United States Army Aviation Center Fort Rucker, Alabama APRIL 2007

United States Army Aviation Center Fort Rucker, Alabama APRIL 2007 Student Handout TITLE: AH-64D Flight Controls System FILE NUMBER: 11-0915-3 PROPONENT FOR THIS STUDENT HANDOUT IS: Commander, Aviation Training Brigade ATTN: ATZQ-ATB-ACD Fort Rucker, Alabama 36362-5000 FOREIGN DISCLOSURE STATEMENT: The materials contained in this student handout have been reviewed by the course developer, in coordination with the USAAVNC foreign disclosure authority. This student handout is releasable to foreign military students without restrictions. D-3

Action: Operate the AH-64D flight control system Conditions: Given an AH-64D helicopter or training device and TM 1-1520-251-10 Standards: In accordance with the TM 1-1520-251-10 The AH-64D utilizes a hydromechanical flight control system with an electrical Back Up Control System (BUCS) to enhance aircraft response during normal and emergency flight operations. All flight control inputs are in response to actions within and outside of the crewstations to provide a stable platform to reduce pilot fatigue and ensure mission success. D-4

A. ENABLING LEARNING OBJECTIVE 1 After this lesson you will: ACTION: CONDITIONS: Identify the characteristics of the primary flight control system Given a written test without the use of student notes or references STANDARDS: In accordance with the TM 1-1520-251-10 1. Learning Step / Activity 1. Identify the flight control system characteristics a. Flight control systems (1) Collective control system (2) Cyclic control system (3) Upper flight controls (4) Directional control system (5) Stabilator control system D-5

b. Primary flight controls (1) Cyclic and collective servoactuators (2) Cyclic and collective mixer assembly (3) Main rotor swashplate (4) Directional servoactuator (5) Tail rotor swashplate (6) Stabilator D-6

c. Collective control system The collective control system provides control input to the main rotor system for vertical control of the helicopter. A collective stick is located to the left of the crew seat in each crewstation. It consists of a grip and friction adjust mechanism mounted in a collective support base. The collective pivots up and down from the base of the canted bulkhead to provide input to the collective control linkage. (1) Collective grip (a) The quick-change grips are identical in both crewstations and are used as hand holds to move the collective stick up and down. (b) The collective grip is divided into two grips, the mission grip and the flight grip. These grips contain switches that are used to provide crew inputs to the various helicopter systems. (2) Friction adjustment twist grip A friction adjust mechanism is located on the stick below the grip. Rotation of the friction grip counterclockwise will cause the friction adjust mechanism to increase friction drag on the friction guide. D-7

d. Cyclic control system The cyclic control system provides mechanical control input to the main rotor system for longitudinal and lateral flight control of the helicopter. (1) The pilot (PLT) and copilot/gunner (CPG) cyclic sticks are interconnected by control rods and are identical in all aspects except that the CPG s cyclic can be manually stowed and unstowed. (a) To unstow the CPG s cyclic, a finger lock release lever must be pulled rearward and the cyclic stick pulled upward until fully extended and locked. (b) To stow the CPG cyclic, the release handle must be pushed forward and the cyclic will collapse to the stow position and lock. (2) The CPG s cyclic is fully functional when stowed or unstowed. It is highly recommended that the CPG s cyclic be in the unstowed position during flight. (3) Movement of the cyclic control stick is transmitted through the longitudinal and lateral mechanical control linkages, longitudinal and lateral servoactuators, and upper flight controls to change the pitch angle of the main rotor blades as required. D-8

e. Upper flight control system (1) The upper flight control system receive and combine control inputs from the collective and cyclic control sticks for vertical, longitudinal, and lateral flight of the helicopter. The mixer is mounted above the main rotor support structure. (2) Upper flight control system components (a) Collective Bellcrank: receives input from the collective servo actuator and transmits vertical motion to the non-rotating swashplate. (b) Lateral Bellcrank: mounted on the collective bellcrank, receives input from the lateral servoactuator and transmits lateral motion to the non-rotating swashplate (c) Lateral Links: transmit the lateral motion from the lateral bellcrank to the nonrotating swashplate. (d) Longitudinal Bellcrank: mounted on the collective bellcrank receives input from the longitudinal servoactuator and transmits longitudinal motion the non-rotating swashplate. (e) Longitudinal Links: transmit the longitudinal motion to the rear longitudinal bellcrank. (f) Rear Longitudinal Bellcrank: transmits longitudinal motion from the longitudinal bellcrank to the torque link. (g) Torque Link: transmits longitudinal motion to the non-rotating swashplate. Also provides a fixed mounting point for the non-rotating swashplate. D-9

(3) Swashplate assembly The swashplate assembly consists of a stationary swashplate and a rotating swashplate. Components of the swashplate assembly are the self-aligning spherical and slider bearing and the double row ball bearing. The swashplate assembly is mounted on the static mast between the mixer assembly and the main rotor head assembly. (4) The swashplate assembly receives control inputs from the two lateral links and the torque link and transmits the control inputs to the main rotor. (5) Scissors assembly The two scissors assemblies transmit rotational power to the rotating swashplate from the main rotor head. (a) Each scissors assembly is made up of a lower arm and upper arm to allow the scissors assembly to shorten or lengthen as the swashplate moves in response to control inputs. (b) In the event one scissors assembly fails, the other can carry the full drive load. (6) Pitch change links The pitch change links transmit control inputs from the swashplate assembly to the main rotor head, controlling the pitch angle of the main rotor blades. D-10

