OCTOBER 1957 UNCU MAINTENANCE INSTRUCTION ARROW DAMPER SYSTEM REPORT NO 70/MAINT 13/1 ENGINEERING DIVISION NAC-CIST I J. H. PARKIN BRANCH ~t~ 111995 ,. SECURITY CLAssr FICATroUN~ASS(FIID &.l!illi.. MAlNTENANCE INSTRUCTIONS DAMPER SYSTEM 70 MAINT 13/1 21 Oct 027 Prepared~ Approved: Approved: Authorized: or aintenance and Reliability Section ENGINEERING DIVISION 9 AVRO AIRCRAFT LIMITED 9 MALTON9 ONTARIO l♦I National Research Council Conseil national de recherches Canada Canada Canada Institute for Scientific and Technical Information J. H. Parkin Brench I lnstitut canadien de l'informetion scientifique et technique Annexe J. H.Perkin Has been: D Downgraded to: ~5! .f.e., .. ~~:':":~:..\1:1.~?..~0 ... 1:\~:.?.~~) .. 0 De-Classified By: (Name) •••••• • ,/_:_~ _ ✓7 (Dept) :·~-;~-~px.-~~ Date:-:$ •• ••••••••••.• Deputy Coordinator Access to Information and Privacy Signature Jiti!:----------UN~s£~}].:_S_Slf_lED _ _ 70/MAINT 13/1 UtJLtrnlTED FOREWORD This report has been prepared to provide information regard ing description and maintenance of the Arrow Damper System. It i s intended that the report shall serve as an indoctr ination document for persons concerned with t his system. The information contained herein, reflects maint enance desi gn objectives as of the date of publication. It mus t be recognized that the Damper System is presently in t he development phase and that significant changes may occur bef ore the system is delivered to the field. It i s intended to maintain the report as an effective document by revising as the development program continues and as more information becomes available . It should be noted that the Auto matic Flight Control p Aut omatic Ground Control Interception and Automatic Ground Control Approach Sub-system will not be installed in t he Arrow I aircraft, although the tie-in between the Damper System and the above mentioned sub-system is covered in this report . - SlflED TABLE OF CONTENTS 70/ MAI NT ~13/1 Chapter ~. Subject Page 1 IN1RODUCTION 1 2 DESCRIPTION 1 3 OPERATION 3 3.1 Pitch Axis 3 J.2 Roll Axi s 12 3.J Yaw Axi s 19 4 GROUND TEST EQUI PMENT 24 4.1 Introduction 24 4.2 Damper Test Set 25 4 . .3 Damper Auxiliary Test Set 30 4 . 4 Pitot - St atic Test Set 32 4.5 "G" Limi t er Test Set 33 4.6 Damper Test Stand 34 4.7 Rate Table .36 5 FIRST LINE TESTING 37 5.1 Test Equi pment Required .37 5.2 Preparati on for Test .37 5.3 Test Procedures 37 5.4 Locat ion of Faulty Components 47 6 SECOND LINE TESTING 62 6.1 Damper Test Stand 62 6. 2 Hate Table 67 APPENDIX I Damper and As sociated Equipment List 70 A FIGURE I FIGURE 2 FIGURE 3 FIGURE 4 FIGURE 5 FIGURE 6 FIGURE 7 FIGURE 8 FIGURE 9 FIGURE 10 FIGUHE 11 FIGURE 12 FIGURE 13 FIGURE 14 FIGURE 15 FIGURE 16 SECRET 70/MAINT 13/1 LIST OF ILLUSTRATIONS Damper and Associated Equipment Location MH-64 Damper, Pitch Axis, Pre-engage, Gear Up or Down MH-64 DamperP Pitch Axis, Engage, Gear Down MH-64 Damper , Pitch Axis, Engage, Gear Up MH- 64 Damper , Pitch Axis, (complete) Shown with Relay in Engage and Gear Up MH-64 Damper, Roll Axis, Pre-engage, Gear Up or Down MH-64 Damper, Roll Axis, Engage, Gear Down MH-64 Damper, Roll Axis, Engage, Gear Up MH-64 Damper, Roll Axis, (complete) Shown with Relay In Engage and Gear Up MH-64 Damper» Yaw Axis UG 6004A- l Damper Test Set UG 6005A-l Damper Auxiliary Test Set MB-1 Pitot-Static Test Set UG 600JA- l Damper Test Stand UG 6002A-l Hate Table Connection of Test Equipment to Aircraft ~:--------------------S'.:E::::'.CR~ET'.:__ ______ _ 70/MAINT 13/1 1. INTRODUCTION 1.1 Movement of t he aircraft 9 s ailerons 9 elevators and . rudder is effected by hydraulic components which are normally controlled by electrical means. There is no provision for direct mechanical control. In . the interests of safety. two independent flying control hydraulic systems are fitted. 1.2 There are three modes of control 9 the normal mode, the automatic riight mode, and the emergency mode . 1.3 In the normal mode, a damping system automatically stabilizes the aircraft i n all three axes 9 and also co-ordinates rudder movement with movement of t he ailerons and elevators . Control of the ailerons and elevators in t hi s mode is by an electrical force transducer fitted in the control column. 1.4 In the automatic flight mode, the damping system is operative in the normal mode 9 but aileron and elevator position is controlled by an Automatic Flight Co ntrol Sub-system (AFCS). The AIBS allows the aircraft to be controlled from the ground for Automatic Ground Control Interception (ADCI) or for Automatic Ground Control Approach (AGGA). It also provide~ certain pilot assist functions by holding attitude ~ any set course and altitude or it may hold a set Mach number by varying t he aircraftVs pitch attitude. The AFCS also provides for automatic navigation by controlling the aircraft according to information fed into a dead reckoning computer by the navigator. 