United nit ed State States s Army Ar my Av iation Warfi Warfi ghting ght ing Center enter Fort Ruck er, Alabama January 2008
UH-60A STUDENT HANDOUT UH-60A Automatic Flight Control System (AFCS) 4748-6
TERMINAL LEARNING OBJECTIVE: ACTION: Identify operational characteristics of the UH-60A/L Automatic Flight Control System (AFCS). CONDITION: In a classroom given appropriate training devices, a list of characteristics, TM 1-1520-237-10, TM 1520-237-10CL, Aircrew Training Manual, and the student handout. (IAW ) TM 1-1520-237-10, TM 1-1520-237-10CL, Aircrew Training Manual, STANDARD: In accordance with (IAW and the student handout. SAFETY REQUIREMENTS REQUIREMENTS:: Use care when operating training aids and/or devices. RISK ASSESSMENT LEVEL: Low ENVIRONMENTAL ENVIRONMENTAL CONSIDERATIONS: CONSIDERATIONS: It is the responsibility of all soldiers and DA civilians to protect the environment from damage. EVALUATION: You must answer 7 out of 10 questions correctly to receive a "GO" on this scoreable unit.
(2) Dynamic stability (short-term stability) is the tendency to resist oscillation. c. Reasons for system. (1) Helicopters are not as stable as fixed-wing aircraft. External forces will cause any aircraft to change attitude, airspeed, or heading. However, fixed-wing aircraft can be designed to return to the desired attitude when the external force is removed (static stability). (2) Helicopters have little or no tendency to return to the desired attitude. Since the rotor head moves with the aircraft, helicopter will assume a new attitude and not the desired one. (3) A helicopter hangs under a rotor head (like a pendulum) and swings or oscillates under rotor head. Dynamic stability prevents porpoise in pitch, rock in roll, and fishtail in yaw. (4) Pilot workload is much higher in a helicopter than in a fixed-wing aircraft. (a) Constant correction is required to maintain attitude, airspeed, and heading. (b) Constant correction is required to minimize oscillation. (c) Instrument flying is extremely difficult. (d) Accuracy of weapons is very poor due to lack of stability. (e) Passengers are uncomfortable because of lack of stability. d. Purpose of the AFCS is to enhance the stability and handling qualities of the helicopter. (1) It provides static stability by holding-(a) Airspeed. (b) Attitude.
(d) FPS provides limited flight control positioning which assists in maintaining helicopter pitch and roll attitudes, airspeed, heading, and turn coordination. LEA RNING STEP/ACTIVITY 2: Identify the operational characteristics of the stabilator system. a. Description. The helicopter has a variable angle of incidence stabilator to enhance handling qualities. The automatic mode of operation positions the stabilator to the best angle of attack for the existing flight conditions. After the pilot engages the automatic mode, no further pilot action is required for stabilator operation. Two stabilator amplifiers receive airspeed, collective stick position, pitch rate, and lateral acceleration information to program the stabilator through the dual electric actuators. b. Operation. The stabilator is programmed to-(1) Align stabilator and main rotor downwash in low speed flight to minimize nose up attitude resulting from downwash. (2) Provide collective coupling to minimize pitch attitude excursions due to collective inputs from the pilot. A collective position sensor detects pilot collective displacement and programs the stabilator for a corresponding amount of movement to counteract for pitch changes. This coupling of stabilator input for collective displacement is automatically phased in between 30 and 60 KIAS (3) Decrease angle of incidence with increased airspeed to improve static stability.