f. Directional control system The directional control system provides control input to the tail rotor system for directional heading (yaw) and anti-torque control of the helicopter through the use of interconnecting control rods. (1) The tops of the directional control pedals also function as the mechanical input to activate the main landing gear brakes (commonly known as toe brakes). (2) The directional control pedals are identical in both crewstations. (3) The directional control pedals can be independently adjusted (crewstation) fore and aft to accommodate individual crew comfort. CAUTION The crewmember feet must be firmly placed on the directional control pedals prior to moving the release mechanism. (a) Adjustment of the directional control pedals is accomplished by the reach adjustment mechanism. Rotating the reach adjust mechanism handle to the left releases the pedals to simultaneously move aft under spring pressure or forward when pushed by the crewmember. (b) Movement of the directional control pedals fore and aft is transmitted through the directional mechanical control linkage, directional servoactuator, bellcrank, swashplate assembly and pitch links to simultaneously change the pitch angle of the four tail rotor blades. D-11

g. Directional control assembly (1) The directional control assembly transmits control movements from the directional servoactuator to the tail rotor. The directional controls are mounted on the static support between the tail rotor gearbox and the tail rotor. (2) Directional control system components (a) Bellcrank assembly The bellcrank transmits control movement from the directional servoactuator to the tail rotor swashplate and prevents rotation of the non-rotating swashplate. The bellcrank is installed between the directional servoactuator support and is connected to the directional servoactuator and the tail rotor swashplate. (b) Swashplate assembly The swashplate assembly is mounted on the tail rotor gearbox static support. It is made up of a non-rotating swashplate and rotating swashplate. 1) Non-rotating swashplate The non-rotating swashplate receives linear control inputs from the bellcrank assembly and transmits these inputs to the rotating swashplate. 2) Rotating swashplate The rotating swashplate receives input from the non-rotating swashplate, converts it to a rotating input, and transmits it to the tail rotor blades. (c) Drive links 1) The two drive links (primary and secondary) transmit rotational power to the rotating swashplate from the tail rotor head. D-12

2) Each drive link is made up of two arms to allow the link to shorten or lengthen as the swashplate moves in response to control inputs. 3) In the event one drive link fails; the other can carry the full drive load. (d) Pitch links The four non-adjustable pitch change links transmit control movements from the swashplate assembly to the tail rotor blades. The pitch links are mounted between the rotating swashplate assembly and the pitch horns of the tail rotor blades. D-13

h. Hydraulic Servoactuators: (1) All control movements are assisted by hydraulic servoactuators except for the Stabilator. Stabilator movement is provided by two tandem linear electrical actuators mounted back-toback. (2) Upper flight control system combines control inputs received from the collective, longitudinal, and lateral hydraulic servoactuators for vertical, longitudinal, and lateral flight control of the helicopter. (3) The primary hydraulic system provides hydraulic power to the collective, cyclic, and directional control servoactuators. (4) The utility hydraulic system provides redundant hydraulic power to the servoactuators and utility hydraulic system components. (5) The utility accumulator will provide emergency hydraulic power to the servoactuators, should both the primary and utility hydraulic systems fail. D-14

i. Stabilator control system The stabilator control system is used to optimize the pitch attitude of the airframe at various airspeeds. At low airspeeds, a selectable Nap-of-the-Earth/Approach (NOE/A) mode increases over-the-nose visibility for the aircrew. (1) The stabilator control system consists of: (a) Stabilator The stabilator is a wing shaped airfoil attached to two pivot points at the base of the vertical stabilizer. The stabilator improves longitudinal handling characteristics and forward visibility of the helicopter for improved flightcrew effectiveness. Gurney flaps are installed on the trailing edge to improve highspeed handling. (b) Stabilator actuator The stabilator actuator positions the stabilator. The actuator is mounted to the bottom of the stabilator and to a bulkhead attachment in the tail boom. (c) Stabilator transducer The stabilator transducer converts the mechanical motion of the horizontal stabilator into an electrical signal that is supplied to the Flight Management Computer (FMC), which supplies information to the Multipurpose Displays (MPDs) for a reference to the crewmembers. It is mounted to the base of the vertical stabilizer and linked mechanically to the horizontal stabilizer. D-15

(2) Collective Nose Up/Nose Down/Reset (NU/ND/Reset) switch The NU/ND/Reset three position switch allows the crew to operate the stabilator in the manual mode. The switch is mounted on the collective flight control grip in each crewstation and is used to activate manual operation or reset automatic mode following manual operation. (a) NU: Pushing the switch aft causes the trailing edge of the stabilator to move upward. (b) ND: Pushing the switch forward causes the trailing edge of the stabilator to move downward. (c) Reset: Pushing the switch down will reset the stabilator into the auto mode. D-16

j. The stabilator has three operational modes: automatic, Nap-of-the-Earth/Approach (NOE/A), and manual. (1) A digital degree of the stabilator position is displayed on the ENG SYS page. (a) In the auto-mode (default mode on power-up), the FMC will automatically position the stabilator based on collective position, airspeed (KTAS), and pitch rate. In this mode, the limits of travel are 5 trailing edge up to +25 trailing edge down. (b) An AUTO STAB advisory is presented on the Up-Front Display (UFD) only if the automatic mode fails. (2) The NOE/A mode can be activated through the MPD ACFT UTIL page on the MPD. The NOE/A mode is defaulted OFF on power up. In the NOE/A mode, the FMC positions the stabilator to +25 trailing edge down. If airspeed exceeds 80 KTAS, the FMC will revert to the auto-mode. D-17

