is indicated by the relationship between a moveable “runway” pointer and fixed lateral deviation scale located on the lower center instrument face. When a valid localizer signal is being received, deviation left of the beam centerline is indicated by displacement of the pointer to the right of scale center. Deviation right of the beam centerline is indicated by displacement of the pointer to the left of scale center. Amplification of localizer data from the navigation receiver permits increased pointer sensitivity within the CAT II approach window. Because this increase in sensitivity makes tracking the localizer more difficult throughout the entire approach, the expanded localizer pointer should be referenced for position assessment only, until established on final approach. During final approach within the CAT II window, pointer displacement within the lateral deviation scale indicates localizer deviation within approximately 0.25° of the beam centerline, or within 33 feet of the runway centerline. Radio Altitude Indication Refer to the description of radio altimeter systems. Test Switch The attitude (ATT) test switch, located on the lower left instrument bezel, initiates an attitude self-test function. When pressed, the attitude sphere should be positioned to indicate approximately 20° right bank and 10° positive pitch, and the ATT warning flag should appear across the instrument face. Refer to the appropriate Operating Manual or Airplane Flight Manual for test procedures specific to the system installed in your airplane. GA Light The go-around light illuminates when the go-around mode has been selected DH Light The decision height light illuminates when the airplane descends below the selected decision height as set on the radio altitude indicator. Mode Annunciators (5”) Ten/twelve annunciators indicate which vertical and horizontal modes are engaged with the flight director.



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Turn and Bank Indication Rate-of-turn is indicated by the relationship between a moveable pointer and fixed scale located on the lower center instrument face of 5" ADIs. Pointer displacement over the left or right scale markings indicates a standard rate turn (2-minute/3° per second) in that direction. A conventional inclinometer is installed on the lower instrument bezel of ADIs to provide visual indication of turn coordination. The inclinometer is comprised of a ball contained in a sealed, silicone liquid filled, glass tube and responds to gravitational and centrifugal forces acting on the airplane. The tube is curved and mounted such that the ball will rest in the center lowest position when the airplane is in coordinated flight. In uncoordinated flight, the ball will move from the center to the outside of a turn (indicating a skid) or the inside of a turn (indicating a slip). Angle-of-Attack Indication (5” ADIs) Angle-of-attack (AOA) is indicated on 5” ADIs by the relationship between a moveable “speed command” pointer and fixed FAST/SLOW scale commonly located on the left side of the instrument face. The circular speed command pointer is positioned by signals received from the AOA transmitter. The pointer is calibrated such that its position relative to the scale corresponds with the AOA indicator on the upper left instrument panel. The FAST and SLOW scale markings correspond to the .4 and .8 AOA indicator markings respectively. The scale center marking corresponds to the .6 AOA indicator marking and represents the optimum landing approach speed (1.3 times stalling speed) for the current airplane configuration. Refer to the description of AOA systems for complete detail. Failure Warning Flags Depending on ADI installation, various failure warning flags will appear across the instrument face whenever electrical power is removed from the instrument or when the corresponding indication becomes invalid. Typically, 4” ADIs provide warning flags for attitude (ATT) and flight director (FD) indication failure only, while most 5” ADIs also provide warning flags for localizer (LOC), glideslope (GS), rate-of-turn (RT), and angleof-attack (SPD) indication failure.

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Cessna Citation II Technical Manual HSI Display Features The following summary describes the functional characteristics and operational features which are common to mechanical HSIs typically installed in the Cessna Citation II. Refer to the appropriate operating manual for specific capabilities of the system installed in your airplane. Compass Card and Heading Index A 360° rotating white compass card indicates airplane heading when referenced to the white triangular heading index in the upper center of the display. The compass scale is divided into 5° increments with 10° division markings being twice as long as the 5° markings. Fixed 45° index markings are positioned adjacent to the scale. Compass heading is referenced to magnetic north. Airplane Symbol The airplane symbol, located in the center of the instrument face, provides visual indication of the airplane’s position in relation to the course deviation bar. Alignment of the airplane symbol with the course deviation bar simulates alignment of the airplane’s flight path to the centerline of the selected navigation course or localizer. Course Deviation Indication Lateral deviation from the centerline of a selected navigation course or localizer is indicated by the relationship between a moveable bar and fixed deviation scale located in the center of the instrument. The deviation scale consists of two filled white circles evenly spaced on each side of the airplane symbol. The outer circles reference full scale deviation while the inner circles reference half scale deviation. Alignment of the course deviation bar with the airplane symbol represents alignment of the airplane’s flight path with the centerline of the selected navigation course or localizer. Heading Bug An orange heading bug is manually rotated about the compass card by the heading (HDG) select knob on the lower right corner of the instrument bezel. Once set, the heading bug rotates with the compass card. The heading bug functions to indicate desired heading and provides selected heading reference for autopilot steering.



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Course Pointer The course pointer is manually rotated about the compass card by the course (CRS) select knob on the lower left corner of the instrument bezel. Once set, the course pointer rotates with the compass card and provides indication of the desired navigation course to be flown. Depending on HSI installation, the selected course may be indicated in the upper left corner of the display. To/From Indicator A filled white triangle, pointing either toward the head (to) or the tail (from) of the course pointer, functions to provide visual indication that the selected course will take the airplane to or from the selected navaid or waypoint. The to/from indicator is not displayed during ILS operation or when an invalid navaid or waypoint signal is received. Distance to Station (DME) Indication Depending on HSI installation, DME information in nautical miles may be indicated in the upper right corner of the display. Glideslope Indication Glideslope deviation is indicated by the relationship between a moveable pointer and fixed vertical deviation scale commonly located on the right side of the instrument face. The deviation scale consists of two filled white circles evenly spaced above and below a filled white diamond. The outer circles reference full-scale deviation while the inner circles reference half-scale deviation. Alignment of the green pointer with the center diamond represents alignment of the airplane’s glide path with the glideslope centerline. Failure Warning Flags Depending on HSI installation, various failure warning flags will appear across the instrument face whenever electrical power is removed from the instrument or when the corresponding indication becomes invalid. Typically, warning flags are provided for navigation (NAV) and heading (HDG) indication failure, and vertical gyro (VERT) failure.

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Cessna Citation II Technical Manual Electronic Flight Instrumentation System (EFIS) Refer to the appropriate operating manual for specific capabilities of the system installed in your airplane. Autopilot/Flight Director System (AP/FD) Refer to the appropriate operating manual for specific capabilities of the system installed in your airplane. Limitations Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.



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Chapter 8 Hydraulics and Landing Gear Table of Contents Overview ............................................... 8-1 Landing Gear System Description ........... 8-1 Nose Gear ............................................ 8-2 Main Gear ............................................ 8-6 Hydraulic Actuators........................... 8-8 Uplock Sequence Actuators .............. 8-9 Left Main Gear Safety Switch .............. 8-10 Landing Gear Control and Position Indicator ................................ 8-12 Landing Gear Control Handle .......... 8-12 Landing Gear Warning Horn ................. 8-15 Landing Gear Indicator Light and Warning Horn Test .................. 8-16 Landing Gear Hydraulics ...................... 8-17 Landing Gear Hydraulic Pressure Source ............................ 8-17 Hydraulic Power System Components ...................... 8-17 Hydraulic Fluid Reservoir ................. 8-17 Engine-Driven Hydraulic Pumps ........ 8-21 Hydraulic Filters ............................. 8-21 Hydraulic Firewall Shutoff Valves ..... 8-22 Hydraulic System Indication ................. 8-23 Hydraulic Flow Annunciators ............ 8-23 Hydraulic Pressure Annunciator ....... 8-24 Landing Gear Control Valve............. 8-25


Hydraulics and Landing Gear, continued Table of Contents Landing Gear Operation ....................... 8-26 Landing Gear Retraction ................. 8-26 Landing Gear Extension .................. 8-28 Static (Open Center) Condition ........ 8-28 Landing Gear Extension/Retraction Speeds and Cycle Times ................ 8-28 Auxiliary/Emergency Landing Gear Extension System ............................... 8-30 Auxiliary/Emergency Air Storage Bottle ............................... 8-30 Auxiliary/Emergency Gear Extension System Operation .......................... 8-31 Emergency Hydraulic Dump Valve .... 8-33 Brakes ............................................... 8-34 Power Brake System ...................... 8-34 Touchdown Protection ..................... 8-37 Anti-Skid System ............................ 8-38 Anti-Skid System Test .................... 8-39 Locked Wheel Crossover Protection ...................................... 8-40 Auxiliary/Emergency Braking System .............................. 8-40 Auxiliary/Emergency Braking System Usage ................... 8-41 Parking Brake System .................... 8-41 Limitations .......................................... 8-42 Emergency Procedures......................... 8-42

Cessna Citation II Technical Manual Overview The Cessna Citation II is equipped with retractable tricycle landing gear that is electrically-controlled and hydraulicallyoperated. The main gear assemblies are located in each wing. The nose gear assembly is located in the nose section of the fuselage. Normal gear extension and retraction is activated by the landing gear control handle. Emergency gear extension is provided by a mechanical uplock release “free-fall” system and a pneumatic “blowdown” system. The airplane has an independent hydraulic system for the main gear wheel brakes. Emergency braking, anti-skid and other braking systems are provided. Landing Gear System Description Normal extension and retraction is accomplished by directing hydraulic fluid under engine-driven pump pressure to an actuator at each gear assembly. All three gear assemblies are held in the extended position by mechanical downlock latches internal to each actuator, and held in the retracted position by uplock hooks. Hydraulic pressure releases the downlock latches during gear retraction, and uplock hooks during gear extension. Six microswitches are incorporated; three actuated by the uplock hooks and three actuated by the downlock latches. These microswitches operate in conjunction with the landing gear control, position indication, and warning circuitry.

HYDRAULICS AND LANDING GEAR 12/99


8-1

Nose Gear

The nose gear is of conventional design incorporating a lower piston/fork assembly, upper trunnion assembly, shock strut assembly, torque links, and a hydraulic shimmy damper. The piston/fork assembly provides attachment points and running clearance for the nosewheel. Attachment points for installation of the nose gear to the nose wheel well structure are provided by the trunnion assembly which also houses the shock strut assembly. Upper and lower torque links connect the piston/fork assembly to the trunnion assembly and maintain alignment of the nosewheel.

8-2



Cessna Citation II Technical Manual The nose gear incorporates an air-oil type shock strut assembly which functions to support the weight of the airplane and absorb shock loads encountered during ground operation. The shock strut assembly consists primarily of an inner cylinder retained within a hydraulic fluid filled outer cylinder. The inner cylinder is integral to the piston/fork assembly while the outer cylinder is housed within the trunnion assembly. A floating isolation piston divides the interior of the inner cylinder into a hydraulic fluid filled upper chamber and nitrogen gas or dry air charged lower chamber. A fixed orifice separates the upper chamber of the inner cylinder from the outer cylinder. A tapered metering pin regulates the flow of hydraulic fluid through this orifice in relation to increasing or decreasing load. Under increasing load, hydraulic fluid flows from the outer cylinder to the upper chamber of the inner cylinder. As the strut compresses, the metering pin progressively restricts the orifice. When the rate of hydraulic fluid flow through the orifice is insufficient to absorb compression shocks, the isolation piston is forced downward against gas/air pressure to assume the additional load. Under decreasing load, the isolation piston is forced upward by gas/air pressure and hydraulic fluid flows from the upper chamber of the inner cylinder to the outer cylinder until the pressure on each side of the orifice is equal. The shimmy damper consists 2 of a hydraulic fluid filled outer cylinder assembly and an internal piston assembly, attached to the nose gear 1 such that the piston effectively moves within the cylinder as the nosewheel is turned. Movement of the piston, caused by lateral oscillation of the nosewheel, is dampened by the hydraulic fluid which 1. Shimmy Damper must be forced through ori2. Compensating Chamber fices in the piston as it moves within the cylinder. The shimmy damper also incorporates a compensating chamber which houses a spring-loaded valve that relieves thermal expansion of the hydraulic fluid.



8-3

The nose gear retracts 1 forward into the nose wheel well and, when fully re2 2 tracted, is enclosed by three doors. The two forward doors are linked by push-pull rods to a common torque tube and 4 3 bellcrank assembly such that they operate simultaneously. The torque tube and bellcrank assembly is linked by a single pushpull rod to the trunnion 1. Torque Tube assembly such that the 2. Push-Pull Rods forward doors mechani3. Trunnion Assembly cally open during exten4. Forward Door Hinges sion and retraction, and close following extension or retraction. The aft door is also linked by a single push-pull rod to the trunnion assembly such that it mechanically opens during extension and closes during retraction. A universal joint straightening mechanism centers the nose gear during retraction.

8-4



Cessna Citation II Technical Manual Landing Gear Assemblies

Main Gear

HYDRAULIC OIL CHAMBER FILL PLUG

ACTUATOR

TRUNNION ASSEMBLY

SWIVEL FITTING DRAG BRACE

SHOCK STRUT

MAIN GEAR DOOR LINKAGE

SQUAT SWITCH AIR CHAMBER AIR VALVE

BRAKE LINES

TORQUE LINKS

Nose Gear

AXLE BRAKE ASSEMBLY

STEERING GEARS HYDRAULIC ACTUATOR

SHIMMY DAMPER

FORWARD DOOR LINKAGE

STEEL UNIVERSAL JOINT

DOWNLOCK SWITCH

TRUNNION AFT DOOR LINKAGE

TORQUE LINKS



8-5

Main Gear

The main gear is of conventional design incorporating a lower piston/axle assembly, upper trunnion assembly, shock strut assembly, drag brace, and torque links. The piston/axle assembly provides attachment points and running clearance for the main wheel and wheel brake assembly. Attachment points for installation of the main gear to the front and rear wing spar are provided by the trunnion assembly which also houses the shock strut assembly. The drag brace interconnects the aft upper and lower ends of the trunnion assembly and functions to reduce fore and aft forces on the shock strut assembly. Upper and lower torque links connect the piston/axle assembly to the trunnion assembly and maintain alignment of the main wheel.

8-6



Cessna Citation II Technical Manual Each main gear assembly incorporates an air-oil type shock strut assembly which functions to support the weight of the airplane and absorb shock loads encountered during ground operation. The shock strut assembly consists primarily of an inner cylinder retained within a hydraulic fluid filled outer cylinder. The inner cylinder is integral to the axle/piston assembly while the outer cylinder is housed within the trunnion assembly. A floating isolation piston divides the interior of the inner cylinder into a hydraulic fluid filled upper chamber and nitrogen gas or dry air charged lower chamber. A variable orifice separates the upper chamber of the inner cylinder from the outer cylinder and regulates the flow of hydraulic fluid between these areas in relation to increasing or decreasing load. Under increasing load, hydraulic fluid flows from the outer cylinder to the upper chamber of the inner cylinder. As the strut compresses, the variable orifice is progressively restricted. When the rate of hydraulic fluid flow through the orifice is insufficient to absorb compression shocks, the isolation piston is forced downward against gas/air pressure to assume the additional load. Under decreasing load, the isolation piston is forced upward by gas/air pressure and hydraulic fluid flows from the upper chamber of the inner cylinder to the outer cylinder until the pressure on each side of the orifice is equal. The main gear assemblies retract inward into the wing wheel wells and, when fully retracted, each is partially enclosed by a hinged door which opens during extension and closes during retraction. Push-pull rods mechanically link the doors to their respective trunnions such that they operate simultaneously.



8-7

Hydraulic Actuators Each hydraulic actuator is comprised of a cylinder body with an internal piston and rod assembly. The cylinder body of each actuator is attached to the airframe structure. Each rod end is attached to its associated gear trunnion assembly. Retract and extend ports on each cylinder are plumbed to the landing gear hydraulic system. Separate pneumatic extend ports are plumbed to the pneumatic blowdown system. 5 4

6

2 1 3

1. Cylinder Body 4. Hydraulic Retract Port 2. Piston Rod 5. Hydraulic Extend Port 3. Downlock Indicator 6. Pneumatic Retract Port The nose gear actuator piston “pulls” the nose gear to the extended position and “pushes” it to the retracted position. The main gear actuator piston “pushes” the main gear to the extended position and “pulls” it to the retracted position. The nose gear actuator incorporates a shuttle valve that is normally spring-loaded open to the hydraulic extend port. During auxiliary/emergency gear extension, the shuttle valve is repositioned open to the pneumatic extend port by gas/air pressure. Each main gear actuator incorporates a separate, concentric, pneumatic extension chamber that is always open to the pneumatic extend port. All three gear actuators incorporate integral “ring and groove” type downlock latches which hold the gear assemblies in the fully-extended position when hydraulic pressure is removed. Approximately 300 PSI is required to release these latches.

8-8



Cessna Citation II Technical Manual Uplock Sequence Actuators 4 Independent uplock 1 sequence actuators are provided for each gear assembly to hydraulically release the uplock hooks during normal gear extension. Each consists of an actuator 3 body, a spring-loaded 2 internal piston and rod assembly, and a check valve. The actuator body is attached to the 1. Actuator Body airframe structure. The 2. Uplock Hook uplock hook is linked to 3. Release Cable the rod end. Three 4. Uplock Switch ports are located on each actuator body: the first (pressure inlet) is plumbed to the extend circuit of the landing gear hydraulic system, the second (pressure outlet) is plumbed to the extend port of its associated gear assembly’s hydraulic actuator, the third is plumbed to the retract port of its associated gear assembly’s hydraulic actuator. The check valve is positioned between the pressure inlet port and pressure outlet port such that hydraulic fluid flow to the extend port of the hydraulic actuator is restricted until the uplock hook is released. During normal gear extension, hydraulic pressure applied to the internal piston “pulls” the rod inward thereby releasing the uplock hook. Following uplock hook release, continued inward movement of rod unseats the check valve allowing hydraulic fluid flow to the extend port of the hydraulic actuator. During gear retraction, return hydraulic fluid flow from the extend port of the hydraulic actuator unseats the check valve. During emergency gear extension, the uplock hooks are mechanically-released.



8-9

Left Main Gear Safety Switch A safety switch, located on the left main gear, is installed primarily to prevent inadvertent retraction of the gear during ground operation when the left main gear strut is compressed. Numerous additional functions are listed below.

On Ground, Strut Compressed, Safety Switch Open Enables: 1.

Generator-assisted engine start

2.

Automatic ground cabin depressurization

3.

Wheel brake anti-skid

4.

Locked wheel crossover protection (550-0437 and after)

5.

Thrust reverser deployment

6.

Sperry TAS heater probe low heat level (550-0505 and earlier)

7.

Tailcone pressurization

Disables:

8-10

1.

Landing gear handle movement

2.

Emergency pressurization control valve

3.

Touchdown protection (550-0437 and after)

4.

Air Data Computer (550-0324 and after)

5.

Stick shaker

6.

Optional approach indexer (550-0627 and after)

7.

Air data warning horn (550-0505 and earlier)

8.

Ground Proximity Warning System (550-0376 and after)

9.

Flight data recorder (550-0550 and after)

10.

Cockpit voice recorder (550-0550 and after)

11.

Angle of attack probe heat

12.

Sperry TAS heater probe high heat level (550-0505 and earlier)

13.

Hobbs meter

14.

Davtron digital clock flight time function

15.

Panel light dimming (550-0689 ~ 0698; 550-0703 and after)

16.

Angle of attack indexer dimming (550-0550 and after)



Cessna Citation II Technical Manual In Flight, Strut Extended, Safety Switch Closed Enables: 1.

Landing gear handle movement

2.

Emergency pressurization control valve

3.

Touchdown protection (550-0437 and after)

4.

Air Data Computer (550-0324 and after)

5.

Stick shaker

6.

Optional approach indexer, if nose gear is down and locked (550-0627 and after)

7.

Altitude alert warning horn

8.

Ground Proximity Warning System (550-0376 and after)

9.

Flight data recorder (550-0550 and after)

10.

Cockpit voice recorder (550-0550 and after)

11.

Angle of attack probe heat

12.

Sperry TAS heater probe high heat level (550-0505 and earlier)

13.

Hobbs meter

14.

Davtron digital clock flight time function

15.

Panel light dimming (550-0689 ~ 0698; 550-0703 and after)

16.

Angle of attack indexer dimming (550-0550 and after)

Disables:


1.

Generator-assisted engine start

2.

Automatic ground cabin depressurization

3.

Wheel brake anti-skid

4.

Locked wheel crossover protection (550-0437 and after)

5.

Thrust reverser deployment

6.

Sperry TAS heater probe low heat level (550-0505 and earlier)

7.

Tailcone pressurization


8-11

Landing Gear Control and Position Indication

Landing Gear Control Handle The landing gear hydraulic system is activated by a two-position, wheel-shaped control handle, labeled LDG GEAR - UP/ DOWN, located in the lower left corner of the center instrument panel. Three microswitches are actuated by the control handle: a retract switch, an extend switch, and a selector switch. The retract switch is actuated when the control handle is moved to the “UP” position; the extend switch is actuated when the control handle is moved to the “DOWN” position. When either switch is actuated, a corresponding retract or extend solenoid integral to the landing gear control valve is energized, thereby positioning the valve to permit gear retraction or extension as selected. The control handle is spring-loaded to the selected position and must be pulled outward before it can be moved. Protection against inadvertent retraction of the gear during ground operation is provided by a spring-loaded, solenoidoperated handle lock. The solenoid is energized and deenergized through the left main gear safety switch. When the left main gear strut is compressed, the solenoid is deenergized and the spring-loaded lock prevents the control handle from being moved to the “UP” position. When the left main gear strut is uncompressed, the solenoid is energized by 28 VDC left main bus power through the 2-amp LDG GEAR circuit breaker, and the handle lock is released.

8-12



Cessna Citation II Technical Manual The selector switch is integral to the landing gear position indication and warning system and has two positions (up and down) that correspond with control handle position. When the control handle is in the “UP” position, the selector switch establishes a circuit to the GEAR UNLOCKED light module through the uplock and downlock microswitches. When the control handle is in the “DOWN” position, the selector switch establishes a circuit to the GEAR UNLOCKED light module through downlock microswitches only.



8-13

1. Gear Down and Locked Light Module 1 A light module incorporating three green lenses individually labeled NOSE, LH, and RH, located to the right of the landing gear control handle, provides indication that the 2 landing gear is in the down and locked position. The LH and RH main gear position indicators are illuminated by one bulb each. The NOSE gear position indicator is illuminated by two bulbs wired in parallel for continued operation should one bulb fail. The down and locked position indicators are individually illuminated through the downlock microswitches as each gear assembly reaches its fully-extended position and its associated downlock latch is engaged. 2. Gear Unlocked Light Module A light module incorporating a single red lens labeled GEAR UNLOCKED, located below the gear down and locked light module, provides indication that the landing gear is in transit or that one or all three landing gear assemblies are not in the same position as the landing gear control handle. This indicator is illuminated by two bulbs wired in parallel for continued operation should one bulb fail. During gear retraction, the indicator is illuminated through the selector switch, downlock microswitches, and uplock microswitches when the control handle is positioned to “UP,” the downlock latches are released, and uplock hooks are not engaged. During gear extension, the indicator is illuminated through the selector switch and the downlock microswitches only, when the control handle is positioned to “DOWN” and the downlock latches are not engaged. During extension and retraction, the downlock microswitches provide the ground required for GEAR UNLOCKED indicator illumination only when downlock latches are not engaged. When all three gear assemblies have reached their fully-retracted or fully-extended position the GEAR UNLOCKED indicator should be extinguished. When all three gear assemblies are up and locked, all position indicators should be extinguished.

8-14


CITATION II CHAPTER 8 12/99 R-6/27/96

Cessna Citation II Technical Manual Landing Gear Warning Horn The landing gear warning horn is located in the overhead flight compartment, aft and outboard of the pilot’s station, and provides audible indication that the gear is not down and locked when the airplane is configured for landing. The horn is energized through a pair of throttle position switches, a flap position switch, and an airspeed switch (550-0627 and after), integral to the landing gear position indication and warning circuit. Each throttle position switch is actuated closed when its associated THROTTLE lever is set below approximately 70% N2. The flap position switch is actuated closed by the FLAP lever when set below the T.O. & APPR position (approximately 15°). The airspeed switch, incorporated on airplanes 550-0627 and after only, is closed below approximately 150 KIAS. In combination, these switches and the downlock microswitches will cause the horn to sound when any one or all three gear assemblies are not down and locked under the following conditions: Airplanes 550-0626 and earlier 1. Either or both THROTTLE levers set below approximately 70% N2 regardless of airspeed 2. FLAP lever set below T.O. & APPR position, regardless of airspeed or THROTTLE position Airplanes 550-0627 and after 1. Either or both THROTTLE levers set below approximately 70% N2 when airspeed is below approximately 150 KIAS 2. FLAP lever set below T.O. & APPR position, regardless of airspeed or THROTTLE lever position

A HORN SILENCE button, located right of the control handle, energizes a pair of relays associated with the throttle position switches. When this button is pressed, both relays are energized thereby interrupting the throttle position switch circuit and silencing the warning horn. ▲ HORN SILENCE BUTTON



8-15

Each relay remains energized through an integral latching circuit as long as its associated throttle position switch remains closed and all three gear assemblies are not down and locked. If the horn has been energized through the throttle position switches only, it may also be silenced by advancing the THROTTLE levers above approximately 70% N2, or by increasing airspeed above 150 KIAS (550-0627 and after only). If the horn has been energized through the flap position switch, pressing the HORN SILENCE button, advancing the THROTTLE levers, or increasing airspeed (550-0627 and after only), will not silence the horn. Landing Gear Indicator Light and Warning Horn Test Operation of the landing gear position indication and warning system may be verified by rotating the TEST selector switch, located on the lower left instrument panel, to the “LDG GEAR” position. This energizes a test relay that bypasses the downlock microswitches which normally ▲ TEST SELECTOR SWITCH provide the ground required to illuminate the position indicators and sound the warning horn. In this condition, if the system is functioning normally, the LH, RH, NOSE, and GEAR UNLOCKED position indicators should be illuminated and the warning horn should sound. Operation of the HORN SILENCE button may also be verified during this test. If functioning normally, pressing the button should silence the horn. The position indicators and warning horn normally receive 28 VDC power from the left main bus (550-0550 and after) or right main bus (550-0505 and earlier) bus through the 2-amp LDG GEAR circuit breaker. During system testing, the test relay and GEAR UNLOCKED position indicator receive 28 VDC power from the left main bus (550-0550 and after) or right main bus (550-0505 and earlier) through the 5-amp WARN LTS 1 circuit breaker.

8-16



Cessna Citation II Technical Manual Landing Gear Hydraulics Landing Gear Hydraulic Pressure Source The hydraulic power system provides hydraulic pressure for landing gear retraction and normal extension. The system also supplies hydraulic pressure for operation of the speed brakes and optional thrust reversers when installed. A detailed description of the hydraulic power system is provided below. Hydraulic Power System Components Major components of the hydraulic power system include a hydraulic fluid reservoir, engine-driven hydraulic pumps, solenoid-operated control valve, and annunciators that indicate system operating conditions. Hydraulic Fluid Reservoir The hydraulic fluid reservoir, located within the aft fuselage, stores fluid and supplies the hydraulic pumps. The reservoir is comprised of two cylindrical sections joined end-to-end. The larger cylinder houses the fluid reservoir; the smaller cylinder (neck) primarily houses a fluid level sight gage. A piston divides the larger cylinder into separate fluid-filled and air-filled chambers. On airplanes 550-0180 and after, and airplanes 5500065 and earlier not incorporating SB550-29-1 or -2, the reservoir is considered pressurized. On airplanes 550-0066 through 0179, and airplanes 550-0065 and earlier incorporating SB55029-1 or -2, the reservoir is considered non-pressurized. In pressurized reservoirs, the large reservoir piston is attached to and positioned by a small piston located within the neck. The small piston is positioned by hydraulic pressure. With 1350 to 1500 PSI hydraulic pressure acting on the small piston, the large piston maintains approximately 15 PSI within the fluidfilled chamber. The large piston is spring-loaded to maintain approximately 3 to 4 PSI within the fluid-filled chamber when hydraulic pressure is not available. In non-pressurized reservoirs, the piston is spring-loaded to maintain approximately 3 to 4 PSI within the fluid-filled chamber. The air-filled chamber is vented to facilitate piston movement.



8-17

Hydraulic Distribution System

8-18


© PCW


Cessna Citation II Technical Manual Overpressurization of the fluid-filled chamber is prevented by a relief valve that begins opening at 40 PSI and fully opens at 60 PSI. When open, excess hydraulic fluid is routed through an overboard relief line. On airplanes 550-0482, 0485 through 0698 incorporating SB550-29-06, and airplanes 550-0699 and after, this fluid drains into a plastic EPA bottle within the aft fuselage. On airplanes 550-0484, 0483, 0481 and earlier, and airplanes 550-0482, 0485 through 0658 not incorporating SB550-29-06, this fluid drains overboard through the hydraulic service panel vent mast located on the lower right surface of the aft fuselage.

© PCW

The fluid level sight gage consists of an indicator rod, visible through a window on the reservoir neck, that is attached to and positioned by the reservoir piston. The position of the rod in relation to REFILL, FULL, and OVERFULL markings above the window, and corresponding 38, 125, and 150 IN3 markings below the window, indicates the reservoir fluid level. The 38, 125, and 150 IN3 markings correspond to 0.2, 0.5, and 0.6 gallons respectively. A microswitch is attached to the neck such that it is held open by the indicator rod when it is positioned above REFILL. When the rod is positioned below REFILL, the microswitch closes illuminating the amber [HYD LOW LEVEL] annunciator (550-0550 and after) or [HYD LEVEL LO] annunciator (550-0505 and earlier).



8-19

The reservoir is serviced through a pressure line coupling on the hydraulic service panel. Reservoir capacity is approximately 0.65 gallons with the gear down and locked. The hydraulic system reservoir may be serviced with Skydrol 500A, 500B, 500B-4, 500C, and LD-4; Hyjet, Hyjet W, III, or IVA; or any equivalent phosphate ester based hydraulic fluid. Mixing hydraulic fluids should not impair system operation. Caution: Phosphate ester based hydraulic fluid will attack a wide range of materials, including rubber, copper, various plastics, and paints. Caution: If heated beyond 270°F, Skydrol decomposes into acids and other products that can damage metal structures.

8-20



Cessna Citation II Technical Manual Engine-Driven Hydraulic Pumps Hydraulic pressure is provided by constant-displacement hydraulic pumps that are geardriven by and mounted on the accessory gearbox of each engine. When hydraulic pressure is not required to operate airplane systems, the hydraulic system functions in an “open center” fashion, as the pumps draw hydraulic fluid from the reservoir, circulate it through the hydraulic system, and return it to the reservoir through the return line. When hydraulic pressure is required to operate landing gear or other systems, a bypass valve is energized closed, thereby allowing hydraulic pressure to increase. A check valve is installed in the return line to the reservoir to prevent reverse flow from the reservoir. Check valves, installed downstream from each pump and its associated filter, prevent reverse fluid flow in the event of opposite pump failure. These check valves also incorporate flow switches or flow detectors that illuminate corresponding annunciators to indicate low hydraulic flow/pressure from each hydraulic pump. A restrictor check valve is installed in the pressure line downstream of the right hydraulic pump to prevent reverse fluid flow during servicing. Hydraulic Filters A hydraulic filter is installed in each pump pressure line and in the hydraulic reservoir return line to prevent foreign material from contaminating the hydraulic fluid. The pump pressure line filters have a 3 GPM nominal flow capacity, and incorporate a 100 PSI differential bypass valve that permits continued flow in the event of filter blockage. The reservoir return line filter has a 12 GPM nominal flow capacity and incorporates a 100 PSI differential bypass valve. These filters have a 5-micron nominal rating and a 15-micron absolute rating.



8-21

Hydraulic Firewall Shutoff Valves A hydraulic firewall shutoff valve is installed within the aft fuselage in each hydraulic pump suction line. The hydraulic firewall shutoff valves function to terminate hydraulic fluid flow in the event of an engine fire. Each ball-type valve is motor-operated and electrically-controlled by independent [LH ENG FIRE] and [RH ENG FIRE] annunciator/switches on the glare shield panel fire tray. A transparent, spring-loaded guard is installed over each switch to protect against inadvertent actuation. When either switch is pressed, the corresponding valve is closed by 28 VDC power supplied through its associated 7.5amp LH F/W SHUTOFF or RH F/W SHUTOFF circuit breaker on the left CB panel. When fully closed, the corresponding amber [LH] or [RH] [HYD FLOW LO] annunciator (550-0550 and after) or [L] or [R] [HYD PRESS LO] annunciator (550-0505 and earlier) should be illuminated. Additionally, both [BOTTLE ARMED PUSH] annunciator/switches on the glare shield panel fire tray should be illuminated. A closed valve may be reopened by pressing the corresponding switch to release it from its locked position. Note: The fuel firewall shutoff valve is also closed when the hydraulic firewall shutoff valve is closed. Refer to Chapter 3 for a complete description of the fuel firewall shutoff valve. Refer to Chapter 2 for a complete description of the engine fire protection system.

8-22



Cessna Citation II Technical Manual Hydraulic System Indication Indication of hydraulic system operation is provided by flow, pressure, and fluid level lights located on the annunciator panel. Hydraulic Flow Annunciators Independent amber [LH] and [RH] [HYD FLOW LOW] annunciators (550-0550 and after) or [L HYD PRESS LO] and [R HYD PRESS LO] annunciators (550-0505 and earlier) illuminate to indicate low hydraulic flow from each engine-driven pump. On airplanes 550-0050 through 0063, 0065 and after, the hydraulic fluid flow annunciators are illuminated by independent left and right flow switch/check valves. Each flow switch/check valve is comprised of a sliding magnet that is attached to and moves with the check valve poppet. If the system is operating normally, each annunciator will illuminate when the airplane electrical system is energized, and will remain illuminated until its associated hydraulic pump develops sufficient pressure to unseat the check valve, thereby magnetically opening the switch and extinguishing the corresponding annunciator. On airplanes 550-0064, 0049 and earlier, the hydraulic fluid flow annunciators are illuminated by a combination check valve/flow detector that is comprised of a sliding magnet and two switches. One switch controls illumination of the amber [L HYD PRESS LO] annunciator; the other switch controls illumination of the amber [R HYD PRESS LO] annunciator. The sliding magnet is spring-loaded to a neutral position. When an inlet port pressure differential exists, the magnet is forced toward the area of lowest pressure. If one pump is generating at least 25 PSI more than the other, the magnet closes the switch corresponding to the pump that is generating the lower pressure, thereby illuminating the associated annunciator. If the system is operating normally, the annunciator corresponding to the engine started last will illuminate following initial engine start, and will remain illuminated until its associated hydraulic pump develops sufficient pressure to move the flow detector magnet to the neutral position, thereby magnetically opening the switch and extinguishing the corresponding annunciator.



8-23

Hydraulic Pressure Annunciator Pressurization of hydraulic components is indicated by an amber [HYD PRESS ON] annunciator. This annunciator is illuminated when pressure exceeds approximately 155 PSI, by a hydraulic pressure switch, located upstream of the landing gear control valve.

8-24



Cessna Citation II Technical Manual Landing Gear Control Valve The landing gear control valve functions to direct hydraulic fluid to the hydraulic actuator retract or extend ports. The hydraulic landing gear control valve is operated by “retract” and “extend” solenoids. The control valve contains an internal selector spool that is spring-loaded to a neutral position when the solenoids are not energized. In this position, both the retract and extend ports are connected to the return port so that fluid will not be trapped under pressure within the hydraulic lines. The other selector spool positions depend upon the position of the landing gear control handle. When the control handle is moved to the “UP” position, the retract solenoid is energized, which positions the selector spool to connect the hydraulic inlet port to the retract port and connect the extend port to the return port. When the control handle is moved to the “DOWN” position the extend switch is actuated, energizing the extend solenoid, which positions the selector spool to connect the hydraulic inlet port to the extend port and connect the retract port to the return port. The control handle is spring-loaded to the selected position and must be pulled outward before it can be moved. The landing gear control valve functions with an input power of 18 to 30 VDC. System pressure is regulated by a relief valve that begins to open at approximately 1350 PSI, and prevents hydraulic system pressure from exceeding 1500 PSI. The relief valve also prevents hydraulic fluid flow rate from exceeding 6.6 GPM. Excess fluid released through the relief valve flows into the reservoir return line. During landing gear extension or retraction, the [HYD PRESS ON] annunciator should be illuminated as an indication of adequate hydraulic pressure. Failure of this light to illuminate indicates that hydraulic pressure is insufficient for system operation. Continued illumination after all three landing gear assemblies have reached the selected position indicates malfunction of a hydraulic system component.