1.5 In the emergency mode 9 the hydraulic components for the ailerons and elevators are controlled mechani cally. Yaw stabilization and rudder co-ordination are maintained by an emergency yaw damping system. 1.6 Pilot "feel" at the control column is provided by the damping system in the normal mode 9 and by mechanical means in the emergency mode . 1. 7 A four-way push but ton on the control column allows for aileron and elevator trim. There are no trimmer t abs, trim being achieved by altering the position of the ent ire control surfaces. To r educe elevator tr im drag at high altitudes, provision is made for an automatic up deflection of both ailer ons. 1 0 8 If certain fli ght limitations are exceeded, the ,system automatically changes over to the emergency mode. 2, DESCRIPTION 2,1 Four di s tinct channels comprise t he damping sy stem; the pitch channel 9 which control s the elevators; the roll channel 9 which c~ntrols the ailerons· and the normal and emergency yaw channels, whi ch control ' the rudd er . PAGE 1 A ~ -------------________ .:::_SE'.::C~R::::ET:.:._ ______ _ 70/MAINT 13/1 2.2 The switchi ng controls for the damper system are l ocated on a function selector panel fitted in the front cockpit LH console , and on t he control column. Mounted on the function selector control panel are a POWER ON=OFF switch protected by a guard, an ENGAGE push switch for the normal mode and an EMERGENCY push-switch for engaging the • emergency mode. Note that when the POWER ON-OFF switch is selected on , power is supplied to but does not engage the damper sys tem and the AFCS o In addition , the normal mode of operation must be selected before the emergency mode of operat ion can be selectedo Two push=switches are fitted in the control column handgrip. One switch, when operated , reverts the damper system t o the emergency mode of operation, and the other switch when operated , disengages the AFCS but leaves the damper engaged. - Three indi cator lights fitted on the master warning system panel indicate, r espectively, roll and/or pitch axis disengaged, emergency yaw damping system operative and all damping inoperative. AYIIII AIIICIIAF T LIMITED DIFFERENTIAL SERVO-AILERON ACCELEROMETER - ROLL RA TE LIMITING TRANSFORMER -AILERON POSITION POTENTIOMETER - AILERC (DAMPER TEST DIFFERENTIAL SERVO-ELEVATOR POTENTIC POSITIO GYRO-YAW RATE NORMAL GYRO-YAW RATE EMERGENCY GYRO-PITCH RATE GYRO-ROLL RATE DAMPER AMPLIFIER CA LIBRA TOR MAGNETIC AMPLIFIER (HINGE MOMENT LIMITER) SHORTING PLUG No 2 (DAMPER TEST) ACCELEROMETER -LATERAL ACCELEROMETER-LATERAL ACCELEROMETER - NORMAL ACCELEROMETER -NORMAL SWITCHING FUNCTION SELECTOR PANEL TRANSDUCER-STICK FORCE PITOT HEAD - (NORMAL) LOW PRESSURE SWITCH - FLYING CONTROL HYDRAULICS ~ ---.___ ACCELEROMETER-LATERAL SWITCHING ~ ACCELEROMETER- NORMAL 'G' LIMITER (FORWARD) CALIBRATOR- 'G' LIMITER DAMPER CIRCUIT BREAKER PANEL INDICATOR -CONTROL SURFACE RESPONSE SENSOR -ALPHA/BETA ACCELEROMETER-LATERAL SCHEDULING ACCELEROMETER-LATERAL SWITCHING FIG. l DAMPER AND ASSOCIATED EQUIPMENT LOCATION SECRET ----------------------------A VR Q ARR Q W ► DIFFERENTIAL SERVO-AILERON PITOT HEAD - (EMERGENCY) METER - ROLL RA TE ON POSITION r POTENTIOMETER -AILERON POSITION (DAMPER TEST) POTENTIOMETER - ELEVATOR DIFFERENTIAL SERVO POSITION POTENTIOMETER -ELEY A TOR POSITION (DAMPER TES LOW PRESSURE SWITCH - FLYING CONTROL HYDRAULICS ~ ACCELEROMETER-LATERAL SWITCHING ~ ACCELEROMETER- NORMAL 'G ' LIMITER (FORWARD) CALIBRATOR-'G' LIMITER ~ CIRCUIT BREAKER PANEL )NTROL SURFACE RESPONSE FIG. 1 DAMPER AND ASSOCIATED EQUIPMENT LOCATION LINEAR Q : ACTUATOR - RUDDER TRIM UNIT-RUDDER DIFFERENTIAL SERVO-RUDDER (DUAL) POTENTIOMETER- RUDDER POSITION (INDICATION AND DAMPER TEST) FEEL AND TRIM EMERGENCY RELEASE - ELEVATOR POTENTIOMETER-ELEVATOR DIFFERENTIAL SERVO POSITION POTENTIOMETER-ELEVATOR POSITION (INDICATION AND DAMPER TEST) TRANSFORMER -AILERON POSITION (INDICATION AND DAMPER TEST) /7 DIFFERENTIAL SERVO-AILERON PARALLEL SERVO-AILERON FEEL & TR!M UNIT-ELEVATOR POTENTIOMETER-ELEVATOR PARALLEL SERVO POSITION PARALLEL SERVO - ELEVATOR ACCELEROMETER - NORMAL 'G ' LIMITER (AFT) A !mt:--------- ----- ---_ ___ _:SE'.::G'.:RE:_:T'.___~---- 3ol Pitch Axi s ( Figures 2,3P4 and 5) Command Circui t s 70/MAINT 13/1 3o l ol The pit ch axi s command ci rcuits operate in response to command s i gnals from the control column force transducer i n the normal mode of operation or from the automatic fl i ght control sub-system (AFCS) in the automatic mode of oper ation. On recei ving such signals the command circuits actuate an electro- hydraulic parallel servo unit and two electro- hydraulic differential servo units and these in turn effect the appropriate positioning of the elevat ors . The operation of the command circuits is described in the following paragraphs. 