(3) In the manual mode a stabilator icon is displayed on the FLT page. (4) The manual mode is selectable below 80 KTAS, or is available in the event the automatic mode fails. (a) The NU/ND/Reset switch on the pilot and CPG collective control flight grip is used to activate the manual mode and control the stabilator. (b) The MAN STAB ON advisory is displayed on the UFD during manual stabilator operation. (c) The manual mode limits are -10 to +35. k. Stabilator failure automatic or manual operation (1) The stabilator icon will be displayed in white during manual operation, yellow when the stabilator is detected failed, and red when the True Airspeed (TAS) is above the displayed limit. (2) If a failure occurs and the stabilator position is still valid (known) an icon is displayed on the FLT page with the nominal KTAS for the current stabilator angle. (3) If the position of the stabilator cannot be determined by the System Processor (SP) and the FMC an icon is displayed on the FLT page at an angle of approximately 31 with 90 knots displayed below the icon and a? mark above. (4) The STAB FAIL caution is displayed on the UFD if both automatic and manual modes fail. D-18

CHECK ON LEARNING: 1. Where is the switch located to manually operate the stabilator or to reset the stabilator to the automatic mode from the manual mode? 2. The Stabilator position can be determined in the automatic mode by referring to what page on the Multipurpose Display (MPD)? 3. The stabilator icon in the manual mode is displayed on what MPD page? 4. What symbol is placed above the stabilator icon on the FLT page if the System Processor and the Flight Management Computer (FMC) cannot determine the position of the stabilator. 5. What component allows the crewmember to manually adjust the friction on the collective? D-19

B. ENABLING LEARNING OBJECTIVE 2 After this lesson you will: ACTION: Identify the characteristics of the Force Trim System CONDITIONS: Given a written test without the use of student notes or references STANDARDS: In accordance with the TM 1-1520-251-10 and TC 1-251 1. Learning Step / Activity 1. Identify the flight control system characteristics a. Force trim release switch The force trim release switch is a five-position, spring-to-center, momentary switch used to activate and release: (1) Force trim momentary interrupt (R) (2) Attitude hold (AT) (3) Altitude hold (AL) (4) All modes Disengage (D) (5) Center position Force trim ON b. FMC release switch The FMC release switch will immediately disengage the Stability Command Augmentation System (SCAS) when pressed. This allows the crewmembers to manually disengage all SCAS channels without utilizing the MPD during emergency situations. D-20

c. The force trim system (1) The force trim system provides an artificial force gradient feel to reduce crewmember fatigue by utilizing a feel spring and a magnetic brake assembly. (a) The pilot s controls are the only controls with the force trim spring attached. The CPG has the same feel because his controls are connected to the pilot s. (b) Each of the pilot s control axis; lateral (roll), longitudinal (pitch) and directional (yaw) contain a feel spring and magnetic brake. (c) The force trim may be selected on or off using the ACFT UTIL page FMC TRIM pushbutton. (d) The force trim function is automatically engaged when DC power is available. WARNING When force trim is selected off from the aircraft (A/C) utility page, no force gradient will be available to position the cyclic. Hands off cyclic capability will NOT be available. (2) The force trim may be momentarily selected off using the force trim/hold mode switch. When actioned forward. (a) Releases the magnetic brakes to allow the pilot to establish a new control reference position. (b) Allows the Stability Augmentation Subsystem (SAS) actuator to recenter. D-21

CHECK ON LEARNING: 1. What are the five positions of the force trim switch located on the cyclic in each crewstation? 2. The force trim may be selected on or off using what Multipurpose Display (MPD) page? D-22

C. ENABLING LEARNING OBJECTIVE 3 After this lesson you will: ACTION: Identify the characteristics of the Flight Management Computer CONDITIONS: Given a written test without the use of student notes or references STANDARDS: In accordance with the TM 1-1520-251-10 1. Learning Step / Activity 1. Identify the flight control system characteristics a. Flight Management Computer (FMC) (1) The FMC commands the position of the flight control actuators based upon control inputs from the crewstation Linear Variable Differential Transducers (LVDT), Automatic Roller Detent Decoupler (ARDD), aircraft rate information from the Embedded Global Positioning System Inertial (EGI), Helicopter Air Data System (HADS), radar altimeter, and pressure sensors. (2) The FMC is located in the right hand Extended Forward Avionics Bay (EFAB). (3) The FMC performs the monitoring and Built-In-Test (BIT) functions for the electronic portion of the flight controls and stabilator system. (4) The FMC provides the software logic and interface necessary to perform the following electronic functions for the flight controls system: (a) Attitude Hold 1) Attitude Hold 2) Velocity Hold 3) Position Hold 4) Heading Hold D-23

(b) Altitude Hold 1) Radar 2) Barometric (c) Automatic Stabilator Control (d) Stability Command Augmentation Subsystem (SCAS) b. The FMC continuously monitors: (1) Pilot and CPG collective, cyclic and directional controls position utilizing LVDTs in each control axis, one for each crewstation. (2) Position of the SAS actuator of each of the four servoactuators. (3) Position of the RAM LVDTs in each of the four servoactuators. (4) Position of the stabilator. (5) Pitot static system. c. Processes inputs from: (1) Internal navigation system, air data sensors and radar altimeter via the multiplex bus. (2) Dedicated switches and the MPDs. d. Provides outputs to: (1) Open the SAS solenoid valves to position the servoactuator, resulting in pitch changes of the main and/or tail rotor blades. (2) Release/engage the force trim magnetic brakes. (3) Drive the stabilator actuator to position the stabilator D-24

CHECK ON LEARNING: 1. What component monitors the Automatic Roller Detent Decoupler (ARDD) through the Rotary Variable Differential Transducer (RVDT)? 2. The FMC performs the monitoring and Fault Detection/Fault Isolation (FDI) for what two systems? D-25