8-25

Landing Gear Operation The following paragraphs describe normal landing gear retraction, as well as normal and emergency landing gear extension. The airplane has no emergency landing gear retraction capability. Landing Gear Retraction Landing gear retraction is initiated by positioning the landing gear control handle to “UP.” In this position, the control valve retract solenoid and hydraulic system center valve are energized to permit landing gear retraction. During retraction, the control valve selector spool to is positioned to direct hydraulic fluid to the retract port of each actuator and the center valve closes, illuminating the amber [HYD PRESS ON] annunciator. Once hydraulic pressure releases the downlock latches, the retract side of the gear actuator pistons is pressurized, thereby initiating gear retraction. Release of the downlock latches actuates downlock switches that illuminate the red GEAR UNLOCKED light while the gear is in transit. As each gear assembly reaches its fully-retracted position, it is held in position by a gear uplock hook that engages an uplock roller on the gear trunnion. An uplock switch is actuated closed by each uplock hook. When the landing gear is fully retracted and all three microswitches have closed, the red GEAR UNLOCKED light is extinguished, the hydraulic system center valve opens, and the retract solenoid is deenergized. In this condition, the control valve selector spool returns to its neutral position and the hydraulic system returns to “open center” mode, extinguishing the amber [HYD PRESS ON] annunciator. Note: When the center valve is energized and hydraulic pressure is being supplied to the landing gear, the amber [HYD PRESS ON] annunciator will be illuminated.

8-26



© PCW

8-27



Landing Gear System Schematic (retraction)


Landing Gear Extension Landing gear extension is normally initiated by positioning the landing gear control handle to “DOWN.” In this position, the control valve extend solenoid and hydraulic system center valve are energized to permit landing gear extension. During extension, the control valve selector spool to is positioned to direct hydraulic fluid to the uplock sequence actuators and the center valve closes, illuminating the amber [HYD PRESS ON] annunciator. As sufficient hydraulic pressure builds within these actuators, the uplock hooks are released. After each uplock hook is fully released, fluid is routed to the extend port of its respective landing gear actuator and the gear begins to extend. Individual microswitches, actuated open by the release of the uplock hooks, energize the red GEAR UNLOCKED light while the gear is in transit. As each gear actuator reaches its fully-extended position, internal downlock latches mechanically engage to hold the gear extended. Individual microswitches are actuated closed by the downlock latches, illuminating their respective green NOSE, LH, and RH gear down indicators. When all three gear assemblies are fully down and locked and their respective microswitches are closed, the red GEAR UNLOCKED light is extinguished, the hydraulic system center valve opens, and the extend solenoid is deenergized. In this condition, the control valve selector spool returns to its neutral position and the hydraulic system returns to “open center” mode, extinguishing the amber [HYD PRESS ON] annunciator. Note: When the center valve is energized and hydraulic pressure is being supplied to the landing gear, the amber [HYD PRESS ON] annunciator will be illuminated. Static (Open Center) Condition When normal extension or retraction is complete, the hydraulic fluid contained in both the retract and extend circuits remains static while fluid circulation through the pumps, filters, check valves, and reservoir continues. Landing Gear Extension/Retraction Speeds and Cycle Times On airplanes 550-0626 and earlier, maximum landing gear extended speed is 250 KIAS. On airplanes 550-0627 and after, maximum landing gear extended speed is 262 KIAS. Maximum landing gear retraction speed is 200 KIAS. Normal cycle time for landing gear operation is approximately six seconds.

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© PCW

8-29



Landing Gear System Schematic (extension)


Auxiliary/Emergency Landing Gear Extension System An auxiliary/emergency landing gear extension system is installed for use in the event of hydraulic system failure. Auxiliary/emergency gear extension is provided by a mechanical uplock release “free-fall” system and a pneumatic “blowdown” system. The free-fall system allows mechanical gear release using a T-handle connected by cables to each uplock hook. Pulling the T-handle releases the uplock hooks mechanically. The pneumatic blowdown system is used to assure that the gear is fully extended and locked, and is actuated using a blowdown knob located behind the T-handle. Major components of the system include an auxiliary/emergency air storage “blowdown” bottle and discharge valve, the auxiliary/emergency gear release T-handle, the blowdown knob, and associated cables and plumbing. Auxiliary/Emergency Air Storage Bottle The auxiliary/emergency air storage “blowdown” bottle is located behind the aft divider within the right nose baggage compartment and is pressurized with dry nitrogen or clean, dry compressed air. The bottle supplies pneumatic pressure to operate both the auxiliary/ emergency landing gear blowdown system and the auxiliary/emergency braking system.

1

2 1. Pressure Gage 2. Air Storage Bottle

A bottle pressure gage is visible through an inspection window positioned on an access panel. The access panel is hinged to facilitate servicing and is accessible through the right nose baggage door. The gage is marked with a green arc, denoting the normal system indicating range, from 1800 to 2050 PSIG. The bottle should be serviced when indicated bottle pressure is less than 1800 PSIG. The bottle holds from 75 to 100 cubic inches of air.

8-30



Cessna Citation II Technical Manual The blowdown bottle is equipped with a thermal relief valve that functions to release excessive pressure overboard through a vent line. The relief valve is designed to rupture at approximately 3750 to 4250 PSIG. The thermal relief valve is not reusable; if it ruptures, the valve or the complete bottle assembly must be replaced. Warning: Thermal relief valve rupture will render the auxiliary/ emergency landing gear extension and braking systems inoperable. Note: The blowdown bottle requires hydrostatic testing every five years. Auxiliary/Emergency Gear Extension System Operation Should the landing gear fail to extend hydraulically, requiring operation of the auxiliary/ emergency extension system, the red AUX GEAR CONTROL 2 T-handle should be pulled fully out to release the landing gear uplocks. The T-handle 1 should then be rotated 45° clockwise to unlock. This 1. T-Handle action allows the landing gear 2. Blowdown Knob to free-fall, and also unlocks the red, collar-type pneumatic blowdown knob. To assure that the landing gear is fully down and locked, the blowdown knob should be pulled to actuate a valve that discharges the blowdown bottle. The blowdown knob is connected by cable to the blowdown bottle discharge valve. Pulling the blowdown knob opens the discharge valve, and routes pneumatic pressure directly to the nose and main gear actuator pneumatic extend ports, thereby forcing the gear into the down and locked position. Hydraulic fluid within the gear actuators is returned directly to the hydraulic fluid reservoir by means of a dump valve. Note: Before operating the AUX GEAR CONTROL T-handle, the flight crew should assure that the landing gear control handle is in the “DOWN” position.



8-31


Landing Gear System Schematic (emergency extension)

8-32

© PCW


Cessna Citation II Technical Manual Note: When the blowdown knob is pulled, the blowdown bottle discharge valve latches in the open position and cannot be reset by the flight crew. Therefore, the blowdown system cannot be reused until the discharge valve is reset and the storage bottle is refilled by service personnel. When the blowdown bottle discharge valve is reset to the closed position during servicing, it connects the landing gear actuators to a vent line, allowing gas/air trapped within the gear actuators to be released overboard when the gear is next operated hydraulically. Note: To ensure the highest probability of full extension, the auxiliary/emergency gear extension system should be operated at an airspeed of approximately 150 KIAS with flaps retracted. The landing gear may not fully extend if free-fall landing gear extension is attempted at airspeeds above 200 KIAS. As in normal extension, respective gear down indicators should illuminate when each gear assembly reaches its fully-extended position. Emergency Hydraulic Dump Valve Following auxiliary/emergency (pneumatic) gear extension, hydraulic fluid within the retract side of the landing gear actuators is returned directly to the hydraulic fluid reservoir by means of a dump valve that is connected to the hydraulic gear retract line, the hydraulic system return line, and the pneumatic gear extension line. When pneumatic pressure within the landing gear system exceeds 200 PSI, the dump valve opens, allowing hydraulic fluid to flow through the dump valve into the return line to the reservoir.



8-33

Brakes The left and right wheel brakes are hydraulically operated by independent master cylinders attached to the pilot’s and copilot’s rudder pedals. The brake system is pressurized when either pilot depresses the toe pedals. Interconnect assemblies allow either pilot to operate the brakes with equal authority. The brake system is power-assisted and provides touchdown protection capability. An anti-skid system provides skid and locked wheel crossover protection. A backup pneumatic system can be used for auxiliary/emergency braking in the event of hydraulic brake system failure; on airplanes 5500460 and earlier not incorporating SB550-32-12, backup manual braking is available as well. Parking brake capability is provided by locking the normal brakes. Power Brake System The power brake system is composed of a brake hydraulic fluid reservoir, an electrically-driven hydraulic pump and filter assembly, one or two accumulators, and an anti-skid servo valve. Hydraulic fluid for brake system operation is supplied from 2 a reservoir installed within the nose compartment on the right side of the forward pressure 1 bulkhead. On airplanes 5500281 and after, airplanes 5500039 and earlier, and airplanes incorporating SB55032-8, the brake reservoir is pressurized by cabin pressure, which enters the reservoir 1. Fluid Reservoir through an assembly that 2. Accumulator incorporates a check valve to Pressure Gage prevent hydraulic fluid or fumes from entering the cabin, and a filter to protect the brake fluid from contamination. The reservoir is vented by an overboard vent line connected to a relief valve that incorporates a restrictor orifice. 8-34



Brake System Schematic


© PCW

8-35



The restrictor orifice allows pressurization to accumulate within the reservoir, but is never fully closed, thereby allowing continuous venting of fumes. On airplanes 550-0040 through 0280 and airplanes not incorporating SB550-32-8, the reservoir is not pressurized, and is vented by an overboard line. Pressure for power brake system operation is provided by pressurized hydraulic fluid supplied by an electrically-driven hydraulic pump installed within the left side of the nose compartment. The pump is controlled by a pressure switch that opens when the pressure approaches 1300 PSI and closes when system pressure approaches 900 PSI. The pump supplies hydraulic fluid to the power brake system and to one or two accumulators under the left nose baggage compartment floor. Each accumulator is divided by a floating piston into an air compartment and a fluid compartment. The air compartment is pneumatically pressurized to a nominal pressure of 650 to 700 PSIG, and acts to pressurize the contents of the fluid compartment. The accumulator(s) functions to maintain hydraulic pressure without the need to continuously operate the hydraulic pump, thereby assuring that hydraulic power-assist is immediately available to the brake actuators. Fluid expelled from the accumulator during brake operation is returned to the accumulator by the pump. On airplanes 550-0453 and after, the main gear power brake system includes one 50 cubic inch accumulator. On airplanes 550-0437 through 0452, the main gear power brake system includes two 25 cubic inch accumulators. On airplanes 550-0436 and earlier, the main gear power brake system includes one 25 cubic inch accumulator.

8-36



Cessna Citation II Technical Manual Applying pressure to the brake pedals actuates a piston assembly in each master cylinder. Master brake cylinder piston displacement transfers hydraulic pressure to the anti-skid servo valve. Braking is power-assisted by hydraulic pressure supplied from the accumulator(s). Hydraulic fluid from the accumulator(s) is released by the anti-skid servo valve through the parking brake valve below the flight compartment floor to the main gear brake assemblies. The additional fluid pressure hydraulically amplifies the effect of brake pedal pressure. A color band gage is provided for checking brake hydraulic accumulator pressure. The pressure gage is visible through an inspection window positioned on an access panel. The access panel is hinged to facilitate servicing and is accessible through the right nose baggage door. The gage is marked with a red arc between 0 and 650 PSIG denoting underpressure, a green arc between 650 to 700 PSIG denoting normal operating pressure range, a yellow arc between 700 to 900 PSIG denoting caution range, a green crosshatched arc between 900 to 1350 PSIG denoting precharge range, and a red arc between 1350 to 1500 PSIG denoting overpressure. Note: The accumulator pressure gage denotes pressure ranges only; it is not marked with numerals denoting specific pressures. The power brake system receives 28 VDC left main bus power through a 20-amp SKID CONTROL circuit breaker on the left CB panel. Touchdown Protection On airplanes 550-0437 and after, the touchdown protection system prevents landing with pressure applied to the brakes. The touchdown protection system energizes the power brake anti-skid control valve open at airplane touchdown through the left main gear safety switch. The valve therefore remains closed until the airplane is on the ground, preventing hydraulic fluid from entering the brake actuators and thereby assuring that the brakes are not applied at airplane touchdown.



8-37

Anti-Skid System The anti-skid system is integral to the power brake system and provides maximum braking capability on wet or icy runways. The system consists of two wheel speed transducers, an electronic control box, an anti-skid servo valve, pressure switches, mode switch, circuit breakers, and indicator lights. The wheel speed transducers are installed in the main gear axles. Each transducer generates electrical signals proportional to wheel rotational speed; these signals are then transmitted to the electronic control box. The control box also receives signals from the left main gear safety switch and the brake hydraulic pump pressure switch. The control box averages signals from left and right transducers to generate a composite reference voltage and compares this voltage to left and right transducer voltage. The control box then generates signals that actuate the anti-skid servo valve within the appropriate disc brake assembly to reduce braking pressure as required to prevent wheel skidding. The system detects incipient skids by using a wheel speed transducer to measure the deceleration of each landing wheel, and then prevents skids by reducing the brake pressure in proportion to the deviation of each wheel from normal braking deceleration. The system also modulates brake pressure to maximize braking efficiency. The anti-skid system is activated by positioning the ANTISKID switch on the LDG GEAR control panel to “ON.” If the ANTI-SKID switch is set to “OFF,” power braking should be available without the antiskid function, and the amber [ANTISKID INOP] annunciator should illuminate. If the brake system hydraulic pressure decreases to 900 PSI, the ▲ ANTI-SKID SWITCH brake hydraulic pump pressure switch closes, causing the control box to illuminate the amber [POWER BRAKE LOW PRESS] annunciator (550-0550 and after) or [POWER BRK PRESS LO] annunciator (550-0505 and earlier).

8-38



Cessna Citation II Technical Manual The anti-skid brake system automatically disengages when ground speed falls below approximately 12 knots. The anti-skid system receives 28 VDC left main bus power through the 20-amp SKID CONTROL circuit breaker on the left CB panel. Warning: On airplanes 550-0461 and after and earlier airplanes incorporating SB550-32-12, power brake system failure renders hydraulic braking completely inoperable. If the power brake system fails, auxiliary/emergency pneumatic braking must be used. Note: When the anti-skid system is operating, the pilot should apply maximum braking pressure throughout the braking run. If the pilot attempts to modulate brake pressure while the system is releasing applied brake pressure to avoid a skid, the applied brake signal can be interrupted, resulting in a considerable increase in braking distance. Anti-Skid System Test Anti-skid system ground selftest may be activated by rotating the test selector switch on the lower left instrument panel to the “ANTISKID” position, then returning it to the “OFF” position. The anti-skid test circuit monitors anti-skid system electrical function. If the anti-skid system is functioning normally, the amber ▲ TEST SELECTOR SWITCH [ANTISKID INOP] annunciator will illuminate, then extinguish 3 to 4 seconds after the TEST switch is returned to “OFF.” If an anti-skid system fault is detected, the [ANTISKID INOP] annunciator will remain illuminated. While airborne, the anti-skid test circuit is automatically activated when the landing gear control handle is positioned to “DOWN” (if the ANTI-SKID switch is positioned to “ON”). If antiskid system test detects a fault, the [ANTISKID INOP] annunciator will remain illuminated. Note: If an anti-skid system fault is detected, the ANTI-SKID switch can be positioned to OFF. If the SKID CONTROL circuit breaker is engaged, normal power-assisted hydraulic braking is available without anti-skid protection. HYDRAULICS AND LANDING GEAR 12/99


8-39

Locked Wheel Crossover Protection On airplanes 550-0437 and after, locked wheel crossover protection prevents loss of airplane control caused by unequal wheel rotation rates. When the anti-skid system detects that one main gear wheel is rotating 50% slower than the other, brake pressure to the slow wheel is dumped, allowing wheel speeds to equalize. Locked wheel crossover protection is functional at ground speeds greater than 40 knots. Auxiliary/Emergency Braking System Auxiliary/emergency braking is accomplished pneumatically, by releasing high-pressure nitrogen or dry compressed air to the main gear braking assemblies. During auxiliary/emergency braking, a shuttle valve that is normally spring-loaded open to the hydraulic port is repositioned open to the pneumatic port by gas/air pressure, thereby allowing the brakes to be operated pneumatically instead of hydraulically. Auxiliary/emergency braking is initiated by pulling the EMER BRAKE-PULL valve handle located below the pilot’s instrument panel. This action releases gas/air to the brakes from the same air storage bottle that is used for auxiliary/emergency landing gear deployment. Pneumatic pressure within the auxiliary/ emergency brake valve re▲ EMERGENCY BRAKE VALVE HANDLE turns the auxiliary/emergency brake handle forward upon release, terminating pneumatic flow to the brakes and releasing the brakes by venting air within the brake assemblies overboard. Pulling the auxiliary/emergency brake handle progressively increases brake pressure. For most efficient use of pneumatic braking, the auxiliary/emergency brake handle should be pulled aft until sufficient deceleration is achieved, then should be held in position until the airplane has stopped. Note: Normal braking using the rudder pedals should not be attempted if auxiliary/emergency pneumatic braking is in use. Depressing the rudder pedals opens the shuttle valves, allowing high-pressure air from the brake housing to enter the hydraulic reservoir, possibly rupturing it.

8-40



Cessna Citation II Technical Manual Note: The capacity of a properly serviced air storage bottle is sufficient for at least ten individual brake applications if the landing gear has not been extended pneumatically. Bottle capacity is adequate to provide auxiliary/emergency braking for most conditions even if the landing gear has been extended pneumatically. After use of auxiliary/emergency braking, it is recommended that the engines be shut down and the airplane be towed to the ramp, as there is no gage, light, or other warning device in the cockpit to alert the flight crew when the pneumatic bottle is depleted. Auxiliary/Emergency Braking System Usage On airplanes 550-0460 and earlier incorporating SB550-32-12 and airplanes 550-0461 and after, the auxiliary/emergency braking system must be used if the power brake system fails. On airplanes 550-0460 and earlier not incorporating SB550-3212, the brakes can be applied without power assist and/or the auxiliary/emergency braking system can be used if the power braking system fails. Parking Brake System The parking brake handle, located below the lower left instrument panel, operates a parking brake control valve installed in-line downstream of the master cylinders and upstream of the brake assemblies. The parking brake is set by pressing the rudder toe pedals until sufficient hydraulic pressure has developed in the lines, then pulling the parking brake handle out to close the control valve. With the control valve closed, hydraulic pressure is retained in the lines, thereby holding the brakes in the applied position. Pushing the parking brake handle in opens the control valve, thereby releasing the brakes. Note: The parking brake should not be set if the flight crew suspects that the brakes may be unusually hot. Setting the brake increases cool down time by impeding airflow, and therefore may allow sufficient heat transfer to open the parking brake thermal relief valves and/or melt the thermal relief plugs in the wheel, causing tire deflation.



8-41

Limitations Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.

8-42




Chapter 9 Environmental System Table of Contents Overview ............................................... 9-1 Pressurization Air Source and Selection .. 9-2 Source Selection .............................. 9-2 Air Cycle Machine and Temperature Control System ..................................... 9-6 Air Cycle Machine ............................ 9-6 Temperature Control System ............. 9-8 Temperature Control System Indication and Warning ..................... 9-9 ACM Overpressure Warning and Protection ........................................ 9-9 ACM Overheat Warning and Protection ...................................... 9-10 Pressurization Air Source and Selection ............................................ 9-13 Nacelle Precooler System ............... 9-14 Source Selection ............................ 9-15 Air Cycle Machine and Temperature Control System ................................... 9-18 Air Cycle Machine .......................... 9-18 ACM Overpressure Protection .......... 9-22 ACM Overheat Warning and Protection ...................................... 9-22 Temperature Control System ................ 9-24 Conditioned Air Temperature Warning and Protection .................. 9-25


Environmental System, continued Table of Contents Air Distribution System ........................ 9-27 Distribution Control ........................ 9-27 Blowers ......................................... 9-29 Distribution Ducting and Air Outlets ..................................... 9-30 Cabin Pressurization Control System .... 9-31 Cabin Outflow Valves...................... 9-31 Cabin Pressurization Controller ........ 9-34 Pneumatic Relay ............................ 9-37 Cabin Altitude Limit Valves ............. 9-38 Pressurization System Indication ..... 9-39 Cabin Altitude and Differential Pressure Indicator .......................... 9-39 Cabin Rate-of-Change Indicator ........ 9-39 Cabin Altitude Warning Annunciator ................................... 9-39 Emergency Dump ........................... 9-40 Ambient Air Sources ............................ 9-41 Vapor Cycle Air Conditioning System .... 9-43 Refrigerant Circulation System ........ 9-43 Vapor Cycle Air Conditioning System Protection .......................... 9-45 Vapor Cycle Air Conditioning System Controls and Indicators .................. 9-46 Vapor Cycle Air Conditioning System Protection .......................... 9-46


Environmental System, continued Table of Contents Emergency Oxygen System .................. 9-47 Oxygen Outlets............................... 9-48 Crew Oxygen Masks ....................... 9-49 Passenger Oxygen Masks ............... 9-50 Oxygen System Controls, Malfunction Warning, and Indication ....................... 9-51 Limitations .......................................... 9-53 Emergency Procedures......................... 9-53

Cessna Citation II Technical Manual Overview The environmental system of the Cessna Citation II functions to control cabin pressure, temperature, and ventilation to ensure the comfort and safety of the flight crew and passengers. Engine bleed air is the primary source of air for cabin pressurization and ventilation. An air cycle machine conditions the bleed air for delivery to the cabin. A cabin pressurization control system regulates cabin pressure. During unpressurized operation, ambient air may be used for cabin ventilation. An optional flood cooling system may be installed to enhance ambient air ventilation. An optional vapor cycle air conditioning system may be installed to provide supplemental cabin cooling primarily during ground operations. In this chapter, the environmental system is divided into the following major groupings: pressurization air source and selection, the air cycle machine and temperature control system, the air distribution system (including cabin ventilation and windshield defogging), cabin pressurization control and indication, ambient air sources (including flood cooling and tailcone pressurization), the vapor cycle air conditioning system, and the emergency oxygen system. Depending on airplane unit number, two basic environmental system configurations exist, differing primarily in relation to pressurization air source and selection as well as the air cycle machine (ACM) and temperature control system. For organizational purposes and clarity, these primary differences are described separately according to unit number range.

ENVIRONMENTAL SYSTEM 12/99


9-1

Pressurization Air Source and Selection (550-0482, 0485 and after) Bleed air used for cabin pressurization and ventilation is extracted from the compressor section of each engine through ports located at 4 and 8 o’clock positions on the gas generator case. From these ports, bleed air is primarily routed through independent supply tubes to the ACM located within the aft fuselage. An environmental flow control/shutoff valve in each of these tubes controls the flow of bleed air from each source to the ACM and functions as a check valve to prevent the backflow of bleed air to the opposite source when either engine has failed or is operating at a sufficiently lower RPM. A ground shutoff valve bypasses the right environmental flow control/ shutoff valve permitting increased right engine bleed air flow through the ACM to enhance cabin ventilation during ground operation. During normal operation, all bleed air flows through the ACM en route to the cabin. In an emergency, left engine bleed air may be supplied directly to the cabin through an emergency supply tube and pressurization valve. The environmental flow control/shutoff valves are normallyopen, electrically-actuated closed, and have a nominal flow rate of approximately 6 pounds per minute (PPM) each. The ground shutoff valve is motor-operated, electrically-actuated open and closed, and has a nominal flow rate of approximately 18 PPM. The emergency pressurization valve is normallyclosed and electrically-actuated open. Source Selection All four valves are controlled primarily by the PRESS SOURCE selector switch on the environmental “tilt” panel. The valves are also controlled by various switches that sense bleed air pressure and temperature. The effects of rotating the selector switch to each of its six positions are described in the following paragraphs.

9-2



Cessna Citation II Technical Manual NORM Position During normal operation with both engines operating and the PRESS SOURCE selector switch set to “NORMAL,” both environmental flow control/shutoff valves will be open, permitting bleed air flow through the ACM and into the cabin at a rate of approximately 12 PPM. LH and RH Positions When set to “LH” or “RH,” bleed air is correspondingly supplied by the left engine or right engine only at a rate of approximately 6 PPM. In this condition, the environmental flow control/shutoff valve for the non-selected source is energized closed by 28 VDC left main bus power through the 5-amp NORM PRESS circuit breaker on the left CB panel. OFF Position When set to “OFF,” both environmental flow control/shutoff valves are energized closed through the NORM PRESS circuit breaker. GND Position To enhance cabin ventilation during ground operation, primarily when the right engine is operating only, the PRESS SOURCE selector switch should be set to “GND.” In this position, both environmental flow control/shutoff valves are energized closed, the ground shutoff valve is energized open, and the amber [BLD AIR GND] annunciator is illuminated through the NORM PRESS circuit breaker. With the ground shutoff valve open, right engine bleed air flows through the ACM and into the cabin at a rate of approximately 18 PPM. Selection of this source is inhibited by the left main gear safety switch when the airplane is in flight.



9-3

EMER Position Should bleed air flow through the ACM be insufficient to maintain selected cabin altitude, the PRESS SOURCE selector switch should be set to “EMER.” In this position, the amber [EMERG PRESS ON] annunciator is illuminated, both environmental flow control/shutoff valves are energized closed, and the emergency pressurization valve is energized open. 28 VDC right main bus power is supplied to the valves and the annunciator through the 5-amp EMER PRESS circuit breaker on the left CB panel. Selection of this source is inhibited by the left main gear safety switch when the airplane is on the ground. With the emergency pressurization valve open, unconditioned left engine bleed air is supplied directly to the cabin through the emergency supply tube. The emergency supply tube terminates within a mixing tube below the aft passenger cabin floor where it forms an ejector nozzle. The ejector nozzle produces a suction force that opens a check valve through which cabin air is drawn into the mixing tube. The cabin air mixes with and reduces the temperature of the bleed air prior to entering the distribution system. Additional bleed air temperature reduction is provided by the emergency supply tube itself, which features a “beaded” or “spiral” exterior that increases surface area to maximize heat transfer. A check valve installed in the aft pressure bulkhead prevents cabin pressure backflow through the emergency supply tube during normal pressurized operation. Note: When emergency pressurization is selected, the ACM and temperature control system are disabled. Limited control of cabin temperature may be accomplished using the left THROTTLE lever to regulate bleed air flow; however, excessive engine power reduction can cause an increase in cabin altitude. Note: Emergency pressurization is automatically activated when the temperature of bleed air flow through the ACM exceeds approximately 435°F. Refer to the ACM Overheat Warning and Protection section of this chapter for a complete description of ACM overheat protection.

9-4



© PCW

9-5



Pressurization Air Source Schematic (550-0482, 550-0485 and after)


Air Cycle Machine and Temperature Control System (550-0482, 550-0485 and after) Bleed air is cooled, mixed with uncooled bleed air, and dehumidified to provide conditioned air with the desired temperature to the air distribution subsystem. Major components of the system include an air cycle machine (ACM) containing a precooler, primary and secondary heat exchangers and a cooling turbine, a water separator, a bypass modulating valve, a water ejector nozzle, a fan, and necessary ducting. Air Cycle Machine From the environmental control/shutoff valves or the ground shutoff valve, bleed air is supplied to the ACM, directed to the bypass modulating valve, and passes through the precooler. From the precooler, bleed air passed through the primary heat exchanger and is cooled by heat transfer. After passing through the primary heat exchanger, the bleed air is supplied to the cooling turbine. The cooling turbine essentially consists of an impeller-type compressor and a turbine, mounted on the same shaft. The shaft rotates at approximately 80,000 RPM and its bearings are lubricated by oil drawn by wicks from a sump mounted on the turbine housing. A fan, external to the cooling turbine and used to circulate ambient air for cooling, is also mounted on the shaft.

© PCW

9-6



Cessna Citation II Technical Manual In the cooling turbine, bleed air first enters the compressor, where its pressure and temperature are increased. The air is then directed to the secondary heat exchanger, where it is cooled again by heat transfer. Water, drawn by the water ejector nozzle from the water separator, is sprayed over the secondary heat exchanger to provide additional cooling. A small amount of high velocity air from the secondary heat exchanger output line is used to create the suction required to draw the water from the water separator. From the secondary heat exchanger, the compressed air is directed to the turbine where its temperature and pressure are rapidly reduced by expansion. From the turbine outlet, this super-cooled air is passed through a mixing tube where it is mixed with hot bleed air supplied through the bypass modulating valve. The electrically controlled and operated bypass modulating valve is located in a bypass duct connected between the bleed air inlet and the mixing tube at the outlet side of the cooling turbine. The valve functions to control the temperature of the conditioned air by opening and closing to modulate the flow of hot bleed air to the mixing tube. From the mixing tube, the conditioned air passes through the water separator, which collects moisture from the passing air and forms large droplets that are removed by centrifugal force. This removed moisture is drawn away by the water ejector nozzle previously described. An integral spring-loaded relief valve allows air to bypass the unit should the water separator become obstructed by ice or foreign material. Cooling air for the precooler, primary, and secondary heat exchangers is drawn from within the aft fuselage by the fan that is driven by the cooling turbine shaft. After passing over the heat exchangers, the cooling air is exhausted through an overboard vent on the lower surface of the aft fuselage.



9-7

Temperature Control System Temperature control is primarily achieved by varying the amount of hot bleed air that the bypass modulating valve mixes with cold air from the ACM. Cabin temperature may be set manually or controlled automatically as desired. Automatic 1 2 Temperature Selection Automatic temperature selection is accomplished using the AUTOMATIC temperature control knob on the environmental panel. Rotat- 1. Automatic Temperature Control Knob 2. Manual Mode Toggle Switch ing the knob clockwise for a higher cabin temperature or counterclockwise for a lower cabin temperature operates a potentiometer that establishes a reference voltage corresponding to the selected temperature. The reference voltage is supplied to the temperature control computer, where it is compared with signals from the duct temperature sensor (ACM conditioned air temperature) and the cabin temperature sensor (actual cabin temperature). The temperature control computer then generates a signal to drive the bypass modulating valve open or closed to maintain the desired cabin temperature. The bypass modulating valve receives power from the left main bus through the 5-amp TEMP circuit breaker. Manual Temperature Selection The manual temperature selection mode is selected by rotating the AUTOMATIC temperature control knob fully counterclockwise till it clicks into the “MANUAL” position. In this mode, the three-position manual mode toggle switch controls the position of the bypass modulating valve. The switch is spring-loaded to the “OFF” position. When held in the “MANUAL HOT” position, the bypass valve moves toward open, allowing more hot air to mix with the cooled air. When released, it returns to the “OFF” position, but the bypass valve remains in the selected position. When held in the “MANUAL COLD” position, the bypass valve moves towards closed. The manual mode toggle switch is only usable when the AUTOMATIC temperature control knob is in the “MANUAL” position.

9-8



Cessna Citation II Technical Manual Temperature Control System Indication and Warning When conditioned air temperature at the water separator outlet (550-0550 and after) or inlet (550-0482 and 0485 ~ 0505) exceeds approximately 315°F, a temperature switch functions to illuminate the amber [AIR DUCT O’HEAT] annunciator. Illumination of this annunciator indicates that corrective action should be taken to lower the cabin temperature in order to avoid duct damage. Following [AIR DUCT O’HEAT] annunciator illumination, the TEMP circuit breaker should be pulled and reset, and manual temperature control mode should be selected. The manual mode toggle switch should be held in the “MANUAL COLD” position until the [AIR DUCT O’HEAT] annunciator extinguishes. Automatic temperature control should then be reselected. If the [AIR DUCT O’HEAT] annunciator reilluminates, cabin temperature should be controlled manually for the remainder of the flight. ACM Overpressure Warning and Protection To protect the ACM from overpressurization, a primary and secondary pressure switch are installed in the environmental supply tubing upstream and downstream of the ground shutoff valve respectively. During ground operations with the PRESS SOURCE selector switch in the “GND” position, when bleed air pressure reaches approximately 38 PSI, the primary pressure switch functions to close the ground shutoff valve and extinguish the [BLD AIR GND] annunciator. In this condition, retarding the right THROTTLE lever below approximately 72% N2 should reduce bleed air pressure sufficiently to cause the valve to reopen and the [BLD AIR GND] annunciator to illuminate. Should the primary pressure switch fail, the secondary pressure switch will activate when bleed air pressure reaches approximately 42 PSI, functioning to close the ground shutoff valve, extinguish the [BLD AIR GND] annunciator, and illuminate the amber [ACM O’PRESS] annunciator. In this condition, the valve will remain closed and the [ACM O’PRESS] annunciator will remain illuminated regardless of N2/bleed air pressure reduction or PRESS SOURCE selector switch position until the pressure switch circuit is restored. This may be accomplished by pulling and resetting the NORM PRESS circuit breaker; however, the cause of the malfunction should be identified and repaired before resuming flight operations. ENVIRONMENTAL SYSTEM 12/99


9-9

Note: Illumination of both the [ACM O’PRESS] and [BLD AIR GND] annunciators may indicate that the primary pressure switch has closed, but the ground shutoff valve is stuck open. Should this occur, the cause of the malfunction should be identified and repaired before resuming flight operations. Should the ground shutoff valve malfunction and open during flight, the [BLD AIR GND] and [ACM O’PRESS] annunciators will both be illuminated. In this condition, the PRESS SOURCE selector switch should be set to “RH” to close the left flow control shutoff valve, and the right THROTTLE lever should be retarded below 80% N2 to reduce bleed air pressure. Failure of the [ACM O’PRESS] annunciator to extinguish may indicate that it was illuminated by the secondary pressure switch because the primary pressure switch failed to close. In this condition, the [ACM O’PRESS] annunciator will remain illuminated until the pressure switch circuit is restored. During flight this may be accomplished by first pulling the EMER PRESS circuit breaker to prevent inadvertent activation of emergency pressurization, pulling and resetting the NORM PRESS circuit breaker, and then resetting the EMER PRESS circuit breaker. Note: If the [ACM O’PRESS] annunciator remains illuminated, the PRESS SOURCE selector switch should remain in the “RH” position, right engine N2 should remain below 80%, and the left engine should be operated normally for the duration of the flight. After landing, the cause of the malfunction should be identified and repaired before resuming flight operations. ACM Overheat Warning and Protection To protect the ACM from overheating, an overheat sensor is installed in the bleed air tube between the ACM compressor outlet and secondary heat exchanger inlet. During flight with the PRESS SOURCE selector switch in the “NORMAL” position, when bleed air temperature exceeds approximately 435°F, the overheat sensor functions to close both environmental flow control/shutoff valves, open the emergency pressurization valve, and illuminate the amber [EMERG PRESS ON] annunciator. Power is supplied to the valves and the annunciator through the EMER PRESS circuit breaker. In this condition, bleed air flow through the ACM is interrupted and unconditioned left engine bleed air is supplied directly to the cabin through the emergency pressurization valve. 9-10



Cessna Citation II Technical Manual If bleed air temperature falls below approximately 405°F within 12-seconds of [EMERG PRESS ON] annunciator illumination, normal system operation will be automatically restored. If normal system operation is not automatically restored within this period of time, an emergency pressurization lockout relay will be energized through the EMER PRESS circuit breaker. With this relay energized, both environmental flow control/shutoff valves will remain closed, the emergency pressurization valve will remain open, and the [EMERG PRESS ON] annunciator will remain illuminated until the overheat circuit is reset. This may be accomplished by rotating the PRESS SOURCE selector switch to the “EMER” position to deenergize the emergency lockout relay, waiting one minute, and then reselecting the “NORMAL” position to restore normal operation. Note: If the [EMERG PRESS ON] annunciator remains illuminated, the PRESS SOURCE selector switch should be set to the “EMER” position, the right engine should be operated normally, and the left THROTTLE lever should be used to control cabin temperature for the duration of the flight. After landing, the cause of the malfunction should be identified and repaired before resuming flight operations. During ground operations, the overheat sensor functions as it does in flight; however, the left main gear safety switch prevents the emergency pressurization valve from opening when the airplane is on the ground. During ground operations with the PRESS SOURCE selector switch in the “GND” position, the overheat sensor functions to close the ground shutoff valve and extinguish the [BLD AIR GND] annunciator. As in flight, system operation will be automatically restored if bleed air temperature falls below approximately 405°F within 12-seconds of [EMERG PRESS ON] annunciator illumination. If system operation is not automatically restored within this period of time, the overheat circuit must be reset by rotating the PRESS SOURCE selector switch to the “EMER” position, waiting one minute, and then reselecting the previous position to restore operation. Note: If the [EMERG PRESS ON] annunciator remains illuminated, the cause of the malfunction should be identified and repaired before resuming flight operations.



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Note: Automatic activation of emergency pressurization may occur under such conditions as low airspeed climbs at high altitudes with a low cabin temperature selected. Should this occur, increasing airspeed and selecting a higher cabin temperature after restoring normal operation should prevent reactivation of emergency pressurization. The conditioned and dehumidified air is routed to the distribution system described later in this chapter.