3.1. 2 For ce exer ted on the control column handgrip results in an ou,_tput voltage from the force transducer which is fitted in the control column at the base of the handgri1r 0 The magnitude of this voltage is directly proportion.al, to the force exerted on the control column. The phasing of the voltage depends upon the direction of the motivating force relative to the neutral position of the control column, voltages produced by motion forward being 180° out of phase with voltages produced by mo tion aft. Note that changes in attitude are not commanded by movement of the control column but by the application of force on the handgrip wh i ch deflects the core of the force transducer. Movement of t he control columnp i n effect, follows the positioni ng of the elevators by the command circui ts p but the response of the sy stem is so immediate that atti tude changes appear to result from movement of the control column. 3. L 3 The output vol tage from the transducer is approximately one volt fo r each six pounds of force exerted on the control column hand gr ip . The output of the transducer is restricted by means of ad j ustable mechanical stops which limit the maximum output f rom the amplifier to 6 volts. 3.1 . 4 The output from the transducer amplifier is supplied via a summing network as a command signal to an integrator uni t the output voltage of which increases a t a rate propor t ional to the strength of the command signal. The r ate of increase of the integrator output is such as to prevent excessive r a te of control surf ace movement. The integrat or maintains i ts output even after the command si gnal is withdrawn, and until it is provided with a signal of opposite phase to reduce, cancel or reverse the output according to the strength of the new signal. PAGE 3 I A ij~ __________________ ___'.S~E~CRE~T ______ _ 3,L5 70/MAINT 13/1 The output of t he integr ator is supplied through .a summing ne t work to a servo amplifi er unit which amplifie~ the signal and deter mines its phase by means of a discriminator ci rcuit. From the servo amplifier the signal passes t o the transfer· valve of an electro- hydraulic parallel servo which controls the elevat or movement in magnitude and direction in accordance with the strength and phase of the input signal. The parallel servo effects movement of the elevators via the elevator quadrant P and so a sympathetic movement of the control column takes place. A self- balancing circuit is introduced fr om the parallel servo to the servo amplifier . Any movement of the parallel servo output rod produces a feed-back signal opposi te in phase to the initiating signal and of a magnitude proportional to the extent of the movement of the parallel servo. Th i s feed- back signal is applied t o t he summing point between the integrator and the servo amplifier, where it apposes and presently nullifies the i ntegrator output and, consequently, output from the servo amplifier. When t his nullifying effect occurs p movement of the parallel servo , and therefore of the elevators p ceases. The system is designed , by the provision of a normal accelerometer and air data transducer components, to produce a given "g11 loading in response to a given force applied to the control handgrip , i ndependent of airspeed and altitude. The normal accelerometer provides an output signal of one volt per "g" of loading in the pitch axis. This signal is 180° out of phase with the command signal from the transducer p so that , when t he aircraft commences a manoeuvre imposing a "g" loading, the resulting output of the normal accelerometer progressively nul lifies the command signal . The command signal is also modified by the air data transducer components to compensate for changes in airspeed and altitude. Thus , since the transducer output signal is proportional to "g" loading, the force that must be applied to the control column handgrip to produce a manoeuvre of a given "g" will be constant~ regardless of airspeed and altitude . As described above , force exerted on the control column in a fore or aft motion rel ative to the neutral position will produce elevator movement in proportion to the applied force , and the elevator movement will cease when the integrator output is balanced by the feed- back signal from the parallel servo. Meanwhil e the resultant "g 11 loading will have produced an accelerometer signal to nullify the signal from the t ransducer, and the system will remain balanced until a new command signal appears. If the force, _exerted on the contr ol column ~andgrip is reduced to zero the signal from the acce.l erometer 1.s no longer opposed by a signal from the transducer ampli f i er and therefore acts as a new command signal. PAGE 4 I A 6~------------- ----...'