D. ENABLING LEARNING OBJECTIVE 4 After this lesson you will: ACTION: CONDITIONS: Identify the characteristics of the Stability and Command Augmentation System (SCAS). Given a written test without the use of student references or notes STANDARDS: In accordance with TM 1-1520-251-10 1. Learning Step / Activity 1. Stability and Command Augmentation System a. Controlled by the FMC, the SCAS has three functions: Stability Augmentation System (SAS), Command Augmentation System (CAS), and the hold modes. The SCAS initializes in the on condition for all axis, and is selectable on or off for each axis. b. The Stability and Command Augmentation System (SCAS) augments initial control motion and provides attitude rate damping for stability augmentation. D-26

c. SCAS engagement (1) SCAS is automatically engaged following EGI alignment. When engaged (on), the circle next to the pushbutton label will fill. (2) Each control axis may be manually selected on or off using aircraft utility page FMC PITCH, ROLL, YAW, and collective (COLL) pushbuttons. D-27

d. Stability Augmentation System (SAS) When an outside force causes a deviation in the helicopter attitude, the Inertial Navigation Units (INUs) sense the change and provide a signal to the FMC that is proportional to the rate and direction of change. (1) The FMC provides a command signal to the Electro-Hydraulic Valve (EVH) on the servoactuator. The EHV ports hydraulic pressure to the stability augmentation actuator. The stability augmentation actuator is displaced from the manual servo valve, which causes the power piston to move in a direction to correct the deviation. (a) As the power piston moves, the mechanical feedback acting on the servo input arm forces the manual servo valve to realign with the stability augmentation actuator. (b) This closes the hydraulic ports to the power piston and power piston movement stops. (c) This action is what creates the limiting authority of SAS. (2) The SAS provides rate damping in all axis and turn coordination at airspeeds greater than 40 knots. It also dampens external forces (wind gust) to the airframe to stabilize the helicopter. (a) Rate damping 1) Yaw rate damping is active at airspeeds from zero to 40 knots when accelerating and below 30 knots when decelerating. D-28

2) Lateral and longitudinal rate damping is active at all airspeeds. (b) Turn coordination Turn coordination is a function of roll attitude, airspeed, and sideslip. It is designed to keep the aircraft in trim ( ball centered ) as follows: 1) From cruise down to 30 knots airspeed when decelerating. 2) Above 40 knots when accelerating. (c) External forces (wind gust) damping 1) The FMC compares the INU inputs to the stick position LVDTs and, if there was no stick input, moves the respective servoactuator to counter the uncommanded airframe movement. 2) This function is active when the SCAS is on for that control axis. (3) The SAS ability to move the flight control is limited in each axis. This limit is referred to as authority, and is expressed as a percentage of the servoactuators total travel. (a) Collective, directional, lateral and aft longitudinal authority is 10%. (b) Forward longitudinal authority is 20%. D-29

(4) SAS commands cause the SAS actuator within the servoactuator to reposition to produce the required flight control input. (a) If the required input is large, the SAS cannot compensate, causing the SAS to run out of authority, or saturate. When this occurs, the message SAS SATURAT will be displayed on the UFD along with the advisory tone. (b) To reset the SAS actuator, press and hold the force trim/hold mode switch in the release (R) position for a minimum of 3 seconds. This re-centers the force trim and allows time for the SAS actuator to reposition to within 5% of center. (c) Momentary selection ( bumping ) of the switch re-centers the force trim, but does not allow enough time for the SAS actuator to re-center. (5) When the force trim is released, it causes the SAS command to decay or washout, repositioning the SAS actuator to the center position. This recentering is not instantaneous; it depends on the position of the SAS actuator. D-30

e. Command Augmentation System (CAS) CAS is defined as an increase of response to the flight control movement. The CAS provides a uniform aircraft response for a given control input at all airspeeds. CAS is active in each control axis when the respective axis SCAS is selected on. (1) The FMC monitors the pilot's position LVDTs for CAS functions as follows: (a) When the flight controls are moved by the pilot or CPG via mechanical linkages to the servoactuators and cause the helicopter to respond to the pilot or CPG input. There is a very slight delay in time between the mechanical control inputs in the crewstations and the actual movement of the servoactuator pilot input lever. This is because of the time it takes to close all the necessary mechanical tolerances in the rod end bearings, push pull tubes, and the bellcranks before the mechanical movement reaches the servoactuator. (b) At the same time, movement of the flight controls in the crewstations causes the LVDT's to develop signals proportional to the direction and amount of flight control movement. (c) Once the LVDT is moved, the signal is developed and present at the FMC immediately (electron current flows at the speed of light). The FMC uses the signal from the LVDTs to compute a command signal proportional to the flight control movement. The computation is instantaneous. (d) When the SAS actuator and the manual servo valve are in their neutral position, the hydraulic ports to the power piston D-31

close and stop power piston movement. The power piston will stay in the new position until the pilot makes another manual flight control movement and/or a CAS or SAS signal is applied to the EHV. (2) The squat switch input disables the yaw CAS when the aircraft is on the ground to prevent over controlling during taxi operations. The SAS is not disabled on the ground and will continue to provide rate damping. D-32

f. Hold modes The AH-64D hold modes include; attitude (position, velocity, attitude, heading) and altitude (radar, barometric). These modes are designed to provide limited hands-off flight and decrease pilot workload. The hold modes are engaged/disengaged using the force trim/hold mode switch AT or AL positions, and disengaged using the D position. WARNING Hold modes operate through the SCAS system, and as such provide only limited capability. The system is not capable of maintaining the selected flight condition in all flight conditions. The pilot shall continually monitor the aircraft s flight condition to ensure safe operation. WARNING Fly the aircraft to a stabilized trimmed state before engaging attitude, hover, or altitude hold. Alternatively, fly the aircraft with the trim gradient (FTR switch released) until the aircraft is stabilized at the desired trimmed state. Failure to follow this procedure may result in undesirable and/or unsuspected aircraft responses. D-33

g. Attitude hold Attitude hold is designed to maintain the helicopter in a constant attitude. (1) Attitude hold mode utilizes pitch and roll SCAS to hold X and Y body attitudes and yaw SACS to provide heading hold or turn coordination as appropriate. Attitude hold is active above 30 knots ground speed, depending if the helicopter is accelerating or decelerating. (a) Accelerating, attitude hold is active from 40 knots ground speed up to V NE. (b) Decelerating, attitude hold is active from V NE down to 30 knots ground speed. (c) In addition to ground speed, the following parameters apply: 1) Pitch attitude < ± 30. 2) Roll attitude < ± 60. 3) Pitch and roll rates < 5 /second. If the roll rate is < 3 /second, the reference will be set to zero (wings level). (d) The attitude hold modes can be engaged any time that the SCAS channels and force trim is on. This mode provides position, velocity, attitude, heading hold based on aircraft ground speed and force trim interrupt settings. D-34