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Cessna Citation II Technical Manual Pressurization Air Source and Selection (550-0484, 0483, 0481 and earlier) Bleed air used for cabin pressurization and ventilation is extracted from the compressor section of each engine through ports located at 4 and 8 o’clock positions on the gas generator case. From these ports, bleed air is primarily routed through independent supply tubes to a precooler installed within each engine nacelle. The precoolers reduce the temperature of engine bleed air supplied to various airplane systems, including the ACM located within the aft fuselage. An environmental flow control/shutoff valve in each of these tubes controls the flow of bleed air from each source to the ACM and functions as a check valve to prevent the backflow of bleed air to the opposite source when either engine has failed or is operating at a sufficiently lower RPM. A ground shutoff/pressure regulating valve bypasses the right environmental flow control/shutoff valve, permitting increased right engine bleed air flow through the ACM to enhance cabin ventilation during ground operation. During normal operation, all bleed air flows through the ACM en route to the cabin. In an emergency, left engine bleed air may be supplied directly to the cabin through an emergency supply tube and pressurization valve. The environmental flow control/shutoff valves each have a nominal flow rate of approximately 6 pounds per minute (PPM) and a maximum flow rate of approximately 9 PPM. Flow rate is controlled by a primary solenoid and secondary solenoid integral to each valve. Both solenoids are normally deenergized; nominal flow rate occurs in this condition. Maximum flow rate occurs when the primary solenoid is deenergized and the secondary solenoid is energized. The ground shutoff/pressure regulating valve is motor-operated, electrically-actuated open and closed, has a nominal flow rate of approximately 18 PPM, and incorporates an independent pressure relief valve. The emergency pressurization valve is normally-closed and electrically-actuated open.



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Nacelle Precooler System The nacelle precooler system functions to regulate the temperature of the bleed air used by airplane systems to approximately 500°F. This is accomplished by routing bleed air from the engine through a heat exchanger, where it is cooled by bypass air. The primary components of the nacelle precooler system are heat exchangers, temperature sensor valves, temperature control valves, and tubular plumbing. The temperature sensor valves monitor engine bleed air temperature and send pneumatic signals to open and close the bleed air temperature control valves. When a temperature sensor valve senses that bleed air temperature is more than approximately 500°F, it sends a signal to open the associated temperature control valve and allow more bypass air to pass over the heat exchanger, lowering the temperature of bleed air supplied to distribution tubes within the aft fuselage. Conversely, when bleed air temperature is less than 500°F, the signal from the temperature sensor valve allows the temperature control valve to close, raising the temperature of the bleed air supplied to the environmental and other airplane systems. Overheat switches are located in the bleed air tubes between the nacelle precoolers and the environmental flow control/ shutoff valves. The switches will operate at a temperature of approximately 540°F and cause the amber [L PRECOOL FAIL] or [R PRECOOL FAIL] annunciator to illuminate, indicating excessively hot bleed air from the respective precooler.

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Cessna Citation II Technical Manual Source Selection All four valves are controlled primarily by the PRESS SOURCE selector switch on the environmental “tilt” panel. The valves are also controlled by various switches that sense bleed air pressure and temperature. The effects of rotating the selector switch to each of its seven positions are described below: NORMAL Position During normal operation with both engines operating and the PRESS SOURCE selector switch set to “NORMAL,” both environmental flow control/shutoff valves will be open, permitting bleed air flow through the ACM and into the cabin at a rate of approximately 12 PPM. LH and RH Positions When set to “LH” or “RH,” bleed air is correspondingly supplied by the left engine or right engine only at a rate of approximately 6 PPM. In this condition, the environmental flow control/shutoff valve for the non-selected source is energized closed by 28 VDC left main bus power through the 5-amp NORM PRESS circuit breaker on the left CB panel BOTH HI Position When set to “BOTH HI”, both environmental flow control/ shutoff valves remain open and both secondary solenoids are energized open. In this condition, approximately 18 PPM of bleed air from both engines flows through the ACM into the cabin and the amber [BLEED AIR GND/HI] annunciator is illuminated through the NORM PRESS circuit breaker. OFF Position When set to “OFF,” both environmental flow control/shutoff valves are energized closed through the NORM PRESS circuit breaker.



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GND Position To enhance cabin ventilation during ground operation, primarily when the right engine is operating only, the PRESS SOURCE selector switch should be set to “GND.” In this position, both environmental flow control/shutoff valves are energized closed, the ground shutoff/pressure regulating valve is energized open, and the amber [BLEED AIR GND/ HI] annunciator is illuminated through the NORM PRESS circuit breaker. With the ground shutoff/pressure regulating valve open, right engine bleed air flows through the ACM and into the cabin at a rate of approximately 18 PPM. Selection of this source is inhibited by the left main gear safety switch when the airplane is in flight. EMER Position Should bleed air flow through the ACM be insufficient to maintain selected cabin altitude, the PRESS SOURCE selector switch should be set to “EMER.” In this position, the amber [EMER PRESS ON] annunciator is illuminated, both environmental flow control/shutoff valves are energized closed, and the emergency pressurization valve is energized open. 28 VDC right main bus power is supplied to the valves and the annunciator through the 5-amp EMER PRESS circuit breaker on the left CB panel. Selection of this source is inhibited by the left main gear safety switch when the airplane is on the ground. With the emergency pressurization valve open, unconditioned left engine bleed air is supplied directly to the cabin through the emergency supply tube. The emergency supply tube terminates within a mixing tube below the aft passenger cabin floor where it forms an ejector nozzle. The ejector nozzle produces a suction force that opens a check valve through which cabin air is drawn into the mixing tube. The cabin air mixes with and reduces the temperature of the bleed air prior to entering the distribution system. Additional bleed air temperature reduction is provided by the emergency supply tube itself, which features a “beaded” or “spiral” exterior that increases surface area to maximize heat transfer. A check valve installed in the aft pressure bulkhead prevents cabin pressure backflow through the emergency supply tube during normal pressurized operation.

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Cessna Citation II Technical Manual Note: When emergency pressurization is selected, the ACM and temperature control system are disabled. Limited control of cabin temperature may be accomplished using the left THROTTLE lever to regulate bleed air flow; however, excessive engine power reduction can cause an increase in cabin altitude. Note: Emergency pressurization is automatically activated when the temperature of bleed air flow through the ACM exceeds approximately 435°F. Refer to the ACM Overheat Warning and Protection section of this chapter for a complete description of ACM overheat protection.



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Air Cycle Machine and Temperature Control System (550-0484, 0483, 0481 and earlier) Bleed air is cooled, mixed with uncooled bleed air, and dehumidified to provide conditioned air with the desired temperature to the air distribution subsystem. Major components of the system include an air cycle machine (ACM) containing primary and secondary heat exchangers and a cooling turbine, a water separator, a bypass modulating valve, a water ejector nozzle, a bleed air ejector solenoid valve, a bleed air ejector, and necessary ducting. Air Cycle Machine From the environmental control/shutoff valves or the ground shutoff/pressure regulating valve, bleed air is supplied to the ACM and directed to the bypass modulating valve and bleed air ejector nozzle supply tube. Within the ACM, the bleed air passes through the primary heat exchanger and is cooled by heat transfer. After passing through the primary heat exchanger, the bleed air is supplied to the cooling turbine. The cooling turbine essentially consists of an impeller-type compressor and a turbine, mounted on the same shaft. The shaft rotates at approximately 80,000 RPM and uses airfoil-type bearings that require no lubrication.

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Cessna Citation II Technical Manual In the cooling turbine, bleed air first enters the compressor, where its pressure and temperature are increased. The air is then directed to the secondary heat exchanger, where it is cooled again by heat transfer. Water, drawn by the water ejector nozzle from the water separator, is sprayed over the secondary heat exchanger to provide additional cooling. A small amount of high velocity air from the secondary heat exchanger outlet line is used to create the suction required to draw the water from the water separator. From the secondary heat exchanger, the compressed air is directed to the turbine, where its temperature and pressure are rapidly reduced by expansion. From the turbine, this supercooled air is passed through a mixing tube where it is mixed with hot bleed air supplied through the bypass modulating valve. The pneumatically controlled and operated bypass modulating valve is located in a bypass duct connected between the primary heat exchanger inlet and the mixing tube at the outlet side of the cooling turbine. The valve functions to control the temperature of the conditioned air by opening and closing to modulate the flow of hot bleed air to the mixing tube. From the mixing tube, the conditioned air passes through a water separator, which collects moisture from the passing air and forms large droplets that are removed by centrifugal force. This removed moisture is drawn away by the water ejector nozzle previously described. An integral spring-loaded relief valve allows air to bypass the unit should the water separator become obstructed by ice or foreign material.



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Pressurization Air Source Schematic (550-0484, 0483, 0481 and earlier)

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Cessna Citation II Technical Manual A duct routes cooling air to the ACM primary and secondary heat exchangers from a NACA-type scoop on the dorsal fairing (dorsal scoop). After passing over the heat exchangers, the air is exhausted through an overboard vent on the lower surface of the aft fuselage. In flight, sufficient air is available due to ram effect. During ground ▲ DORSAL SCOOP operations, ram effect is not available to move air over the heat exchangers. To compensate, a bleed air ejector nozzle is installed within the heat exchanger exhaust duct. Bleed air is admitted to the bleed air ejector nozzle by the solenoid valve installed in the ejector nozzle supply tube. The ejector nozzle directs bleed air toward the ambient air exhaust outlet. Bleed air flow exiting through the exhaust outlet creates a suction that draws ambient air through the heat exchangers. Bleed air and conditioned air are then exhausted overboard through an exhaust outlet below the right engine pylon. The shutoff valve is enabled by the left main gear safety switch when the airplane is on the ground, and is disabled by the safety switch in flight. To maximize engine power during takeoff, throttle position switches, brake switches, and a differential pressure switch act together to close the ejector nozzle shutoff valve and thereby disable the ACM bleed air ejector nozzle during takeoff roll. The amber [ACM EJECTOR ON] annunciator illuminates when the ejector nozzle shutoff valve is open.



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ACM Overpressure Protection The ACM is protected from overpressurization by a relief valve incorporated into the ground shutoff/pressure regulating valve. The relief valve, which functions independently of the ground shutoff/pressure regulating valve, opens at approximately 50 PSIG, releasing excess pressure overboard through a pneumatic tube. The valve reseats at approximately 40 PSIG. ACM Overheat Warning and Protection To protect the ACM from overheating, an overheat sensor is installed in the bleed air duct between the ACM compressor outlet and secondary heat exchanger inlet. During flight with the PRESS SOURCE selector switch in the “NORMAL” position, when bleed air temperature exceeds approximately 435°F, the overheat sensor functions to close both environmental flow control/shutoff valves, open the emergency pressurization valve, and illuminate the amber [EMER PRESS ON] annunciator. Power is supplied to the valves and the annunciator through the EMER PRESS circuit breaker. In this condition, bleed air flow through the ACM is interrupted and unconditioned left engine bleed air is supplied directly to the cabin through the emergency pressurization valve. If bleed air temperature falls below approximately 405°F within 12-seconds of [EMER PRESS ON] annunciator illumination, normal system operation will be automatically restored. If normal system operation is not automatically restored within this period of time, an emergency pressurization lockout relay will be energized through the EMER PRESS circuit breaker. With this relay energized, both environmental flow control/shutoff valves will remain closed, the emergency pressurization valve will remain open, and the [EMER PRESS ON] annunciator will remain illuminated until the overheat circuit is reset. This may be accomplished by rotating the PRESS SOURCE selector switch to the “EMER” position to deenergize the emergency lockout relay, and then reselecting the “NORMAL” position to restore normal operation. Note: If the [EMER PRESS ON] annunciator remains illuminated, the PRESS SOURCE selector switch should be set to the “EMER” position, the right engine should be operated normally, and the left THROTTLE lever should be used to control cabin temperature for the duration of the flight. After landing, the cause of the malfunction should be identified and repaired before resuming flight operations. 9-22



Cessna Citation II Technical Manual During ground operations, the overheat sensor functions as it does in flight; however, the left main gear safety switch prevents the emergency pressurization valve from opening when the airplane is on the ground. During ground operations with the PRESS SOURCE selector switch in the “GND” position, the overheat sensor functions to close the ground shutoff/pressure regulating valve. As in flight, system operation will be automatically restored if bleed air temperature falls below approximately 405°F within 12-seconds of [EMER PRESS ON] annunciator illumination. If system operation is not automatically restored within this period of time, the overheat circuit must be reset by rotating the PRESS SOURCE selector switch to the "OFF" or “EMER” position and then reselecting the previous position to restore operation. Note: If the [EMER PRESS ON] annunciator remains illuminated, the cause of the malfunction should be identified and repaired before resuming flight operations. Note: Automatic activation of emergency pressurization may occur under such conditions as low airspeed climbs at high altitudes with a low cabin temperature selected. Should this occur, increasing airspeed and selecting a higher cabin temperature after restoring normal operation should prevent reactivation of emergency pressurization.



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Temperature Control System Temperature control is primarily achieved 1 2 by varying the amount of hot bleed air that the bypass modulating valve mixes with cold air 3 from the air cycle machine. Cabin temperature may be 1. Manual Temperature Control Knob 2. Automatic Temperature Control Knob set manually or 3. Selector Switch controlled automatically. In either mode, pneumatic pressure is used to open the normally closed bypass modulating valve and allow hot bleed air to mix with cold air from the ACM. Major components of the temperature control system are a selector switch, the bypass modulating valve, a solenoid shutoff valve, an automatic temperature control pressure regulator, automatic temperature control knob and manual temperature control knob, cabin temperature, supply temperature, and low limit sensors, and an air duct temperature switch. Automatic Mode Automatic temperature control mode selection is accomplished by positioning the selector switch on the environmental panel to “AUTOMATIC.” When automatic mode is selected, the solenoid shutoff valve energizes, supplying 15 PSI from the automatic control pressure regulator, located on the right side of the aft fuselage, to the automatic cabin temperature selector, the low limit sensor, and the cabin temperature sensor. These sensors, the supply duct temperature sensor, and the temperature selector interact to develop a control pressure that causes the bypass modulating valve to open when sensed cabin temperature is too low or too close when sensed cabin temperature is too high. When duct temperature is less than approximately 35°F, the low limit sensor causes the bypass valve to open, raising duct temperature regardless of cabin temperature or selected temperature.

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Cessna Citation II Technical Manual Note: The automatic temperature control operates on bleed air from the right engine only; the manual system operates on bleed air from both engines. Therefore, the automatic temperature control system cannot be used unless the right engine is operating. Manual Mode When the cabin temperature selector switch is positioned to “MANUAL,” the temperature solenoid shutoff valve is deenergized, routing 23 PSI manual control pneumatic pressure to the bypass modulating valve and manual temperature selector. The manual temperature control knob operates an internal spring-loaded poppet, which determines the amount of control pressure that is allowed to bypass the poppet and vent overboard through the forward pressure bulkhead. This action varies the control pressure applied to the bypass modulating valve, thereby increasing or decreasing cabin temperature as with the automatic system. Therefore, in manual control mode, compensations for changes in cabin temperature must be achieved by rotating the manual temperature control knob. The manual mode is available as a backup should the automatic system fail. Note: When in manual mode, the low-limit sensor is inoperative and ice formation in the water separator/ducting is possible. Ice formation is more likely if the outside relative humidity exceeds 40%. When operating in manual mode, the ACM system should be carefully monitored to detect overheating. Note: The design of the temperature control system is such that insufficient bleed air pressure will cause the temperature control system to operate fully cold. Conditioned Air Temperature Warning and Protection Warning of excessive temperature of conditioned air is provided by an amber [AIR DUCT O’HEAT] annunciator that illuminates to alert the flight crew of a conditioned air overheat. An air duct overheat switch, installed in the air supply duct downstream from the water separator outlet, closes at approximately 315°F and causes the annunciator to illuminate.



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Temperature Control System Schematic (550-0484, 0483, 0481 and earlier)

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Cessna Citation II Technical Manual Air Distribution System The air distribution system directs and controls the flow of pressurization and ambient (fresh) air to the pressurized part of the fuselage. Ambient air is used to augment cabin airflow during ground operations. Major components of the air distribution system include distribution control devices, blowers, distribution ducting, and air outlets. Distribution Control Pressurization and ambient air are introduced into the cabin through flapper-type check valves that function to prevent loss of cabin pressurization through the fresh air duct or during operation of the emergency pressurization system. After passing through the check valves, air enters a ventilation junction box, located below the aft passenger cabin floor at the aft pressure bulkhead. The ventilation junction box functions to control the source of air directed to the overhead duct according to the temperature of the conditioned pressurization air entering the cabin. The ventilation junction box contains a swing-type door, thermoswitch, temperature motor, two limit switches, and the overhead blower. Operation of the overhead blower is described later in this chapter. When pressurization air temperature exceeds approximately 100°F, the thermoswitch closes an electrical circuit, causing the temperature motor to close the swing-type door. With the door closed, pressurization air is prevented from entering the overhead duct and recirculated cabin air is admitted. When pressurization air temperature falls below 100°F, the thermoswitch deenergizes and causes the temperature motor to run in the opposite direction, opening the door, and thereby readmitting pressurization air into the overhead ducts. Door travel is controlled by limit switches that turn off the temperature motor when the door is fully open or fully closed. The motor receives 28 VDC power from the left (550-482, 550-485 and after) or right (550-484, 550-483, 550-0481 and earlier) main bus through the 5-amp TEMP circuit breaker.



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A flexible air duct, connected to the ventilation junction box blower inlet, draws air from the aft baggage compartment to be circulated through the overhead duct network. This arrangement allows the aft divider door to be kept closed, yet still provides a means for the crew and passengers to smell smoke originating in the aft baggage compartment earlier than would otherwise be the case. From the ventilation junction box, conditioned air is passed through the emergency pressurization mixing tube. During emergency pressurization, hot engine bleed air used to pressurize the cabin is released into the mixing tube ejector nozzle. The ejector nozzle produces a suction force that opens a check valve through which cabin air is drawn into the mixing tube. The cabin air mixes with and reduces the temperature of the bleed air prior to entering the distribution system. From the mixing tube, cabin air enters a flow divider. The flow divider, located below the cabin floor, contains a divider vane and motor, and functions to apportion air between the flight compartment and the passenger cabin. Proportions of conditioned air delivered to the cockpit and cabin ▲ CKPIT/CABIN CONTROL are controlled using CKPT/CABIN control, located on the environmental panel. Rotating this switch energizes the motor to move the flow divider vane. Clockwise rotation increases cabin airflow; counterclockwise rotation increases cockpit airflow.

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Cessna Citation II Technical Manual Blowers Two impeller-type blowers are used to enhance cabin air circulation. Both blowers receive 28 VDC power from the left main bus through the 20-amp CABIN FAN circuit breaker.

1

2

The overhead blower is an integral part of the ventilation junction box and functions to force pressurization air, recir1. Overhead Blower Switch culated air, or ambient air 2. Defog Blower Switch through the overhead duct. The overhead blower is controlled by the three-position (HI/ OFF/LOW) FAN OVHD switch on the copilot’s lower instrument panel. The defog blower is located in the underfloor ducting downstream from the flow divider and functions primarily to increase the flow of conditioned air to the forward part of the cabin. The defog blower is controlled by the three position (HI/OFF/LOW) FAN DEFOG switch on the copilot’s lower instrument panel.



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Distribution Ducting and Air Outlets Within the cabin interior, distribution ducting is located in the overhead and at floor and armrest level. Additional distribution ducting is located under the cabin floorboards. Conditioned pressurization air below 100°F, recirculated air, or ambient air is taken from the distribution junction box and circulated through the overhead duct to adjustable outlets in the passenger cabin and flight compartment. Passenger air outlets are located over each seat position and are fully adjustable from no flow to maximum flow. Air from ▲ PASSENGER AIR OUTLET the aft baggage compartment is also supplied to the overhead duct to maximize the crew’s ability to detect smoke in the baggage compartment. Because of this method of smoke detection, air outlets in the flight compartment are configured so they can not be fully shut off. Airplanes with optional vapor cycle air conditioning systems use the overhead duct to circulate air conditioned air; these airplanes are equipped with overhead outlets that have higher flow rates. Air from the upper branch of the flow divider is routed to a main and auxiliary plenum. The main plenum supplies air to the footwarmer and armrest manifolds on the left side of the passenger cabin. The auxiliary plenum supplies the corresponding right side components. Passenger footwarmer manifolds are located along the outboard cabin walls at floor level. The footwarmer manifolds are assembled in segments, each segment including several outlet holes. The armrest manifolds are located along the outboard cabin walls at passenger seat armrest level. Air outlet holes are located beneath the armrest cover assemblies. The armrest and footwarmer manifolds are connected together by ducting at several points.

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Cessna Citation II Technical Manual Air from the lower branch of the flow divider is drawn through the defog fan and routed through underfloor ducting to the forward part of the pressurized cabin. In the flight compartment, conditioned air is routed to cockpit footwarmer manifolds, located on both sides of the cockpit, to the windshield and side window defog vents, and to optional supplementary ducts. In all airplanes, the underfloor ducting supplies air to the footwarmer and armrest warmer associated with the passenger seat immediately forward of the main entrance door. On airplanes 550-0482 and 550-0485 and after, the underfloor ducting also supplies air to warm the cabin door seal. On these airplanes and in airplanes 550-0010 through 550-0049 not incorporating SB550-21-1, the underfloor ducting is also connected to the left footwarmer/armrest manifold by a duct located aft of the main entrance door and to the right footwarmer/ armrest manifold by another duct located immediately aft of the flight compartment divider. Cabin Pressurization Control System The pressurization control system provides for passenger comfort by allowing the selection of a desired cabin altitude and rate-of-change during ascent or descent. The cabin is pressurized using engine bleed air as previously described. Cabin pressurization is regulated using outflow valves that open to allow pressurized air to exit the cabin, raising cabin altitude; and close to retain pressurized air in the cabin, lowering cabin altitude. Major components of the system include two cabin outflow valves, a pneumatic relay, two cabin altitude limit valves, a depressurization (dump) toggle valve, a pressurization source selection system, cabin altitude controls and indicators, and associated circuitry and plumbing. Cabin Outflow Valves Two cabin outflow valves, both mounted on the aft pressure bulkhead below the passenger cabin floor, vent pressurization air overboard to maintain the selected cabin altitude or pressure differential in reference to the ambient air pressure. The maximum pressure differential of 8.8 PSI is primarily determined by the structural limitations of the airplane’s pressurized center section. The system is designed to maintain a cabin altitude of 8,000 feet at airplane altitudes of up to 43,000 feet.



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Both cabin outflow valves function primarily to regulate the outflow of pressurization air from the cabin based on pneumatic signals received from the controller through the pneumatic relay. These signals establish a reference pressure that is applied to one side of an internal poppet valve. The other side of the poppet valve is exposed to actual cabin pressure. The poppet valve is spring-loaded closed and suspended by a flexible diaphragm between these two pressures such that the outflow of pressurization air increases when cabin pressure exceeds reference pressure, and decreases when reference pressure exceeds cabin pressure. In this way, the outflow valve modulates to maintain the selected cabin altitude or to effect a desired change in cabin altitude at a selected rate. Positive pressure relief and negative pressure relief functions are also provided by the outflow valves. Both functions override the controller and the pneumatic relay. Positive pressure relief is provided by a Schrader-type valve suspended by a flexible diaphragm between the reference pressure control chamber and an ambient pressure chamber that is vented to the atmosphere. The valve is spring-loaded closed and factory preset to open when the differential between reference pressure and ambient pressure exceeds the nominal differential of approximately 8.6 PSID. Should this occur, the release of control pressure to the atmosphere through the open valve would allow the poppet valve to modulate toward open, increasing pressurization air outflow. In this condition, cabin pressure is maintained at the nominal differential and cabin rate-of-change follows that of the airplane. Protection against exceeding the airplane’s negative pressure structural limits is provided by a flexible diaphragm exposed to cabin pressure on one side and ambient pressure on the other. Should ambient pressure exceed cabin pressure, as in during rapid descent, this diaphragm would raise and lift the poppet valve open allowing ambient pressure to enter the cabin until both pressures become approximately equal.

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Cabin Pressurization Control System Schematic


Cabin Pressurization Controller The cabin pressurization controller, located on the environmental panel, incorporates two knobs: one for selecting cabin altitude and the other for selecting cabin rate-of-change. The cabin altitude selector features an outer CABIN scale and an inner ACFT scale, both calibrated in feet X 1000 and visible through a window at the top center of the selector face. Both scales rotate simultaneously as the selector is rotated. The selected altitude is indicated by the alignment of each scale with the twelve o’clock position relative to the selector. The CABIN scale indicates the cabin altitude the controller is set to maintain. The ACFT scale indicates the maximum altitude to which the airplane may ascend without causing the selected cabin altitude to be exceeded. The cabin rate selector is marked with an arrow for position reference only. Cabin Pressurization Controller The pressurization controller establishes desired cabin altitude and rate of climb by modulating reference air pressure to the pneumatic relay. The controller body is divided into three chambers: cabin pressure, rate pressure, and reference pressure. Cabin air enters the cabin pressure chamber through a filtered orifice. The cabin pressure chamber houses an absolute bellows. Rotating the cabin altitude selector mechanically compresses or extends this bellows to a position that sets the controller to maintain the selected cabin altitude. The rate pressure chamber houses a rate spring secured to the absolute bellows on one side and a rate diaphragm on the other. The rate diaphragm separates the rate chamber from the reference chamber. Airflow passage between these two chambers is regulated by a needle valve that sets the controller to provide the desired cabin rate-of-change according to the position of the CABIN RATE selector.

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Cessna Citation II Technical Manual Clockwise rotation of the CABIN RATE selector decreases airflow through the needle valve and increases cabin rate-ofchange; counterclockwise rotation increases airflow through the needle valve and decreases cabin rate-of-change. In the twelve o’clock position, indicated cabin rate-of-change should be approximately 500 FPM. The rate chamber is ported to a small tank, installed below the airplane floor, which supplies auxiliary volume to the rate chamber to provide greater accuracy in cabin rate-of-change control. Should rate pressure exceed cabin pressure, a check valve will permit airflow from the rate chamber to the cabin pressure chamber. The reference pressure chamber houses a metering valve and follower spring linked to the rate diaphragm. The chamber is ported to cabin pressure, the airplane suction supply, and the cabin outflow valve. Cabin pressure enters the reference chamber through a filtered orifice. The metering valve regulates the flow of cabin pressure to airplane suction to produce the reference pressure, which is then applied to the pneumatic relay. When the metering valve is modulating towards closed, the flow of cabin pressure to airplane suction is reduced and reference pressure is increased. Conversely, when the metering valve is modulating towards open, the flow of cabin pressure to airplane suction is increased and reference pressure is reduced. With a desired cabin altitude and rate-of-change selected, changes in cabin pressure cause the absolute bellows to expand or contract. As it does, the metering valve is repositioned to maintain the correct reference pressure. Airflow between the rate chamber and the reference chamber produces a pressure differential across the rate diaphragm, which further repositions the metering valve to provide the correct cabin rateof-change. Increasing cabin altitude generates an increasing pressure differential between the cabin and reference pressure chambers, causing the rate diaphragm to move, which routes reference air to the pneumatic relay.



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Before takeoff, the cabin altitude selector should be set to 500 feet above the planned cruise altitude on the inner ACFT scale or 500 feet above the destination field pressure altitude on the outer CABIN scale, whichever is greater. The CABIN RATE selector should be positioned to provide a rate-of-change which will result in the cabin reaching the altitude indicated on the CABIN scale as the airplane reaches the corresponding altitude indicated on the ACFT scale. During the takeoff roll, when the airplane is on the ground and the left main gear safety switch is closed, throttle advancement beyond approximately 85% N2 closes a solenoid valve that traps cabin air pressure within the auxiliary volume tank for reference by the cabin pressurization controller, closes a solenoid valve that removes suction to the outflow valves, and closes another solenoid valve that allows the cabin to prepressurize to 60 feet below field altitude at a fixed 500 FPM rate of change. After liftoff, the safety switch functions to open this valve, thereby restoring cabin rate of change control. As the airplane climbs, the absolute bellows expands and contracts as minute changes in cabin pressure are sensed. Expansion and contraction of the bellows is resisted by the pressure differential across the rate diaphragm such that the sum of these forces repositions the metering valve to apply the correct reference pressure to the pneumatic relay. During climb, increasing reference pressure causes the outflow valves to be modulated toward the closed position such that the selected cabin rate-of-change is maintained to the selected altitude. As the airplane reaches the planned cruise altitude at the selected rate and levels off, the pressure differential across the rate diaphragm equalizes and the flow of cabin pressure to airplane suction becomes steady. In this condition, reference pressure becomes essentially constant and cabin altitude stabilizes. If required to ascend beyond the altitude indicated on the ACFT scale, the controller should be reset to a higher altitude to maximize passenger comfort and to prevent unscheduled differential pressure control by the outflow valve. If required to descend below the altitude indicated on the ACFT scale, resetting the controller is normally not required unless the descent will result in airplane altitude being less than the selected cabin altitude. In this case, the controller should be reset to a lower cabin altitude to maximize passenger comfort and to prevent unpressurized operation.

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Cessna Citation II Technical Manual Before descent to landing, the cabin altitude selector should be set to 500 feet above the destination field pressure altitude on the inner ACFT scale and the CABIN RATE selector should be positioned to provide a rate-of-change that will allow the cabin to reach the altitude selected on the CABIN scale as the airplane reaches the corresponding altitude indicated on the ACFT scale. During descent, decreasing reference pressure causes the outflow valve to modulate toward open such that the cabin depressurizes at the selected cabin rate-of-change until the airplane reaches the altitude indicated on the CABIN scale. Below this altitude, the outflow valve will be fully open, the cabin will be unpressurized, and cabin rate-of-change will follow airplane rate-of-change until touch-down. Note: The desired cabin altitude should be set as early as practical to provide the lowest cabin rate-of-change. Rate-ofchange should be adjusted as necessary during ascent or descent so that the cabin reaches the altitude indicated on the CABIN scale at approximately the same time that the airplane reaches the altitude indicated on the ACFT scale. Note: To calculate the approximate field pressure altitude, add 100 feet to the field elevation for each 0.10 inHg that the field altimeter setting is over 29.92 inHg. Subtract 100 feet from the field elevation for each 0.10 inHg that the field altimeter setting is below 29.92 inHg. Pneumatic Relay The pneumatic relay is the primary control device for the outflow valves. The pneumatic relay amplifies reference pressure from the cabin pressurization controller by mixing it with suction air flow generated by an ejector-type pump installed in the left engine environmental supply tube. The pneumatic relay contains four chambers separated by two spring-loaded diaphragms. The upper diaphragm is exposed on the upper side to rate pressure and on the lower side to reference pressure. The lower diaphragm is exposed on the upper side to reference pressure and on the lower side to cabin pressure. Both diaphragms are connected to a metering valve that controls suction air flow to create the amplified reference pressure which is then used to modulate (control) the outflow valves.



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Cabin Altitude Limit Valves Two cabin altitude limit valves, mounted on the aft pressure bulkhead to the right of the outflow valves, serve as backups to the pressurization controller. Each altitude limit valve contains an evacuated bellows within a chamber that is open to cabin pressure through an inlet port, and a spring-loaded poppet valve within a chamber that is connected to the cabin outflow valve reference line. The poppet valve is normally held closed by pressure from a valve spring. If a malfunction causes outflow valve control suction to exceed normal limits, thereby opening the outflow valves excessively, the reduction in cabin pressure allows the altitude limit valve bellows to expand, unseating the poppet valve. With the poppet valve open, cabin pressure enters the outflow valve reference line, increasing outflow valve reference pressure, thereby modulating the outflow valves toward the closed position. The cabin altitude limit valves function to prevent cabin altitude from exceeding approximately 13,000 feet.

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Cessna Citation II Technical Manual Pressurization System Indication Pressurization system operation status is indicated on the cabin altitude and differential pressure indicator, the cabin rate-ofchange indicator, and by the cabin altitude warning annunciator.

1

2

3 1. Cabin Altitude Scale 2. Differential Pressure Scale 3. Cabin Rate of Change Indicator

Cabin Altitude and Differential Pressure Indicator The cabin altitude and differential pressure indicator, located on the environmental panel, is a combination gage having an outer CABIN ALT scale denoting cabin altitude from 0 to 45,000 feet in feet X 1000, and an inner DIFF PRESS scale denoting 0 to 9 PSI differential pressure between the cabin and the atmosphere. The DIFF PRESS scale features a green arc between 0 and 8.7 PSI, denoting the differential pressure range within the normal operating limitations of the system. A red line at 8.8 PSI denotes system overpressure. Cabin Rate-of-Change Indicator The cabin rate-of-change indicator denotes the rate of cabin pressure change from 0 to 6,000 feet per minute in FT/MIN X 1000. Cabin Altitude Warning Annunciator A red [CAB ALT 10000 FEET] annunciator, activated by a barometric pressure switch, illuminates to indicate that the cabin altitude has exceeded 10,000 feet and the use of supplemental oxygen is required.



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Emergency Dump The cabin can be rapidly depressurized by using the emergency dump (depressurization toggle) valve. The guarded operating lever for this valve is labeled EMER DUMP and is located on the environmental panel. Activation of the emergency dump valve applies suc▲ EMERGENCY DUMP VALVE tion that opens the pressurization outflow valves, thereby releasing cabin pressure and allowing cabin altitude to equalize with airplane altitude.

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Cessna Citation II Technical Manual Ambient Air Sources For the purposes of this chapter, ambient air is used for cabin and flight compartment ventilation during ground operation, to provide a positive pressure differential in the aft fuselage while in flight, and to act as the heat exchange medium for air-to-air heat exchangers. Other uses are discussed elsewhere in this manual. On airplanes 550-0484, 0483, 0482 and earlier and on airplanes 550-0627 and after, ambient ventilation air for the cabin is routed from the NACA-type scoop on the dorsal fairing to the cabin fresh air check valve through ducting. On airplanes 5500482, 0485 through 0626, ambient ventilation air is taken from the tailcone. On some of these airplanes the cabin fresh air check valve, attached to the aft pressure bulkhead, is mounted at the end of a duct. On other airplanes no duct is used. An optional flood cooling system is available on airplanes 550-0356 and after. This system is primarily used on the ground, but may also be used at flight altitudes below 10,000 feet. The system functions to supply a mixture of ambient air and conditioned air directly to the passenger cabin, bypassing the normal ▲ FLOOD COOLING SWITCH air distribution system. Flood cooling is activated by a two-position FLOOD COOLING switch on the tilt panel. Positioning the switch to “ON” activates the system by operating an electric actuator in a flow divider and an electric motor/fan. The divider diverts the flow of conditioned air to the axial fan, where it is mixed with ambient air before entering the cabin through a grille at the top of the rear pressure bulkhead. The fan and actuator receive 28 VDC from the right main bus through the 20-amp FLOOD COOLING circuit breaker located in the aft fuselage electrical power junction “J” box. In flight, the aft fuselage (tailcone) is pressurized (relative to outside pressure) to prevent ingestion of external fluids. This is accomplished using ram effect air taken in to the aft fuselage through the dorsal scoop. ENVIRONMENTAL SYSTEM 12/99


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On the ground, the tailcone pressurization system is disabled to prevent pressure transients in the environmental system during the takeoff roll. On airplanes 550-0593 and after (and earlier airplanes incorporating SB550-21-23), the tailcone pressurization system incorporates a valve in the inlet duct from the dorsal scoop that, during ground operation, is energized closed to prevent air from entering the aft fuselage. On airplanes 550-0592 and earlier, a valve in the aft fuselage skin is energized open to allow air to vent overboard until the airplane is airborne. Both types of valve are actuated by thermal expansion of an enclosed fluid, so operating time for the valve will vary according to initial conditions. A 60-second time delay before valve actuation is incorporated on airplanes 550-0550 and after. The pressurization valve receives power from the left main bus through a 5-amp circuit breaker located in the “J” box. The circuit breaker is labeled TAIL PRESSURIZATION or TAIL BUMP PRESS, depending on service bulletin incorporation. On airplanes 550-0482, 0485 and after, ambient air from the aft fuselage is drawn by an ACM-driven fan that directs it through ducting, where the air absorbs heat from both ACM heat exchangers and the precooler before it is dumped overboard through an outlet located on the lower side of the aft fuselage below the engine pylon. On airplanes 550-0484, 0483, 0481 and earlier, ambient air to cool the ACM heat exchangers is routed through ducting from the dorsal scoop to the ACM. Moisture drains are located at the bottom of the ambient air inlet duct and on the bottom of the aft pressure bulkhead check valve duct attach connection. These moisture drains direct accumulated water into the lower aft fuselage area, where it exits through skin drain holes.