.::S~EC~R::ET:__ ______ _ 7D/J1AINT 13/1 3.1.9 (continued) Being of equal strengt h but of opposite phase to the ?riginal command signal~ its appearance at the integrator input results in the cancelling of the integrator output. The feed- back signal from the parallel servo thereupon f i nds itself ' unopposea and returns the parallel servo and therefor e t he elevator s and the control column to a position wher e the 11 g" l oading is removed f rom the ax:fs 0 3.1.10 A slight time lag occurs between the r eceipt of a signal by the integrator and the subsequent output by the integrator of a signal to the servo amplifier. 'While t his delay has a useful effect in preventing the transmission to the control column via the parallel servo of short- term corrective movements of the elevators (see Sub paras 3. 1.12 through 3. 1.16) it would also have the undesirable effect of delaying the response of the elevators to command signals but for the action of the differential servos. Command signals which are fed to the integrator are also fed to two differen_t i al servos connected into the elevator control l i nkages. These electro- hydraulic units operate the elevators immediately in response to the command signals. Elevator movement is then taken over by the parallel servo as output becomes available from the integrator . As the command signal circuit to the differential servos does not incorpora te an integrator » the differential servos return to their original position as soon as the command signal is nullified by the action of the accelerometer. This allows them to return to their mai n task of damping uncontrolled movements abou~ the pi tch axi s as described in sub paras 3. 1.13 and 3. 1.19. 3.1.11 The operation of the command circuits under comr1and signals from the AFCS in the aut omatic mode of operation is identical with that in the normal mode . In the aut matic mode the command signal from t he AFCS substitutes for the signal from the transducer. 3.1.12 Selecting the landing gear down introduces certain changes in the command circuits. The main effects of such action are to interrupt the command circuit to the differential servos and to prevent the accelerometer output an~ the air~pee? compensation from acting upon the command si?11~lll which i s supplied directly to the parallel servo amplifier~ by passing the i ntegrator . These command circuit changes are effected by two fader units and resul t in elevator movement becoming directly proportional to control column movemen t without regard to airspeed and altitude. Other effects resulting from selecting the l anding gear down ar~ the re routi ng of the position tr~m circuit ~hrough a point immediately preceding the integrator instead of t hrough the trim mo tor (see sub paras 3. 1.24 and 3. l .25) ll and t he PAGE 5 I A SECRET 70/MAINT 13/1 3.Ll2 (continued) shorting out of the integrator in the pitch rate gy-ro circuit (see sub-para. 3.1.19) Sensing Circuit 3.1.13 A pitch rate gyro is incorporated in the system as a sensing device to_correct uncontrolled movement about the pitch axis. On detecting movement the gyro produces a signal which is proportional to the rate of pitch, rather that to aircraft displacement, so that comparatively quick corrective action can be achieved automatically before the displacement becomes significant. 3.1.14 The output signal from the gyro is supplied to the circuits of the differential filld parallel servos to correct the uncontrolled movement and maintain stable flight. 3.1.15 In order to prevent signals from the pitch rate gyro from opposing co mmand signals for pitch changes, an integrator is incorporated in the sensing circuit the function of which is to nullify output si gnals from the gyro resulting from steady state rates of rotation about the pitch axis. 3.1.16 Signals resulting fro m movement of the aircraft about its pitch axis at a rate of 4 cps or more are referred to as short ter m si gnals. Signals produced by movement of less than 4 cps are called long term signals. An operating time lag in the gyro circuit integrator prevents the nulling out of short term si gnal s which are supplied only to the differential servos. It also permits long term signals to appear in the parallel servo circuit, thus initiating corrective movements of the servos. Movement of the parallel servo produces sympathetic movement of the control column via the mechanical linkage of the elevator quadrant so that the balance of the command circuit is maintained. Corrections applied by the differential servos, however, do not produce movement of the control column, the mechanical linkage of the hydraulic jack acting to prevent such an eff ect. 3.1.17 Short term signals resulting fro m movements exceeding 4 cps about the pitch axis effect