(4) Provide sideslip to pitch coupling to reduce susceptibility to gusts. When the helicopter is out of trim in a slip or skid, pitch excursions are also induced as a result of the canted tail rotor and downwash on the stabilator. Lateral accelerometers sense this out of trim condition and signal the stabilator amplifiers to compensate for the pitch attitude change (called lateral to sideslip to pitch coupling). Nose left (right slip) results in the trailing edge programming down. Nose right produces the opposite stabilator reaction. (5) Provide pitch rate feedback to improve dynamic stability. The rate of pitch attitude change of the helicopter is sensed by a pitch rate gyro in each of the two stabilator ampli fiers and is used to position the stabilator to help dampen pitch excursions during gusty wind conditions. A sudden pitch up due to gusts would cause the stabilator to be programmed trailing edge down a small amount to induce a nose-down pitch to
e. Position indicators (2). (1) Description. Two identical stabilator position indicators allow the pilot and copilot to monitor stabilator position. (2) Location. The indicators, are located on the instrument panel (just below the pilot's and copilot's Airspeed Indicators). (3) Operation. They contain OFF flags that are removed from view when power is applied and an indicator pointer that displays stabilator angles between 10 up and 45 down. °
°
NOTE: Although the indicators are marked 10 up and 45 down, do not confuse this with the actual stabilator travel of 9 up and 39 down. °
°
°
°
f. Stabilator position placards. (1) Description. These lighted decals show maximum airspeed in relation to stabilator position for stuck stabilator. (2) Location. A lighted decal is located beside each stabilator indicator. (3) Function. Decals serve to limit airspeed for varying stabilator angles. They are marked with degrees on the left side and airspeed on the right side. g. Amplifiers (2). (1) Description. The two identical stabilator control amplifiers contain electronic components, pitch rate gyros used in automatic stabilator control, and relays that allow manual control. (2) Purpose. Stabilator control amplifiers provide lateral acceleration, airspeed discrete signals, and pitch rate signals to the SAS/FPS computer and automatic stabilator control and provide relays that allow manual control of the stabilator.
(2) Purpose. Airspeed and air data transducers supply airspeed signals to the stabilator amplifiers for the stabilators automatic mode operation and the SAS/FPS computer. (3) Location. Airspeed and air data transducers are located on the cockpit bulkheads (forward of and below the instrument panel). The air data transducer is on the right side of the helicopter; the airspeed transducer is on the left side. (4) Operation. The airspeed transducer supplies an airspeed signal to the Number 1 stabilator control amplifier. The air data transducer supplies airspeed signals to the Number 2-stabilator-control amplifier. The air data transducer also supplies airspeed and pressure altitude signals to the command instrument system. Both transducers supply airspeed signals to the SAS/FPS computer and receive air pressure inputs from the instrument pitot-static system.
(1) Description. The two lateral accelerometers are Identical. (2) Purpose. They provide the stabilator control amplifiers with DC electrical signals that represent the relationship of the helicopter bank angle to its turn rate. (3) Location. The lateral accelerometers are mounted in the cabin ceiling. The Number 1 accelerometer is located on the left side of the helicopter, and the Number 2 accelerometer is located on the right side. On some aircraft, they are located on the left and right sides of the stabilator amplifiers. (4) Operation. (a) The Number 1 accelerometer receives excitation from the left side of the helicopter and supplies signals to the Number 1 amplifier. (b) The Number 2 accelerometer receives excitation from the right side of the helicopter and provides signals to the Number 2 amplifier. k. Control panel. (1) Description. The stabilator/flight control panel contains switches and relays used to control the stabilator system. (2) Location. The panel is located on the center of the lower console. (3) Switches. (a) Cyclic mounted stabilator slew-up switch. The preferred method of manually slewing the stabilator is to use the cyclic mounted stabilator slew-up switch. NOTE: Use of the cyclic mounted stabilator slew-up switch should be announced to the crew to minimize cockpit confusion. (b) MAN SLEW. This switch is spring loaded to the center (OFF) lever lock switch. It allows the pilot to
WARNING: MAKE SURE ALL PERSONNEL AND EQUIPMENT ARE CLEAR OF THE STABILATOR BEFORE APPLYING ELECTRICAL POWER TO THE HELICOPTER. (a) The automatic mode disengages if positions of the stabilator actuators disagree by an amount that depends on forward airspeed. Shutdown occurs at 10 of error at 0 knots and 4 of error at 150 knots. (b) The automatic mode also disengages when the MAN SLEW switch is moved to the UP or DN position or the cyclic mounted stabilator slew up switch is used. °
°
(2) Operation (general). (a) Both stabilator amplifiers receive AC and DC power when electrical power is applied to the helicopter. They, in turn, supply VDC to the airspeed and air data transducers, collective stick position transducers, lateral accelerometers, and actuator feedback potentiometers. (b) Airspeed and air data transducers produce DC output signals that increase for forward airspeeds between 30 and 180 knots. (c) Collective stick position transducers produce a DC output signal that is proportional to the collective stick's position. (d) A centered collective equals 0 volts. A down collective results in a positive signal; an up collective causes the output to go negative. (Output voltage is 1.34 volts per inch of stick displacement from the center.) (e) Pitch rate signal. 1. Each stabilator amplifier contains a pitch rate gyro. These produce DC signals that represent the helicopters rate of pitch attitude change. 2. Pitch rate signals also are routed out of each stabilator amplifier through filters. 3. The number 1 signal is used by the SAS amplifier. The number 2 signal is used by the SAS/FPS computer. b. Manual mode operation.