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Cessna Citation II Technical Manual Vapor Cycle Air Conditioning System The optional vapor cycle air conditioning system is electricallycontrolled, thermostatically-regulated, and uses Freon (R-12) as a refrigerant. This system consists of one cockpit evaporator, two cabin evaporators, and either a nose baggage compartment or aft fuselage mounted condenser/compressor with associated controls, wiring, and plumbing. The optional vapor cycle air conditioning system provides effective cockpit and cabin cooling. It is used primarily during ground operations, but may also be operated at flight altitudes up to 18,000 feet. The vapor cycle air conditioning system can be used alone or in conjunction with the ACM. Refrigerant Circulation System The refrigerant circulation system functions to activate and control the vapor cycle that reduces the temperature of cabin air. Major components of the system include a compressor, condenser, condenser blower, receiver-dryer, and three evaporator modules. A compressor/condenser unit is located in either the nose baggage compartment or aft fuselage. The receiverdrier unit is installed within the compressor/condenser unit. One evaporator unit is located within the cockpit and two additional evaporator units are located within the cabin. The Freon (R-12) refrigerant used in this system is normally in a gaseous state at standard atmospheric temperatures and pressures. Within specific ranges of temperature and pressure, however, the state of Freon may be transformed between liquid and gas. This characteristic of Freon is critical to understanding the vapor cycle, because in the transformation from gas to liquid (condensation), heat is emitted; and in the transformation from liquid to gas (evaporation), heat is absorbed. Compressor The compressor functions to provide the pressure and suction that circulates Freon through the condenser, the receiver-dryer, and the evaporator modules during air conditioning system operation. The compressor is belt-driven by means of a pulley attached to an electric motor. The same motor also drives an axial fan that provides airflow through the condenser. This airflow provides a cooling effect that condenses the hot gaseous Freon from the compressor into a liquid.



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When air conditioning is not required, the compressor is idle. When air conditioning is required, the compressor initiates the vapor cycle by drawing low-pressure, low-temperature Freon gas through a suction tube and delivering high-pressure, hightemperature Freon gas to the condenser. Condenser The condenser is mounted in proximity to the compressor and functions to transform the high-pressure, high-temperature Freon gas into a high-pressure, low-temperature liquid. Condensation occurs as heat energy in the Freon gas passing through the condenser coils is transferred to cooling fins which are exposed to lower temperature airflow provided by a compressor-driven axial fan. The cooled, high-pressure, liquid Freon is then routed to the receiver-dryer. Receiver-Dryer The receiver-dryer, installed in the high-pressure tube between the condensers and the evaporator modules, functions to remove moisture from the liquid Freon when the air conditioning system is operating. Moisture removal is critical not only in the prevention of corrosion damage, but in the prevention of refrigerant circulation blockage caused by thermal expansion valve freeze-up. Normal operation of the air conditioning system for several minutes followed by the loss of cooling airflow may indicate that freeze-up has occurred. Evaporator Modules One evaporator module is installed within the cockpit, and is accessed by removing floor panels located behind the pilot’s seat. Two additional evaporator modules are installed within the aft baggage compartment in proximity to the aft pressure bulkhead. Each module contains a thermal expansion valve, evaporator coil, drain tube, and evaporator blower. The evaporator modules transform high-pressure liquid Freon into a lowpressure, low-temperature gas, completing the vapor cycle that reduces the temperature of the cabin air.

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Cessna Citation II Technical Manual High-pressure liquid Freon enters each evaporator module through its respective thermal expansion valve that reduces pressure. Each expansion valve incorporates a variable orifice that is regulated automatically through a temperature-sensing bulb attached to the Freon suction tube near the evaporator outlet. When the temperature of the Freon gas leaving the evaporator is too high, the orifice constricts to provide increased cooling. When this temperature becomes too low, the orifice opens to reduce cooling. From the low-pressure side of the expansion valve, reduced pressure liquid Freon is routed through the evaporator coil where it is transformed into a gas. In the transformation from liquid to gas, heat is absorbed from the cabin air as it is drawn through each evaporator coil by its respective blower. The refrigerated cabin air is then forced by the evaporator blowers into the conditioned air distribution tubing. Cooled air from the forward (cockpit) evaporator is introduced into the cabin through armrest-level vents in the flight compartment. Air that has been cooled by passing over the aft evaporators is introduced into the passenger cabin and flight compartment through the overhead duct. As heat is absorbed from the cabin air, moisture accumulates on the evaporator coil and collects in the lower portion of each evaporator module. Drain tubes carry this moisture to forward and aft heated drain assemblies located below the cabin floor, which automatically control drainage of moisture from the airplane. The condensate drain valves within these assemblies incorporate a two-stage orifice that provides maximum drainage during ground operation, and reduced drainage during pressurized flight to minimize loss of cabin pressure. Vapor Cycle Air Conditioning System Protection Vapor cycle air conditioner compressor protection is provided by low-pressure, high-pressure, and suction switches and a 150-amp current limiter. The low and high-pressure switches are mounted on fittings installed on the compressor housing, or on a fitting assembly that is installed in proximity to the compressor.



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Compressor power is routed through an anti-cycle relay. If compressor pressure exceeds 350 PSI, the high-pressure switch opens, thereby removing power from the anti-cycle relay. The anti-cycle relay then interrupts current to the recycling timer, thereby shutting down the compressor. A suction switch is installed into a fitting on the compressor housing. If compressor suction drops below 9.5 PSI, the suction switch opens, thereby removing power from the anti-cycle relay. The anti-cycle relay then interrupts current to the recycling timer, thereby shutting down the compressor. On airplanes 550-0505 and earlier incorporating SB550-21-16, the low-pressure switch is disabled. On airplanes 550-0505 and earlier not incorporating SB550-21-16, if compressor pressure drops below 32 PSI, the low-pressure switch opens, thereby removing power from the anti-cycle relay. The anti-cycle relay then interrupts current to the recycling timer, thereby shutting down the compressor. Vapor Cycle Air Conditioning System Controls and Indicators Airplanes equipped with vapor cycle air conditioning have a FREON AIR CONDITIONER control panel mounted on the copilot’s instrument panel. This air conditioning control panel includes a rotary switch with four positions, labeled OFF, FAN FWD, FAN ALL, and COMP and a FAN SPEED toggle switch with HI and LO positions. The FAN FWD position energizes only the flight compartment blower, the FAN ALL position energizes the all the blowers, and the COMP position energizes the compressor and all the blowers. The selected blowers will run at the speed selected by the FAN SPEED toggle switch. Vapor Cycle Air Conditioning System Protection The compressor motor will automatically shut down should motor current exceed 350 amps, condenser Freon discharge pressure exceed 350 PSIG, or Freon condenser suction pressure fall below 10 PSIG.

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Cessna Citation II Technical Manual Emergency Oxygen System The oxygen system is designed to provide emergency breathing oxygen for the crew and passengers in the event of a pressurization system emergency occurring at flight altitudes above 10,000 feet, or if the cabin fills with smoke. The standard oxygen system includes a 22 cu-ft oxygen bottle that will provide emergency oxygen for crew and six passengers for up to 15 minutes; an optional system incorporates a 64 cu-ft bottle that will provide emergency oxygen for crew and six passengers for up to 50 minutes. Emergency oxygen system operation is initiated by an altitude pressure switch installed in the pressurized cabin area. At a cabin altitude of 12,900 to 14,000 feet, the altitude pressure switch energizes a solenoid valve that allows oxygen to flow to the oxygen distribution system. Oxygen pressure of approximately 70 PSI automatically opens passenger oxygen mask storage compartment doors, thereby dropping the masks. Both oxygen systems provide distribution plumbing for the flight compartment and passenger cabin, individual outlets for each airplane occupant (location depending on seating configuration), an oxygen bottle, regulator, filler valve, pressure gage, and control switch. In all installations, the system regulator is assembled directly to the oxygen bottle and functions to reduce bottle pressure to a lower, constant supply pressure. The regulator incorporates a mechanically-operated supply shutoff valve and ports for a supply tube, filler tube, pressure gage tube, and overboard discharge indicator tube. The pressure regulator outlet ports in the 64 cu-ft system are installed in slightly different positions than the ports in the 22 cu-ft system. On airplanes 550-0255 and after and earlier airplanes incorporating SB550-35-2, the standard or optional oxygen bottle is installed on the aft fuselage compartment structure using mounting brackets. On airplanes 550-0254 and earlier not incorporating SB550-35-2, the standard or optional oxygen bottle is installed below the right nose baggage compartment floor using mounting brackets.



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All oxygen bottles require a U.S. Department of Transportation (D.O.T.) designation that identifies bottle specification and service pressure. The D.O.T. designation also identifies the bottle’s life limitation and hydrostatic testing requirements. The 22 cu-ft bottle bears a DOT-3AA 1800 designation and requires hydrostatic testing to 167% of its service pressure every five years. There is no life limitation for this bottle unless failure occurs during hydrostatic testing. The 22 cu-ft bottle is charged to 1800 PSI under standard atmospheric conditions. Two types of 64 cu-ft bottles are available. The lightweight bottle bears a DOT-3HT 1850 designation; the fiber-wound bottle bears a DOT-3FC1850 designation. Both types of bottle require hydrostatic testing to 167% of service pressure every three years. Life limitation is twenty-four years from date of manufacture. All oxygen bottles should be filled with breathing oxygen that conforms to the requirements of MIL-0-27210, Type 1. On airplanes and 550-0255 and after and earlier airplanes incorporating SB550-35-2, the oxygen filler valve is located inside the tailcone baggage compartment door. On airplanes 550-0254 and earlier not incorporating SB550-35-2, the oxygen filler valve is accessible through the right nose baggage compartment door. The oxygen filler valve incorporates a filter and a protective cap. A check valve, installed in-line between the regulator and the filler valve, prevents the escape of bottle pressure from the filler tube or its connections. Oxygen Outlets Oxygen outlets for the pilot and copilot are located on the pilot’s and copilot’s side consoles in the flight compartment. Up to eight oxygen outlets may be installed in the overhead passenger cabin. Due to differences in seating configurations, the location of the passenger cabin outlets will vary between airplanes. Each outlet incorporates a spring-loaded valve that prevents oxygen flow from the outlet unless a hose assembly is connected.

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Cessna Citation II Technical Manual Warning: Permit no smoking when using oxygen. Materials that will not normally flash in the atmosphere will readily burn or explode in the presence of concentrated oxygen. Oil, grease, soap, lipstick, lip balm, and other fatty materials constitute a serious fire hazard when in contact with oxygen. Be sure hands and clothing are oil free before handling oxygen equipment. Crew Oxygen Masks Two types of crew oxygen masks are available in the Citation II. The standard oxygen mask is a diluter demand type with integral oxygen regulator and microphone. Each oxygen regulator includes a lever that allows selection of diluter demand (NORMAL) or demand (100% OXY) modes. The demand mode should be selected to insure adequate supplemental oxygen at altitudes above 20,000 feet. The standard mask qualifies as quick donning when it is worn with the head strap around the neck. The optional crew oxygen mask is a quick-donning sweep-on type with a regulator and microphone attachment. This mask is a diluter demand type with pressurized flow (100% oxygen) selectable by placing the regulator in the “EMER” (demand) position. The EMER position should be selected to insure adequate supplemental oxygen at cabin altitudes above 20,000 feet. To conserve oxygen, the regulator may be set to “NORMAL” if cabin altitude is below 20,000 feet. To qualify as quickdonning, the mask must be properly stowed in its retainer. Either mask should be set to the 100% oxygen (“100% OXY” or “EMER”) position when it is used for smoke protection.



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Oxygen mask microphones are operated using a twoposition toggle switch on the pilot’s and copilot’s side consoles. Setting this switch to “MIC OXY MASK” energizes the mask microphone; setting the switch to “MIC HEAD SET” energizes the headset microphone. The selected microphone may then be used for transmission by depressing the microphone button on the control wheel. Passenger Oxygen Masks Passengers are provided oronasal type oxygen masks that deform to seal around the nose and mouth area. Each mask consists of a face plate, economizer bag, plastic supply tube, and a lanyard cord with pintle pin attached. The pintle pins are installed to prevent oxygen loss from unused masks. After oxygen masks are deployed, the lanyard cord must be pulled to withdraw the pintle pin and thereby initiate oxygen flow to each mask. Passenger oxygen masks provide a constant flow rate of 4.5 liters per minute.

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Cessna Citation II Technical Manual Oxygen System Controls, Malfunction Warning, and Indication On airplanes 550-0501 and after, the rotary, three-position OXYGEN CONTROL VALVE switch is located on the pilot’s side console. When this switch is set to the center “NORMAL” position, oxygen automatically flows to both the cockpit and cabin. When the switch is rotated counterclockwise to the “CREW ONLY” position, oxygen flow is limited to the cockpit. The switch is rotated clockwise to the “MANUAL DROP” position to manually deploy passenger oxygen masks if the automatic mask deployment system should fail. On airplanes 550-0500 and earlier, two-position toggle switches labeled OXYGEN PRIORITY VALVE and PASS OXY MASKS are located on the pilot’s side console. When the OXYGEN PRIORITY VALVE switch is set to the upper “NORMAL” position, oxygen automatically flows to both the cockpit and cabin. When the switch is set to the “CREW ONLY” position, oxygen flow is limited to the cockpit. The PASS OXY MASKS switch is repositioned from “NORMAL” to “MANUAL DROP” to manually deploy passenger oxygen masks if the automatic mask deployment system should fail.



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The oxygen pressure gage is located on the right instrument panel. The gage provides visual indication of bottle pressure and is marked with a yellow arc from 0 to 400 PSI, a green arc from 1600 to 1800 PSI, and red line at 2000 PSI. When fully charged and stabilized at approximately 70°F, indicated pressure should be 1800 PSI for the 22 cu-ft bottle and 1850 PSI for the 64 cu-ft bottle; however, indicated pressure will vary with ambient temperature. Either bottle will require recharging if indicated pressure falls below 300 PSI. Evidence of oxygen bottle overpressure is provided by an indicator disc (originally green in color). Should overpressure occur, a high-pressure rupture fitting within the regulator releases bottle pressure through the overboard discharge indicator tube. When oxygen bottle pressure exceeds 2850 ± 150 PSI (at 70°F/21°C), the disc is “blown out,” thereby providing visual indication that oxygen was discharged overboard. On airplanes with oxygen bottles located in the nose section, the indicator disc is located on the lower right surface of the nose. On airplanes with oxygen bottles located in the aft fuselage, the indicator disc is located on the lower left surface of the tailcone.

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Cessna Citation II Technical Manual Limitations Refer to the applicable airplane manufacturer’s FAA approved airplane flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable airplane manufacturer’s FAA approved airplane flight manual or approved manual material (supplementary checklist) as revised, for procedural information.



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Chapter 10 Ice Protection Systems Table of Contents Overview ............................................. 10-1 Anti-Ice Systems ................................. 10-2 Pitot-Static Anti-Ice .............................. 10-2 Windshield Anti-Ice .............................. 10-4 Bleed Air Windshield Anti-Ice ........... 10-4 System Operation .......................... 10-8 Malfunction Warning and Protection ............................... 10-9 Bleed Air Windshield Rain Removal .............................. 10-10 Alcohol Windshield Anti-Ice ............ 10-11 Surface Deice System ....................... 10-13 System Operation ........................ 10-15 Ice Detection ............................... 10-18 Engine Ice Protection......................... 10-19 Inboard Wing Leading Edge Anti-Ice ................................ 10-20 System Operation ........................ 10-22 Malfunction Indication................... 10-22 Limitations ........................................ 10-23 Emergency Procedures....................... 10-23

Cessna Citation II Technical Manual Overview This chapter describes the ice protection equipment required on the Cessna Citation II for flight in icing conditions. Anti-ice systems are designed to prevent the formation of ice and should be activated prior to entering icing conditions. For this purpose, electrically-powered heating elements are installed in the pitot tubes, static ports, and angle-of-attack sensor (if installed). Protection against windshield icing is accomplished primarily using engine bleed air, with alcohol used as a backup anti-ice system for the pilot’s windshield. Engine ice protection is accomplished by bleed air heating of induction air inlet components and electrical heating of the inboard wing leading edges forward of each engine. Deice systems are designed to remove ice which has accumulated. For this purpose, pneumatically-operated boots are attached to the leading edges of the stabilizers and the outboard leading edge of each wing. The Cessna Citation II is approved for flight in icing conditions as defined by the FAA only when the following ice protection equipment is installed and checked operational before flight: Anti-Ice Heated Pitot Tubes Heated Static Ports Heated Windshield Backup Windshield Alcohol System Engine Ice Protection Deice Wing and Stabilizer Deice Boots Note: Refer to the FAA-approved Master Minimum Equipment List (MMEL) for conditions and limitations specific to the ice protection equipment installed in your airplane.

ICE PROTECTION SYSTEMS 12/99


10-1

Anti-Ice Systems For organizational purposes, this section is divided into pitotstatic anti-ice, windshield anti-ice, and engine anti-ice. All antiice systems must be activated when operating in visible moisture at indicated outside air temperatures (IOAT) between +4°C (39°F) and -30°C (-22°F). Pitot-Static Anti-Ice

▲ PITOT TUBE

▲ STATIC PORTS

The pitot tubes and static ports are protected against icing by integral, electrically-powered heating elements which are controlled by the PITOT & STATIC switch on the lower left instrument panel. When this switch is in the upper (on) position, 28 VDC power is supplied to the heating elements of the pilot’s pitot tube and static ports from the left main bus through the 7.5amp LH PITOT STATIC circuit breaker, and supplied to the copilot’s pitot tube and static ports from the right main bus through the 7.5-amp RH PITOT STATIC circuit breaker. Independent left (pilot’s) and right (copilot’s) current sensors monitor the flow of current to their associated heating elements. On airplanes 550-0550 and after, these current sensors control the illumination of an amber, three lens [LH] [RH] [P/S HTR OFF] annunciator. On airplanes 550-0028 and earlier, the current sensors control the illumination of a single pitot heat off/fail light on the left instrument panel; on airplanes 550-0029~0505, this light is replaced by an amber, single lens [P/S HTR OFF] annunciator.

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Cessna Citation II Technical Manual When the PITOT & STATIC switch is in the upper “on” position and all heating elements are functioning normally, the annunciator should be extinguished. When this switch is in the “on” position and at least one heating element is inoperative, the annunciator should be illuminated. When the switch is in the “OFF” position or when no ▲ PITOT-STATIC HEAT SWITCH power is being supplied to the heating elements, the annunciator should also be illuminated. On airplanes 550-0550 and after, the corresponding [LH] (pilot’s) or [RH] (copilot’s) lens will illuminate in conjunction with the [P/S HTR OFF] lens to indicate which heating elements are malfunctioning. To minimize battery drain and prevent overheating of the elements during ground operation, the PITOT & STATIC switch should remain in the “OFF” position except for system testing. To test the system prior to flight, the pitot tube covers should be removed (if installed), the PITOT & STATIC switch should be positioned to “on” for a period of 30-seconds and then returned to “OFF.” If the system is functioning normally, the [P/S HTR OFF] annunciator should not have illuminated when the PITOT & STATIC switch was in the “on” position, the pitot tubes should be hot, and the static ports should be warm. To reduce the risk of severe burns when checking pitot tube heat, physical contact with the tube should be minimized and grasping the tube with any more than a light grip should be avoided. Checking static port heat in high ambient temperatures is best accomplished using the back of a finger to compare the temperature of each static port to that of the surrounding fuselage skin. Refer to the appropriate Operating Manual or Airplane Flight Manual for test procedures specific to the system installed in your airplane. Caution: To prevent overheating of the elements, ground operation of the pitot-static ice protection system is limited to 2 minutes.



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Windshield Anti-Ice

2 3

1

1. Rain Removal Augmenter Door 2. Alcohol Spray Tubes 3. Bleed Air Discharge Nozzle Shroud Protection against windshield icing is accomplished primarily using engine bleed air, with alcohol used as a backup anti-ice system for the pilot’s windshield. In addition to ice protection, the bleed air windshield anti-ice system provides rain removal and external defogging capabilities. Bleed Air Windshield Anti-Ice The bleed air windshield anti-ice system directs engine bleed air against the windshield to prevent the formation of ice. Major components of the system include a bleed air control valve, heat exchanger, automatic temperature controls, temperature and pressure sensors, manual flow controls, and bleed air discharge nozzles. Bleed air used by the system is tapped from the supply tubes between the compressor section of each engine and the environmental flow control/shutoff valves within the aft fuselage. On airplanes 550-0484, 0483, 0481 and earlier, the bleed air is routed through a precooler within each engine nacelle prior to entering the aft fuselage. On airplanes 550-0482, 0485 and after, availability of bleed air to the system is continuous when either or both engines are operating and is not influenced by the position of the PRESS SOURCE selector. On airplanes 5500484, 0483, 0481 and earlier, however, setting the PRESS SOURCE selector to “LH” or “RH” correspondingly results in bleed air being made available to the system from the left engine or right engine only.

10-4



Cessna Citation II Technical Manual The bleed air tapped from each source is routed through independent supply tubes to a common cross fitting within the aft fuselage. A check valve in each of these tubes prevents the backflow of bleed air to either engine when the opposite engine has failed or is operating at a sufficiently lower RPM. From the cross fitting, bleed air is routed through the bleed air control valve to the heat exchanger, each also located within the aft fuselage.

4

1

1. Heat Exchanger 2. Bleed Air Control Valve

3

2

3. Cross Fitting 4. Exit Duct

The bleed air control valve is normally-open, electrically-actuated closed, and controlled primarily by the W/S BLEED switch on the lower left instrument panel. When this valve is open, bleed air is routed through the heat exchanger to the rest of the system.



10-5

The heat exchanger is an air-to-air type unit which functions to reduce bleed air temperature to that required for system operation. The heat exchanger is installed within an air duct through which cooling airflow passes and conducts heat from the bleed air. On airplanes 550-0482, 0485 and after, airflow enters this duct through a screened inlet on the right side of the tailcone, and exits through a NACAtype exhaust scoop on the left side of the tailcone below the engine pylon. On airplanes 550-0484, 0483, and 0481 and earlier, airflow enters this duct through flush-mounted, NACAtype intake scoops on the dorsal fairing, and exits through a NACA-type exhaust scoop on the left side of the tailcone below the engine ▲ AIRFLOW OVERBOARD EXHAUST SCOOP pylon. The automatic temperature controller maintains the required bleed air temperature by modulating the position of an electrically-actuated air control valve, located in the heat exchanger exit duct, which regulates ambient airflow through the heat exchanger. Input signals are provided to the controller by the W/S BLEED switch and by temperature sensors located in the bleed air supply tubing: one downstream of the heat exchanger, another upstream of the nozzles.

10-6



Cessna Citation II Technical Manual From the heat exchanger, bleed air is routed through a pair of mechanically-actuated control valves which regulate bleed air flow to the discharge nozzles. Each valve is located within the nose section and manually positioned by its associated LH or RH WINDSHIELD BLEED AIR control on the copilot’s lower instrument panel. Each discharge nozzle ▲ DISCHARGE NOZZLE VALVE is enclosed in an aerodynamic shroud and comprised of a manifold which supplies an array of outlet tubes that direct bleed air against the windshield. Each shroud is fitted with a hinged augmenter door for rain removal. Both augmenter doors are mechanically-operated by a single PULL RAIN control, located below the copilot’s instrument panel. The left (pilot’s) shroud also houses the alcohol dispersal nozzle. Rotating the WINDSHIELD BLEED AIR controls clockwise progressively increases bleed air flow to the discharge nozzles; counterclockwise rotation progressively decreases bleed air flow to the discharge nozzles. When windshield rain removal is required, the WINDSHIELD BLEED AIR controls should be rotated fully-clockwise to “MAX.” When windshield ice protection or rain removal is not required, the WINDSHIELD BLEED AIR controls should be rotated fully-counterclockwise to “OFF.”



1

2

1. Windshield Bleed Air Controls 2. Pull Rain Control Knob

10-7

System Operation The system is activated by the three-position (HI/OFF/LOW) W/S BLEED switch on the lower left instrument panel. When this switch is set to “HI” or “LOW,” the bleed air control valve is deenergized open and 28 VDC power is supplied from the left main bus (550-0550 and after) or right main bus (550-0505 and ▲ W/S BLEED SWITCH earlier) to the automatic temperature controller through the 5-amp W/S BLEED TEMP circuit breaker. When “HI” is selected, the temperature controller modulates the position of the air control valve to maintain bleed air temperature at approximately 138°C. When “LOW” is selected, the temperature controller modulates the position of the air control valve to maintain bleed air temperature at approximately 127°C. When set to “OFF,” the temperature controller is deactivated and the bleed air control valve is energized closed. On airplanes 550-0550 and after, 28 VDC left main bus power is supplied to the bleed air control valve through the 5-amp W/S BLEED circuit breaker. On airplanes 550-0505 and earlier, 28 VDC right main bus power is supplied to the bleed air control valve through the 5-amp WINDSHIELD BLEED AIR circuit breaker. Before activating the system, the WINDSHIELD BLEED AIR controls should be rotated clockwise and the PULL RAIN control, located below the copilot’s instrument panel, should be pushed fully-in. When windshield ice protection is required, the W/S BLEED switch should be set to “HI” when IOAT is below -18°C, or “LOW” when IOAT is above -18°C. Normal system operation is indicated by an increase in air noise as bleed air is discharged from the nozzles.

10-8



Cessna Citation II Technical Manual Bleed Air Windshield Anti-Ice Schematic

© PCW

Malfunction Warning and Protection Malfunction warning and protection is provided by a pressure switch and overheat sensor, located in the bleed air supply tubing downstream of the heat exchanger, and an amber [W/S AIR O’HEAT] light located on the annunciator panel. The pressure switch functions to illuminate the [W/S AIR O’HEAT] annunciator when the W/S BLEED switch is set to “OFF” and bleed air pressure is at least 5 PSI, indicating failure of the bleed air control valve to energize closed. The overheat sensor functions to illuminate the [W/S AIR O’HEAT] annunciator and energize the bleed air control valve closed when bleed air temperature exceeds 146°C, indicating failure of the automatic temperature controls to regulate ambient airflow through the heat exchanger. Should this condition occur, normal system operation should be restored automatically when bleed air temperature no longer exceeds 146°C. Note: If normal system operation is not restored automatically within 60-seconds of [W/S AIR O’HEAT] annunciator illumination, bleed air flow to the windshield should be reduced to the minimum necessary to maintain sufficient visibility. This can be accomplished by setting the W/S BLEED switch to “LOW” or “OFF” and/or rotating the WINDSHIELD BLEED AIR controls to “OFF.” ICE PROTECTION SYSTEMS 12/99


10-9

Operation of the overheat sensor and [W/S AIR O’HEAT] annunciator may be verified by rotating the TEST selector, located on the lower left instrument panel, to the “W/S TEMP” position and setting either W/S BLEED switch to “HI” or “LOW,” In this condition, the [W/S AIR O’HEAT] annunciator should illuminate if the sensor is functional. When illuminated ▲ TEST SELECTOR SWITCH by the overheat sensor during system operation or testing, the [W/S AIR O’HEAT] annunciator receives 28 VDC power from the left main bus through the 5amp W/S BLEED circuit breaker (550-0550 and after), or from the right main bus through the 5-amp WINDSHIELD BLEED AIR circuit breaker (550-0505 and earlier). When illuminated by the pressure switch, the [W/S AIR O’HEAT] annunciator receives 28 VDC power from the left main bus (550-0550 and after) or right main bus (550-0505 and earlier) through the 5-amp W/S BLEED TEMP circuit breaker. Bleed Air Windshield Rain Removal Rain removal is provided by the windshield anti-ice system and the augmenter doors on each discharge nozzle shroud. When bleed air flow from discharge nozzles is insufficient to clear the windshield of heavy rain, the augmenter doors can be opened to provide increased airflow over the windshield. Both augmenter doors are mechanicallyoperated by a single PULL RAIN control, located below the copilot’s instrument panel. When windshield rain removal is required, the WINDSHIELD BLEED AIR controls should be rotated fully-clockwise to “MAX,” the PULL RAIN control should be pulled fully-out, and the W/S BLEED switch should be set to “LOW.” Note: Difficulty may be encountered opening the augmenter doors at airspeeds above 175 KIAS, or when the W/S BLEED switch is set to “LOW” prior to operating the PULL RAIN control. 10-10



Cessna Citation II Technical Manual Alcohol Windshield Anti-Ice The alcohol windshield anti-ice system provides a backup to the bleed air windshield anti-ice system for the pilot’s windshield only. Major components of the system include an alcohol reservoir, an electrically-operated pump, and a six-tube dispersal nozzle.

1

3 2

1. Fluid Level Sight Gage 2. Alcohol Reservoir 3. Augmenter Door Linkage The alcohol reservoir is located behind the aft divider within the right nose baggage compartment. A sight gage on the upper reservoir permits fluid level inspection. The sight gage is visible through an inspection window positioned on an access panel. The access panel is hinged to facilitate servicing and is accessible through the right baggage door. If fluid is not visible in the sight gage, the reservoir should be replenished. Reservoir capacity is 0.5 U.S. gallons TT-I-735 isopropyl alcohol only.



10-11

The electrically-operated pump, also located behind the aft divider within the right nose baggage compartment, delivers alcohol under positive pressure to the dispersal nozzle. The pump incorporates an integral fluid filter through which alcohol passes before delivery to the dispersal nozzle. To prevent audio interference during system operation, a radio noise filter is installed in the electrical circuit to the pump. The dispersal nozzle, enclosed within the left (pilot’s) shroud, incorporates six spray tubes which distribute alcohol over the pilot’s windshield. The system is activated by the two-position (ON/OFF) W/S ALCOHOL ANTI-ICE switch on the lower left instrument panel. When this switch is positioned to “ON,” 28 VDC power is supplied to the pump from the right main bus (550-0550 and after) or left main bus (5500505 and earlier) through the 5-amp W/S ALCOHOL circuit breaker. With the pump ener▲ W/S ALCOHOL ANTI-ICE SWITCH gized, alcohol is drawn from the reservoir and delivered to the dispersal nozzle. With the alcohol reservoir serviced to capacity, maximum continuous operation endurance is approximately 10 minutes. Note: If failure of the bleed air windshield anti-ice system necessitates activation of the alcohol windshield anti-ice system, icing conditions should be exited as soon as practicable. During preflight inspection, the alcohol spray tubes should be checked for general condition and cleanliness, and the reservoir level should be checked full. Operation of the system can be tested before flight by positioning the W/S ALCOHOL ANTIICE switch to “ON” until alcohol is observed flowing from all six spray tubes.

10-12



Cessna Citation II Technical Manual Surface Deice System The surface deice system functions to remove ice accumulations from the leading edges of the stabilizers and outboard leading edge of each wing. The electrically-controlled, pneumatically-operated system consists of inflatable rubber deice boots, a pneumatic pressure regulator, three control valves, a timer module, two pressure switches, and associated controls, plumbing and circuitry.

1

1. Outboard Leading Edge Boot 2. Vertical Stabilizer Boot 3. Horizontal Stabilizer Boot

2

3



10-13

The deice boots are essentially fabric-reinforced rubber sheets containing built-in, spanwise inflation tubes. Each boot is bonded by adhesive to the leading edge of the surface being protected and features a conductive coating which discharges static electricity. The boots are normally held deflated against their respective leading edge surfaces by suction. When inflated by air pressure, the change in boot contour breaks up ice accumulations to facilitate removal by normal in-flight air forces. Air pressure for boot inflation and suction for boot deflation is provided by engine bleed air supplied from the same cross fitting that supplies the windshield anti-ice system. From this cross fitting, bleed air is routed through the pneumatic pressure regulator which functions to reduce bleed air pressure to approximately 23 PSIG. From the regulator, bleed air is routed through a cross fitting where its flow is divided into three paths which independently supply the stabilizer, left wing, and right wing control valves, also located within the aft fuselage. Bleed air is continuously supplied to the control valves whenever either or both engines are operating. Refer to the Bleed Air Windshield Anti-Ice section of this chapter for a description of the bleed air source. The three electrically-actuated control valves function to control the application of suction (when deenergized closed) or pressure (when energized open) to their associated boots as determined primarily by the timer module. When deenergized closed, each control valve functions as an ejector, producing approximately 5.5 inHg of ▲ DEICE BOOT OVERBOARD VENT TUBES suction by directing bleed air through an overboard vent tube. When energized open, the overboard vent tube is blocked and bleed air inflation pressure is directed to the boots. The timer module, located within the flight compartment left side console, functions to energize and deenergize the control valves sequentially. The pressure switches, one located in the stabilizer boot supply line, the other located in the right wing boot supply line, function to illuminate the white [SURFACE DEICE] (550-0550 and after) or [SURF DEICE] (550-0505 and earlier) annunciator when inflation pressure is at least 20 PSI. 10-14



Cessna Citation II Technical Manual System Operation The system is activated by the three-position SURFACE DEICE switch on the lower left instrument panel. When this switch is momentarily actuated to the upper (on) position, 28 VDC power is supplied from the right main bus (550-0550 and after) or left main bus (550-0505 and earlier) to the timer module through the 5▲ SURFACE DE-ICE SWITCH amp SURFACE DE-ICE circuit breaker, thereby initiating a two-sequence deice cycle. During the first sequence, the timer energizes the stabilizer control valve open directing inflation pressure to the stabilizer boots for approximately six seconds. Full inflation and annunciator illumination normally occur within approximately two seconds. During the second sequence, the timer deenergizes the stabilizer control valve closed and energizes the left and right wing control valves open directing inflation pressure to the wing boots for approximately six seconds. The annunciator will extinguish momentarily between sequences and illuminate when the wing boots are fully-inflated. Full deflation of the stabilizer boots normally occurs within approximately twelve seconds following completion of the first sequence. Full deflation of the wing boots normally occurs within approximately twelve seconds following completion of the second sequence. Upon completion of the cycle, the timer module and control valves are deenergized, the annunciator is extinguished and suction is applied to all of the boots. Each momentary actuation of the SURFACE DE-ICE switch to “on” results in one complete cycle. Though the inflation sequences last approximately twelve seconds combined, the additional time required for the deflation of all boots results in one complete cycle actually lasting approximately twenty-four seconds. System activation may be repeated as necessary allowing twenty-four seconds between cycles. Note: If the boots fail to deflate or if cycle termination is desired, momentary actuation of the SURFACE DE-ICE switch to the lower “RESET” position overrides the timer module and immediately deenergizes all three control valves closed.



10-15

Surface Deice System Schematic (boots deflated)

© PCW

The system should be activated when ice accumulations of at least 1/4” to 1/2” are observed on the leading edge of either outboard wing. Activation of the system with accumulations of less than 1/4” may result in ice bridging on the wing. Accumulations of greater than 1/2” may exceed the system’s ice removal capabilities. Operation and/or testing of the system at indicated outside air temperatures (IOAT) below -40°C (-40°F) may result in boot cracking or failure of the boots to fully-deflate. During preflight inspection, the deice boots should be checked for general condition and cleanliness. Operation of the system can be tested before flight by momentarily actuating the SURFACE DE-ICE switch to the upper “on” position and visually confirming normal inflation and deflation of the wing boots as well as illumination of the annunciator.

10-16



Cessna Citation II Technical Manual Surface Deice System Schematic (first cycle: empennage boot inflation)

© PCW

Surface Deice System Schematic (second cycle: wing boot inflation)

© PCW



10-17

Ice Detection A wing ice inspection light is installed on the left side of the fuselage, forward of the wing. The light illuminates the upper surface and leading edge of the wing so that these surfaces can be checked for ice accumulation during night operations. The light is controlled by the WING INSP switch on the lower left instrument panel. ▲ WING ICE INSPECTION LIGHT When this switch is positioned to “ON,” 28 VDC power is supplied to the light from the right main bus (550-0550 and after) or left main bus (550-0505 and earlier) through the 5-amp WING INSP circuit breaker.

10-18



Cessna Citation II Technical Manual Engine Ice Protection Engine ice protection is accomplished by bleed air heating of induction air inlet components and electrical heating of the inboard wing leading edges forward of each engine.

1

2

1. T2 Probe Supply Line 2. Inlet Lip Bleed Air Valve 3. Inlet Stator Vane Supply Line

3

Each engine’s bleed air antiice system operates independently of the other and uses compressor discharge air (P3) to heat the inlet nose cone, core inlet guide vanes (stators), T2 probe, and inlet lip. During engine operation, bleed air is supplied continuously to the inlet nose cone through the hollow interior of the N1 shaft, and to the T2 probe through an external line. Inlet lip bleed air is tapped from the primary bleed air supply tubes. These supply tubes extract bleed air from the compressor through ports located at 4 and 8 o’clock positions on the gas generator case. Core inlet stator bleed air is ported directly from the compressor through a dedicated external line.