the operation of the differential servos only. These signals are not supervised by th~ gyro circuit integr ator because of its opera ting time lag, and, for the same reason, they do not affect the parallel servo. Thus short term signals are effectively blocked out of the paral lel servo circuit where they would produce undesirable control reactions. The differential servos are positioned continuously by -the short term signals to correct short term movements automatically and maintain stable fli ght. PAGE 6 SECRET 70/MAI NT 13/1 3.1.18 When a cor rective movement is initi a t ed and t he aircraft co mmences to return to its orieinal attitude, t he signal produced by t he gyro will be of the opposite phase. In the case of short term movement the differential servos will t hen be moved in the opposite direction 9 positioning · the elevators to limit the correction. In the case of l ong term movements , the s ignal cancels t he output of the parallel servo integrator, thus permitting the ser vo feed back signal to operate the amplifier and return t he elevators to the neutral position. 3.1,19 'When t he landing gear is selected down and t he command cir cuits are operative as described in sub para 3 , 1.12, the gyro circuit integrator is short ed out. In t his condi t ion steady state pitch rate signals, in add ition to long term and short term signals , are supplied t o the di fferential servos. This action aids s tability in the pitch axis at low airspeeds. Trim Circuits 3.1.20 Three trim circuits are incorporated in t he pitch axis damper channel, via. an a cceleration trim circuit 9 a pre s sure trim circuit and an artificial feel circuit. Acceleration D:im Circuit 3.1.21 The acceleration trim ci rcuit can be used to trim t he aircraft into a "g" manoeuvre. The circuit is controlled by t he trim swi tch on the control column handgrip. Operating t he trim switch provides a power supply to a motor-positioned potentiometer incor porated i n the tr i m w1it. The voltage picked off this potentiometer is supplied to the summing point of signals fro m th e normal accelerometer and signals from the force tr ansducer or the AFCS. The trim signal oper ates in the command circuits in thesame way a s does a command signal. 3.1.22 When the trim switch is operated , the movement of t he aircraft about the pitch axis produces signals from the rate gyro and t he normal accelerometer. The s ignal fro m t he pitch rate gyro being a steady st ate s i gnal is nullifi ed by the filter integrator, Note that 9 due t o t he inability of the parallel servo integrator to follow 9 short t er m s i gnals continue to be provided by the pitch r ate gyro t o the differential servos only . The output of t he nor mal acceler ometer nullifies the trim signal with t he same results as when a command signal i s nullified. Until the trim switch is again oper a ted , transducer signal s will add to or sub_tract from the standing signal from the trim potentiometer to effect further movement of the elevators. PAGE 7 9 A SECREI' 70/MAINT 13/1 3.1. 23 Wh~n the_pitch axis damper channel is disengaged~ the trim switch will govern the feel and trim unit of the artificial feel circuit instead of the position trim circuit. See sub para 3.1.33. 3.1.24 When the landing gear is selected down, the power supply from the trim switch is supplied directly to the parallel servo integrator. This prevents the normal accelerometer and the airspeed compensation from acting on the trim signal these functions being by- passed in the command signal circ~its. 3.1.25 The output of the integrator is supplied to t he summing point of command signals from the for ce transducer and the parallel servo feed-back circuit . The trim signal w;i.ll effect movement of the parallel servo until the circuit is balanced by the feed-back signal. Subsequent command signals will t hen add to or subtract from the standing trim signal to effect further movement of the elevators. Pressure Trim Circuit 3. 1.26 The pressure trim circuit is operative automatically when the pitch axis channel . of the damper system is disengaged. The function of the circuit is to simulate command signals rel a tive to the position of the elevators. These signal s are supplied to the input of the integrator, By t his means, when the sy stem is re-engaged, the output of the integrator will corr espond with the position ~f the control column thus ensuring smooth re-engagement, 3,1.27 The operation of the circuit is such that during emergency mode operation~ the normal feed- back signal produced through movement of the output rod of the parallel servo will continue to be applied to the summing point preceding the servo amplifier. This signal wil l eff ect the operation of the amplifier and in turn the transfer valve of the parallel servo . Note that in the emer gency mode of operat ion ~ action by the transfer valve does not i nfluence the output rod, which is not subjected to hydraulic pressure during emergency mode operation. The operation of the transfer valve results in a pressure differential between the two hydraulic metered supply lines. This action moves a . pressure sensing ram which is conne~ted to t he cor~ of a t~im transformer. The signal from the trim t r ansformer is supplied to the integrator which provides an output si gnal to nullify the feed - back signal. Due to the inherent time lag of the integrator a portion of the trim transformer_ signal by:pas~es the integrator to speed response. See also Disengage Circuits, sub para 3. 