WARNING: COVERS ON PITOT TUBES WOULD CAUSE THE STABILATOR TO REMAIN IN THE TRAILINGEDGE DOWN POSITION WITH NO CAUTION LIGHTS OR AURAL WARNING. LEA RNING STEP/ACTIVITY 4: Identify the operational characteristics of the stability augmentation system (SAS). a. Analog stability augmentation system. The analog system (SAS 1) is one of two SAS systems. It operates independently of the SAS/FPS computer and provides the aviator with redundancy. b. Digital stability augmentation system (SAS 2). SAS 2 is operated by the SAS/FPS computer. It is independent of SAS 1 and provides stability in the same axis using the same actuators. c. Purpose--provides dynamic stability in the pitch, roll, and yaw axes. d. Components. (1) Actuators (3). (a) Description. Three actuators are provided for the pitch, roll, and yaw channels. (b) Purpose. Actuators link SAS electronic components to the helicopters mechanical flight control system. (c) Location. The actuators are mounted on the transmission deck at the pilot assist servo. (d) Operation. 1. Hydraulic pressure. An electrohydraulic servo control flapper valve allows 3,000 psi hydraulic pressure from the Number 2 hydraulic system, or backup system, to operate the actuator. 2. Electrical inputs. Electrical inputs are supplied by the analog SAS amplifier and digital SAS/FPS computer. The actuators respond to electrical inputs and hydraulic pressure by moving control linkages that change rotor blade angles without moving cockpit controls.
(2) SAS Amplifier. (a) Description. It contains a rate gyro that serves as a sensor for yaw SAS. (b) Purpose. It processes aircraft sensor signals to develop command signals that are applied to SAS actuators when SAS 1 is engaged. (c) Location. The SAS amplifier is located on the floor of the electronic compartment's "tunnel" (below the instrument panel).
(b) SAS 1 switch energizes the SAS amplifier for SAS 1 coils, and the SAS 2 switch energizes circuitry in the SAS/FPS computer for SAS 2 coils. (c) SAS 1 and SAS 2 each have a 5-percent authority for a total of a 10- percent authority. (d) Failure of SAS 1 will not be indicated visually but may be known by the lack of aircraft stability. (e) Failure of sensors controlling SAS 2 will cause the failure advisory panel on the flight control panel to illuminate. (5) Indicators. In case of loss of actuator pressure, or if both SAS 1 and SAS 2 are off, the SAS OFF caution will appear. LEA RNING STEP/ACTIVITY 5: Identify the operational characteristics of the digital automatic flight controls system (AFCS/SAS2). a. Description. The digital AFCS will provide the following: (1) Cyclic stick and pedal trim. (2) Stability augmentation (SAS 2). (3) Autopilot functions (FPS).