(e) The UFD displays the advisory ATT HOLD and a rounded cornered box is displayed around the airspeed on the MPD flight page. (f) If the hold mode is disengaged, the advisory tone will sound, the UFD advisory ATT HOLD will blank, and the box around the airspeed changes to white, flashes for three seconds, then blanks. (g) Failure of the hold mode will have the same indications as disengagement, plus the UFD will display the advisory ATTHLD FAIL. D-35

h. Position hold Position hold is designed to hold the helicopter over a selected point on the ground. (1) Below 5 knots ground speed, the position hold mode uses pitch and roll SCAS to maintain precise aircraft positioning over a selected geographical location. (2) Having position hold engaged in combination with altitude hold provides three-dimensional hover hold. (3) A new position is automatically captured whenever the pilot momentarily releases the force trim or stops providing control inputs within the allowable ground speed and vertical rate. (4) The pilot is provided with UFD advisories, along with an advisory tone, upon position hold disengagement or a failure. (5) If the helicopter drifts more than one rotor diameter (48 feet) from the engagement point, the advisory tone will sound and the UFD will display the advisory HOVER DRIFT. i. Heading hold Heading hold is designed to maintain the helicopter at a constant heading. (1) Heading hold mode is automatically engaged when attitude hold is on and the: (a) Pedals displaced < 0.10 inches from the trim position. (b) Yaw rate is < 3 /second. (c) Roll cyclic displaced < 0.25 inches from trim position. (d) Aircraft roll is <± 3 from level. D-36

Pitch and Roll cyclic displacement > 0.25 in. Collective displacement > 0.50 in. Pedal displacement depends on the Attitude Hold submode as follows: Position hold > 0.10 in. Velocity hold > 0.20 in. Attitude hold > 0.30 in. NOTE: Flight Control Breakout Values NOTE: There is no pilot symbology or tone indications of heading hold engagement or disengagement. j. Velocity hold (1) Velocity hold is designed to maintain the helicopter at a constant velocity and utilizes pitch and roll SCAS to hold X and Y body velocities and heading hold to maintain heading. Velocity hold is active between 5 and 40 knots, depending if the helicopter is accelerating or decelerating. (a) Accelerating, velocity hold is active from 5 up to 40 knots ground speed. (b) Decelerating, velocity hold is active from 30 down to 5 knots ground speed. This is to avoid the hold mode from engaging and disengaging at one set ground speed. D-37

(2) Having velocity hold engaged in combination with altitude hold provides three-dimensional velocity hold. (3) A new position is automatically captured whenever the pilot momentarily releases the force trim or stops providing control inputs within the allowable ground speed and vertical rate. (4) The MPD symbology during velocity hold consists of a velocity vector. The velocity vector indicates the magnitude and direction of the captured velocity. D-38

k. Altitude hold (radar and barometric) Altitude hold is provided through the collective SCAS and may be selected at any velocity. There are two modes of operation for altitude hold. (1) Radar altitude hold is engaged if the EGI s ground speed is less than 40 knots and the radar altimeter is less than 1428 feet. (2) At speeds greater than 40 knots the barometric altitude submode is automatically designated. Barometric altitude hold is also used if the radar altimeter is off or failed. (3) The Altitude hold submode may be engaged at any airspeed. Vertical velocity must be less than 200 fpm at a hover, or 400 fpm at cruise for the mode to engage. NOTE: This is not a terrain following mode. The radar altimeter provides only distance from the ground directly below the aircraft and does not provide any approaching terrain variation information. (4) If significant drift from the commanded altitude occurs, an ALT DRIFT advisory message will be displayed on the UFD accompanied by an advisory tone. The drift threshold in barometric hold is 100 feet. In radar altitude hold, the threshold is linear from 5 feet deviation at 10 feet to 100 feet deviation at 1400 feet. (5) A "home plate" symbol is placed next to the rate of climb indicator and indicates the value of the altitude being held. If the altitude mode disengages for any reason an advisory tone is initiated, and the home plate symbol will flash white for 3 seconds before being removed. (6) The UFD displays the advisory RAD HOLD or BAR HOLD, depending on the mode engaged. (7) If the helicopter deviates significantly from the reference altitude, the advisory tone will sound and the UFD will display the advisory altitude (ALT) DRIFT. D-39

(8) If the hold mode is disengaged, the advisory tone will sound, the UFD advisory RAD HOLD or BAR HOLD will blank, and the home plate symbol changes to white, flashes for three seconds, then blanks. (9) Failure of the hold mode will have the same indications as disengagement, plus the UFD will display the advisory RADHLD FAIL or BARHLD FAIL. (10) Attitude hold will automatically disengage when: (a) Collective is displaces more than 0.50 inch from the reference position. (b) Rotor speed is < 97% or > 104%. (c) Either engine torque exceeds 100%. (d) Either engine TGT exceeds 867 C. D-40