10-19

Availability of bleed air to the inlet lip and core inlet stators is through independent pressure regulating/shutoff valves which are electrically-actuated closed, and controlled primarily by the LH and RH ENGINE anti-ice switches on the lower left instrument panel. Although these valves are electrically considered normally-open, they are pneumatically considered normallyclosed and require at least 8 PSI bleed air pressure to open. Additionally, each inlet lip valve requires that its corresponding THROTTLE lever be positioned above 60% N2 to open. The stator valve incorporates a position switch and the inlet lip incorporates a temperature switch, each associated primarily with system malfunction indication. Inboard Wing Leading Edge Anti-Ice The inboard wing leading edge anti-ice system operates in conjunction with the engine bleed air anti-ice system and uses electrically-heated panels to prevent the formation of ice on the upper wing surface forward of the engines. Each removable panel features a highly-polished exterior that ▲ INBOARD WING ANTI-ICE PANEL forms a 61” section of its associated inboard wing leading edge. Five independent, spanwise heating elements, a high temperature switch, low temperature switch, and a temperature sensor are bonded to the interior of each panel. A Kevlar insulation shield provides a thermal barrier between the heated panel and the wing structure. An independent temperature control circuit is provided for each panel to maintain operating temperatures between 54°C and 78°C nominal. Each circuit includes a temperature controller, control relay, and power relay which function to regulate the flow of current to the heating elements. With the system activated, current flow to the heating elements will occur only when the control relay and power relay are energized closed.

10-20



Cessna Citation II Technical Manual The temperature sensor in each panel is positioned where the highest operating temperatures exist. The temperature controller cycles the system on and off according to input signals supplied by the temperature sensor. To accomplish this, the temperature controller energizes and deenergizes the control relay as a function of sensed temperature. When approximately 54°C is sensed, the control relay is energized closed, thereby supplying current to the heating elements. When approximately 78°C is sensed, the control relay is deenergized open, thereby interrupting current flow to the heating elements. The low temperature switch in each panel is also positioned where the highest operating temperatures exist, while the high temperature switch is positioned where lower operating temperatures exist. In the event of temperature controller failure, the high temperature switch energizes and deenergizes the power relay as a function of sensed temperature. When approximately 74°C is sensed, the power relay is deenergized open, thereby interrupting current flow to the heating elements. When approximately 68°C is sensed, the power relay is energized closed, thereby supplying current to the heating elements. Note: Although the control relay and power relay appear to have overlapping temperature activation ranges, the location of the temperature sensor and high temperature switch ensures that neither senses the same temperature simultaneously. 28 VDC power is supplied to each panel by its associated left or right main bus through a 175 amp current limiter. Each heating element is provided with a circuit breaker and current sensor. The five current sensors for each panel are wired in series and independently monitor the flow of current to their associated heating elements. The current sensors and the low temperature switch are associated primarily with system malfunction indication.



10-21

System Operation The engine bleed air and inboard leading edge anti-ice systems are simultaneously activated by the LH and RH ENGINE anti-ice switches on the lower left instrument panel. When these switches are in the upper (on) position, 28 VDC power is supplied to the inboard wing leading edge heating elements and the ▲ ENGINE ANTI-ICE SWITCHES ignition system. After a five second time delay, the pressure regulating shutoff valves for each inlet lip and core inlet stator will be deenergized open if the THROTTLE levers are positioned above 60% N2 and at least 8 PSI bleed air pressure is available. If the system is functioning normally, each heated leading edge panel will draw approximately 150 amps and the consumption of bleed air for inlet lip and core inlet stator heating will increase indicated ITT and decrease engine RPM. Malfunction Indication Indication of engine ice protection system malfunction is provided by the amber [LH] and [RH] [ENG ANTI-ICE] annunciators (550-0550 and after) or [L ENG ICE FAIL] and [R ENG ICE FAIL] annunciators (550-0505 and earlier). Illumination of the corresponding annunciator(s) will occur under the following conditions when the LH and RH ENGINE anti-ice switches are in the upper (on) position: 1. by current sensor when at least one heating element is inoperative 2. by low temperature switch when leading edge temperature is below approximately 16°C 3. by high temperature switch when leading edge temperature is above approximately 74°C 4. by temperature controller when temperature sensor malfunctions 5. by position switch when core inlet stator valve fails to open 6. by temperature switch when inlet lip temperature is below 104°C

10-22



Cessna Citation II Technical Manual Engine Ice Protection Schematic

© PCW

Note: Illumination of these annunciators during the five seconds following system activation is normal before the pressure regulating shutoff valves for each inlet lip and core inlet stator open. Limitations Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material, markings and placards, or any combination thereof for all limitations. Emergency Procedures Refer to the applicable airplane manufacturer’s FAA approved flight manual or approved manual material (supplementary checklist) as revised, for procedural information.



10-23


Chapter 11 Limitations Table of Contents Overview ............................................. 11-1 Airspeed Limitations ............................ 11-2 Operating Limitations ........................... 11-3 Weight Limitations.......................... 11-3 Center of Gravity Limits .................. 11-3 Takeoff and Landing Limitations ...... 11-4 Flight Load Factor Limitations ......... 11-4 Enroute Limitations ........................ 11-5 Approved Operations ...................... 11-5 Engine Operating Limitations ................ 11-6 Engine Fan .................................... 11-6 Battery and Starter Cycle Limitations ................................. 11-7 Battery Limitation ........................... 11-7 Prolonged Ground Operations .......... 11-7 Oil Limitations .................................... 11-8 Approved Oils ................................ 11-8 Fuel Limitations .................................. 11-9 Approved Fuels .............................. 11-9 Fuel Temperature and Density Limitations ....................... 11-10 Maximum Fuel Imbalance ............. 11-10


Limitations, continued Table of Contents Hydraulic Fluid Limitations ................. 11-10 Flight Crew Limitations ...................... 11-11 Cabin Limitations .............................. 11-11 Pressurization Differential .................. 11-11 Pressurization Source Selector ........... 11-11 Icing Limitations ................................ 11-12 Thrust Reversing Limitations .............. 11-13 Oxygen System ................................. 11-13 Autopilot ........................................... 11-14 HF/ADF System ................................ 11-14 Baggage Limitations .......................... 11-14 Baggage Compartment Weight Limitations........................ 11-14 Baggage Compartment Volume Limitations ....................... 11-14

Cessna Citation II Technical Manual Overview This chapter provides a comprehensive listing of operational limitations for the safe operation of the Citation II airplane, its engines, systems, and equipment. Note: The limitations given in this section are for training purposes only. Consult your Pilot’s Operating Handbook for limitations specific to the year, model and serial number of your airplane.

LIMITATIONS 12/99


11-1

Airspeed Limitations Airspeed Maximum Operating Speed MMO (Mach) Above 28,000 Feet Above 30,500 Feet Maximum Operating Speed VMO (Knots) 14,000~28,000 Feet Below 30,500 Feet 14,000~30,500 Feet (11,000 LB ZFW) Below 14,000 Feet

550-0627 and After

550-0626 and Earlier

550-0626 and Earlier

(not incorporating SB550-32-14)

(incorporating SB550-32-14)

N/A 0.705 Mach

0.705 Mach N/A

0.705 Mach N/A

N/A 262 KIAS N/A

277 KIAS N/A 262 KIAS

277 KIAS N/A 262 KIAS

N/A

262 KIAS

262 KIAS

Maneuvering Speed VA

Per Sec II of FAA Approved Airplane Flight Manual

Maximum Flap Extended Speed VFE (Knots) 15° Flaps 40° Flaps

202 KIAS 176 KIAS

202 KIAS 176 KIAS

202 KIAS 176 KIAS

Maximum Landing Gear Operating Speed VLO (Knots) Extend Retract

250 KIAS 200 KIAS

176 KIAS 176 KIAS

250 KIAS 200 KIAS

Maximum Landing Gear Extended Speed VLE (Knots)

262 KIAS

176 KIAS

277 KIAS

Maximum Speed Brake Operating Speed VSB (Knots)

No Limit

No Limit

No Limit

Minimum Controllable Airspeed VMCA (Knots)

77 KIAS

77 KIAS

77 KIAS

Minimum Controllable Ground Speed VMCG (Knots)

62 KIAS

62 KIAS

62 KIAS

Maximum Tire Ground Speed

165 KIAS

165 KIAS

165 KIAS

Autopilot Operation Above 14,000 Feet Below 14,000 Feet Above 30,500 Feet Below 30,500 Feet

N/A N/A 262 KIAS/0.705 Mach 262 KIAS/0.705 Mach

277 KIAS/0.705 Mach 262 KIAS N/A N/A

277 KIAS/0.705 Mach 262 KIAS N/A N/A

11-2



Cessna Citation II Technical Manual Operating Limitations Weight Limitations 550-0550 ~ 550-0626: Maximum Ramp Weight ................... 13,500 LBS Maximum Takeoff Weight ................. 13,300 LBS Maximum Landing Weight ................ 12,700 LBS Maximum Zero Fuel Weight .............. 11,000 LBS (550-0505 and earlier: 9500 LBS – standard, 11,000 LBS optional) 550-0627 and after: Maximum Ramp Weight ................... 14,300 LBS Maximum Takeoff Weight ................. 14,100 LBS Maximum Landing Weight ................ 13,500 LBS Maximum Zero Fuel Weight .............. 11,000 LBS Note: Maximum takeoff and landing weights may be additionally restricted due to altitude, temperature and field length. Center of Gravity Limits 550-0626 and earlier: Forward Limit: 8540 LBS or less ............................... 276.10 inches aft of reference datum. 13,300 LBS or less ............................ 279.80 inches aft of reference datum. 12,500 LBS or less ............................ 279.20 inches aft of reference datum. Aft Limit: ............................................ 285.8 inches aft of reference datum. 550-0627 and after: Forward Limit: 8540 LBS or less ............................... 276.10 inches aft of reference datum. 14,100 LBS or less ............................ 280.40 inches aft of reference datum. Aft Limit: 14,100 LBS or less ............................ 285.80 inches aft of reference datum. Note: It is the responsibility of the pilot to ensure that the airplane is loaded properly. Refer to Weight and Balance Data Sheet for proper loading instructions. LIMITATIONS 12/99 R-6/9/00


11-3

Takeoff and Landing Limitations Maximum Altitude................................... 14,000 Feet Max Tailwind Component (<0626) ......... 10 Knots Crosswind Component (>0627) ............. 23 Knots Maximum Runway Water/Slush Accumulation ..................... 0.4 Inches Maximum Ambient Temperature ........................................... ISA + 39°C (130°F) Minimum Ambient Temperature ........................................... -54°C (-65°F) Note: Autopilot and yaw damper must be OFF for takeoff and landing. Vertical navigation system must be OFF below 500 feet AGL. Flight Load Factor Limitations 550-0626 and earlier at 13,300 LBS maximum takeoff weight Flaps Up ........................................... + 3.8G, -1.52G Flaps Down ....................................... + 2.0G, 0.0G Landing ............................................. + 3.5G 550-0627 and after at 14,100 LBS maximum takeoff weight Flaps Up ........................................... + 3.8G, -1.52G Flaps Down ....................................... + 2.0G, 0.0G Landing ............................................. + 3.38G at 13,500 LB landing weight Note: These accelerations limit the angle-of-bank in turns and severity of pullup maneuvers. Note: This airplane is certificated in the normal category. The normal category is applicable to aircraft intended for non-aerobatic operations. Aerobatic maneuvers and spins are prohibited. No intentional stalls are permitted above 25,000 feet or at any altitude with engine speeds between 61.0% and 65% N1.

11-4



Cessna Citation II Technical Manual Enroute Limitations 550-0626 and earlier not incorporating SB55-54-4: Maximum Operating Altitude ............ 43,000 Feet Temperature Limits ........................... ISA +39°C* Generator Load Up to 25,000 Feet .......................... 400 Amps Above 25,000 Feet ........................ 250 Amps incorporating SB55-54-4: Maximum Operating Altitude ............ 43,000 Feet Temperature Limits ........................... ISA +39°C* Generator Load Up to 35,000 Feet .......................... 400 Amps Above 35,000 Feet ........................ 325 Amps 550-0627 and after: Maximum Operating Altitude ............ 43,000 Feet Temperature Limits ........................... ISA +39°C* Generator Load Up to 35,000 Feet .......................... 400 Amps Above 35,000 Feet ........................ 325 Amps *Note: Maximum enroute operating temperature limit is ISA +39°C ambient temperature adjusted for ram rise or indicated outside air temperature (IOAT), whichever is less. Approved Operations The Citation II is approved for the following types of operation when the required equipment is installed and operational as defined within the Federal Aviation Regulations: 1. 2. 3. 4.

VFR day VFR night IFR day including Category I and Category II approaches IFR night including Category I and Category II approaches

5. Flight into known icing conditions

LIMITATIONS 12/99 R-6/9/00


11-5

Engine Operating Limitations Number of Engines ........................... 2 Engine Manufacturer ........................ Pratt & Whitney Canada, Inc. Engine Model .................................... JT15D-4 Engine Type ...................................... Medium-bypass, axial-flow turbofan Engine Bypass Ratio ......................... 2.7 to 1 Engine Thrust Rating ........................ 2500 LBS each

Thrust Setting

100%=32,760

100%=15,904

N2

N1

Time Limit Minutes

Maximum ITT C

RPM

%

RPM

%

Oil Pressure PSIG (2)

Oil Temp. C

5

700 (4)

31,450

96

16,540

104 (6)

70 - 85 (5)

10 - 121

Maximum Continuous

Continuous

680

31,450

96

16,540

104 (6)

70 - 85

0 - 121

Maximum Cruise

Continuous

670

31,450

96

16,540

104 (6)

70 - 85

0 - 121

Idle

Continuous

580

16,000 (min)

49.0 (3)

---

---

35 (min)

-40 - 121

Starting (6)

---

(1)

---

---

---

---

---

-40 (min)

Transient (<0626) Acceleration (>0627)

-----

700 (4) 700

31,450 31,450

96 96

16,540 16,540

104 104

(5) ---

0 - 121 0 - 121

Takeoff

1. 2.

3. 4. 5. 6.

Maximum ITT limited to 2-seconds during engine start. Normal oil pressure is 70 to 85 PSIG at engine speeds above 60% N2. Oil pressures under 70 PSIG are undesirable, and are allowed only under emergency conditions in order to complete a flight. Oil pressures below 35 PSIG are unsafe and require engine shut down, or landing as soon as possible using minimum power required to sustain flight. Idle turbine RPM is 49, ±0.5% with ignition on. A minimum decrease of 0.5% will be noted with ignition off. ITT indications in excess of 700°C during takeoff or in excess of 680°C for more than 5 minutes require reference to the Engine Maintenance Manual. The maximum transient oil pressure can be 95 PSIG for 90-seconds. Refer to the appropriate thrust setting charts for percent fan RPM (N1) setting.

Engine Fan To ensure accurate fan speed thrust indication, the fan must be inspected for damage prior to each flight.

11-6



Cessna Citation II Technical Manual Battery and Starter Cycle Limitations Whether powered by battery, external power unit, or cross start with generator assist, starter operation is limited to three engine start attempts per 30-minute period, with a minimum 30-second rest period between cycles. Battery cycling is limited to three engine start attempts per hour. Battery Limitation 1. If battery limitation is exceeded, a deep cycle, including a capacity check, must be accomplished to detect possible cell damage. Refer to Chapter 24 of the Maintenance Manual for procedure. 2. Three generator assisted cross starts are equal to one battery start. 3. If an external power unit is used for start, no battery cycle is counted. 4. Use of an external power source with voltage in excess of 28 VDC or current in excess of 1000 amps, may damage the starter. Note: Starting ITT exceeding 500°C should be investigated in accordance with Maintenance Manual. Note: If the BATT O’HEAT (BATT O’TEMP, 550-0627 and after) annunciator illuminates during ground operation, do not take off until after the proper maintenance procedures have been accomplished. Prolonged Ground Operations Continuous engine ground static operation up to and including five minutes at takeoff thrust is limited to ambient temperatures not to exceed ISA + 39°C. Continuous ground operation of the starter-generator above 325 amps is prohibited. Limit ground operation of pitot/static heat to two minutes to preclude damage to the pitot/static heater. Operation in the GND bleed mode at power settings greater than 70% N2 for the right engine is prohibited.

LIMITATIONS 12/99


11-7

Oil Limitations Approved Oils Mobile Jet Oil II or 254, Exxon Turbo Oil 2380, Aeroshell Turbine 560 or 500, Castrol 5000, and Royco Turbine Oil 560 or 500. In addition, The engine should be serviced with approved synthetic oils listed in the most current revision of P&WC SB 7001. Caution: When changing from an existing lubricant formulation to a “third generation” lubricant formulation (Aero Shell/Royco Turbine Oil 560 or Mobile Jet 254) the engine manufacturer strongly recommends that such a change should only be made when an engine is new or freshly overhauled. For additional information on use of third generation oils, refer to the engine manufacturers pertinent oil service bulletins. Note: Do not mix types or brands of oil. Should it be necessary to replenish oil consumption loss when oil of the same brand (as contents in tank) is unavailable, then the following requirements apply: 1. The total quantity of added oil does not exceed two US quarts in any 400-hour period. 2. If it is required to add more than two US quarts of dissimilar oil brands, drain and flush complete oil system and refill with an approved oil in accordance with Engine Maintenance Manual instructions. Should oils of non-approved brands or of different viscosities become intermixed, drain and flush complete oil system and refill with an approved oil in accordance with Engine Maintenance Manual instructions. Note: Minimum starting oil temperature is -40°C.

11-8



Cessna Citation II Technical Manual Fuel Limitations Approved Fuels The following approved fuels comply with the latest revision of Pratt & Whitney Canada Specification 204 and Pratt & Whitney Canada Service Bulletin 7144R14. Approved fuels are JET A, JET A-1, JET B, JP-4, JP-5, or JP-8, all with 0.15% PFA55MB anti-icing additive in solution. When preblended fuel is not available, anti-icing additives conforming to MIL-I-27686E (Ethylene Glycol Monomethyl Ether (EGME)) or MIL-I-85470 (Diethylene Glycol Monomethyl Ether (DIEGME)) specifications such as “Prist” may be introduced directly into the nozzle fuel stream during servicing. Concentrations of less than 0.06% (20 fluid ounces of additive per 260 gallons of fuel or more) may be insufficient to prevent fuel system icing or microbiological contamination. Conversely, concentrations of more than 0.15% (20 fluid ounces of additive per 104 gallons of fuel or less) could cause damage to internal components of the fuel system or erroneous fuel quantity indications. Caution: EGME and DIEGME are aggressive chemicals and should not exceed 0.15% of fuel volume. Improperly handled, these materials will damage the epoxy primer and sealants used in the fuel tanks, O-ring seals, and any part of the airplane’s exterior finish with which it comes in contact. Warning: Anti-icing additives containing EGME or DIEGME are harmful if inhaled, swallowed, or absorbed through the skin, and will cause eye irritation. Refer to all instructions and warnings regarding toxicity and flammability before using these materials. All grades of aviation gasoline (AVGAS) conforming to MIL-G5572 specifications are approved for use under emergency circumstances only. If used during flight, boost pumps should be activated and airplane altitude should not exceed 18,000 feet. Use of AVGAS is limited to no more than 3500 US gallons or 50 hours of engine operation during any period between engine overhaul. For record keeping purposes, 1 hour of engine operation may be considered equivalent to 70 US gallons.

LIMITATIONS 12/99


11-9

Fuel Temperature and Density Limitations Approved Fuel Types Jet A, A-1, A-2, JP-5, -8

Jet B, JP-4

Aviation Gasoline

Minimum Fuel Temperature (Takeoff) (Starting)

-40°C -40°C

-54°C -54°C

-54°C -54°C

Maximum Fuel Temperature

+50°C

+50°C

+32°C

Maximum Altitude

43,000’

43,000’

18,000’

0.81

0.79

0.73

Fuel Control Density (Adjustment for Optimum Engine Acceleration)

Maximum Fuel Imbalance Maintaining fuel load symmetry during servicing is unnecessary; however, the maximum permissible asymmetry is 200 LBS during normal flight operations and 600 LBS in an emergency. Hydraulic Fluid Limitations The only approved hydraulic fluids are Skydrol 500A, B, B-4, C, or LD-4 or Hyjet W, Hyjet III, IV, or IVA.

11-10



Cessna Citation II Technical Manual Flight Crew Limitations Minimum flight crew required for Category I operations is one pilot who holds a C-500 type rating and who satisfies requirements of FAR 61.58 for two-pilot operation, and one copilot who holds a multi-engine rating and satisfies requirements of FAR 61.55. Category II operation requires a pilot and copilot who both satisfy requirements of FAR 61.3. Cabin Limitations For takeoff and landing, all seats must be upright and outboard. The seat adjacent to the emergency exit must be fully tracked toward the rear of the airplane to ensure unobstructed access to the emergency exit. To meet smoke detection criteria, the cabin (OVHD) fan must be operating any time the aft cabin privacy curtain is closed. If the fan is inoperable, the curtain must remain open unless the toilet is in use. Pressurization Differential Normal (both valves) .............................. 0.0 to 8.8 PSI ±0.1 PSI Pressurization Source Selector On airplanes 550-0481 and earlier, 0483 and 0484, operation in BOTH HI mode is not approved for takeoff, landing or at high power settings.

LIMITATIONS 12/99


11-11

Icing Limitations All anti-ice systems must be activated when operating in visible moisture at indicated outside air temperatures (IOAT) between +4°C (39°F) and -30°C (-22°F). The surface deice system should be activated when ice accumulations of at 1/4” to 1/2” are observed on the leading edge of either outboard wing. Activation of the system with accumulations of less than 1/4” may result in ice bridging on the wing. Accumulations greater than 1/4" may exceed the system's ice removal capabilities. Operation and/or testing of the system at IOAT below -40°C (-40°F) may result in boot cracking or failure of the boots to fully deflate. The aircraft must be clear of all deposits of snow, ice, and frost adhering to the lifting and control surfaces immediately prior to takeoff. Prolonged flight in severe icing conditions should be avoided as this may exceed the capabilities of the aircraft ice protection systems. Note: Isopropyl alcohol conforming to TT-I-735 should be used for windshield ice protection.

11-12



Cessna Citation II Technical Manual Thrust Reversing Limitations During landing roll, reverse thrust power must be reduced to idle (thrust reverser levers at the idle reverse detent position) when airplane speed reaches 60 KIAS. Maximum reverse thrust is limited to 94% N1 at ambient temperatures above -18°C or 92% N1 at ambient temperatures below -18°C. Maximum allowable thrust reverser deployed time is 15 minutes in any 1-hour period. Deployment of thrust reversers is prohibited when the aircraft is operating on sod, dirt, or gravel runways. The drag chute may not be released while thrust reversers are deployed. Oxygen System The standard diluter demand oxygen mask qualifies as a quick-donning mask only if it is positioned around the neck. The optional crew oxygen mask is a sweep-on diluter demand mask with selectable pressure breathing. The sweep-on mask qualifies as a quick-donning mask only if it is properly stowed. Note: Headsets, eyeglasses or hats worn by the crew may interfere with the quick-donning capabilities of the optional oxygen masks.

LIMITATIONS 12/99


11-13

Autopilot During autopilot operation, either the pilot or copilot must be seated in the flight compartment with seat belt fastened. The autopilot torque monitor must be functionally tested; if torque monitor functional test is not successful and/or if the [AP TORQUE] annunciator does not illuminate, autopilot operation is prohibited above 14,500 feet. Continued autopilot operation is prohibited following abnormal operation or malfunctioning prior to corrective maintenance. HF/ADF System The ADF bearing information may be erratic when keying the HF transmitter. Should this occur, disregard the ADF bearing during periods of transmission. Baggage Limitations Baggage Compartment Weight Limitations Maximum nose baggage compartment load .................................. 350 LBS Maximum cabin baggage compartment load .................................. 400 LBS Maximum tailcone baggage compartment load .................................. 200 LBS Baggage Compartment Volume Limitations Maximum nose baggage compartment volume ............................. 17 cubic feet Maximum cabin baggage compartment volume ............................. 34 cubic feet Maximum tailcone baggage compartment volume ............................. 13 cubic feet

11-14




Chapter 12 Emergency Procedures Table of Contents Overview ............................................. 12-1 Engine Emergency Procedures .............. 12-2 Engine Fire ......................................... 12-9 Inadvertent Thrust Reverser Deployment ...................................... 12-10 Electrical System Abnormal Procedures (550-0626 and earlier) ..... 12-13 Loss of Both Generators ................... 12-16 Battery Overheat ............................... 12-18 AC Power Failure ............................... 12-20 Autopilot Hardover ............................. 12-21 Environmental System Abnormal Procedures......................... 12-22 Emergency Descent ........................... 12-23 Spins ............................................... 12-26 Ditching ............................................ 12-27 Forced Landing ................................. 12-28 Electrical System Abnormal Procedures (550-0627 and after) ....... 12-29 Battery Overheat ............................... 12-32 Loss of Both Generators ................... 12-36 AC Power Failure ............................... 12-38 Environmental System Abnormal Procedures......................... 12-41 Emergency Descent ........................... 12-43 Autopilot Hardover ............................. 12-45

Cessna Citation II Technical Manual Overview This chapter provides procedures recommended by the manufacturer for various emergency situations that may be encountered during operation of the Cessna Citation II. The procedures are per the Operating Manual(s) for the corresponding serial number range(s) of airplanes, however, their use in this manual is for training purposes only. The appropriate Operating Manual, FAA approved Airplane Flight Manual (AFM), Pilot’s Check List, and/or related publications should be refered to for normal, abnormal, and emergency procedures specific to your airplane. A thorough understanding of the airplane’s systems and their interrelationships is essential to successfully respond to emergency situations. It is suggested that the corresponding chapter(s) in this manual be referenced for specific systems descriptions. Immediate action or “commit to memory” items for each procedure are identified by bold type within a shaded box. Notes, Cautions, and Warnings are provided to amplify the procedures.

EMERGENCY PROCEDURES 12/99


12-1

Engine Emergency Procedures Engine Failure or Fire During Takeoff (without thrust reversers) SPEED BELOW V1 – TAKEOFF SHOULD NORMALLY BE ABORTED: 1. Brakes .......................................... AS REQUIRED 2. Throttles ....................................... IDLE 3. Speed Brakes................................ EXTEND IF ENGINE FIRE: 4. Accomplish Engine Fire Procedures. IF ENGINE FAILURE: 4. Accomplish Engine Failure/Precautionary Shutdown Procedure. Note: To obtain maximum braking performance from the anti skid system, it is required that the pilot apply continuous maximum effort (no modulation) to the brake pedals. Note: The takeoff field lengths assume that the pilot has maximum effort applied to the brakes at the scheduled V1 speed during the aborted takeoff. SPEED ABOVE V1 – TAKEOFF SHOULD NORMALLY BE CONTINUED: 1. After establishing a positive rate of climb, retract landing gear (and climb at V2, airplanes 550-0626 and earlier). 2. At 400 feet, retract the flaps at V2 + 10 and accelerate to VENR. IF ENGINE FIRE: 3. Accomplish Engine Fire Procedures. IF ENGINE FAILURE: 3. Accomplish Engine Failure/Precautionary Shutdown Procedure.

12-2



Cessna Citation II Technical Manual Engine Failure or Fire During Takeoff (with thrust reversers) SPEED BELOW V1 – TAKEOFF SHOULD NORMALLY BE ABORTED: 1. 2. 3. 4.

Brakes .......................................... Throttles ....................................... Speed Brakes................................ Thrust Reverser .............................

AS REQUIRED IDLE EXTEND DEPLOY on unaffected engine.

5. Reverser Indicator Lights .................CHECK illumination of ARM, UNLOCK, and DEPLOY lights. 6. Thrust Reverser ................................REVERSE power on the unaffected engine. IF ENGINE FIRE: 7. Accomplish Engine Fire Procedures. IF ENGINE FAILURE: 7. Accomplish Engine Failure/Precautionary Shutdown Procedure. Note: To obtain maximum braking performance from the antiskid system, it is required that the pilot apply continuous maximum effort (no modulation) to the brake pedals. Note: The takeoff field lengths assume that the pilot has maximum effort applied to the brakes at the scheduled V1 speed during the aborted takeoff. SPEED ABOVE V1 – TAKEOFF SHOULD NORMALLY BE CONTINUED: (550-0627 and after) 1. After establishing a positive rate of climb, retract landing gear. 2. At 400 feet, retract the flaps at V2 + 10 and accelerate to VENR. IF ENGINE FIRE: 3. Accomplish Engine Fire Procedures. IF ENGINE FAILURE: 3. Accomplish Engine Failure/Precautionary Shutdown Procedure. EMERGENCY PROCEDURES 12/99


12-3

Engine Failure/Precautionary Shutdown 1. Throttle (affected engine) .................OFF Any one or more of the following indications might suggest a precautionary shutdown: abnormal or rising ITT, engine vibration, fluctuating or abnormally high or low N1 or N2, abnormal oil pressure or temperature, or erratic fuel flow. Circumstances will normally dictate whether to continue to operate the engine with possible further damage or shut it down. 2. Ignition (affected engine) .................OFF 3. Engine Synchronizer ........................OFF 4. Generator (affected engine) ............OFF 5. Electrical Load .................................REDUCE as required. Airplanes 550-0114 and earlier, 0127, 0296 and after, and 0115~0126 and 0128~0295 incorporating SB55054-4: 400 amps maximum up to 35,000 feet, 325 amps maximum above 35,000 feet. Airplanes 550-0115~0126 and 0128~0295 not incorporating SB550-54-4: 400 amps maximum up to 25,000 feet, 250 amps maximum above 25,000 feet. 6. Fuel Crossfeed .................................AS REQUIRED Do not exceed asymmetric fuel load of 600 LBS (550-0626 and earlier) or 200 LBS (550-0627 and after). IF NO FIRE: 7. Firewall Shutoff .................................OPEN 8. Fuel Boost Pump ..............................ON Note: If no fire hazard exists, leave firewall shutoff open and turn boost pump ON to prevent damage to engine fuel pump. If engine windmills with firewall shutoff CLOSED or with no indication of oil pressure, refer to engine maintenance manual. 9. Refer to Emergency Restart, One Engine or Single Engine Approach and Landing Procedures.

12-4



Cessna Citation II Technical Manual Engine Failure During Coupled Approach 1. Power (operating engine) .............. INCREASE as required. Only a small power increase will be required to maintain approach speed and correct rate of descent. 2. Autopilot and Yaw Damper ............ OFF 3. Airspeed ....................................... VREF + 10 KIAS Accelerate to VREF + 10 before raising flaps. 4. Rudder Trim .................................. TRIM toward operating engine. The yaw change will be relatively small since the operating engine is at an approach power setting. 5. Flaps ........................................... T.O. & APPR 6. Throttle (affected engine) .................OFF IF ENGINE FIRE: 7. Accomplish Engine Fire Procedure. 8. Passenger Advisory Lights ..............PASS SAFETY 9. Passenger Seats ..............................CHECK full upright, outboard and positioned aft or forward to clear exit doors. 10. Seats, Seat Belts and Shoulder Harnesses ........................SECURE Check Seats locked in the desired position. Check seat belts snug and shoulder harnesses latched to the buckle. 11. Fuel Crossfeed .................................CHECK 12. Ignition (operating engine) ...............ON 13. Landing Gear ...................................DOWN and LOCKED 14. Anti-Skid...........................................CHECK ON 15. Annunciator Panel ............................CHECK With one engine shut down by the throttle, the appropriate [OIL PRESS WARN] and [HYD FLOW LOW] annunciators (550-0550 and after), or [OIL PRESS LO] and [HYD PRESS LOW] annunciators (550-0505 and earlier), and [GEN OFF] annunciator will be illuminated. If low fuel pressure causes automatic boost pump activation prior to shut down by the throttle, the appropriate [FUEL BOOST ON] annunciator will also be illuminated. If automatic boost pump activation does not occur prior to shut down by the throttle, the appropriate [FUEL LOW PRESS] (550-0550 and after) or [FUEL PRESS LO] (5500505 and earlier) annunciator will be illuminated instead. EMERGENCY PROCEDURES 12/99


12-5

With one engine shut down by the firewall shutoff switch, the appropriate [OIL PRESS WARN], [HYD FLOW LOW], and [FUEL LOW PRESS] annunciators (550-0550 and after), or [OIL PRESS LO], [HYD PRESS LOW], and [FUEL PRESS LO] annunciators (550-0505 and earlier), and the [F/W SHUTOFF] and [GEN OFF] annunciators will be illuminated. If the boost pump is automatically activated by low fuel pressure, the appropriate [FUEL BOOST ON] annunciator will also be illuminated. Should this occur, the corresponding FUEL BOOST switch should be positioned to OFF. If the [MASTER WARNING] light is flashing, it should be extinguished to reduce distraction. 16. Flaps ...............................................LAND (when landing assured) At the pilot’s discretion, flaps may be left at T.O. & APPR or lowered to LAND. If T.O. & APPR flaps are used, maintain VREF + 10 KIAS or “on speed” angle-of-attack (if optional AOA indicator installed). LAND flaps are used under most conditions since little pitch change is encountered when they are selected and touchdown speed can be reduced. 17. Airspeed ..........................................VREF 18. Pressurization ..................................CHECK ZERO DIFFERENTIAL Passing approximately 500 feet above ground level (AGL), check the cabin differential pressure near zero. If it is in excess of about one half PSI, select a higher cabin altitude and adjust RATE to ascend the cabin. Differential pressure should be at zero for landing. Any pressure existing at touchdown will be dumped by the outflow valves (actuated by the left main gear squat switch) and may cause discomfort. If landing above 12,000 feet pressure altitude, turn the OXYGEN CONTROL VALVE to CREW ONLY and turn pressurization bleed air OFF to preclude passenger mask deployment. 19. Speed brakes ..................................RETRACTED prior to 50 feet. Note: Do not allow Turbine RPM (N2) to be less than 49%.

12-6



Cessna Citation II Technical Manual Emergency Restart – One Engine FOLLOWING SHUTDOWN, WITH STARTER ASSIST: 1. Throttle .............................................OFF 2. Generator .........................................GEN 3. Firewall Shutoff .................................CHECK OPEN 4. Ignition .............................................ON 5. Start Button ......................................PRESS momentarily. Generator cross start is disabled with weight off the left main gear squat switch to preclude generator damage from excessive N2 on the operating engine. 6. Throttle .............................................IDLE at 8-10% N2 7. Engine Instruments ..........................MONITOR Maximum start ITT 700°C for two seconds. 8. Ignition .............................................NORM 9. If Start Does Not Occur ....................PRESS starter disengage switch. FOLLOWING SHUTDOWN – WINDMILLING WITH AIRSPEED ABOVE 200 KIAS: 1. Throttle .............................................OFF 2. Firewall Shutoff .................................CHECK OPEN 3. Ignition .............................................ON 4. Boost Pump .....................................ON Associated engine ignition and boost pump switches must be selected ON since automatic sequencing and selection of these functions does not occur when the start button is not utilized. 5. Throttle .............................................IDLE With airspeed maintained above 200 KIAS, throttle should be brought to IDLE. An N2 of 8-10% is not required. 6. Engine Instruments ..........................MONITOR After engine stabilizes: 7. Boost Pump and Ignition ..................NORM It may be necessary to select the associated generator RESET position momentarily to reinstate the generator following a windmilling airstart. Maximum start ITT 700°C for two seconds. 8. Generator .........................................GEN



12-7

Emergency Restart – Two Engines 1. Ignition ......................................... BOTH ON 2. Boost Pumps ................................ BOTH ON Engine ignition and boost pump switches must be selected ON since automatic sequencing and selection of these functions does not occur when the start button is not utilized. 3. Throttles ....................................... IDLE Throttles remain at idle for attempted immediate light-off. 4. If Altitude Allows ........................... INCREASE AIRSPEED to 200 KIAS. Possibilities of immediate start are increased if airspeed is over 200 KIAS. 5. Firewall Shutoff .................................CHECK OPEN 6. All Anti-ice Switches ........................OFF They are turned OFF to minimize engine bleed air loss. IF NO START IN TEN SECONDS: 7. Either Start Button ............................PRESS momentarily. Attempt a starter assist restart if altitude and time permit. Maximum Glide – Emergency Landing (550-0626 and earlier) 1. Airspeed ..........................................BEST GLIDE AT 9500 LBS - 120 KIAS. Increase speed 3 KIAS for each 500 LB increase in weight. 2. Flaps ...............................................UP 3. Speed Brakes ..................................RETRACT 4. Landing Gear ...................................UP 5. Transponder ....................................EMERGENCY 6. ATC

...............................................ADVISE

7. Passenger Advisory Switch .............PASS SAFETY 8. Shoulder Harness ............................SECURE 9. Landing Gear, Speed Brakes, and Flaps .........................................AS REQUIRED for landing anticipated.