(c) Power. The computer is powered by the Number 2 AC primary bus and Number1 and Number 2 primary DC busses. (2) Vertical gyros. Two vertical gyros, located in the nose electronic compartment, produce attitude signals for pitch and roll. (3) ASN 43 compass system directional gyro. This gyro is located in the nose electronic compartment and produces a heading signal. (4) Airspeed and air data transducers. (a) Location. These transducers are located in the forward section of the cockpit, above the tail rotor pedals, with airspeed on the left side and air data on the right side. (b) Operation. They operate from the pitotstatic system and are powered by the stabilator system. (c) Signal use. 1. Yaw trim. 2. Pitch FPS--airspeed hold. 3. Yaw FPS--automatic turn coordination logic 4. Yaw SAS 2. (5) Rate gyros. Rate gyros in the pitch, roll, and yaw channels control the rate of any change. (6) Lateral accelerometers. These are used to produce signals when the helicopter slips or skids. (7) Collective stick position transducers. These are located at the flight controls mixer. They produce a signal that represents collective stick's position. The signals are used for yaw trim, collective to yaw coupling, and pitch FPS airspeed hold.
(b) SAS amplifier circuits modify the rate signal to provide desired aircraft and system response. (c) SAS 1 gain depends on the engage condition of SAS 2. If SAS 2 is ON, SAS 1 operates at a normal gain. SAS 1 and SAS 2 supply the actuator with signals to stabilize the helicopter. If SAS 2 is OFF, the SAS 1 amplifier doubles its gain. This provides larger signals for a given aircraft movement to help compensate for the loss of SAS 2. (d) The servo valve driver supplies current to operate the SAS actuator. LEA RNING STEP/ACTIVITY 6: Identify the operational characteristics of the trim system. a. Trim actuators are connected to flight control linkages in a manner that allows them to move cockpit controls. (1) Pitch--to fore and aft cyclic stick linkage. (2) Roll--to lateral cyclic stick linkage. (3) Yaw--to pedal linkage. b. Pitch actuator is hydraulic. (1) It receives 1,000 psi from pilot assist module. (2) Pressure is supplied only if trim is ON and computer detects no pitch trim malfunction. c. Roll and yaw actuators are electromechanical. d. FPS provides 100 percent control authority using trim actuators. e. All actuators operate at a limited rate (about 10 percent per second). They are self-limited and also limited by computer. f. All actuators contain override springs which--
(b) Actuator remains very close to trimmed position. (c) Actuator holds the cyclic stick fixed. j. Cyclic stick trim release button causes computer to release trim. (1) Hydraulic pressure is removed from pitch trim actuator. (2) Feedback signal is zeroed. (3) Stick is free to move. (4) Trimmed or referenced position becomes changed. (5) Stick is trimmed to position where button is released. (6) Cyclic stick trim switch changes computers command signal to actuator. (a) Actuator will drive. (b) Stick will drive at about 0.4 inches per second. k. Roll trim actuator has a built-in servo system. (1) System holds actuator and cyclic stick fixed when trim is ON and not released. (2) It is referenced through clutches and centering springs in actuators when trim is OFF or released. (3) Computer supplies command to the actuator when cyclic stick trim switch is used. NOTE: When the cyclic trim switch is slewed left and right, while on the ground and with FPS on, the cyclic will return to center. l. Yaw trim actuator has a built-in servo system.
(1) Pedals or cyclic stick become free in affected axis trim and FPS is inoperative. (2) During pitch failure, actuator hydraulic pressure is removed. (3) During roll or yaw failure, the actuator clutch is disengaged. (4) Computer senses failure when actuator position signal (feedback) does not agree with command signal. (a) Trim failure advisory light will be ON. (b) Trim failure and flight path stabilization caution lights will be ON. n. Computer failure cause affected trim axi s to shut down. (1) Pitch channel. Flight indications are the same as for an actuator failure. (2) Roll or yaw. (a) Stick or pedals will be free. (b) Trim failure and flight path stabilization caution lights will be ON. (c) Computer (CPTR) failure advisory light will be ON. (d) Trim and FPS in affected axis will be inoperative.