1.38. PAGE 8 3.1.28 3.1.29 3.1.30 3.1.31 3.1.32 3.1.33 8 • 70/MAINT 13/1 Incorporated in the trim circuit feed - back transformer a~sem?ly are two switches. One switch completes the supply cir cm. t to the feel and trim 'Unit of the artificial feed circuit (see sub para 3.1.31.) The other switch functions to prevent the channel engaging if the elevators are moving. Artificial Feel Circuit The purpose of the artificial feel circuit is to prevent a sudder change in feel at the control column if the pitch axis damper channel ais=engage~. The artificial feel circuit consi s ts of tin artificial feel and trim unit which comprises an electrically operated actuator and a spring. Operating the actuator incr eases or decreases the load on the spring dependent upon the direction of rotation of the actuator . One end of the unit is attached to the aircraf t structure . The other end is attached to the elevator quadrant, adjacent to the point of attachment of the parallel servo. Exerting a force on the control column handgrip closes a switch which completes a circuit from the actuator to the parallel servo. When the parallel servo operates to execute the command signal the trim actuator receives a power supply. This action increases or decreases the loading of the spring according to the position of the quadrant. If the pitch axis damper suddenly reverts to emergency 9 the trim unit ensures that the load r equi red on the quadrant by the cables from the control column balances t he load that was being imposed on the quadrant by the parallel servo. An additional switch is fitted in parallel wi th the control column switch to serve as a safety precaution against overstressing the airframe should an automatic mode failure occur. This second switch opens when the normal acceleration exceeds :t o. 5 "g" from straight and level fligh t condition of the 1 "g". Therefore "g" in excess of thi s figure applied by the automatic mode is against the feel spring. Gonsequentlyj if an automatic mode failure or disengagement occurs the feed spring re- asserts itself+and reduces the 11g 11 loading imposed by the elevators to = 0.5 "g" or less of 1 11 g 11 flieht. The feel and trim unit serves also to trim the elevators via movement of the elevator quadrant when the damper channel is disengaged. The actuator in this condition is controlled by the control column trim actuation switch w~ich i~ transferred automatically 9 from the acceleration trim motor to the trim ac tuator upon disengagement of the damper channel. PAGE 9 I A .zjA(; _________________ --..'::SE~C~RE~T~-------- 3.1. 34 3.1.35 3.1.36 3.1.37 3. 1 .38 70/MAINT 13/1 In the event of a seizure in the f eel and trim unit , movement of the elevator quadrant would not be possible. To meet this contingency 9 a rele ase circuit is incorporated in the feel and trim unit linRage t o the quadrant . The circuit comprise~ a control switch fitted in the front cockpit and a solenoid , the plunger of which is ~ttached to a link pin in.the feel :111d trim linkage t o the quadrant. Operating the switch energizes the solenoid which extracts the link pin. This action per mi ts the feel and trim uni t to pivot freely with movement of the quadrant , Disengage Circuits The pitch axis disengage circuit operates automatically if the normal acceleration exceeds 5.0 = 505 ngn. The circuit is controlled by an electronic "g" limiter which, when operated, completes supply circuits to the differential servos centering circuit and the parallel servo disengage circuit. Note that the power supply to the damper system components is uninterrupted in the emergency mode . The pitch axis disengage circuit also oper ates if the yaw axis monitor acts due t o excessive skid, side slip or transverse acceleration. Operation of the disengage circuits reverts the channel to the pre-engage condition and illumi nates the R- P AXIS indicator li ght on the master warning panel. In the pre engage condition the pressure t rim circuit is operative, and the acceleration trim actuation switch is transferred to the trim actuator of the feel trim circuit. When the channel