12-8



Cessna Citation II Technical Manual Engine Fire (engine fire annunciator/switch illuminated) 1. Throttle (affected engine) .............. IDLE IF ANNUNCIATOR REMAINS ILLUMINATED: 2. Engine Fire Annunciator/Switch ...................... LIFT COVER and PUSH. Cuts off fuel to engine, hydraulic fluid supply to enginedriven pump, trips the generator field, positions a valve to allow both bottles to be fired into the affected engine and illuminates the bottle armed lights. 3. Either Illuminated Bottle Armed Annunciator/Switch ........... PUSH 4. Ignition .............................................NORM If ignition is ON, return the switch to NORM. 5. Throttle (affected engine) .................OFF 6. Reduce Electrical Load ....................AS REQUIRED Airplanes 550-0114 and earlier, 0127, 0296 and after, and 0115~0126 and 0128~0295 incorporating SB55054-4: 400 amps maximum up to 35,000 feet, 325 amps maximum above 35,000 feet. Airplanes 550-0115~0126 and 0128~0295 not incorporating SB550-54-4: 400 amps maximum up to 25,000 feet, 250 amps maximum above 25,000 feet. 7. Boost Pump .....................................OFF If pump is ON, return the switch to OFF. IF FIRE WARNING ANNUNCIATOR ILLUMINATED AFTER 30-SECONDS: 8. Remaining Illuminated Bottle Armed Annunciator/Switch ..............PUSH 9. Land as soon as possible. IF ANNUNCIATOR EXTINGUISHED AND SECONDARY INDICATIONS ARE NOT PRESENT: 2. Land as soon as possible.



12-9

Inadvertent Thrust Reverser Deployment Inadvertent Deployment of Thrust Reversers During Takeoff SPEED BELOW V1 – TAKEOFF SHOULD BE ABORTED: 1. Brakes .......................................... AS REQUIRED 2. Throttles ....................................... IDLE 3. Speed Brakes................................ EXTEND 4. Thrust Reversers ........................... BOTH DEPLOY 5. Reverser Indicator Annunciators .....CHECK illumination of ARM, UNLOCK and DEPLOY annunciators. 6. Thrust Reversers ..............................REVERSE power on both engines. SPEED ABOVE V1 – TAKEOFF SHOULD NORMALLY BE CONTINUED: 1. Emergency Stow Switch ................ ACTUATE on affected engine. 2. After establishing a positive rate-of-climb, retract landing gear. Do not exceed 125 KIAS until thrust reverser stows. 3. At 400 feet, retract flaps at V2 + 10 and accelerate. Do not exceed 200 KIAS after thrust reverser stows. 4. Land as soon as practical. IF THRUST REVERSER WILL NOT STOW: 5. Thrust Reverser Circuit Breaker .......CHECK in 6. Throttle (affected engine) .................CUTOFF 7. Airspeed ..........................................MAINTAIN 150 KIAS or below. 8. Refer to Abnormal Procedures for landing.

12-10



Cessna Citation II Technical Manual Inadvertent Deployment of Thrust Reversers In Flight 1. Reverser Indicator Annunciators .... CHECK illumination of ARM, UNLOCK and DEPLOY annunciators. 2. Affected Throttle ........................... CHECK idle 3. Emergency Stow Switch ................ ACTUATE on affected engine. 4. Airspeed ....................................... REDUCE to 125 KIAS (115 KIAS with flaps extended) or below. After thrust reverser stows, do not exceed 200 KIAS. 5. Reverser Indicator Annunciators: UNLOCK and DEPLOY Annunciator ......................................EXTINGUISHED ARM and HYD PRESS ON Annunciator ......................................ILLUMINATED Note: If thrust reverser is stowed, engine may be operated normally. Thrust reverser cannot be used during landing if emergency stowed. 6. Land as soon as practical. IF THRUST REVERSER WILL NOT STOW: 7. Thrust Reverser Circuit Breaker .......CHECK in 8. Throttle (affected engine) .................CUTOFF 9. Airspeed ..........................................MAINTAIN 150 KIAS or below. 10. Refer to Single Engine Approach and Landing Procedures.



12-11

Thrust Reverser UNLOCK Annunciator Illuminated In Flight 1. Emergency Stow Switch ................ ACTUATE on affected engine. 2. Thrust Reverser Levers .................. CHECK thrust reverser levers at stowed (full forward) position. IF ANNUNCIATORS WILL NOT EXTINGUISH: 3. Thrust Reverser Circuit Breaker .......CHECK in 4. Maintain 200 KIAS or below. 5. Land as soon as practical. Thrust Reverser ARM Annunciator Illuminated In Flight (550-0626 and earlier) 1. Thrust Reverser Levers .................... CHECK stowed (full forward) 2. Emergency Stow Switch ..................VERIFY OFF IF ARM ANNUNCIATOR STILL ILLUMINATED: 3. Airspeed ..........................................MAINTAIN 200 KIAS or below. 4. HYD PRESS ON Annunciator ...........CHECK IF HYD PRESS ON ANNUNCIATOR ILLUMINATED (T/R ISOLATION VALVE IS OPEN): 5. Affected Thrust Reverser Circuit Breaker .................................PULL 6. Land as soon as possible (affected T/R should be inoperative). IF HYD PRESS ON ANNUNCIATOR NOT ILLUMINATED (PROBABLE PRESSURE SWITCH PROBLEMS): 5. Land as soon as practical. Note: With a thrust reverser circuit breaker pulled, the emergency stow system of the opposite reverser is deactivated.

12-12



Cessna Citation II Technical Manual Electrical System Abnormal Procedures (550-0626 and earlier) Electrical Fire or Smoke 1. Oxygen Masks and Oxygen MIC Switches ................... AS REQUIRED. Ensure selector is on 100% oxygen when masks are used. Ensure oxygen MIC switch is in MIC OXY MASK position. 2. Pressurization Source Selector: .......NORM (airplanes 550-0481, and 550-0485~0626) Pressurization Source Selector: .......BOTH HI (airplanes 550-0481 and earlier, 550-0483, and 550-0484) KNOWN SOURCE OF FIRE: 3. Isolate faulty circuit. Pull circuit breaker to remove power from faulty equipment. UNKNOWN SOURCE OF FIRE: 3. Flood Lights .....................................FULL BRIGHT 4. Battery Switch ..................................EMER Will have COMM 1, NAV 2, copilot’s HSI and cockpit floodlights after generators turned off. 5. Generators .......................................OFF With the battery switch in the emergency position and the generators off, power is supplied for approximately 30-minutes to COMM 1, NAV 2, overhead floodlights and copilot’s HSI.



12-13

Caution: When landing with emergency power (Battery SwitchEMER and both generators off), the following are not available: a. The landing gear normal operation is not available. The landing gear must be lowered using the blowdown system and the landing gear warning lights will not illuminate. b. The flaps will not operate. If not previously in landing position, a flaps inoperative landing must be made. c. The anti-skid/power brake system is inoperative. Only the emergency brake system is available. d. The engine anti-ice valves will be open. Refer to anti-ice on thrust charts. e. The outside air temperature gage is not reliable, so use caution when applying power (except for go-around where ground temperatures can be used). f. All engine instruments except the vertical tape N1 will be inoperative. The vertical tape N1 will indicate erratically below approximately 50% N1, but will give reliable indications above 50% N1. 6. MIC Selector ....................................EMER/COMM 1 Must be in the EMER/COMM 1 position to transmit when operating on emergency battery, and pilot must wear headset to receive. 7. Receiver Select ................................COMM 1 to HDPH (550-0356~0626) (required only if AUTO SELECT is OFF) 8. All Electrical Switches ......................OFF 9. Windshield Bleed Air Manual Valves ..................................OFF With electrical power lost, the windshield bleed air shutoff valve will fail open. The bleed air manual valves are closed to prevent an excessive volume of high temperature air from reaching the windshield. 10. DC Power RH Bus ............................PULL No. 1,2,3 CB’s 11. RH CB Panel Circuit Breaker (LH Panel) ........................................PULL 12. AC Inverter No. 1 Circuit Breaker (LH Panel) ........................................PULL 13. Land as soon as practical (within 30-minutes).

12-14



Cessna Citation II Technical Manual IF SEVERITY OF SMOKE WARRANTS: 13. Initiate Smoke Removal and/or Emergency Descent procedures. Land as soon as possible. WHEN LANDING ASSURED: 14. LH Generator ...................................ON 15. Landing Gear ...................................DOWN 16. Flaps ...............................................LAND 17. Airspeed ..........................................VREF Note: Right thrust reverser (if applicable) will be inoperative. IF FIRE OR SMOKE STARTS AGAIN: 18. LH Generator ...................................OFF Note: Anti-skid system will be inoperative. Power brakes will be available until accumulator discharges. Multiply landing distance by 1.6. Be prepared to use the emergency brake system.



12-15

Loss of Both Generators 1. Generators .......................................RESET then GEN Attempt to reset both generators. IF ONLY ONE GENERATOR COMES ON: 2. Electrical Load .................................REDUCE as required. Airplanes 550-0114 and earlier, 0127, 0296 and after, and 0115~0126 and 0128~0295 incorporating SB55054-4: 400 amps maximum up to 35,000 feet, 325 amps maximum above 35,000 feet. Airplanes 550-0115~0126 and 0128~0295 not incorporating SB550-54-4: 400 amps maximum up to 25,000 feet, 250 amps maximum above 25,000 feet. IF NEITHER GENERATOR COMES ON: 2. Flood Lights .....................................FULL BRIGHT 3. Battery Switch ..................................EMER With the battery switch in emergency position and the generators off, power is supplied for approximately 30minutes to COMM 1, NAV 2, overhead floodlights, voltmeter and copilot’s HSI. On airplanes with single EFIS, with the battery switch in the emergency position and the generators off, power is supplied for approximately 30-minutes to COMM 1, NAV 2, Copilot’s HSI and DG 2. and copilot’s audio panels, copilot’s attitude indicator, voltmeter, cockpit floodlights and standby gyro. Airplanes 550-0682 and after also have the RH pitot static heater on the emergency bus. On airplanes with dual EFIS, with the battery switch in emergency position and the generators off, power is supplied for approximately 30-minutes to COMM1, NAV 2, copilot’s RMI, NAV 2 repeater indicator, pilot’s and copilot’s audio panels, voltmeter, cockpit floodlights and standby gyro. Airplanes 550-0682 and after also have the RH pitot static heater on the emergency bus.

12-16



Cessna Citation II Technical Manual Caution: When landing with emergency power (Battery SwitchEMER and both generators off), the following are not available: a. The landing gear normal operation is not available. The landing gear must be lowered using the blowdown system and the landing gear warning lights will not illuminate. b. The flaps will not operate. If not previously in the landing position, a flap inoperative landing must be made. c. The Anti-skid/power brake system is inoperative. Only the emergency brake system is available. d. The engine anti-ice valves will be open. Refer to anti-ice on thrust charts. e. The outside air temperature gage is not reliable, so use caution when applying power (except for go-around where ground temperatures can be used). f. All engine instruments except the vertical tape N1 will be inoperative. The vertical tape N1 will indicate erratically below approximately 50% N1, but will give reliable indications above 50% N1. 4. MIC Selector ....................................EMER/COMM 1 Must be in EMER/COMM 1 position to transmit while operating on emergency battery and headset worn to receive. Volume is controlled at the radio since normal amplification is bypassed. 5. Receiver Select ................................COMM 1 to HDPH (550-0356 and after) Required only if Auto-Select is OFF 6. Windshield Bleed Air Manual Valves ..................................OFF With electrical power lost, the windshield bleed air shutoff valve will fail open. The bleed air manual valves are closed to prevent an excessive volume of high temperature air from reaching the windshield. 7. Land as soon as practical. WHEN LANDING ASSURED: 7. Battery Switch ..................................BATT 8. Landing Gear ...................................DOWN 9. Flaps ...............................................LAND 10. Airspeed ..........................................VREF EMERGENCY PROCEDURES 12/99


12-17

Battery Overheat BATT O’TEMP Annunciator Illuminated (temp between 145° and 160°F) 1. Battery Switch .............................. EMER In EMER position the battery will be disconnected from the generators and will no longer be charged by them. All electrical equipment will continue to receive power since the generators are still on line. The Emergency DC bus is powered by the battery. Battery voltage may now be read with the voltage selector in BATT and generator bus voltage with the voltage selector in LH GEN or RH GEN. Individual generator voltages can be read by selecting one (LH or RH) GEN and turning the other generator off. 2. Amperage ..................................... NOTE decrease 3. If battery voltage is 1 volt less than generator voltage in 30seconds to 2-minutes, monitor battery overheat annunciator for possible change. In thirty seconds to two minutes after disconnect, battery voltage should read at least one volt less than the generators. Rotate the voltage selector to LH GEN and RH GEN position to read generator voltage. Battery voltage will be indicated when the voltage selector is in the BATT position. IF BATT O’TEMP (BATT O’HEAT, 550-0505 and earlier) ANNUNCIATOR EXTINGUISHED: 4. Battery Switch ..................................BATT IF NO AMP DECREASE OR BATT O’TEMP (BATT O’HEAT, 550-0505 and earlier) ANNUNCIATOR FLASHES: 4. Flood Lights .....................................FULL BRIGHT 5. Generators .......................................OFF Since the battery has continued to overheat, it may be because the battery is still being charged through a failed battery relay. Turning both generators OFF trips the generators and opens the power relays, isolating each generator from its bus. The BATT O’TEMP annunciator will extinguish immediately when the generators are turned off if the battery relay is not stuck. With the battery in the emergency position and the generators off, power is supplied for approximately 30-minutes to COMM 1, NAV 2, overhead floodlights and copilot’s HSI. 12-18



Cessna Citation II Technical Manual Caution: When landing with emergency power (Battery SwitchEMER and both generators off), the following are not available: a. The landing gear normal operation is not available. The landing gear must be lowered using the blowdown system and the landing gear warning lights will not illuminate. b. The flaps will not operate. If not previously in the landing position, a flap inoperative landing must be made. c. The Anti-skid/power brake system is inoperative. Only the emergency brake system is available. d. The engine anti-ice valves will be open. Refer to anti-ice on thrust charts. e. The outside air temperature gage is not reliable, so use caution when applying power (except for go-around where ground temperatures can be used). f. All engine instruments except the vertical tape N1 will be inoperative. The vertical tape N1 will indicate erratically below approximately 50% N1, but will give reliable indications above 50% N1. IF NORMAL DC POWER LOST(BATTERY RELAY NOT STUCK): 6. Generators .......................................GEN (BATT O’TEMP annunciator will illuminate until battery cools). 7. Battery Switch ..................................OFF Caution: With the battery switch off, all power from the emergency bus will be removed. After landing, refer to maintenance manual for proper maintenance procedures as damage to the battery may have occurred. 8. Land as soon as practical IF NO DC POWER LOST (BATTERY RELAY STUCK): 6. Mic Selector .....................................EMER/COMM 1 (headphones required to receive audio) 7. Receiver Select ................................COMM 1 to HDPH (required only if Auto Select is OFF) (550-0356 and after), 8. Windshield Bleed Air Manual Valves ..................................OFF 9. DC Power LH and RH Bus ...............PULL No. 1,2,3 CB’s 10. Land as soon as practical. EMERGENCY PROCEDURES 12/99


12-19

WHEN LANDING ASSURED: 11. DC Power LH and RH Circuit Breakers ...............................RESET 12. Landing Gear ...................................DOWN 13. Flaps ...............................................LAND 14. Airspeed ..........................................VREF Caution: After landing, refer to maintenance manual for proper maintenance procedures as damage to the battery may have occurred. AC Power Failure Both INVERTER 1 FAIL and INVERTER 2 FAIL – Annunciators Illuminated (Airplanes with dual AC Busses – 550-0550~0626): 1. Inverter 1 and Inverter 2 Circuit Breakers ...............................RESET 2. Battery Switch ..................................EMER If the inverters will not come back on line after the circuit breakers have been reset, complete the flight by using the copilot’s air-driven attitude indicator or the standby gyro horizon (if installed). Placing the battery switch to EMER will provide AC power from the copilot’s C-14D static inverter to power the copilot’s compass system and NAV 2. AC Power and/or Distribution Failure (AC FAIL Annunciator Illuminated After MASTER WARNING Has Been Reset, INVERTER FAIL 1 or 2 Annunciators Extinguished – Airplanes 550-0550~0626): 1. Check the right sub-circuit breaker panel for disengaged AC BUS circuit breaker(s)

12-20



Cessna Citation II Technical Manual Caution: If circuit breaker(s) is/are disengaged, operate with loss of bus as re-engagement may result in further damage to the electrical system. Note: Depending on which buses have been lost, the following equipment will be lost: a. 26 VAC Bus 1: pilot’s ADI, pilot’s RMI, pilot’s HSI (HDG and NAV flags in view), pilot’s rate-of-turn indicator, copilot’s RMI compass card and NAV 1 bearing pointer, and ADF 1. b. 115 VAC Bus 1: pilot’s flight director, autopilot, yaw damper, radar, pilot’s attitude gyro, and VNAV computer/ controller. c. 26 VAC Bus 2: NAV 2, copilot’s ADI (optional dual flight director installation), copilot’s HSI, copilots RMI, pilot’s RMI compass card and NAV 2 bearing pointer, and ADF 2 (optional). Operation of the following equipment can be reinstated by placing the BATT switch to EMER: NAV 2, bearing pointer and DG of copilot’s HSI. d. 115 VAC Bus 2: Copilot’s flight director (optional), air data computer and pilot’s altimeter. The auto air data computer and pilot’s altimeter. The auto pilot will only operate in basic autopilot modes due to loss of valid signal from the air data computer. Autopilot Hardover 1. Autopilot/Trim Disengage Switch ......................... PRESS Press switch on either yoke. Flight director modes remain selected. 2. Maximum Altitude Losses during Autopilot Malfunction:


550-0161 and earlier

550-0162 and after

a. Cruise

450 feet at 43,000 feet


b. Climb



c. Maneuvering


d. ILS

37 feet (autopilot must be off at 100 feet)


34 feet (autopilot must be off at 90 feet)

12-21

Environmental Systems Abnormal Procedures Rapid Decompression 1. Oxygen Masks .............................. DON Check oxygen selector on 100%. 2. Emergency Descent ...................... AS REQUIRED 3. Pass Oxygen ................................. ENSURE passengers are receiving oxygen. Visually check mask drop when cabin reaches 13,500 ± 600 feet If masks are not down, drop them by the PASS OXY MASK switch on the left console (OXYGEN CONTROL VALVE, 550-0550~0626). 4. Oxygen MIC Switches ................... MIC OXY MASK Switch to MIC OXY MASK in order to use microphone in oxygen mask. 5. Transponder ....................................EMER 7700 6. Refer to use of Supplemental Oxygen Procedures in the Abnormal Procedures.

12-22



Cessna Citation II Technical Manual Emergency Descent 1. Throttles ....................................... IDLE 2. Speed brakes ................................ EXTEND 3. Initiate moderate bank. 4. Airplane Pitch Attitude .................. 15° NOSE DOWN 5. Passenger Advisory Lights ..............PASS SAFETY 6. Maximum Airspeed ..........................VMO/MMO MMO (above 28,000 feet) ..................0.705 MACH VMO (14,000-28,000 feet) .................277 KIAS VMO (below 14,000 feet) ...................262 KIAS Use reduced speed if structural damage has occurred. 7. Transponder ....................................EMERGENCY 7700 Environmental System Smoke or Odor 1. Oxygen Mask and Oxygen MIC Switches ...................................AS REQUIRED. Oxygen selector on 100% and MIC oxygen switch in MIC OXY MASK position in order to use MIC in oxygen mask. 2. Cabin (ovhd) Fan .............................OFF 3. Defog Fan ........................................OFF Both cabin and cockpit defog fans off to prevent further circulation of smoke through the aircraft and possibly identify them as the source. 4. Pressurization Source Selector ............................................ISOLATE source by selecting LH. Note: Pressurization source selector must remain in each position long enough to allow adequate system purging to determine the source of smoke. If smoke has not begun to clear in a minute, switch to another source. IF SMOKE CONTINUES: 5. Pressurization Source Selector ........RH (allow time for smoke to dissipate) IF SMOKE STILL CONTINUES (AIR CYCLE MACHINE SEAL MAY BE LEAKING): 6. Pressurization Source Selector ........EMER (control cabin pressure with LH throttle) EMERGENCY PROCEDURES 12/99


12-23

Smoke Removal Note: No action is normally required; however, if smoke is intense: 1. Oxygen Mask ...................................DON. Check oxygen selector is on 100%. 2. PASS OXY MASKS ...........................MANUAL DROP (OXYGEN CONTROL VALVE, 550-0550~0626) 3. CREW OXY PRIORITY Valve ............CHECK NORMAL (550-0505 and earlier) 4. Ensure passengers are receiving oxygen. Visually check mask drop when cabin reaches 13,500 ± 600 feet. If masks are not down, drop them by PASS OXY MASKS or OXYGEN CONTROL VALVE (550-0550~0626) switch on the left console. 5. Oxygen Mic Switches ......................MIC OXY MASK Switch must be in this position to use microphone in the oxygen mask. 6. Passenger Advisory Annunciator .....PASS SAFETY 7. Cabin Altitude Selector ....................SET to higher cabin altitude. 8. Pressurization Source Selector ........BOTH HI (550-0481 and earlier, 0483, and 0484) 9. Emergency Dump Switch ................DUMP This switch manually opens the normal dump valve to rapidly depressurize the airplane, allowing the smoke to clear. All smoking material should be extinguished. 10. Refer to use of Supplemental Oxygen Procedures in the Abnormal Procedures. IF SMOKE PERSISTS OR IT CANNOT BE VERIFIED THAT THERE IS NO FIRE: 11. Land as soon as possible.

12-24



Cessna Citation II Technical Manual Overpressurization 1. Cabin Altitude Selector ....................SET to higher cabin altitude. Cabin altitude will be descending and differential pressure will be increasing. Attempt to decrease the differential pressure by calling for a higher cabin altitude forcing the outflow valve to open. 2. Rate Control .....................................INC Go to full increase to program the controller to work as rapidly as possible. 3. Pressurization Source Selector ........LH or RH Attempt to control cabin pressure with the appropriate throttle by reducing power, thereby letting a smaller amount of air into the aircraft to pressurize the cabin. IF UNABLE TO CONTROL: 4. Oxygen Masks .................................DON Check oxygen selector on 100%. 5. PASS OXY MASK .............................MANUAL DROP (OXYGEN CONTROL VALVE, 550-0550~0626) 6. CREW OXY PRIORITY Valve ............CHECK NORMAL (550-0505 and earlier) 7. Assure passengers are receiving oxygen. Visually check mask drop when cabin reaches 13,500 ± 600 feet. If masks are not down, drop them by the PASS OXY MASKS or OXYGEN CONTROL VALVE (550-0550~0626) manual switch on the left console. 8. Oxygen MIC Switches .....................MIC OXY MASK Switch to MIC OXY MASK in order to use microphone in oxygen mask. 9. Passenger Advisory Light ................PASS SAFETY 10. Pressurization Source Selector ........OFF 11. Descend IF STILL OVERPRESSURIZED: 12. Emergency Dump Switch ................DUMP This switch manually opens the normal dump valve to rapidly depressurize the airplane. All smoking material should be extinguished. 13. Refer to the Use of Supplemental Oxygen Procedures in the Abnormal Procedures.



12-25

Spins Intentional spins are prohibited and were not conducted during flight tests of the aircraft. Should a spin occur, the following procedures are recommended: 1. Power to idle on both engines. 2. Neutralize yoke and apply full rudder opposite the direction of rotation. 3. Approximately 1/2 turn of spin after applying rudder, push yoke forward. 4. Remove rudder input as rotation slows so that rudder is centered when rotation stops. 5. Pull out of the dive with smooth steady control pressure. 6. Indicated airspeed, or angle of attack if installed, should be closely monitored during the pullout to avoid a secondary stall.

12-26



Cessna Citation II Technical Manual Ditching Good crew coordination is essential to the success of any ditching. Radio contact should be attempted giving aircraft identification, position, heading, altitude, and the transponder set on 7700 and the locator beacon set on EMER. Passengers should be briefed and don life jackets keeping them uninflated until outside the airplane. Plan on an approach to parallel any uniform swell pattern and attempt to touch down along a wave crest or just behind it. If the surface wind is very strong or the water surface rough and irregular, ditch into the wind on the back side of a wave. Gear should be left up with flaps in the LAND position. The LDG GEAR circuit breaker can be pulled to silence the gear warning horn. Speed should be maintained at VREF with the rate of descent at 200-300 feet per minute. Ditch while power is available if possible, so that the most desirable approach can be made. Touchdown should be slightly nose high and throttles cut off just before water contact. Passengers and crew exit through the emergency escape hatch inflating life jacket when clear. 1. Radio ...............................................MAYDAY Identify airplane, position, heading altitude and IAS. 2. Transponder ....................................7700 3. Locator Beacon ...............................EMER 4. Pressurization Source Selector ........OFF Prevents water from entering through bleed valves. 5. Passenger Advisory Switch .............PASS SAFETY Check aft facing seats full aft and all seats upright and outboard. 6. Passenger Life Jackets ....................ON Life jackets should not be inflated until outside airplane. 7. Gear ...............................................UP 8. Flaps ...............................................40° 9. Speed ..............................................VREF 10. Rate of Descent ...............................200-300 feet per min.



12-27

11. Plan an approach to parallel any uniform swell pattern and attempt to touchdown along a wave crest or just behind it. If the surface wind is very strong or the water surface rough and irregular, ditch into the wind on the backside of a wave. Airplane should touch down nose high with a minimum rate of descent. 12. Throttles ...........................................CUTOFF just prior to contact. 13. Emergency Exit ................................OPEN Forced Landing All considerations for a successful forced landing are similar to those for ditching. Attempt to establish radio contact, squawk the emergency code, and brief passengers. The approach should be made with gear DOWN, flaps in LAND position, speed VREF, and a 200-300 feet per minute rate of descent. If possible, establish an abeam position with gear extended and altitude sufficient to enable a safe landing to be made in the event of power loss. Just before touchdown, place throttles in cut-off and turn off the battery. Touchdown should be made in a normal landing attitude and emergency braking employed if necessary. 1. Radio ...............................................MAYDAY Identify airplane, position, heading altitude and IAS. 2. Transponder ....................................7700 3. Locator Beacon ...............................EMER 4. Passenger Advisory Switch .............PASS SAFETY Brief passengers as thoroughly as possible. 5. Gear ...............................................DOWN 6. FIaps ...............................................40° 7. Speed ..............................................VREF 8. Rate of Descent ...............................AS REQUIRED to effect touchdown in selected landing area 9. Throttles ...........................................CUTOFF just prior to contact 12-28



Cessna Citation II Technical Manual Electrical System Abnormal Procedures (550-0627 and after) Electrical Fire or Smoke 1. Oxygen Masks .............................. DON and 100% oxygen 2. Oxygen MIC Switches ................... AS REQUIRED Ensure selector is on 100% oxygen when masks are used. Ensure oxygen MIC switch is in MIC OXY MASK position. 3. Pressurization Source Selector ............................................NORM KNOWN SOURCE OF FIRE: 4. Isolate faulty circuit. Pull circuit breaker to remove power from faulty equipment. UNKNOWN SOURCE OF FIRE: 4. Flood Lights .....................................FULL BRIGHT 5. Battery Switch ..................................EMER 6. Generators .......................................OFF With the battery switch in the emergency position and the generators off, power is supplied for approximately 30-minutes to COMM 1, NAV 2, overhead floodlights, copilot’s HSI and DG 2, copilot’s attitude indicator, the voltmeter and both audio panels. Additionally, the standby gyro battery provides at least 30-minutes of power to operate the standby gyro indicator. On airplanes 550-0682 and after, the RH pitot-static heater is also on the emergency bus. In dual EFIS installations, the CP HSI and CP ADI are not on the emergency bus, and the CP RMI is on the emergency bus. In dual EFIS installations when emergency battery power only is available, attitude information is provided only by the standby gyro horizon.



12-29

Caution: When landing with emergency power (Battery SwitchEMER and both generators off), the following are not available: a. The landing gear normal operation is not available. The landing gear must be lowered using the blowdown system and the landing gear warning lights will not illuminate. b. The flaps will not operate. If not previously in landing position, a flaps inoperative landing must be made. c. The anti-skid/power brake system is inoperative. Only the emergency brake system is available. d. The engine anti-ice valves will be open. Refer to anti-ice on thrust charts. e. The outside air temperature gage is not reliable, so use caution when applying power (except for go-around where ground temperatures can be used). f. All engine instruments except the vertical tape N1 will be inoperative. The vertical tape N1 will indicate erratically below approximately 50% N1, but will give reliable indications above 50% N1. 7. Windshield Bleed Air Manual Valves ..................................OFF or MINIMUM for clear vision through the windshield. With electrical power lost, the windshield bleed air shutoff valve will fail open. The bleed air manual valves are closed to prevent an excessive volume of high temperature air from reaching the windshield. 8. DC Power RH Bus ............................PULL No. 1,2,3 CB’s 9. RH CB Panel Circuit Breaker (LH Panel) ........................................PULL 10. AC Inverter No. 1 Circuit Breaker (LH Panel) ........................................PULL 11. Land as soon as practical (within 30-minutes) IF SEVERITY OF SMOKE WARRANTS: 11. Initiate Smoke Removal and/or Emergency Descent procedures. Land as soon as possible.

12-30



Cessna Citation II Technical Manual WHEN LANDING ASSURED: 12. LH Generator ...................................ON 13. Landing Gear ...................................DOWN 14. Flaps ...............................................LAND 15. Airspeed ..........................................VREF Note: Right thrust reverser (if applicable) will be inoperative. IF SMOKE OR FIRE STARTS AGAIN: 16. LH Generator ...................................OFF Note: Anti-skid system will be inoperative. Power brakes will be available until accumulator discharges. Multiply landing distance by 1.6. Be prepared to use the emergency brake system.



12-31

Battery Overheat BATT O’TEMP Annunciator Illuminated (temp between 145° and 160°F) 1. Battery Switch .............................. EMER In EMER position the battery will be disconnected from the generators and will no longer be charged by them. All electrical equipment will continue to receive power since the generators are still on line. The Emergency DC bus is powered by the battery. Battery voltage may now be read with the voltage selector in BATT, and generator bus voltage with the voltage selector in LH GEN or RH GEN. Individual generator voltages can be read by selecting one (LH or RH) GEN and turning the other generator off. 2. Amperage ..................................... NOTE decrease 3. If battery voltage is 1 volt less than generator voltage in 30seconds to 2-minutes, monitor battery overheat annunciator for possible change. In thirty seconds to two minutes after disconnect, battery voltage should read at least one volt less than the generators. Rotate the voltage selector to LH GEN and RH GEN position to read generator voltage. Battery voltage will be indicated when the voltage selector is in the BATT position. IF BATT O’TEMP ANNUNCIATOR EXTINGUISHED: 4. Battery Switch ..................................BATT

12-32



Cessna Citation II Technical Manual IF NO AMP DECREASE OR BATT O’TEMP ANNUNCIATOR FLASHES: 4. Battery Switch ..................................EMER 5. Flood Lights .....................................FULL BRIGHT 6. Generators .......................................OFF Since the battery has continued to overheat, it may be because the battery is still being charged through a failed battery relay. Turning both generators OFF trips the generators and opens the power relays, isolating each generator from its bus. In standard single EFIS installations, the emergency DC bus is now powered by the battery, with COMM 1, NAV 2, voltmeter, pilot’s and copilot’s audio panels, copilot’s HSI and DG 2, copilot’s attitude indicator and cockpit floodlights available to the crew. In optional dual EFIS installations, the emergency DC bus is powered by the battery with COMM 1, NAV 2, copilot’s RMI, the NAV 2 repeater indicator, pilot’s and copilot’s audio panels, voltmeter, and cockpit floodlights available to the crew. On airplanes 550-0682 and after the RH pitot-static heater is also on the emergency bus (single and dual EFIS). Emergency battery power will be available for approximately 30-minutes under normal conditions. Additionally, the standby gyro battery provides at least 30-minutes of power to operate the standby gyro indicator. The BATT O’TEMP light will extinghish immediately when the generators are turned off if the battery relay is not stuck.



12-33

Caution: When landing with emergency power (Battery SwitchEMER and both generators off), the following are not available: a. The landing gear normal operation is not available. The landing gear must be lowered using the blowdown system and the landing gear warning lights will not illuminate. b. The flaps will not operate. If not previously in the landing position, a flaps inoperative landing must be made. c. The Anti-skid/Power Brake system is inoperative. Only the emergency brake system is available. d. The engine anti-ice valves will be open. Refer to anti-ice on thrust charts. e. The outside air temperature gage is not reliable, so use caution when applying power (except for go-around where ground temperatures can be used). f. All engine instruments except the vertical tape N1 will be inoperative. The vertical tape N1 will indicate erratically below approximately 50% N1, but will give reliable indications above 50% N1. IF NORMAL DC POWER LOST (BATTERY RELAY NOT STUCK): 7. Generators .......................................GEN (BATT O’TEMP light will come back on until battery cools). 8. Battery Switch ..................................OFF 9. Land as soon as practical. Caution: After landing, refer to maintenance manual for proper maintenance procedures as damage to the battery may have occurred. IF NO DC POWER LOST (BATTERY RELAY STUCK): 7. Windshield Bleed Air Manual Valves ..................................OFF or MINIMUM for clear vision through windshield. 8. DC Power LH and RH Bus ...............PULL No. 1,2,3 CB’s 9. Land as soon as practical.

12-34



Cessna Citation II Technical Manual WHEN LANDING ASSURED: 10. DC Power LH and RH Circuit Breakers ...............................RESET 11. Landing Gear ...................................DOWN 12. Flaps ...............................................LAND 13. Airspeed ..........................................VREF Caution: After landing, refer to maintenance manual for proper maintenance procedures as damage to the battery may have occurred.



12-35

Loss of Both Generators 1. Generators .......................................RESET then GEN Attempt to reset both generators. IF ONLY ONE GENERATOR COMES ON: 2. Electrical Load .................................REDUCE as required. 400 amps maximum up to 35,000 feet, 325 amps maximum above 35,000 feet. High load items are engine antiice (wing leading edge), the autopilot computer and the radar. IF NEITHER GENERATOR COMES ON: 2. Flood Lights .....................................FULL BRIGHT 3. Battery Switch ..................................EMER With the battery switch in emergency position and the generators off, power is supplied for approximately 30minutes to COMM 1, NAV 2, overhead floodlights, voltmeter and copilot’s HSI. On airplanes with single EFIS, with the battery switch in the emergency position and the generators off, power is supplied for approximately 30-minutes to COMM 1, NAV 2, Copilot’s HSI and DG 2. and copilot’s audio panels, copilot’s attitude indicator, voltmeter, cockpit floodlights and standby gyro. Airplanes 550-0682 and after also have the RH pitot static heater on the emergency bus. On airplanes with dual EFIS, with the battery switch in emergency position and the generators off, power is supplied for approximately 30-minutes to COMM1, NAV 2, copilot’s RMI, NAV 2 repeater indicator, pilot’s and copilot’s audio panels, voltmeter, cockpit floodlights and standby gyro. Airplanes 550-0682 and after also have the RH pitot static heater on the emergency bus.

12-36



Cessna Citation II Technical Manual Caution: When landing with emergency power (Battery SwitchEMER and both generators off), the following are not available: a. The landing gear normal operation is not available. The landing gear must be lowered using the blowdown system and the landing gear warning lights will not illuminate. b. The flaps will not operate. If not previously in the landing position, a flap inoperative landing must be made. c. The Anti-skid/power brake system is inoperative. Only the emergency brake system is available. d. The engine anti-ice valves will be open. Refer to anti-ice on thrust charts. e. The outside air temperature gage is not reliable, so use caution when applying power (except for go-around where ground temperatures can be used). f. All engine instruments except the vertical tape N1 will be inoperative. The vertical tape N1 will indicate erratically below approximately 50% N1, but will give reliable indications above 50% N1. 4. Windshield Bleed Air Manual Valves ..................................OFF or MINIMUM for clear vision through windshield. With electrical power lost, the windshield bleed air shutoff valve will fail open. The bleed air manual valves are closed to prevent an excessive volume of high temperature air from reaching the windshield. 5. Land as soon as practical. WHEN LANDING ASSURED: 6. Battery Switch ..................................BATT 7. Landing Gear ...................................DOWN 8. Flaps ...............................................LAND 9. Airspeed ..........................................VREF



12-37

AC Power Failure Both INVERTER 1 FAIL, INVERTER 2 FAIL and AC FAIL Annunciators Illuminated SINGLE EFIS INSTALLATION: 1. Inverter 1 and Inverter 2 CB’s ..........RESET 2. Battery Switch ..................................EMER If the inverters will not come back on the line after the circuit breakers have been reset, complete the flight by using the copilot’s attitude indicator or the standby gyro horizon. Placing the battery switch to EMER will provide AC power from the copilot’s C-14D static inverter to power the copilot’s compass system and NAV 2. The EFIS system will be inoperative with electrical system failure. With BATT in EMER, NAV 2 and compass information will be displayed on the copilot’s HSI, and attitude information on the copilot’s attitude indicator and the standby gyro. OPTIONAL DUAL EFIS INSTALLATION: 1. Inverter 1 and Inverter 2 CB’s ..........RESET 2. Battery Switch ..................................EMER If the inverters will not come back on the line after the circuit breakers have been reset, complete the flight by using the standby gyro horizon. Placing the battery switch to EMER will provide AC power from the copilot’s C-14D static inverter to power the DG 2 and the copilot’s RMI, NAV 2, and the NAV 2 repeater indicator. AC Power and/or Distribution Failure (AC FAIL Annunciators Illuminated After MASTER WARNING Has Been Reset, INVERTER 1 and 2 Annunciators Extinguished) 1. Check the right sub-circuit breaker panel for disengaged AC BUS circuit breaker(s). Caution: If circuit breaker(s) is/are disengaged, operate with loss of bus as re-engagement may result in further damage to the electrical system.