LEA RNING STEP/ACTIVITY 7: Identify the operational characteristics of the flight path stabilization system (FPS). a. It provides the autopilot functions, which are to --
d. Pitch channel provides attitude hold (below 60 Kts) and attitude & airspeed hold (above 60 Kts)(airspeed hold overrides attitude hold), by trimming cyclic stick to position required to maintain pilot’s desired attitude and airspeed.. NOTE: Airspeed hold goes OFF for 30 seconds after collective stick is moved through the 60-percent position below 100 knots. e. Yaw channel provides heading hold or automatic turn coordination/automatic turn logic by driving pedals as required to maintain pilots desired heading or balanced flight. It holds heading unless pedal trim is released or automatic turn coordination/automatic turn logic is engaged. (1) Pedal trim is released when airspeeds are less than 60 knots by pedal microswitches or when airspeeds are greater than 60 knots by pedal microswitches and the cyclic trim release. (2) Automatic turn coordination/automatic turn logic operates only at airspeeds above 60 knots and engages when the pilot slews (by use of the cyclic stick trim switch) left or right about ½ inch and a roll attitude of about 1.5 degrees or more. f. FPS failures. (1) Directional gyro failure disables heading hold. (a) Pedal trim and automatic turn coordination remains functional. (b) Flight path stabilization caution light is ON. (c) Gyro failure advisory light is ON. (2) Airspeed sensor failure disables automatic turn coordination. (a) Flight path stabilization caution light is ON. (b) Airspeed failure advisory light is ON.
DIGITAL AUTOMATIC FLIGHT CONTROL UNITS EFFECTS OF MALFUNCTIONS
STABILATOR SYSTEM
R/S pitot tube
ver 1.1 11/00
COLLECTIVE STICK POSITION TRANSDUCERS (on Mechanical Mixing Unit)
AIR DATA TRANSDUCER
80 KTS or LESS: Amps usehigher of thetwo A/S signals. 80KTS & ABOVE: Eachampuses it's own A/S signal. (if oneamp is sensing below 80 with theother sensing above 80, thelow oneprograms for80 KTS)
STABILATORCONTROLS MANSLEW UP PNL LTS
AUTO CONTROL
TEST
O F F
ON DN SAS1
AUTOFLIGHTCONTROL SAS 2 TRIM
R E S E T
to both stab. amps.
ON
ON
ON
BOOST ON
R E S E T
ON
FAILURE ADVISORY CPTR SAS 2
ACCL CLTV
TRIM RGYR
A/S GYRD
POWER ONRESET
#2 Stab. Amp. pitch rate gyro Stab position Limit & A/S signals
#1 Stab. Amp. pitch rate gyro
FPS
PNL LTS
position commands
R E S E T
MAN SLEW: Sprungto off position, selects manual mode when movedup or down,alows manual positioning of stabilator to any position.
L/S pitot tube
USE OF EXCHANGED A/S SIGNALS
#2 Lateral Accelerometer
NO.2 ACTUATOR (moves only whenconvertedpower is present via#2 DCPRI bus)
2
position & limit feedback
NO.1 ACTUATOR (can be moved with battery power via DC ESS bus)
1
#1Lateral Accelerometer
9 deg
TEST: Functions@ 60KTS& below, signals #1amp. to retracts #1 actuator, fault monitoring detects miscompare (IAW range chart) between #1/#2 actuators and switches to manual mode (audible & visual warnings activate).
STABILATOR
AUTOCONTROL: Switches from manual to automatic mode.
AIRSPEED TRANSDUCER
0 deg RANGEOF MOTION
both actuators: @48 deg.