is in the pre-engage condition the output of the force transduc er.11 the normal accelerometer and t he pitch rate gyro is r outed t o t he trim unit to ensure smooth re-engagement. A trim unit arnplifier 9 fitted in the circuit i mmedi ately preceding the trim unit 9 discriminates upon t he phase of the applied signals to effect the pr oper direction of rotation of t he motor driven potent iometer of the trim unit. The effect of this provision i s such t hat any unbalanced signals appearing in the co mmand circuit initiate the oper ation of the trim unit and are balanced out by the resulti ng output signal from the trim unit. This ensures that t he channel will be re-engaged into a balanced cir cuit and as described in sub-paras 3. 1. 26 and 3,1.27 9 in agreement with t he prevailing position of the el evators. But for this provision.11 an unbalanced signal .11 r esulting for example from failure of the gyro circuit i nte grator to nullify a pitch rate signal.11 would produce unexpected and undesirable movement of the elevators. PAGE 10 li~---------------------S_E_C_R_ET _________ _ 3.1.39 70/MAINT 13/1 Note that if an unbalanced signal is being nulled by the trim unit when re-engagement occures, the effect is equivalent to introducing an out-of-trim condition on the axis. This is due to the standing trim signal at the summing point of signals from the force transducer and normal accelerometer. In this condition, to regain level flight with the control column in neutral , the trim actuation switch must be operated to cancel t he trim signal. PAGE 11 ... p IJ ~ :I: °' ~ C )> ~ .,, m '!" .,, G LIMITER RUDDER MONITOR~~ PITCH RATE ,___,.TO PARALLEL SERVO DISENGAGE CKT. ~-- TO DIFFERENTIAL SERVO CENTERING CKT. UP :::j n :I: )> X NORMAL -----, ACCEL. p; .,, ;,II 'l' ____J TRIM ~ 11 MOTOR )> C) !" G) > I I STICK ;,11 FORCE C .,, 0 i I ~ z PILOT'S BUTTON UP ON ON TRIM UP FADER □ SCHEDULED GAINS PRESSURE TRIM PARALLEL SERVO STICK --=--=--=--=--=--=?1- FORCE LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO "' a, "" I 6 = I f-" TO HYDRAULIC JACK ASSEMBLY TRIM CTUATOR ~ ~ ~ c::i .. "S .. .. .. .... .. ... .. .. .,, .... .... .... !I: .... .... .., .. ~ ~ ~ c::i c:: ~ z - ~ ' i .,, !5 w :t :c b, ,,,. C > :t -a m '!O -a :::j n :c > >< p; m z C) > C) !" C) m > ;Ill C 0 ~ 21 PILOT'S BUTTON G LIMITER ~--- TO PARALLEL SERVO DISENGAGE CKT. □ SCHEDULED GAINS .___ "' 0, 0 ' ,.., 0 = ' ,.., TO DIFFERENTIAL SERVO ..__ ___ CENTERING CKT. RUDDER MONITOR t------' PITCH RATE NORMAL ----, ACCEL. TRIM MOTOR STICK FORCE ~ UP LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO TO HYDRAULIC JACK ASSEMBLY DN,:i,,.FADER FADER PARALLEL SERVO STIC~~ ~ FOR~ • - TRIM CTUATOR CLOSE~ WHEN STICK FORCE IS APPLIED - - OPENS WHEN I INTER o A- - A Z EXCEEDS O.SG INC. :t:.. ~ =-:i ~ :t:.. =-:i =-:i ~ ~ ... "' .. ... ... .... .. .., .. ... .... .... " .... 11:: .... .... .., ... T§ '· "TI j5 ~ ~ ::z: °' ~ C > ~ "'Cl m ?" "'Cl =i n ::z: > X pi m z C) > C) .!" C) m > ;,a C "'Cl G LIMITER RUDDER t-------' MONITOR PITCH RATE .-----.-TO PARALLEL SERVO DISENGAGE CKT. ~-._TO DIFFERENTIAL SERVO CENTERING CKT. NORMAL ACCEL. 1-----~ TRIM MOTOR STICK FORCE STEP ~I PILOT'S ~------BUTTON q'c 4 n DN UP FADER FADER □ SCHEDULED GAINS "' a> = ' 6 = ' .... CLOSES WHEN STICK FORCE - - IS APPLIED PARALLEL SERVO '-•...J FORCE SWITCH LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO TO HYDRAULIC JACK ASSEMBLY TRIM ACTUATOR1 OPENS WHEN I INTER 0 A- - A z EXCEEDS 0.5 G INC. :ti. 'II::: ~ ~ :ti. ~ ~ ~ ~ ... "' .. .. ... .... .. .., .. ... .,, " ... .... ii: .... " .,, .. T§ ... ~ 3:: :c b. "" C )> 3:: "'Cl m ?O "'Cl =i n :c )> X pi n 0 3:: "'Cl ... m -i .!!! "' :c 0 ~ :e =i :c ;,11 m ... )> ~ z m z Ci) )> c;, m )> z C c;, m ~ PILOT'S c: BUTTON "'Cl TO PARALLEL SERVO ~ DISENGAGE CKT. LIM~TER ~ TO DIFFERENTIAL SERVO CENTERING CKT. RUDDER MONITOR i---~ PITCH RATE NORMAL---~ ACCEL. TRIM MOTOR • STICK FORCE STEP ~ UP FADER ENG PRE □ SCHEDULED GAINS PRESSURE TRIM PARALLEL SERVO I FORCE STICK SWITCH FORCE "" ENG " LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO "' "' I ,... 0 "' I ,... TO HYDRAULIC JACK ASSEMBLY TRIM !ACTUATOR CLOSES WHEN ~ ! OPENS WHEN 7 STICK FORCE -- 0 ~-Al EXCEEDS ~~ ~.. IS APPLIED 0.5 G INC. PRE ~ 'Ill::: ~ c:::, ~ ~ ~ c:::, ~ .. "' .. .. .. ' .. '> .. .. " " ... ' ~ ' " .,, ., T§ D~:-------------------~S'..::'.EC~R~E~T--------~ 70/MAINT 13/1 49 A 3.2 Roll Axis (Figure 6,7 1 8 and 9) Command Circuits 3.2.1 3.2.2 3.2.3 T~e roll axis command circuits operate in response to command signals from a force transducer in the normal mode of operati~n or from t he_Automatic Flight Control Sub-system (AFCS) _in the automatic mode of operation. On receiving such signals, the command circuits actuate . an electro ~draulic_parallel servo unit and two electro-hydraulic differential servo