12-38



Cessna Citation II Technical Manual Note: Depending on which buses have been lost, the following equipment will be lost: a. 26 VAC Bus 1: pilot’s ADI, pilot’s RMI, pilot’s HSI (HDG and NAV flags in view), pilot’s rate-of-turn indicator, copilot’s RMI compass card and NAV 1 bearing pointer, and ADF 1. b. 115 VAC Bus 1: pilot’s flight director, autopilot, yaw damper, radar, pilot’s attitude gyro, and VNAV computer/ controller. c. 26 VAC Bus 2: NAV 2, copilot’s ADI (optional dual flight director installation), copilot’s HSI, copilot’s RMI, pilot’s RMI compass card and NAV 2 bearing pointer, and ADF 2 (optional). Operation of the following equipment can be reinstated by placing the BATT switch to EMER: NAV 2, bearing pointer and DG of copilot’s HSI. d. 115 VAC Bus 2: Copilot’s flight director (optional), air data computer and pilot’s altimeter. The auto air data computer and pilot’s altimeter. The autopilot will only operate in basic autopilot modes due to loss of valid signal from the air data computer.



12-39

EFIS Symbol Generator Failure (Standard Two-Tube Installation) RED X AND SG FAIL ON EADI OR BOTH DISPLAYS BLANK: 1. The EFIS system is inoperative. Refer to the copilot’s attitude indicator, the standby gyro horizon and the copilot’s HSI for attitude and compass information. Symbol Generator Overtemperature (SG HOT Light On) (Standard EFIS Installation) 1. EFIS 1, EADI 1 and EHSI 1 DC Circuit Breakers .........................PULL The SG HOT annunciator has probably illuminated due to an overtemp condition caused by failure of the Symbol Generator internal fan. Continued use of the Symbol Generator without the fan may lead to its failure. Removing power from the symbol generator will allow it to cool, but restoring power will likely result in another overtemp indication. Consideration should be given to leaving the circuit breaker disengaged, using the copilot’s flight instruments to complete the flight and, if necessary, restoring power to the symbol generator for the approach and landing. Pulling the EADI and EHSI circuit breakers will enable the display tubes to cool. Refer to the copilot’s attitude indicator, the standby gyro horizon and the copilot’s HSI for attitude and heading information.

12-40



Cessna Citation II Technical Manual Environmental Systems Abnormal Procedures Overpressurization 1. Cabin Altitude Selector ....................SET to higher cabin altitude. Cabin altitude will be descending and differential pressure will be increasing. Attempt to decrease the differential pressure by calling for a higher cabin altitude forcing the outflow valve to open. 2. Rate Control .....................................INC Go to full increase to program the controller to work as rapidly as possible. IF STILL OVERPRESSURIZED: 3. Pressurization Source Selector ........LH or RH Attempt to control cabin pressure with the appropriate throttle by reducing power, thereby letting a smaller amount of air into the aircraft to pressurize the cabin. IF UNABLE TO CONTROL: 4. Oxygen Masks .................................DON. Check oxygen selector on 100%. 5. PASS OXY MASK .............................MANUAL DROP 6. Ensure passengers are receiving oxygen. Visually check mask drop has occurred. 7. Oxygen MIC Switches .....................MIC OXY MASK Switch to MIC OXY MASK in order to use microphone in oxygen mask. 8. Passenger Advisory Light ................PASS SAFETY 9. Pressurization Source Selector ........OFF 10. Descend IF STILL OVERPRESSURIZED: 11. Emergency Dump Switch ................DUMP This switch manually opens the normal dump valve to rapidly depressurize the airplane. All smoking material should be extinguished.



12-41

Rapid Decompression CAB ALT 10,000 FEET LIGHT WILL ILLUMINATE: 1. Oxygen Masks .............................. DON Check oxygen selector on 100%. 2. Emergency Descent ...................... AS REQUIRED 3. Pass Oxygen ................................. ENSURE passengers are receiving oxygen. Visually check mask drop when cabin reaches 13,500 ± 600 feet If masks are not down, drop them by placing the OXYGEN CONTROL VALVE on the left console to MANUAL DROP. 4. Oxygen MIC Switches ................... MIC OXY MASK Switch to MIC OXY MASK in order to use microphone in oxygen mask. 5. Transponder ....................................EMER 7700

12-42



Cessna Citation II Technical Manual Emergency Descent 1. Throttles ....................................... IDLE 2. Speed brakes ................................ EXTEND 3. Initiate moderate bank. 4. Airplane Pitch Attitude .................. 15° NOSE DOWN 5. Passenger Advisory Lights ..............PASS SAFETY 6. Maximum Airspeed ..........................VMO/MMO MMO (above 30,500 feet) ..................0.705 MACH VMO (below 30,500 feet) ...................262 KIAS Use reduced speed if structural damage has occurred. 7. Transponder ....................................EMERGENCY 7700 IF DESCENT INTO ICING CONDITIONS IS REQUIRED: 8. Throttles ...........................................AS REQUIRED Maintain sufficient power for anti-icing (engine anti-ice lights remain OFF). Environmental System Smoke or Odor 1. Oxygen Mask ...................................DON and 100% 2. Oxygen MIC Switches .....................AS REQUIRED Oxygen selector on 100% and MIC oxygen switch in MIC OXY MASK position in order to use MIC in oxygen mask. 3. Cabin (OVHD) Fan ...........................OFF 4. Defog Fan ........................................OFF Both cabin and cockpit defog fans off to prevent further circulation of smoke through the aircraft and possibly identify them as the source. 5. Pressurization Source Selector ............................................ISOLATE source by selecting LH. IF SMOKE CONTINUES: 6. Pressurization Source Selector ........RH (allow time for smoke to dissipate) IF SMOKE STILL CONTINUES (AIR CYCLE MACHINE SEAL MAY BE LEAKING): 7. Pressurization Source Selector ........EMER (control cabin pressure with LH throttle)



12-43

Smoke Removal Note: No action is normally required; however, if smoke is intense: 1. Oxygen Mask ...................................DON Check oxygen selector is on 100%. 2. OXYGEN CONTROL VALVE ............MANUAL DROP 3. Ensure passengers are receiving oxygen. Visually check mask drop has occurred. 4. Oxygen Mic Switches ......................MIC OXY MASK Switch must be in this position to use microphone in the oxygen mask. 5. Passenger Advisory Light ................PASS SAFETY 6. Cabin Altitude Selector ....................SET to higher cabin altitude. Selecting a higher cabin altitude will cause the outflow valves to open and increase the rate of airflow to clear the smoke. 7. Emergency Dump Switch ................DUMP This switch manually opens the normal dump valve to rapidly depressurize the airplane, allowing the smoke to clear. All smoking material should be extinguished. IF SMOKE PERSISTS OR IT CANNOT BE VERIFIED THAT THERE IS NO FIRE: 8. Land as soon as possible.

12-44



Cessna Citation II Technical Manual Autopilot Hardover 1. Autopilot/Trim Disengage Switch ......................... PRESS Press switch on either yoke. Flight director modes remain selected. 2. a. b. c.

Maximum Altitude Losses during Autopilot Malfunction: Cruise ..............................................550 feet at 43,000 feet Climb ...............................................300 feet at 17,000 feet ILS ...............................................34 feet (autopilot must be off at 90 feet)

Maximum Glide – Emergency Landing 1. Airspeed ..........................................BEST GLIDE AT 9500 LBS-120 KIAS. Increase speed 3 KIAS for each 500 pounds increase in weight. 2. Flaps ...............................................UP 3. Speed brakes ..................................RETRACT 4. Landing Gear ...................................UP 5. Transponder ....................................EMERGENCY, 7700 6. ATC

...............................................ADVISE

7. Passenger Advisory Switch .............PASS SAFETY 8. Shoulder Harness ............................SECURE 9. Landing Gear, Speed brakes, and Flaps .........................................AS REQUIRED for landing anticipated



12-45


Appendix 1 Standard Operating Procedures Table of Contents Standard Operating Procedures Overview ......................................... A1-1 Engine Start ................................... A1-2 Taxiing ........................................... A1-3 Takeoff ........................................... A1-4 After Takeoff Climb ......................... A1-5 Cruise ............................................ A1-6 Descent.......................................... A1-7 Before Landing ............................... A1-8 Landing (without thrust reversers) .... A1-9 Landing (with thrust reversers) ....... A1-10 After Landing ................................ A1-11 Shutdown ..................................... A1-11 Standard Operating Procedures (crew briefings) Before Takeoff .............................. A1-12 Takeoff ......................................... A1-12 Climb ........................................... A1-13 Cruise .......................................... A1-14 Descent........................................ A1-15 Precision Approach ........................ A1-16 Precision Missed Approach ............ A1-18 Non-Precision Approach ................. A1-19 Non-Precision Missed Approach ...... A1-21 Visual Approach ............................ A1-22 Landing ........................................ A1-23

Cessna Citation II Technical Manual Overview This chapter provides standard operating procedures as well as the abbreviated cockpit checklist in expanded form. The standard operating procedures are per the Operating Manual, however, their use in this manual is for training purposes only. The appropriate Operating Manual, FAA approved Airplane Flight Manual (AFM), Pilot’s Check List, and/or related publications should be referred to for procedures specific to the year, model and serial number of your aircraft. For each expanded phase-of-flight procedure, the call, action and response from the Pilot in Command and the Second in Command, (hereafter referred to as PIC and SIC) is delineated. It is the responsibility of the pilot in command to ensure that the aircraft is safely loaded and properly configured for flight. It is also necessary to understand and utilize the various graphs and tables which outline the performance characteristics of your airplane, its weight and balance data, and the forms which are used to determine load placement. All Notes, Cautions, and Warnings are provided to amplify each procedure.

STANDARD OPERATING PROCEDURES 12/99


A1-1

Engine Start Either engine may be started first. If the door is secured prior to battery start initiation, it is recommended that the left engine be started first. Spool up will be slightly faster due to less line loss because the battery is mounted on the left side of the tailcone compartment. Due to foreign object ingestion hazard, the left engine should not be running during boarding or deplaning. If last minute boarding and use of BLEED AIR GND is anticipated, the right engine should be started first. With external power in use, the GEN switches can be off until starting is complete. It may not be possible to bring the generators on the line until the external power unit is removed. In any case, electrical equipment should not be turned on until both GEN OFF lights are extinguished. An overcurrent and overvoltage protection system is provided during use of an external power unit (EPU). The control unit monitors the external power unit voltage and will deenergize the external power relay if the voltage is above 32.5 volts. During an engine start using the external power unit, a signal is applied by the current to the control unit. If the signal indicates more than 1200 amps after two seconds the control unit will deenergize the external power relay and terminate the start. External power cannot be reapplied to the airplane until the current has been interrupted after the start termination for the current protection or until the voltage is reduced below 32.5 volts for the voltage protection. Should automatic start sequencing not terminate, the boost pump, ignition and associated lights will remain on. The starter, however, will discontinue cranking due to speed sensing which governs at approximately 40 to 43 percent N2. Depressing the STARTER DISENGAGE button will terminate the automatic start sequence. This button is illuminated any time the PANEL LIGHT CONTROL master switch is ON. Prior to taxiing, the Second in Command (SIC) will complete the Takeoff Data Card insuring the latest information is used for data computation. Computations for takeoff thrust setting, V1, VR, V2, and VENR speeds, takeoff field length, and climb thrust are based on the runway length available, runway gradient, field temperature, field pressure, wind, icing conditions and runway condition.

A1-2


CITATION II APPENDIX 1 12/99

Cessna Citation II Technical Manual Taxiing Taxiing on one engine may be advisable at light weights to reduce brake wear, particularly in very cold weather when idle thrust is relatively high. Turning capability into the live engine is reduced however, and consideration should be given to the direction of anticipated turns in deciding which engine to operate. Peak exhaust velocity to generate the necessary thrust will be higher on one engine. Maneuvering in close quarters may dictate use of both engines. Ground operations in visible moisture with an outside ambient air temperature from -30°C to +4°C require that ENGINE ANTI-ICE be “ON” and the engines run at or above 65% N2 one minute out of every four. Note: The anti-ice system must be checked prior to takeoff if flight into icing conditions is expected. Approximately 70% turbine speed is required to provide adequate engine bleed air to extinguish the ENGINE ANTI-ICE light in two minutes or less. Note: When operating in visible moisture and ambient air temperature is between +4°C and -30°C, position ground idle switch to “HIGH”, turn pitot and static heat and engine LH and RH anti-ice systems “ON”. If temperature is above 18°C, turn W/S BLEED air switch to “LOW”. If temperature is -18°C or below, turn W/S BLEED air switch to “HI”. Check W/S bleed air valves “MAX”. For sustained ground operation, the engines should be operated one out of every four minutes at 65% turbine RPM or above.



A1-3

Takeoff Monitoring the engine instruments, apply power slowly while referencing the cabin vertical speed indicator. Very rapid thrust application can cause a pressure surge due to increased airflow into the cabin. A rolling takeoff may be used with sufficient runway available, but it should be remembered that Flight Manual takeoff field length data and takeoff N1 settings assume a static runup. Directional control is normally maintained with nose gear steering and rudder; and upwind (wing down) aileron in crosswind conditions. It is suggested that the copilot perform the engine instrument monitoring function and set the throttles enabling the pilot to direct his full attention to airplane control. N1 should be closely observed, and throttle corrections made as necessary to ensure symmetrical thrust application. Large differential power changes, particularly at the higher thrust settings, can induce significant yaw. It is recommended that the copilot verbally state when takeoff thrust is set, a cross-check of airspeed indicators at 70 knots is made, and when reaching V1 and VR. Positive back pressure is required to rotate the Citation II and it should be accomplished precisely at Vr. Early or late rotation will degrade takeoff performance. It should be done smoothly, however, so that a decrease in airspeed does not occur. Should an emergency situation occur at a speed below V1, takeoff should normally be aborted. Proceed with a normal takeoff should the emergency situation occur at a speed above V1. Single engine rotation is approximately 7° to 10° pitch attitude. Normal rotation angle is 10 to 12° nose up with both engines operating. Procedures for abort and single engine takeoff are outlined in Chapter 12 – Emergency Procedures.

A1-4



Cessna Citation II Technical Manual After Takeoff Climb Using indicated temperature and the MULTI ENGINE NORMAL CLIMB thrust setting graph in your FAA Approved Airplane Flight Manual (or in the Abbreviated checklist), determine climb N1. During climb, observe the differential pressure/cabin altitude and cabin vertical speed gages for proper programming and comfortable rate. Periodic checks of time to climb remaining, cabin altitude and rate of cabin ascent will provide the required information to determine any adjustments necessary. As an example, passing 20,000 feet with a cabin altitude of 4000 feet and an estimated climb time remaining of 10 minutes to 35,000 feet (8000 feet cabin altitude), would require a cabin climb rate of 400 feet-per-minute to attain planned cruise and cabin altitudes concurrently. With RATE set too low, maximum differential pressure may be reached before cruise altitude. This takes control of the system away from the crew because the outflow valve will relieve as necessary to maintain maximum differential. A RATE setting too high may be uncomfortable and will result in programmed cabin altitude being reached before cruise flight level. A thorough understanding of DIFF PRESS/CABIN ALT gage interpretation will aid the crew in smooth operation of the pressurization system.



A1-5

Cruise Climb thrust is normally maintained upon level off until acceleration to the desired cruise mode takes place. Thrust is then adjusted to the appropriate setting. If engine RPM does not automatically synchronize at desired cruise setting, turn the engine synchronizer switch to “OFF”, allowing the synchronizer actuator to center; roughly synchronize the engines with the throttles and turn the synchronizer switch to “FAN” or “TURB”. For the maximum range case, thrust necessary to maintain optimum angle of attack diminishes with fuel burn off because of increased performance and lower airspeed requirements as weight decreases. Although the Citation II is not operationally restricted in rough air, flight in severe turbulence should be avoided. If severe turbulence is encountered, it is recommended that the igniters be turned “ON” and airspeed maintained at approximately 180 KIAS. Maintain a constant attitude, avoid abrupt or large control inputs, and do not chase airspeed and altitude indications. Use of the autopilot in the SOFT RIDE mode is recommended. A comfortable cabin temperature is normally maintained with the AUTO TEMP SELECT in the 12 to 2 o’clock position. During daylight, the crew environment may not be an accurate reference to cabin comfort level due to solar heating taking place through the wide expanse of cockpit windows. An approximate indication of airflow warmth into the cabin can be determined by placing a hand over an open crew foot warmer outlet. The foot warmers are an extension from the same source as the cabin under floor ducting and can be used as a reference for AUTO TEMP SELECT adjustments to maintain a comfortable cabin. The DEFOG FAN should be turned on and foot warmers closed approximately 15 minutes before descent to reduce condensation on the windshield and cockpit side windows.

A1-6



Cessna Citation II Technical Manual This is particularly important when a rapid let down into an area of high humidity is anticipated after cold soaking at altitude. The engine bleed air anti-ice must be activated when operating in visible moisture at temperatures of from -4°C to -30°C indicated by OAT and any time icing is occurring. The pitot and static anti-ice is normally operated during flight. Detailed instructions of the engine anti-ice and surface deice systems are found in Chapter 10 – Ice Protection and in the FAA Approved Airplane Flight Manual. Descent Once destination altimeter setting is known, field pressure altitude can be determined because each .10 inches of mercury deviation from 29.92 equates to 100 feet difference between field elevation and pressure altitude. An altimeter setting above standard gives a pressure altitude below field elevation and the inverse is also true. As an example, descending to a field elevation of 350 feet with a reported altimeter of 29.77 would result in a field pressure altitude of 500 feet. The cabin altitude should be set to 700 feet to insure depressurization prior to touchdown. Rate is normally adjusted to give a 300-500 FPM cabin rate of descent. Monitor the differential pressure/cabin altitude and cabin vertical speed gages. A high cabin altitude and low differential pressure indicates an insufficient rate of descent and depressurization will occur when cabin and airplane altitude are identical. High cabin descent rates will occur when cabin and airplane altitude are identical. High cabin descent rates may be uncomfortable and may result in programmed cabin altitude being reached well before landing. Optimum comfort is realized by spreading cabin descent required over the majority of airplane let down time.



A1-7

Before Landing Consistently comfortable and safe landings are best achieved from a stabilized approach. The point at which the airplane should be stabilized with airspeed at VREF to VREF + KIAS, full flaps, and the desired descent rate is normally coincident with commencing the final descent to landing. Under instrument conditions, this usually occurs at the final approach fix inbound. During visual approaches, this would be a point approximately equal to a turn on to base leg, adjusted for the altitude difference between the traffic pattern and field elevation. After passing the instrument approach fix outbound or nearing the airport traffic area, airspeed should be reduced below 202 KIAS and the flaps extended to the APPR (15°) position. Approaching the final instrument fix inbound (one dot from glideslope intercept on an ILS), or a downwind abeam position, extend the landing gear below 176 KIAS. At the point where final descent to landing is begun, extend FULL FLAPS, establish the desired vertical rate, and adjust power to maintain VREF to VREF + 10 KIAS indicated airspeed. Power management during the approach/landing phase is relatively easy in the Citation II because an N1 setting in the 60-65% range will normally result in desired indicated airspeeds for the various configurations. Depending on air traffic control requirements, thrust necessary for the entire approach can often be set during descent keeping in mind that fan (N1) RPM will decrease slightly for a fixed throttle setting with a decrease in altitude for indicated airspeed. Using a sea level airport with zero wind at a typical landing weight (10,000 LBS), a throttle setting that results in about 60% N1 in close will give approximately level flight indicated airspeeds of 160 knots clean and 140 with flaps APPR. Gear extended, flaps FULL and commencing an average descent (500 FPM) will result in approximately VREF airspeed. Higher field elevations, landing gross weights and/or headwind component will require a greater power setting.

A1-8



Cessna Citation II Technical Manual For maneuvering prior to final approach, minimum airspeeds of VREF + 30, VREF + 20 and VREF + 10 should be maintained clean, flaps APPR and flaps LAND respectively to provide an adequate margin above stall. Speed control on final should be precise for optimum landing performance and this is best accomplished by establishing Vref airspeed well before crossing the threshold. In gusty wind conditions, it is recommended that one half the gust factor in excess of 5 knots be added to Vref. Approaching within approximately 50 feet of airport elevation, power should be gradually reduced to counter the acceleration induced by ground effect. Wind velocity and direction will dictate the rate at which the throttles are retarded. In very high surface headwind conditions, as an example, it may be necessary to maintain at or near approach power until close to touchdown. With a tailwind, a fairly rapid power reduction may be necessary in the final descent to landing phase for accurate speed control. In ground effect, where induced drag is reduced, leaving approach power on will cause the airplane to float to a longer touchdown than desired. Retarding the throttles gradually in the final descent will normally result in idle thrust being reached just before touchdown. Landing (without thrust reversers) Touchdown, preceded by a slight flare, should occur on the main wheels. Check thrust at idle and extend the speedbrakes while lowering the nosewheel. With the nosewheel on the runway, the drag chute (if installed) may be deployed by depressing the safety button on the drag chute handle and then pulling up firmly to full travel (approximately 27°). This will require approximately 35 LBS force. Braking should be commenced according to runway length available to reduce brake wear. Normally with excess runway, braking is begun after aerodynamic deceleration to below 80 KIAS takes place. Apply smooth, gradually increasing pressure until a comfortable turn off speed is reached. For maximum braking performance, immediately after touchdown and wheel spin-up, apply brakes firmly and hold to full stop (do not modulate brake pedals).



A1-9

Landing (with thrust reversers) Touchdown, preceded by a slight flare, should occur on the main wheels. Check thrust at idle and extend the speed brakes while lowering the nose, apply wheel brakes and deploy the thrust reversers. The airplane pitches slightly upward during the deployment of the reversers. Therefore, slight nosedown elevator pressure should be used during thrust reverser deployment especially at high speeds such as a refused actuation of the thrust reversers to eliminate the possibility of FOD and improve directional control. To avoid possible jamming of the throttle lockout cams, do not exceed approximately 15 LBS force on the thrust reverser levers during deployment. Check illumination of the thrust reverser lights. Caution: On any airplane, do not attempt to restow reversers and take off once reversers have started to deploy. On airplanes not incorporating SB550-78-03, throttle linkage damage may occur resulting in loss of power or flameout. Once the thrust reversers are deployed, move the thrust reverser levers aft to a maximum reverse thrust. For convenience, “Stops” have been installed on the thrust reverser levers and are set to provide 92% fan speed (N1) + 2% or 2% at -18°C at sea level. This will allow the pilot to keep his attention on the landing rollout instead of diverting his attention on the landing rollout instead of diverting his attention to the reverse power settings, except in an abnormal ambient temperature condition. At 60 KIAS, return the thrust reverser levers to the idle reverse detent position leaving the thrust reversers deployed for aerodynamic drag. Thrust reversing and braking should be commenced according to runway length. Normally, with excess runway, braking is begun after thrust reverser deceleration is below 60 knots. For maximum braking performance, immediately after touchdown and wheel spin-up, apply brakes firmly and hold to full stop (do not modulate brake pedals ). The thrust reversers should not be used for touch and go landings; a full stop landing should be made once the reversers are selected.

A1-10



Cessna Citation II Technical Manual After Landing It is recommended that the checklist be delayed until the airplane is clear of the runway. Turn off anti-collision and recognition lights. The rotating beacon (if installed) may be left “ON”. Do not advance the throttles while the thrust reversers are being stowed. On airplanes not incorporating SB550-78-03 the automatic engine power retard system will activate causing misrigging of the throttle linkage system. This would result in only partial takeoff power or possibly a flameout if the throttle was placed in idle position. On airplanes incorporating SB550-78-03 there is danger of the throttle being rapidly returned to idle position, which could cause injury. To avoid activating the automatic retard system, do not advance the primary throttle after moving the reverse thrust lever to stow until the UNLOCK light is out. Shutdown Always check cabin differential pressure at zero before opening the door. Any pressure existing due to malfunction of the left main gear squat switch or outflow valves could cause the door to open rapidly presenting a hazard to personnel in the vicinity. For deplaning at night, the battery switch may be left in BATT to make available all cabin lighting until passengers and cabin baggage are disembarked. Turning the EXTERIOR WING INSP LIGHTS switch “ON” provides additional illumination in front of the cabin door. An illuminated courtesy light switch located on the forward door post is wired to the hot battery bus and turns on the emergency exit lights and one aft baggage compartment light. When securing the airplane, install the engine and pitot tube covers. Check the BATT, passenger advisory and courtesy light switches off. Closing the door extinguishes integral courtesy light switch illumination. All doors and the nose avionics compartment can be key locked. A locking pin can be installed in the internal emergency exit door handle to prevent access from the outside. This pin must be removed prior to flight.



A1-11

Before Takeoff SIC

PIC Crew Briefing: The pilot should review crew coordination with respect to flap setting, ice protection procedures, takeoff power setting, "V" speeds and airspeed limits, as well as normal and emergency procedures.

Call:

“Before Takeoff checklist”

ACTION: Complete Before Takeoff checklist Call: “Before Takeoff checklist complete”

Takeoff SIC

PIC Call:

“Set takeoff power” ACTION: Check engine instruments Call: “Takeoff power set” Call: “Airspeed alive” ACTION: 70 knots crosscheck (both airspeed indicators) At: Call:

V1 “V1”

At: Call:

VR “Rotate”

ACTION: Positive back pressure should be applied and a rotation to an approximate 9° pitch attitude begun

Note: Should an emergency situation occur at a speed below V1, takeoff should normally be aborted. Proceed with a normal takeoff should the emergency situation occur at a speed above V1. Single engine rotation is approximately 7° to 10° pitch attitude. Procedures for abort and single engine takeoff are outlined in Chapter 12 – Emergency Procedures.

A1-12



Cessna Citation II Technical Manual Climb SIC

PIC At:

Call:

Call:

Positive rate of climb “Positive rate”

ACTION: Call: ACTION: Call:

Retract gear “Gear selected up” Check indication “Gear indicating up”

At: Call:

V2 + 10 KIAS “V2 + 10 400 feet”

Call:

“Flaps indicating up”

At: Call:

VENR “Best climb”

“Gear up”

ACTION: Retract flaps Call: “Flaps Up”

Crew Briefing: Using both indicated temperature and thrust setting graph (see Abbreviated Checklist), determine climb N1.

At: Call:

VENR “Set climb power” ACTION: Check Call: “Climb power set”

Call:

“Climb checklist” ACTION: Complete climb checklist Call: “Climb checklist complete”

At: 1.5% N1 or 1% N2 ACTION: Select FAN or TURB switch as required Call: “Engine Sync set” ACTION: Check engine instruments within limits Call: “Engine Sync set”



A1-13

Climb (continued) PIC At: Call:

SIC

Transition Altitude “29.92 set. Transition Altitude checklist” Call: “29.92 set” ACTION: Complete Transition Altitude checklist Call: “Transition Altitude checklist complete”

Cruise PIC Crew Briefing: If engine RPM does not automatically synchronize at desired cruise setting, turn the engine synchronizer switch to OFF, allowing the synchronizer actuator to center. Roughly synchronize the engines with the throttles and turn the synchronizer switch to FAN or TURB as required.

Call:

SIC

“Cruise checklist” ACTION: Complete Cruise checklist Call: “Cruise checklist complete”

Note: Although the engine is not operationally restricted in rough air, flight in severe turbulence should be avoided. If severe turbulence is encountered, it is recommended that the igniters be turned ON and airspeed maintained at approximately 180 KIAS. Maintain a constant airspeed, avoid abrupt or extended control inputs, and do not chase airspeed and altitude indications. Use of the autopilot in the SOFT RIDE mode is recommended. Caution: Do not operate deice boots when indicated OAT is below -40°C (-40°F). Note: The engine bleed air anti-ice must be activated when operating in visible moisture at temperatures from +4°C to -30°C indicated OAT and any time icing is occurring.

A1-14



Cessna Citation II Technical Manual Descent SIC

PIC Crew Briefing: When practicable, review approach and missed approach procedures. Determine N1 and V2 for use in the event of a missed approach. Set proper NAV frequencies and required heading and course information. Check runway requirements based on gross weight and destination field information.

Call:

“Descent checklist” ACTION: Complete Descent checklist Call: “Descent checklist complete” At: Call:

Call:

Transition altitude “18,000 ft”

“Set altimeter. Transition Level checklist” ACTION: Set altimeter Call: “Altimeter set” ACTION: Complete Transition Level checklist Call: “Transition Level checklist complete”



A1-15

Precision Approach SIC

PIC Call:

“Approach checklist” ACTION: Complete Approach checklist Call: “Approach checklist complete”

At: Call:

Airspeed below 202 KIAS “Flaps 15°” ACTION: Select 15° flaps Call: “Flaps selected 15° Call: “Flaps indicate 15°”

Call:

“Localizer captured” Call:

“Localizer captured”

At:

One dot from glideslope intercept “One dot to go”

Call: Call: Call:

“Gear down” “Before Landing checklist”

At: Call:

Glideslope intercept “Glideslope captured”

ACTION: Extend gear Call: “Gear indicates down” ACTION: Complete Before Landing checklist Call: “Before Landing checklist complete”

Call: Call:

“Glideslope captured”

“Flaps 40°” ACTION: Select 40° flaps Call: “Flaps selected 40°” Call: “Flaps indicate 40°”

A1-16



Cessna Citation II Technical Manual Precision Approach (continued) SIC

PIC

Call:

Call:

Call:

At: Call:

1,000 ft above DH “1,000 ft to minimums”

At: Call:

500 ft above DH “500 ft to minimums”

At: Call:

200 ft above DH “200 ft to minimums”

At:

Point of visual contact “Runway at ____ o’clock”

“Check”

“Check”

“Check”

Call: Call:

“Looking” At: Call:

Decision height “Minimums”

Note: Missed approach procedures must be executed immediately if either of these parameters have not been met at Decision Height: 1. a normal approach to the intended runway cannot be established or, 2. adequate visual reference cannot be maintained.



A1-17

Precision Missed Approach SIC

PIC At: Call:

DH “Minimums. Missed approach”

Call: “Missed approach” ACTION: Set go-around power and rotate to 10° Call: “Flaps 15°” ACTION: Select 15° flaps Call: “Flaps selected 15°” Call: “Flaps indicate 15°” At: Call: Call:


“Gear up” ACTION: Retract gear Call: “Gear selected up” Call: “Gear indicates up”

A1-18



Cessna Citation II Technical Manual Non-Precision Approach SIC

PIC Call:


At: Call:

Airspeed below 202 KIAS “Flaps 15°” ACTION: Select 15° flaps Call: “Flaps selected 15° Call: “Flaps indicate 15°”

At: Call:

Call: Call:


Localizer/course capture “Localizer/course capture”

“Gear down” “Landing checklist”


Call:

“Localizer/course capture”

At: Call:

Approach to FAF “_____ minutes/ _____ miles to FAF”

ACTION: Extend gear Call: “Gear selected down” Call: “Gear indicates down” ACTION: Complete Landing checklist Call: “Landing checklist complete”

A1-19

Non-Precision Approach (continued) SIC

PIC At: Call:

FAF “Final fix”

ACTION: Begin adequate descent rate (approx. 1,000 fpm) Call:

Call:

Call:

Call:

Call: “Final fix” ACTION: Start timing, set minimum descent altitude, and check altimeters

“Flaps 40°” ACTION: Call: Call: Call:

Select 40° flaps “Flaps selected 40°” “Flaps indicate 40°” “Altimeters check”

At: Call:

1,000 ft above MDA “1,000 ft to minimums”

At: Call:

500 ft above MDA “500 ft to minimums”

At: Call:

200 ft above MDA “200 ft to minimums”

At:

Point of visual contact “Runway at _____ o’clock”

“Check”

“Check”

“Check”

Call: Call:

“Looking” At: Call:

MDA “Minimums”

Note: Missed approach procedures must be executed immediately if either of these parameters have not been met at Minimum Descent Altitude: 1. a normal approach to the intended runway cannot be established or, 2. adequate visual reference cannot be maintained. A1-20



Cessna Citation II Technical Manual Non-Precision Missed Approach SIC

PIC At: Call:

MDA “Minimums. Missed approach”

Call: “Missed approach” ACTION: Disconnect autopilot (if engaged), set goaround power and rotate to 10° Call: “Flaps 15°” ACTION: Select 15° flaps Call: “Flaps selected 15°” Call: “Flaps indicate 15°” At: Call: Call:


“Gear up” ACTION: Retract gear Call: “Gear selected up” Call: “Gear indicates up”



A1-21

Visual Approach PIC Call:

SIC


At: Call:

At: Call:

Airspeed below 202 KIAS “Flaps 15°”

ACTION: Select 15° flaps Call: “Flaps selected 15° Call: “Flaps indicate 15°”

Abeam touchdown point “Gear down” ACTION: Extend gear Call: “Gear selected down” Call: “Gear indicates down”

At: Base turn Call: “Flaps 40°” ACTION: Set fuel flow to 400lb/ENG., begin descent Call:

A1-22

“Before Landing checklist”


ACTION: Select 40° flaps Call: “Flaps selected 40°” Call: “Flaps indicate 40°”

ACTION: Complete Before Landing checklist Call: “Before Landing checklist complete”


Cessna Citation II Technical Manual Landing SIC

PIC At:

Point of visual contact ACTION: Disengage autopilot and yaw damper

ACTION: Monitor VREF and annunciator panel, verify landing gear and flap indication At:

100 ft above touchdown Call: “100 feet” ACTION: Retract speedbrakes (as required) At: Call: At: Call:

50 ft above touchdown “50 feet”

Touchdown “Extend speedbrakes” ACTION: Extend speedbrakes Call: “Speedbrakes extended”

Call:

“Deploy thrust reversers” ACTION: Deploy thrust reversers Call: “Two deployed” At: 60 KIAS Call: “60 knots” ACTION: Reverser levers to idle reverse



A1-23


Appendix 2 Flight Profiles Table of Contents Flight Profiles Normal Takeoff ............................... A2-1 ILS Flight Director/ Autopilot Approach .......................... A2-2 Non-Precision Approach ................... A2-3 Radar Approach .............................. A2-4 Circling Approach ............................ A2-5 Steep Turns .................................... A2-6 Acceleration and Deceleration .......... A2-7 High Rate of Descent Recovery ........ A2-8 Emergency Descent ......................... A2-9 Visual Approach and Landing ......... A2-10 Flaps Up Landing .......................... A2-11 Single Engine ILS Approach Landing, and Go-Around ................. A2-12 Takeoff Engine Failure After V1 ....................................... A2-13

Roll Out

Airspeed ................. V2 + 20 KIAS MIN.

Climb power ........... SET

Accelerate

Flaps ...................... RETRACT

After-takeoff/ climb checklist ........ COMPLETE

4 Takeoff

Takeoff power .......... SET (prior to 60 KIAS)

Airspeed .................. 70 KIAS

2

Close-In Turn

Flaps ............... T/O & APPR.

Airspeed .......... V2 + 20 KIAS

Bank angle ....... 30° MAX

Rolling Takeoff

Takeoff power ....... SET

Rotate smoothly to 10°

Rate of climb ........ POSITIVE

Gear ..................... UP

3

Straight Climbout

Airspeed ............ V2 + 20 KIAS

Climb power ...... SET

Flaps ......... RETRACT (400 ft or obstacle clearance altitude)

After-takeoff/ climb checklist ........ COMPLETE

A2-1



1

Normal Takeoff

FLIGHT PROFILES 12/99

5

Approach Preparations

Radar Vectors

Airspeed .......... VREF + 20 KIAS MIN.

Flaps ............... T/O & APPR.