CYCLIC SLEW: Selects manual mode& moves stab.up only. (stabilator up movement stops when switch is released)
singleactuator: @ 35deg. (may be less depending on stab position when other actuator waslost) 0
10 STAB POS
STAB POS
DEG
DEG
Copilot
Pilot (Used for checks & rigging)
39 deg
stabilator actuator position
0
40 30
airspeed
150
Miscompare Range Chart
FLIGHT PATH STABILIZATION SYSTEM MAIN SLEW UP O F F
AUTO CONTROL
TEST
R E S E T
ON SAS 1
SAS 2
TRIM
ON
ON
ON
ON
R E S E T
FPS ON
FAI LU RE
BOOST
ver 1.1 10/97
Input sensor B,I,M,P, and Q
STABILATOR CONTROLS
CPTR S A S 2 TRIM RGYR
A DV IS ORY ACCL CLTV A/S
GYRO
R E S E T
TRIM FAIL
FLT PATH STAB
POWERON RESET
fault monitoring / advisory
SAS 2 TRIM FPS S E N S O R S
1
GYRO
SAS VALVE
COMP RATE
V E RT
D IR
AIR SPEED
COLL STICK
LAT ACCEL
2
1 SAS2 sensor signals: Q,G,J,K,N,O,H, and R 2 TRIM sensor signals: E and B
3
TRIM ACT
FAN FAIL PROC A
3 FPS sensor sign als: B,E,J,K,N, and O
GND
FAN TEST
NORM PROC B
J139 J12
J111
NOTE: sig nals A,F,L,M, and R are for fault monitoring only
J121
Au to mati c Fl igh t Co nt ro l In put sig nals A B C D E F G H I
air da ta tran sdu cer s ign al air speed transducer signal #1 lateral accelerometer signal #2 lateral accelerometer signal #1 collective position sensor #2 collective position sensor r ol l r at e g yr o s ig nal #2 yaw rate gyro signal #1 pitch rate gyro signal (#1 stab amp)
J #2 pitc h rat e gyr o si gnal (#2 stab amp ) K heading signal (gyro magnetic compass) L pilots vert. gyro pitch signal (att. ind. sys.) M pilots vert. gyro roll signal (att. ind.sys.) N copilots vert. gyro pitch signal (att. ind. sys.) O copilots vert. gyro roll signal (att.ind.sys.) P #1 y aw r at e g yr o s ig nal (f ro m s as am pl if er ) Q #1 filtered lateral acceleromtere signal (#1 stab amp) R #2 filtered lateral acceleromtere signal (#2 stab amp)
STABILATOR SYSTEM SIGNALS IN # 1 CLTV # 1 ACCL A/S X-DUCTER
#1 Stab. Amp . pitch rate gyro
SIGNALS OUT
SIGNALS OUT
#1 CLTV #1 ACCL A/S X-DUCER PITCH RATE GYRO
#2 CLTV #2 ACCL AIR DATA PITCH RATE GYRO
2
TEST FAULT MONITORING CIRCUIT STABILATOR CONTROL PANEL STAB. MODE SELECT (AUTO) STAB. MAN. SLEW STAB. TEST
1
SIGNAL S IN #1 PITCH RATE GYRO #2 VERTICAL GYRO #1 ACCL. A/S X-DUCER
TEST DRIVE / ACT ONLY STABILATOR
SIGNALS OUT
PITCH RATE GYRO VERTICAL GYRO
YAW RATE GYRO
#2 Stab. Amp. pitch rate gyro
TEST FAULT MONITORING CIRCUIT
SAS / FPS SYSTEM SAS 1 AMP.
SIGNALS IN # 2 CLTV # 2 ACCL A/D X-DUCER
#1 ACCL. A/S X-DUCER YAW RATE GYRO
SAS ACTUATORS
TRIM ACTUATORS PITCH
ROLL
YAW
SIGNAL OUT
PITCH
fault monitoring / advisory
ROLL
TRIM SIGNALS IN
SAS 2 YAW
PILOT'S CYCLIC AND PEDAL TRIM CONTROL SWITCHES
#1 CLTV A/S X-DUCER
TRIM
(COPILOT'S) #1 VERTICAL GYRO
NOTE: #1 AND #2 ACCL (FILTERED) AIR DATA X-DUCER #1 AND #2 ACCL YAW RATE GYRO FILTERED SIGNALS A RE PROCESSED IN THE STABILATOR AMP ROLL RATE GYRO
A/S X-DUCER #1 CLTV #1 ACCL. #1 PITCH RATE GYRO
FPS
ASN 43 COMPASS (HEADING)
#2 PITCH RATE GYRO GYRO COMP RATE
AIR SPEED
TRIM ACT
SAS VALVE
VERT
DIR
COLL STICK
LAT ACCEL
FAN FAIL PROCA GND
FAN TEST
NORM PROCB
J139 J12
J111
D-31
J121