CHECK ON LEARNING: 1. What is Attitude hold designed to accomplish? 2. The best procedure to follow after receipt of a SAS SATURAT message is to fly the helicopter to the desired state (hover, velocity, or attitude), then re-center the Stability Augmentation Subsystem (SAS). How do you re-center the SAS from the crewstations? 3. What are the three functions of Stability and Command Augmentation System (SCAS)? D-41

E. ENABLING LEARNING OBJECTIVE 5 After this lesson you will: ACTION: Identify the characteristics of the Backup Control System (BUCS) CONDITIONS: Given a written test without the use of student notes or references STANDARDS: In accordance with the TM 1-1520-251-10 1. Learning Step / Activity 1. Introduction to BUCS D-42

a. The BUCS is an emergency fly-by-wire backup control system which permits continued controlled flight in the event of jammed or severed mechanical controls in any or all of the four control axis. (1) All other current helicopter backup systems require a transfer of controls to the non-flying pilot. (2) BUCS is continuously monitored to protect the crew from any failures. b. BUCS engagements (1) BUCS can be engaged for two reasons: (a) If the FMC senses a mistrack between both the pilot and CPG s LVDT inputs and the servoactuator RAM LVDT outputs. This is called a dual mistrack and will cause an automatic BUCS engagement. If there is only a mistrack between the pilot s LVDT and the CPG s LVDT, this will cause a BUCS FAIL message to be displayed and no BUCS engagement will occur. (b) If a control linkage gets jammed, the crewmember will have to decouple the ARDD and manually engage the BUCS. (2) If the mechanical controls are severed, the FMC automatically engages BUCS after sensing a mistrack between the PLT/CPG control stick LVDTs and the RAM A / RAM B LVDTS on the hydraulic servoactuators. (3) If a jam has occurred, the FMC will engage BUCS when one of the crewmembers disconnects his controls from the mechanical control linkage by decoupling via the ARDD. D-43

c. Flight control servoactuators (1) The flight control system contains four servoactuators, one for each control axis. The hydraulic servos consist of two internally independent systems (primary system and utility system). (2) The primary and utility hydraulic systems provide approximately 3000 psi of pressure to the servoactuators. The primary side of the servoactuator has EHV that allows the FMC to affect the flight controls. Failure of the primary hydraulic system results in the loss of all FMC SCAS (SCAS, CAS, and BUCS). (3) An emergency hydraulic backup system is provided through the Utility accumulator. The utility accumulator can provide emergency hydraulic power to the utility side of the flight control servoactuators in the event of a dual hydraulic system failure. This function can be manually initiated by either crewmember. NOTE: If the utility hydraulic fluid level is low, the low level shutoff valve will close, isolating the directional servoactuator from the utility hydraulic system. In this condition, the utility accumulator cannot provide emergency hydraulic power to the directional servoactuator when the emergency function is selected. The tail rotor will remain fixed pitch. (4) The servoactuators provide hydraulic assistance for the crewstation controls. (a) Full authority motion in response to manual inputs. (b) Limited authority motion in response to electrical stabilization commands (only if primary hydraulic system is operational). (c) Full authority motion in response to electrical BUCS commands only through the primary side of the servoactuator. (5) The servoactuators have internal RAM LVDTS that provide SAS actuator and piston position to the FMC via shielded cables. D-44

d. Linear Variable Differential Transducer (LVDT) The LVDT is a potentiometer that translates control motion into an electrical signal that is utilized by the FMC to provide SCAS, CAS, BUCS, and hold mode functions. Each axis of the mechanical controls has one LVDT attached for a total of eight. The four hydraulic servoactuators each house three LVDTS (2-RAM, 1- SAS) for a total of twelve. D-45

e. ARDD The ARDD provides the means to physically disconnect a crewstation flight control from its associated linkage and is located by each flight control. The Rotary Variable Differential Transducer (RVDT) is connected to the ARDD in such a way that it can detect an ARDD when it is decoupled. It is monitored by the FMC to ensure proper operation and to detect control axis breakout. The ARDD is resettable with a ¼ drive socket wrench. D-46

f. BUCS control (1) In the BUCS mode the mechanical controls are physically decoupled and control is now maintained through an electronic control loop consisting of the following components: (a) FMC (b) LVDT (c) Electro-Hydraulic Valves (d) Actuator Solenoids (2) Many BUCS critical components are continuously monitored for condition and status. These components are the: (a) FMC (b) Flight control actuator (c) ARDD and RVDT (d) Linear Variable Differential Transducers (e) Tracer Wires (3) The FMC provides the monitoring, BUCS engagement command logic, and electronic processing to allow engagement of BUCS. It outputs the commands necessary to the flight control actuators. D-47

(a) (b) In the normal flight mode, the FMC executes pre-programmed control laws, provides automatic stabilator control, air data processing, BUCS tests, and Continuous BIT. In BUCS mode, the FMC provides the above plus BUCS engagement and control on a per axis basis, and command of other axis in normal mode. After a BUCS engagement only successful completion of IBIT will allow BUCS to disengage and normal operation to resume. (4) The flight control servoactuators have a SAS sleeve that is used to augment the mechanical control of the actuator during normal operation for SAS and CAS. The FMC provides command augmentation by driving the EHV, which in turn drives the SAS sleeve. This allows hydraulics to flow resulting in power piston motion. Such motion results in the mechanical spool following the SAS sleeve until the spool aligns with the sleeve. Once aligned, hydraulic flow is prevented and the RAM stops. Mechanical stops on the SAS sleeve limits authority to 10% (20% pitch forward). During BUCS operation the mechanical spool is locked allowing the SAS sleeve 100% authority. The servoactuators incorporate a shear pin to allow actuator motion in the event of a control jam between the crewstation and the actuator. When the servoactuators are commanded to move, while in BUCS, the mechanical spool is locked out by the BUCS plunger and can not follow the SAS sleeve. This will cause the shear pin connected between the power piston and the mechanical spool to break allowing free movement of the actuator from the jammed controls. D-48