Hdg ................. ON

Nav ................. ARM

Apr .................. ARM

Alt ................... ON

3 FOR TRAINING PURPOSES ONLY

2

ILS Flight Director/Autopilot Approach

A2-2

1

Approach procedure ............... REVIEW

Go-around procedure .............. REVIEW

Airspeed bug ......................... SET VREF

Avionics ................................. CHECK

Before landing checklist ......... COMPLETE

Localizer ......... TRACK

Gear ................ DOWN

Flaps ............... LAND

Hdg ................. LIGHT EXTINGUISHED

Glideslope ....... CAPTURE

Alt ................... ON

6

Localizer ..... CAPTURE

Missed Approach

Go-around button ...... PUSH

Rate of climb ....... POSITIVE

Autopilot .................. OFF

Gear .................... UP

Go-around light ......... ON

Airspeed .............. V2 + 10 KIAS

Go-around power ..... SET

Command bars ......... 7

Rotate ..................... 10°

Flaps ................... RETRACT (400 ft or obstacle clearance altitude)

Flaps ....................... 15°

Mode selector ...... REPROGRAM

7 CITATION II APPENDIX 2 12/99

5 4

Autopilot/FD (Coupled Approach) ............ BEGINS DESCENT

LOC and GS Glideslope ........... COUPLED

Alt ....................... OFF

Airspeed ....... VREF MIN.

Landing

Autopilot .......... DISCONNECT

Do not "duck under"

2

Descent ........ 1,000 FPM RECOMMENDED

1




Avionics ................................. CHECK


Procedure Turn

Flaps ................... T/O & APPR.

Airspeed .............. VREF + 20 MIN.

4

3

Descent ........ LEVEL

Airspeed ....... VREF + 20 MIN.

Gear ............. DOWN

Flaps ............ LAND

Missed Approach

Autopilot ................... DISENGAGE

Runway in Sight

Go-around power ....... SET

Circling approach ....... INITIATE

Attitude .................... GO-AROUND

Airspeed ....... VREF

Flaps ........................ 15°

Rate of climb ............ POSITIVE

Gear ......................... UP

Minimum descent altitude

6 Note: Maximum use of flight director/autopilot if desired: Press APR button for VOR approach

A2-3

Press NAV button for LOC only approach Press BC button for back course localizer approach

Maintain normal descent to landing



5

Non-Precision Approach



Flaps .............. T/O & APPR.

Airspeed ......... VREF + 20 KIAS MIN.

Intercept final approach at 30° to 45°


2


At Glideslope Intercept

Airspeed .......... VREF

Gear ................ DOWN

Flaps ............... LAND


3





Avionics ................................. CHECK


Radar Approach

A2-4 1

Turn to Final

3

Abeam "Key" Point

Bank angle ....... 30° MAX.

Timing ............. BEGIN

Flaps ............... LAND

MDA ................ MAINTAIN

Airspeed .......... VREF + 10 MIN.

15 SECONDS


5

Normal descent to landing

Airspeed .......... VREF + WIND

Circling Approach


4

2

At Radio Fix

Gear ................ DOWN

Flaps ............... T/O & APPR.

Airspeed .......... VREF + 20 MIN.

45°

6

Missed Approach

Go-around power ........ SET

Flaps ......................... T/O & APPR.

Gear .......................... UP

Flaps ......................... RETRACT ON FLAP/SPEED SCHEDULE

A2-5

Caution: The visual cues used when flying normal traffic patterns can be misleading due to the lower altitudes associated with circling approaches. Common mistakes are flying the downwind leg too close to the runway, beginning the downwind-to-final turn too soon, and descending below the MDA too early.

Runway in Sight

Circling approach ....... INITIATE

Maintain visual reference to runway

Turn 45° to runway centerline





Avionics ................................. CHECK



1

Steep Turns

A2-6 1


1

Airspeed ................ 200 KIAS

2

Bank angle ............. 45° (increase thrust through 30°)

3

Altitude .................. MAINTAIN


3

Airspeed bug ........ VREF

Airspeed ............... STABILIZE AT 250 KIAS

Elevator trim ......... SET

Climb power ......... SET

Acceleration and Deceleration


1

ACCELERATE

MANEUVER 2

3

Airspeed ............... VMO

VMO aural warning sounds

Power ................... IDLE

Speed brakes ....... DEPLOY

Timing .................. BEGIN

Altitude ................. MAINTAIN

DECELERATE

3

A2-7

Airspeed ............... VREF + 30

Timing .................. STOP

STABILIZE



Airspeed .......... STABILIZE JUST BELOW VMO

Airspeed bug ........... VREF

Gear ........................ DOWN

Flaps ....................... LAND

2

Power ...................... IDLE

Rate of descent ....... CHECK

3


Recovery .............................. INITIATE

Go-around power ................... SET

Attitude ................................ GO-AROUND

Engine acceleration time ....... CHECK

Altitude lost .......................... CHECK

4


Warning: Attempting a landing from a high rate of descent is extremely hazardous. Selecting the proper flare height to arrest the descent rate at touchdown while airspeed is rapidly decreasing would be difficult if not impossible. Idle thrust approaches and high rates of descent near the ground must be avoided.

Maneuver .................... COMPLETE

Missed approach ......... EXECUTE

High Rate of Descent Recovery

A2-8 1

Passengers ................................... CHECK

1

Rapid Decompression

Emergency Descent

2

Roll wings level

ATC ......................... ADVISE

Transponder ............ 7700 EMERGENCY

Altimeter setting ...... OBTAIN

Altitude .................... CALL

MEA ....................... CONFIRM

Oxygen masks ............................... DON (100%)

Emergency descent ...................... INITIATE

Pass. O2 valve ............................... MANUAL DROP

MIC OXY MASK switch .................... SELECTED

Rapid decompression checklist ..... COMPLETE

Throttles ......................................... IDLE

Speed brakes ................................. DEPLOY

Moderate bank ................................ INITIATE

Pitch ................................................ 15° NOSE DOWN

Emergency descent checklist .......... COMPLETE

5

2000 FT Above Desired Altitude

Rate of descent ....... REDUCE SMOOTHLY

6

1000 FT Above Desired Altitude

Speed brakes .......... RETRACT

7

A2-9

Note: If structural damage is suspected, limit airspeed. Speedbrakes may be used depending on type of damage.

14,000 or MEA

Level off

Crew oxygen ......... NORMAL

Ignition ................. AS REQUIRED

Airspeed ............... LONG-RANGE CRUISE

Determine new course of action



4

Emergency Descent

R-2/8/01


3

2

Bank angle .............. 30° MAX.

Final approach ........ CLEAR

Flaps .............. LAND

Airspeed ................. VREF + 10 KIAS MIN.

Fuel flow ......... SET (400 LB/ENG)

Descent .......... BEGIN

Base Turn





Avionics ................................. CHECK



1

Abeam Runway Midpoint

Flaps .............. T/O & APPR.

Airspeed ......... VREF + 20 KIAS MIN.

Abeam Touchdown Point

4

Landing Assured

Gear ............... DOWN

5

Touchdown


Airspeed .................. VREF

Speed brake ............ DEPLOY


Thrust reverser ........ AS REQUIRED

Brakes ..................... AS REQUIRED

Visual Approach and Landing

A2-10 3

3

Base Turn


Descent ........... BEGIN (300-500 FPM)

Bank angle ....... 30° MAX.



Airspeed bug ......................... SET NEW VREF + 15/20

Avionics ................................. CHECK


Flaps Up Landing


R-6/9/00

2

Rollout Airspeed .......... NEW VREF

Altitude ............ CALL

Stabilized in slot

1

Gear ................ DOWN

Airspeed .......... New VREF + 10 MIN.

5

4


Missed Approach

Go-around power ......... SET

Attitude ...................... GO-AROUND

Rate of climb .............. POSITIVE

Gear ........................... UP

A2-11



2

Airspeed ............ VREF + 10

Gear .................. DOWN

Engine failure checklist ..................... COMPLETE

Approach procedure ......................... REVIEW

Go-around procedure ........................ REVIEW

Airspeed bug ................................... SET VREF

Avionics ........................................... CHECK

Single engine landing checklist ......... COMPLETE


3

1

Flaps ................. T/O & APPR.

Airspeed ............ VREF + 10

When Runway Assured

Flaps ........... LAND

4

Go-Around

Go-around power ......... SET

GA button ................... PRESS

Flaps .......................... T/O & APPR.

Rate of climb .............. POSITIVE

Gear ........................... UP

Flaps .......................... RETRACT ON FLAP/SPEED SCHEDULE

Single Engine ILS Approach, Landing, and Go-Around

A2-12



Roll Out

Airspeed ............................... VENR

Flaps .................................... UP

Climb power ......................... SET

Engine failure checklist .......... COMPLETE

After takeoff checklist ........... COMPLETE

5

Airspeed ............ V2 + 20 KIAS

Flaps ................. T/O & APPR.

Bank angle ......... 30°

Takeoff

Takeoff power .......... SET (prior to 60 KIAS)

Airspeed .................. 70 KIAS

2

Airspeed .......... V1/VR

Attitude ........... 10° NOSE UP

4

3

Straight Climbout

Airspeed ............................ V2 + 10 KIAS

Flaps ................................. UP (400 ft. or Airspeed ............................ ACCEL. TO VENR

Obstacle Clearance Altitude)

Engine Failure

Rudder ................... AS REQUIRED

Wings ..................... LEVEL

Climb power ...................... SET

Rate of climb .......... POSITIVE

Engine failure checklist ....... COMPLETE

Gear ....................... UP

After takeoff checklist ........ COMPLETE

A2-13



1

Close-In Turn

Takeoff Engine Failure After V1


R-6/9/00

6

Stall Recovery — Clean

A2-14

Airplane Configuration

1

Airspeed ........... DECELERATE to 160 KIAS

Power ............... 50% N1


2

Maintain altitude with pitch trim to .6 AOA

Flaps ............... UP

Gear ............... UP

5

Airspeed ......... ACCELERATE to 160 KIAS

Maintain altitude with back pressure at an approximate rate of 1 per second — PITCH UP

4

At first indication of an imminent stall (buffet or shaker):

Power .................... TAKEOFF THRUST

Release back pressure to reduce pitch to 7° NOSE UP

Airspeed ................ ACCELERATE

R-7/3/02


Stall Recovery — Departure

R-2/8/01



1

n

Flaps ............... T/O & APPR.

n

Gear ............... UP

5


n

At airspeed ..... VREF + 10 KIAS

n

Flaps ................. T/O & APPR.

n

Flaps ............... UP

n

Power ............... 50% N1

n

n

Bank angle ........ 20° either direction


3

n

Maintain altitude with back pressure at an approximate rate of 1 per second — PITCH UP

4 2

n

Maintain altitude with pitch trim to .6 AOA

At first indication of an imminent stall (buffet or shaker): n


n

Bank angle ............. ROLL WINGS LEVEL

n


n

Airspeed ................ ACCELERATE

A2-15



n

Stall Recovery — Landing

A2-16


1



Flaps ................. T/O & APPR.

Power ............... 65% N1

3

Flaps ............... FULL DOWN

Gear ............... DOWN

5

Gear ................. DOWN

2

Maintain altitude with pitch trim to .6 AOA Maintain altitude with back pressure at an approximate rate of 1° per second

Flaps ............... DOWN

Power ............. REDUCE to 65% N1

Maintain altitude with pitch trim

4

At positive rate indication on VSI .................. GEAR UP

At airspeed ..... VREF + 10 KIAS

Flaps ............... UP


At first indication of an imminent stall (buffet or shaker):

R-7/3/02



Flaps ...................... T/O & APPR.


1

Accelerate aircraft with takeoff power

Prior to V1, ABORT the takeoff

3

Maintain centerline orientation

Advise tower of the abort

Power .................... REDUCE to IDLE

Brakes ................... APPLY

Speed brakes ........ DEPLOY

Thrust reversers or drag shoot ......... DEPLOY

5

Exit runway as instructed by tower

A2-17



2

Flaps ............... T/O & APPR.

Rejected Takeoff

R-2/8/01




Appendix 3 Glossary Table of Contents Glossary of Abbreviations and Terminology .................................A3-1 Weight and Balance Terminology ........A3-1 Performance and Flight Planning Terminology .................. A3-5 Airspeed Terminology ........................... A3-7 Meteorological Terminology ................ A3-11 Powerplant Terminology ..................... A3-13

Cessna Citation ll Technical Manual Glossary of Abbreviations and Terminology The following glossary is a reference of abbreviations and terminology applicable to airplane operation, non-specific to a single aircraft manufacturer or type. Weight and Balance Terminology Standard Empty Weight The empty weight of an airplane in standard configuration as documented by the manufacturer, including the weight of unusable fuel, full engine oil, and full operating fluids. Basic Empty Weight The standard empty weight of an airplane plus all optional equipment installed, used as a basis for loading data determination. Payload The total weight of the crew, passengers, baggage, cabinet contents, and cargo. Maximum Payload The maximum permissible weight of the crew, passengers, baggage, cabinet contents, and cargo as determined by structural limitations. Usable Fuel The weight of that portion of the total fuel load which is available for consumption as determined in accordance with applicable regulatory standards. Unusable Fuel The weight of that portion of the total fuel load which is not available for consumption as determined in accordance with applicable regulatory standards, including the weight of residual fuel. Residual Fuel The weight of all undrainable fuel remaining onboard after the airplane has been defueled for weighing purposes in accordance with specified procedures.

APPENDIX 3 12/99


A3-1

Maximum Useful Load The maximum permissible combined weight of payload and usable fuel as determined by structural limitations associated with ground handling. Equivalent to maximum ramp weight minus basic empty weight. Ramp Weight The basic empty weight of an airplane plus payload and usable fuel prior to engine start. Ramp weight must not exceed the maximum ramp weight. Maximum Ramp Weight The maximum permissible ramp weight of an airplane as determined by structural limitations associated with ground handling. Equivalent to maximum takeoff weight plus the weight of fuel consumed during engine start, taxi, run-up, and takeoff roll to VR. Zero Fuel Weight The ramp weight of an airplane excluding the weight of usable fuel. Zero fuel weight must not exceed the maximum zero fuel weight. Maximum Zero Fuel Weight The maximum permissible ramp weight of an airplane excluding the weight of usable fuel as determined by associated structural limitations. Any weight in excess of this value must be fuel. Takeoff Weight The weight of an airplane upon lift-off from the runway. Takeoff weight must not exceed the maximum takeoff weight. Maximum Takeoff Weight The maximum permissible takeoff weight of an airplane as determined by associated structural limitations. Landing Weight The weight of an airplane upon runway touchdown. Landing weight must not exceed the maximum landing weight.

A3-2



Cessna Citation ll Technical Manual Maximum Landing Weight The maximum permissible landing weight of an airplane as determined by associated structural limitations. Maximum Weight The maximum permissible weight of an airplane as determined by associated design, structural, and performance limitations. Center of Gravity (CG) An imaginary point at which the weight of an airplane is considered to be concentrated, used to determine proper weight and balance for flight. CG Limits The extreme fore and aft limits of CG range within which an airplane must be operated at a specific weight. Reference Datum An imaginary vertical plane perpendicular to an arbitrary point along the airplane’s longitudinal axis from which all horizontal distances are measured to determine proper weight and balance for flight. Arm

The horizontal distance in inches from the reference datum to the center of gravity of an object or component. Arm measurements are expressed in positive (+) inches aft of the datum and negative (-) inches forward of the datum.

Moment A measurement of the rotational force about the airplane’s CG, obtained by multiplying the weight of an object or component by its respective arm (weight x arm = moment). CG Arm The horizontal distance in inches from the reference datum to the airplane’s center of gravity, obtained by dividing the total sum of all moments by the total sum of their respective weights (total moment ÷ total weight = CG Arm).

APPENDIX 3 12/99


A3-3

Approved Loading Envelope Those combinations of airplane weight and center of gravity which define the limits beyond which loading is not approved. Station The horizontal distance in inches from the reference datum to any position along the airplane’s longitudinal axis. Jack Points Specific points on the airplane identified by the manufacturer as suitable for supporting the airplane on jacks. Leveling Points Specific points on the airplane identified by the manufacturer as suitable for leveling the airplane. Tare

A3-4

Any weight indicated by a scale before a load is applied.



Cessna Citation ll Technical Manual Performance and Flight Planning Terminology Accelerate-Go Distance The distance required to accelerate an airplane to takeoff decision speed (V1) and upon experiencing an engine failure, continue accelerating to takeoff rotation speed (VR), then rotate, climb, and accelerate to engine failure speed (VEF) before reaching 50 feet AGL for the purposes of obstacle clearance. Takeoff Flight Path The minimum takeoff climb gradient required to clear obstacles beyond the end of the runway, measured horizontally from the distance at which the airplane has reached 50 feet AGL, and vertically from the surface of the runway. Accelerate-Stop Distance The distance required to accelerate an airplane to takeoff decision speed (V1) and upon experiencing an engine failure, bring the aircraft to a complete stop on the remaining runway using maximum effective braking. Maximum Effective Braking The maximum amount of braking pressure which can be applied to the toe brakes without locking the wheels. Clearway An area beyond the end of a runway not less than 300 feet on either side of the extended center line of the runway, at an elevation no greater than the end of the runway, clear of all fixed obstacles, and under the control of the airport authorities. Balked Landing An aborted landing. Balked Landing Transition Speed The minimum speed at which transition to a balked landing climb should be attempted from 50 feet AGL.

APPENDIX 3 12/99


A3-5

Demonstrated Crosswind The maximum 90° crosswind component under which adequate control of the airplane during takeoff and landing was actually demonstrated during certification testing. Maneuvering Fuel The usable fuel available for all airplane configurations, provided the maximum sideslip duration is not exceeded. Aerobatic Maneuver An intentional maneuver involving an abrupt changing of the airplane attitude, abnormal attitude, or abnormal acceleration, beyond the requirements for normal flight. Fix

Any geographic location which can be identified by visual reference or radio navigational aids.

Route Segment A specific portion of a route identified by geographic or radio navigational fixes. Minimum Obstacle Clearance Altitude (MOCA) The minimum airplane altitude permissible between fixes which meets necessary obstacle clearance requirements for a specific route segment. Minimum En Route Altitude (MEA) The minimum airplane altitude permissible between radio navigational fixes which assures adequate signal reception while meeting necessary obstacle clearance requirements for a specific route segment.

A3-6



Cessna Citation ll Technical Manual Airspeed Terminology IAS

Indicated Airspeed: Airplane speed as displayed on an airspeed indicator.

CAS

Calibrated Airspeed: Airplane speed as displayed on an airspeed indicator corrected for instrument error. CAS is equal to true airspeed in standard atmosphere at sea level.

TAS

True Airspeed: Airplane speed relative to undisturbed air as displayed on an airspeed indicator corrected for altitude, temperature, and compressibility.

KIAS Indicated Airspeed expressed in knots. KCAS Calibrated Airspeed expressed in knots. KTAS True Airspeed expressed in knots.

APPENDIX 3 12/99

GS

Ground Speed: Airplane speed relative to the ground.

M

Mach: Airplane speed expressed in numerical proportion to the speed of sound under standard atmospheric conditions (e.g., Mach 1 is equivalent to the speed of sound (approximately 642 KTS at sea level)).

G

Acceleration Force: A measurement of force expressed in numerical proportion to the force of gravity (e.g., 1 G is equivalent to the force gravity acting on an object at rest (approximately 32ft./sec2), which in turn determines the weight of the object). An acceleration force of 2 Gs acting on an airplane is therefore equivalent to doubling its actual weight.

V

Velocity: A measurement of airplane speed relative to specific operating limitations.

V1

Takeoff Decision Speed: The maximum speed to which an airplane may accelerate before deciding to continue or abort takeoff according to accelerate-stop and accelerate-go performance limitations.


A3-7

A3-8

VR

Takeoff Rotation Speed: The speed at which an airplane may rotate according to take-off weight and takeoff performance limitations.

V2

Takeoff Safety Speed: The speed to which an airplane must accelerate after rotation before reaching 50 feet AGL for the purposes of obstacle clearance.

VEF

Engine Failure Speed: The speed to which an airplane must accelerate after rotation before reaching 50 feet AGL for the purposes of obstacle clearance with one engine inoperative.

VX

Best Angle-of-Climb Speed: The speed at which an airplane will deliver the greatest gain in altitude in the shortest possible distance over ground.

VXSE

Best Single-Engine Angle-of-Climb Speed: The speed at which an airplane will deliver the greatest gain in altitude in the shortest possible distance over ground with one engine inoperative.

VY

Best Rate-of-Climb Speed: The speed at which an airplane will deliver the greatest gain in altitude in the shortest possible period of time.

VYSE

Best Single-Engine Rate-of-Climb Speed: The speed at which an airplane will deliver the greatest gain in altitude in the shortest possible period of time with one engine inoperative.



Cessna Citation ll Technical Manual VMCA

Air Minimum Control Speed: The minimum speed at which the airplane is directionally controllable in flight with one engine inoperative as established by FAA certification procedures. Certification is based on the following conditions: One engine inoperative and secured, operative engine set for takeoff power, 5° bank towards the operative engine, gear up, flaps in takeoff/approach position, and rearward most C.G. At certain variations of aircraft weight and operating altitude, stall conditions can be encountered at speeds above VMCA as established by the certification procedure described above. Under these conditions, stall speed (VS) must be regarded as the air minimum control speed.

APPENDIX 3 12/99

VS

Stalling Speed: The minimum steady speed at which the airplane is controllable in flight.

VSO

Landing Configuration Stalling Speed: The minimum steady speed at which the airplane is controllable in flight when configured for landing.

VSSE

Intentional One-Engine-Inoperative Speed: A speed limitation above both VMCA and VS established to provide a margin of lateral and directional control when one engine is intentionally failed for the purposes of pilot training. Intentional failing of one engine below this speed is not recommended.

VMO

Maximum Operating Limit Speed: A speed limitation which may not be deliberately exceeded in normal flight operations. Also expressed as MMO in reference to the equivalent Mach limitation.

VNO

Maximum Structural Cruising Speed: A speed limitation which may not be exceeded except with caution in smooth air conditions.

VNE

Never Exceed Speed: A speed limitation which may not be exceeded under any condition.


A3-9

A3-10

VA

Maneuvering Speed: The maximum speed at which application of full available aerodynamic control will not structurally over stress the airplane.

VF

Design Flap Speed: The maximum airplane speed at which the wing flaps may be safely extended to a specific position.

VFE

Maximum Flap Extended Speed: The maximum speed at which the airplane may be safely operated with the wing flaps extended to a specific position.

VLE

Maximum Landing Gear Extended Speed: The maximum speed at which the airplane may be safely operated with the landing gear extended.

VLO

Maximum Landing Gear Operating Speed: The maximum airplane speed at which the landing gear may be safely extended or retracted.

VMCG

Ground Minimum Control Speed: The minimum speed at which the airplane is directionally controllable on the ground with one engine inoperative as established by FAA certification procedures.



Cessna Citation ll Technical Manual Meteorological Terminology °C

Degrees Celsius: A measurement of temperature relative to a scale which indicates the freezing point of water as 0° and the boiling point of water as 100° under standard atmospheric conditions. Equivalent to (°F - 32) x 0.566.

°F

Degrees Fahrenheit: A measurement of temperature relative to a scale which indicates the freezing point of water as 32° and the boiling point of water as 212° under standard atmospheric conditions. Equivalent to (°C X 1.8) + 32.

Dry Adiabatic Lapse Rate The rate at which air temperature decreases with increasing altitude under standard atmospheric conditions without losing or gaining heat energy (approximately 3°C/5.5°F per 1000 feet). Standard Lapse Rate The average rate at which air temperature decreases with increasing altitude without losing or gaining heat energy (approximately 2°C/3.6°F per 1000 feet), used to determine freezing levels in the atmosphere relative to surface air temperatures. IOAT Indicated Outside Air Temperature: A measurement of static air temperature as displayed on an OAT indicator not corrected for instrument error. OAT

Outside Air Temperature: A measurement of true outside air temperature obtained from ground meteorological sources, or by correcting IOAT for the compressibility effects of airspeed and altitude, and used as a basis for airplane performance determination.

Temperature Compressibility Effects IOAT error corresponding to airspeed and altitude.

APPENDIX 3 12/99


A3-11

ISA

International Standard Atmosphere: A standardized reference measurement of atmospheric conditions used to determine airplane performance in nonstandard conditions, assuming the following factors: 1. Air is a dry, perfect gas. 2. Atmospheric pressure at sea level is 29.92 inHg/ 1013.2 mb. 3. Air temperature at sea level is 15°C/59°F. 4. Air temperature decreases with increasing altitude by approximately 3°C/5.5°F per 1000 feet to 35,750 feet and zero above that altitude.

Pressure Altitude A measurement of altitude above the Standard Datum Plane, a theoretical level where atmospheric pressure is equal to standard sea level pressure. Equivalent to indicated altitude when the altimeter is set to 29.92 inHg/ 1013.2 mb. Density Altitude A measurement of pressure altitude used to determine airplane performance in nonstandard atmospheric conditions. Equivalent to true altitude under standard atmospheric conditions. Indicated Altitude An indicated measurement of altitude above the atmospheric pressure to which the altimeter is set. Equivalent to true altitude when the altimeter is correctly set to the local reported pressure. True Altitude An indicated measurement of altitude above Mean Sea Level (MSL), a level where atmospheric pressure is equal to sea level pressure as computed from station pressure corrected for nonstandard conditions, assuming the indicating altimeter is correctly set and zero instrument error. Equivalent to pressure altitude and density altitude under standard atmospheric conditions.

A3-12



Cessna Citation ll Technical Manual Station Pressure A measurement of actual atmospheric pressure (barometric pressure) at field elevation. Local Reported Pressure A measurement of atmospheric pressure corresponding to sea level pressure as computed from station pressure corrected for nonstandard conditions. Used to determine altimeter correction settings required to indicate true altitude. Powerplant Terminology Igniter A device used to start the burning of the fuel/air mixture in a combustion chamber. Impeller The main rotor of a radial compressor, which increases the velocity of the air which it pumps. Plenum A duct, housing or enclosure used to contain air under pressure. Stator A row of stationary airfoils that direct the airflow between the rows of rotor blades. Turbine A rotating device actuated either by reaction or impulse (or a combination of both), and used to transform some of the kinetic energy of the exhaust gases into shaft horsepower to drive the compressor(s) and accessories. Turbofan A type of gas turbine that converts heat energy into core and bypass thrust.

APPENDIX 3 12/99


A3-13


Appendix 4 Annunciators Table of Contents Annunciators Airplanes (550-0550 and after) ....... A4-1 Annunciators Airplanes (550-0550 and earlier) .... A4-5 Airplanes (550-0482, 550-0485~550-0505 ...................... A4-5

Cessna Citation ll Technical Manual Annunciators Citation II advisory lights (annunciators) are designed to provide an easily interpreted representation of both normal and abnormal system conditions. The majority of annunciators are collocated on an annunciator panel located on the center instrument panel. Annunciator Panel Airplanes 550-0550 and after

AC FAIL

GEN OFF LH

RH

BATT O’TEMP

INVERTER FAIL 1

CABIN ALT 10000 FT

EMERG PRESS ON

2

SPEED BRAKE EXTEND

1

BLD AIR GND

OIL PRESS WARN

FUEL LOW LEVEL

FUEL LOW PRESS

LH

LH

LH

RH

POWER BRAKE LOW PRESS

RH

FUEL FLTR BYPASS LH

RH

RH

FUEL BOOST ON LH

RH

AIR DUCT O’HEAT

ANTI SKID INOP

F/W SHUT OFF

ACM O’PRESS

DOOR NOT LOCKED

LH

HYD FLOW LOW LH

RH

HYD LOW LEVEL HYD PRESS ON

ENGINE ANTI-ICE LH

RH

SURFACE DEICE

P/S HTR OFF LH

RH

W/S AIR O’HEAT

RH

KEY:

ACM O’PRESS

Red

requires immediate attention, hazardous condition exists.

Amber requires attention, possible dangerous condition exists. 2

BLD AIR GND

APPENDIX 4 12/99 R-6/9/00

4

3

SPARE

White

SPARE


safe or normal configuration, routine action.

A4-1

Annunciator Test If the system is functioning normally, all annunciators and the MASTER WARNING lights should illuminate when the TEST selector knob on the lower left instrument panel is positioned to "ANNUN" and the BATT switch is in the "ON" position.

AC FAIL

BATT O'TEMP

Steady illumination: battery temperature over 145°F; flashing: battery temperature over 160°F. Illumination of either annunciator triggers the master warning system, which will illuminate the MASTER WARNING light.

CABIN ALT 10,000 FT

Cabin pressure altitude is above 10,000 feet. Illumination of either annunciator triggers the master warning system, which will illuminate the MASTER WARNING light.

OIL PRESS WARN

Oil pressure is below safe limits (35 PSI) in left or right engine. Illumination of light also triggers the master warning system, which will illuminate the master warning light.

LH

RH

FUEL LOW LEVEL LH

LH

RH

HYD FLOW LOW LH

Fuel quantity in left and or right tanks has reached a level of 169 to 219 pounds.

RH

FUEL LOW PRESS

A4-2

AC power bus voltage is above 130 VAC or below 90 VAC. Illumination of light triggers the master warning system, which will illuminate the MASTER WARNING light.

Low fuel supply pressure to engine-driven pump. Primary pump failure will automatically initiate boost pump operation as long as FUEL BOOST switch is in the NORM position. Light remaining on indicates failure of both pumps. Left and/or right hydraulic system flow is below approximately 0.35 to 0.55 gallons per minute.

RH



Cessna Citation ll Technical Manual

ENGINE ANTI-ICE LH

RH

SURFACE DEICE

GEN OFF LH

RH

INVERTER FAIL 1

Left and/or right engine inlet, stator, or inboard wing anti-ice has failed.

2

EMER PRESS ON BLD AIR GND

Illuminates twice during the 12- second surface deice boot cycle to indicate proper boot inflation pressure. Tail boot inflation cannot be checked visually from the cockpit. Left and/or right generator is not connected to the airplane bus. Illumination of both annunciators triggers the master warning system, which will illuminate the MASTER WARNING light. Number 1 or 2 inverter output voltage is above 130 VAC or below 90 VAC. Inverter failure triggers AC FAIL annunciator. Resetting MASTER WARNING will extinguish the AC FAIL annunciation unless both INVERTER FAIL lights are illuminated.

Emergency pressurization has been manually selected or automatically activated by an air cycle machine overheat. High flow rate of bleed air has been selected from the right engine for ground operation of the air conditioner. Power brake hydraulic pressure is low.

POWER BRAKE LOW PRESS

FUEL FLTR BY PASS LH

APPENDIX 4 12/99 R-6/9/00

Bypass on the AMBER fuel low pressure light pressure is low in left and right systems.

RH


A4-3

FUEL BOOST ON LH

RH

HYD LOW LEVEL HYD PRES ON P/S HTR OFF LH

Hydraulic reservoir is low. Hydraulic system is pressurized. Left and/or right pilot heat is off.

RH

SPEED BRAKE EXTEND

Left and right speedbrakes are fully extended.

AIR DUCT O'HEAT ACM O'PRESS

Ventilation duct temperature exceeds safe limits.

ANTI SKID INOP DOOR NOT LOCKED

Anti skid system is inoperative.

F/W SHUTOFF LH

Air cycle machine pressure is over 42 psi.

Cabin door, aft compartment access door lock, or either nose baggage door lock is/are not locked. Left and/or right fuel and hydraulic shutoff valves are closed.

RH

W/S AIR O'HEAT

A4-4

Electric power has been applied to the left and/ or right fuel boost pump.

Bleed air to the windshield exceeds safe temperature limits, or >5 PSI in the plumbing with the system off.


CITATION II APPENDIX 4 12/99 R-6/9/00

Cessna Citation ll Technical Manual Annunciators Citation II advisory lights (annunciators) are designed to provide an easily interpreted representation of both normal and abnormal system conditions. The majority of annunciators are collocated on an annunciator panel located on the center instrument panel. Annunciator Panel Airplanes 550-0550 and earlier

AC FAIL

BATT O’HEAT

ANTI-SKID INOP

W/S AIR O’HEAT

PWR BRK PRESS LO

AIR DUCT O’HEAT

HYD PRESS ON

DOOR NOT LOCKED

HYD LEVEL LO

P/S HTR OFF

1

2

CAB ALT 10,000 FT

L HYD PRESS LO

R HYD PRESS LO

L OIL PRESS LO

R OIL PRESS LO

BLEED AIR GND/HI

L ENG ICE FAIL

R ENG ICE FAIL

L GEN OFF

R GEN OFF

EMER PRESS ON

L F/W SHUTOFF

R F/W SHUTOFF

L FUEL PRESS LO

R FUEL PRESS LO

R PRECOOL FAIL

L FUEL BOOST ON

R FUEL BOOST ON

FUEL FILT BYPASS

L FUEL LEVEL LO

R FUEL LEVEL LO

ACM EJECTOR ON

3

SURF DEICE

L PRECOOL FAIL

4

SPD BRAKE EXTENDED

Airplanes 550-0482, 550-0485~550-0505

1

KEY:

ACM O’PRESS

Red

requires immediate attention, hazardous condition exists.

Amber requires attention, possible dangerous condition exists. 2

BLD AIR GND

APPENDIX 4 12/99 R-6/9/00

4

3

SPARE

White

SPARE


safe or normal configuration, routine action.

A4-5

Annunciator Test If the system is functioning normally, all annunciators and the MASTER WARNING lights should illuminate when the TEST selector knob on the lower left instrument panel is positioned to "ANNUN" and the BATT switch is in the "ON" position. AC FAIL

BATT O'HEAT

AC power bus voltage is above 130 VAC or below 90 VAC. Illumination of light triggers the master warning system, which will illuminate the MASTER WARNING light. Steady illumination: battery temperature over 145°F; flashing: battery temperature over 160°F. Illumination of either annunciator triggers the master warning system, which will illuminate the MASTER WARNING light.

CABIN ALT 10,000 FT

Cabin pressure altitude is above 10,000 feet. Illumination of either annunciator triggers the master warning system, which will illuminate the MASTER WARNING light.

L HYD PRESS LO

Left and/or right hydraulic system flow is below approximately 0.35 to 0.55 gallons per minute.

R HYD PRESS LO L OIL PRESS LO R OIL PRESS LO ANTI-SKID INOP

A4-6

Oil pressure is below safe limits (35 PSI) in left or right engine. Illumination of light also triggers the master warning system, which will illuminate the master warning light.

Anti-skid system is inoperative.

W/S AIR O'HEAT

Bleed air to the windshield exceeds safe temperature limits, or >5 PSI in the plumbing with the system off.

BLEED AIR GND/HI

High flow rate of bleed air has been selected from the right engine for ground operation of the air conditioner.


CITATION II APPENDIX 4 12/99 R-6/9/00


ACM O'PRESS L ENG ICE FAIL

Air cycle machine pressure is over 42 psi. (550-0482, 550-0485~550-0505) Left and/or right engine inlet, stator, or inboard wing anti-ice has failed.

R ENG ICE FAIL L GEN OFF R GEN OFF PWR BRK PRESS LOW

Left and/or right generator is not connected to the airplane bus. Illumination of both annunciators triggers the master warning system, which will illuminate the MASTER WARNING light. Power brake hydraulic pressure is low.

AIR DUCT O'HEAT

Ventilation duct temperature exceeds safe limits.

EMER PRESS ON

Emergency pressurization has been manually selected or automatically activated by an air cycle machine overheat.

L F/W SHUTOFF

Left and/or right fuel and hydraulic shutoff valves are closed.

R F/W SHUTOFF L FUEL PRESS LO R FUEL PRESS LO HYD PRESS ON

APPENDIX 4 12/99 R-6/9/00

Low fuel supply pressure to engine-driven pump. Primary pump failure will automatically initiate boost pump operation as long as FUEL BOOST switch is in the NORM position. Light remaining on indicates failure of both pumps. Hydraulic system is pressurized.


A4-7

DOOR NOT LOCKED

Cabin door, aft compartment access door lock, or either nose baggage door lock is/are not locked.

ACM EJECTOR ON

Air cycle machine ejector on during ground operation with both throttles below 85% or with pressure applied to wheel brakes. (5500550 and earlier)

BLD AIR GND

High flow rate of bleed air has been selected from the right engine for ground operation of the air conditioner. (550-0482, 5500485~550-0505)

L PRECOOL FAIL

Air from precooler above 282°, +6 or -6°C. (550-0481 and earlier, -0483, -0484) Spare on (550-0482, 550-0485~550-0505)

R PRECOOL FAIL L FUEL BOOST ON

Electric power has been applied to the left and/or right fuel boost pump.

R FUEL BOOST ON HYD LEVEL LO P/S HTR OFF

SURF DEICE

SPD BRAKE EXTENDED

A4-8

Fluid in the hydraulic reservoir is low.

Left and/or right pilot heat is off.

Illuminates twice during the 12- second surface deice boot cycle to indicate proper boot inflation pressure. Tail boot inflation cannot be checked visually from the cockpit. Left and right speedbrakes are fully extended.




FUEL FILT BYPASS L FUEL LEVEL LO

Left, right or both fuel filters approaching, or actually being bypassed due to fuel filter restriction. Fuel quantity in left and or right tanks has reached a level of 169 to 219 pounds.

R FUEL LEVEL LO

APPENDIX 4 12/99


A4-9

Cessna Citation II TM [PWD] Dec 1999

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