g. During normal operation, the ARDD assembly is part of the mechanical linkage. If the mechanical flight controls become jammed, the ARDD provides the means to physically disconnect a crewstation flight control from its associated linkage. The RVDT is connected to the ARDD in such a way that it can detect a disconnect. It is monitored by the FMC to ensure its proper operation and to detect control axis breakout. (1) The drive cam input arm connects to the pilot s and CPG s flight controls (Cyclic, Collective, directional controls). (2) The out put arm is connected to the flight control rods. (3) The rocker arm and its roller, maintains the input arm and output arm in relationship with each other. (4) The load arm, and its roller, holds the rocker arm in its engaged or disengaged position. (5) The spring assembly maintains pressure on the load arm and output arm to hold the rocker arm in position. (6) The reset cam is used to reset the ARDDs back to the engaged position. (7) The reset spring pulls the rocker arm back into the normal position when the reset cam is rotated to the reset position. (8) The RVDT monitors the position the output arm and the rocker arm. The RVDT provides two signals to the FMC to indicate whether the ARDDs is engaged or disengaged. D-49

h. Operation (1) When pressure is applied to the input arm and the output arm movement is restricted (flight control jam), the ARDD will disengage. (a) The rocker arm roller is forced out of the detent on the input arm. (b) The roller forces the rocker arm down, compressing the spring assembly. (c) The load arm roller rides up the rocker arm until the roller is over centered. (d) The load arm roller then forces the rocker arm away from the input arm. (e) The rocker arm also rotates the shaft of the RVDT. The RVDT then provides an electrical signal to the FMC acknowledging a decoupled action. (f) The input arm is now free to move about its center axis. (2) The ARDDs are re-settable with a ¼ inch square socket drive. D-50

i. The tracer wire is a third wire in the twisted pair wire cables linking the RVDTs and LVDTs to the FMC. Its function is to ensure the integrity of the BUCS associated wiring. A broken tracer wire can prevent BUCS engagement on a per axis basis. It cannot cause BUCS to disengage if the tracer fails after BUCS is engaged. D-51

j. Types of BUCS engagements (1) Jammed control axis (a) Upon sensing a jammed control axis the crewmember applies sufficient force to decouple the ARDD freeing the mechanical controls. (b) BUCS activates when the RVDT indicates to the FMC that the ARRD has decoupled. (c) Full control is assumed after 3 second easy on. This aids the crewmember in aircraft control after forcing the controls to decouple the ARDD. (2) Jam in CPG stick (a) With the CPG controls locked above the ARDD, he cannot decouple his ARDD. (b) PLT must decouple his control via his ARDD. (c) RVDT activation causes FMC to engage BUCS in the affected axis. (d) The PLT LVDT commands the system. (e) Full control assumed after 3 second easy on. D-52

(f) Jam between or aft of crewstations 1) In a jam between or aft of crewstations, force on the control for the affected axis from either crewstation will decouple the respective ARDD. 2) BUCS engaged upon RVDT activation. 3) The LVDT in the decoupled crewstation will command the system. 4) If BUCS is initiated by the CPG, control can be transferred to the PLT by decoupling his ARDD. 5) If BUCS is initiated by the PLT, the CPG can take control by decoupling his ARDD and activating his BUCS select trigger. (g) Jam engagement techniques 1) Make an aggressive application of force in the axis that is jammed. If more than one axis is jammed, decouple the axis with highest priority first. 2) After decoupling the control center it, and do what comes naturally, fly the aircraft. D-53

k. Severed control axis (1) Upon sensing a control mistrack between the PLT control position LVDT and the Ram A actuator LVDT and between the CPG control position LVDT and the Ram B actuator LVDT the FMC engages BUCS in that axis. This is a simultaneous mistrack between both crewmember LVDT inputs to the FMC and the servoactuator LVDT outputs. (2) Severance between crewstations (a) In a severance between crewstations the CPG controls are free to move to mechanical stop. During a severance the BUCS FAIL warning is issued to the crew. This is the only warning associated with the flight controls. (b) The PLT controls are still connected normally to the flight controls, and BUCS engagement is not authorized, as there is no mistrack between the PLT LVDT and actuator Ram A LVDT. (c) If the PLT is incapacitated or otherwise unable to fly the aircraft, the CPG can engage BUCS using his BUCS select trigger and fly using BUCS. Control cannot be transferred back to the PLT. CAUTION The effect of the CPG assuming control with the BUCS select trigger is to transfer control of the helicopter from a flight control that still has integrity to a non-redundant electronic flight control. This shall only be done if the PLT is incapable of flying the helicopter. D-54

l. Severance aft of crewstations (1) Both PLT and CPG flight controls are free to move to mechanical stops. (2) BUCS engagement authorized as soon as PLT/Ram A LVDT and CPG/Ram B mistrack simultaneously. Either crewmember can fly the aircraft with the PLT LVDT commanding the actuator. (3) Full control is assumed within 1 second of engagement. (3) Severance Engagement Technique: Fly the aircraft. D-55

m. BUCS logic for crewmember in control (1) If the PLT decouples his ARDD first he has control through his LVDTs. The CPG can obtain control by decoupling his ARDD and pressing his BUCS trigger. (2) If the CPG decouples first he has control through his LVDTs. The PLT can obtain control by decoupling his ARDD. (3) If a severance mistrack occurs with both sticks mistracking both Ram LVDTs, the PLT LVDTs have control. ARDD/RVDT signals are ignored. The CPG can gain control by pressing his BUCS trigger. (4) If the CPG activates BUCS through his trigger, BUCS engages under CPG LVDT control if a stick mistrack occurs and ARDD/RVDT signals are ignored. n. BUCS ON procedures (1) Pilot and CPG establish communication (2) CPG extend cyclic if stowed (3) Crew coordinate in which crewstation and axis BUCS is active and transfer control if appropriate. (4) LAND AS SOON AS POSSIBLE (5) Handling qualities identical to SAS mechanical controls in BUCS axis (6) SAS and hold modes still available in non BUCS axis (7) Pilot crewstation retains force trim o. BUCS FAIL notification (1) When the FMC detects a BUCS failure an output discrete is enabled signaling the SP (2) The SP initiates a warning message and master warning light (3) BUCS FAIL is presented on the UFDs D-56

(4) Detailed failure message is sent to the MPD and Maintenance Data Recorder (MDR) p. BUCS failure procedures (1) Avoid rapid control inputs (2) Pilot and CPG establish communication (3) CPG extend cyclic if stowed (4) Crew should attempt to reset the BUCS FAIL by toggling the appropriate SAS channel on the A/C UTIL page (5) If BUCS FAIL remains on- LAND AS SOON AS POSSIBLE q. Power on reset (1) System is tested to determine if BUCS was previously engaged in any axis (2) If BUCS was engaged then Power on BIT (PBIT) is bypassed and BUCS is re-engaged in that axis (3) If BUCS was not engaged in any axis normal PBIT is run. r. Power on BIT (1) A self check of the FMC to determine if it is capable of normal operation (2) The main processor, RAM, EPROMS, and dual port RAM are tested. (3) PBIT is bypassed if BUCS was previously engaged s. BUCS modes of operation (1) One or more axis has BUCS engaged (2) All other axis continue in normal mode (3) Successful completion of IBIT is the only way to exit BUCS t. BUCS safety considerations (1) Minor degradation in flight control operation while in BUCS. While in BUCS, SCAS functions are not available in the affected axis. The aircraft response in that axis is identical to FMC off flight. (2) Continuous monitoring of the BUCS critical components/functions is degraded while SAS is selected OFF. (3) Continuous monitoring of some BUCS critical components/functions will not occur if SAS is selected off by either the crew or due to the FMC detecting a fault with SAS. (4) Crew Contention (Pilot and CPG fight for the controls and decoupling occurs.) (a) Case 1. Pilot s ARDD decouples in the affected axis (most likely as pilot station ARDDs require less force to decouple). BUCS will engage in the affected axis and track the pilot s LVDT. (b) Case 2. BUCS engages without a decoupling of the ARDD could occur in the event the crew were fighting over the controls and the CPG BUCS select switch had failed on, most likely in the YAW axis. D-57

(5) Crew continues to operate with a BUCS FAIL indication. In this case there is no backup for the flight controls. There would possibly be no degradation of flight control operation. If there were jam or severance in the affected axis, probable loss of control of the aircraft could occur. The crew is required to LAND AS SOON AS POSSIBLE. The ARDD must be reset, if required, and IBIT passed prior to next flight. (6) Flight control failure as crew continues to operate in BUCS ON when returning to base. BUCS is a simplex system meaning single point failures can adversely affect the operation of BUCS. In order to minimize exposure time, the crew must LAND AS SOON AS POSSIBLE. The ARDD must be reset and IBIT passed prior to next flight. (7) BUCS fails to engage when required when the hard stop is encountered during ARDD decoupling. (a) Case 1. In the event the pedals jam at NULL (centered), it is possible that the hard stops are encountered in both directions (this is still under review). (b) Case 2. In the event of a control jam if the crewmember does not perform the necessary action (decouple in the opposite direction even if that is not the direction he wants to go) control authority in the affected axis is lost. It is imperative that crew-training stress that breakout can be done in either direction and immediate control movement be made to correct aircraft attitude. (8) CPG experiences loss of force trim while in BUCS. The force trim brakes for all axis are connected to the Pilots flight controls. When the CPG engages BUCS by decoupling an ARDD, he is disconnected from the Pilots flight controls and the force trim brakes. The CPG may experience small-uncommanded movements in the collective. WARNING When BUCS is engaged do not release the flight controls until the flight has been completed and the main rotor has come to a complete stop. Force trim may or may not be available in the BUCS ON axis. CAUTION A Breakout of the ARDD at the base of the collective control will eliminate the normal mass of the control system and may cause the collective to move slightly in response to rotor system vibration. This slight movement will be detected by the LVDT and can produce an unwanted heave application. Increasing collective friction can eliminate this characteristic. If the CPG were to release the control in which BUCS was engaged, it could move to a hard stop with obvious results. D-58

CHECK ON LEARNING 1. The Back Up Flight Control System (BUCS) is an emergency fly-by-wire backup control system that permits continued controlled flight in the event of what two causes? 2. What component in the flight control system provides the means to physically disconnect a crewstation flight control from its associated linkage? 3. What is the purpose of the Linear Variable Differential Transducer (LVDT)? 4. What three hydraulic systems have an impact on the hydraulic servos? 5. Failure of the primary hydraulic system will have what effect on the flight control system? D-59

F. ENABLING LEARNING OBJECTIVE 6 After this lesson you will: ACTION: Identify the characteristics of the flight control system Built-In-Test (BIT) functions CONDITIONS: Given a written test without the use of student notes or references STANDARDS: In accordance with the TM 1-1520-251-10 1. Learning Step / Activity 1. Identify the functions of the flight control initiated Built in Test a. Flight control BIT functions (1) The FMC implements Power-on Built-in Test (PBIT), Continuous Built-in Test (CBIT), and Initiated Built-in Test (IBIT) routines. Faults confirmed during flight are reported to the MDR and stored in Non-Volatile Memory (NVM). (2) The crewmember can attempt to reset faults by toggling the SCAS switches. Confirmed faults are cleared by this procedure only if the failure no longer exists D-60