units which effect the appropriate positioning of the ailerons. The operation of the command circuits is described in the following paragraphs . In order to reduce elevator trim drag at high altitude both ailerons are deflected up when the aircraft reach;s a pressure altitude of 45,000 feet. This is done by an actuator fitted on the aileron quadrant in such a manner that it drives the quadrant segments apru:-t. The actuator is control l ed by a pressure switch adjusted :Lto ,, closearid complete the actuator supply circuit at 45~000 feet. To provide an operating differential, the switch is adjusted to re- open at 42,000 feet. The operation of this circuit does not affect the positioning of the ailerons by the damper system. For ce exerted on the control column handgrip in a left or right motion, produces an output vol tage from the force transducer which is fitted in t he control column at the base of the hand grip. The magnitud e of this voltage is directly proportional t o the force exer ted on the handgrip. The phasing of the voltage depends upon the direct i on of the motivating force relative to the neutral position of the control column, voltages resulting from left motion being 18o0 out of phase with voltages r esulting from right motion. Note that changes in attitude are not commanded -by movement of the control column but by the application of force on the handgrip which deflects the core of the force transducer. Movement of t he control column, in effect , follows the positioning of the ailer ons by the command circuits , but the response of the system is so immediate that attitude changes appear to result from movement of the control column. The roll rate produced is proportional to the output voltage from the transducer 0.05 volt producing 1°/sec. rate of roll O The maximum o~tput of the transducer is limited by means of mechanical stops to p .vol t 1 with a forc e of 20 lb. exerted on t he handgrip, which produces a roll rate of 120 degrees per second. PAGE 12 ~ SECREI' ANW~:---------===---~ 3.2.5 3.2.6 3.2.7 3.2.8 3.2.9 70/MAINT 13/1 The re~ulting output from the t ransducer is supplied via a summing network as a command signal to an integrator unit th~ output voltage of which increases at a rate directly proporti~nal to the strength of the command signal. The rate of incr ease of the integrator output is such as to prevent excessive control surface movement. The integl'ator maintains its output even after the command signal is withdrawn and until it is provided with a signal of opposite phase to ' reduc~p cancel or reverse the output according to the strength of the new signal. The output of the integrator is supplied through a summing network to a servo amplifier unit. This unit, in addition to its function of amplifying signalsp incorporates a discriminator circuit to determine the phasing of the applied signal. From the servo amplifier the signal passes to the transfer valve of an electro-hydraulic parallel servo which controls the aileron movement in magnitude and direction in accordance with the strength and phase of the input signal. The parallel servo effects movement of the ailerons via the aileron quadrantP and so a sympathetic movement of the control ' column takes place. A self-balancing or feed-back circuit is introduced from the parallel servo to the servo amplifier. Any movement of the parallel servo output rod produces a feed-back signal opposite in phase to the initiating signal and of a magntitude pro portional to the extent of the movement of the parallel servo. This feed-back signal is applied to the summing point between the integrator and the servo amplifier p where it opposes and presently nullifies the integrator output andp consequently, output from the servo amplifier. When this nullifying ~ffect occursp movement of the parallel servo, and therefore of the ailerons, ceases. The sy stem is designedp by the provision of a roll rate gJTO and air data transducer componsnts, to produce a given rate of roll in response to a given force and degree of movement applied t o the control column handgrip independent of airspeed and altitude. The roll rate gyro provides an output signal of 0.05 volt per unit r ate of roll. This signal ~s 180° out of phase with the command si~n~l. The gyr~ signal and command signals are also modified by the a ir. dat~ transducer components to compensate for changes 1n airspeed and altitude. The compensation function is preset_to result in a decrease of gain at t he extremes of ~ow and high . a irspeeds and is constant for all other airspeeds. This action aids staQility in the roll axis by limiting over- control of the ailerons. PAGE 13 ~---------------___ ____!S~E~C~RE~T:__ ______ _ 70/MAINT 1.3/1 .3.2. 9 (continued) .3 .2. 10 .3.2. 11 Thus 9 since the transducer output signal i s proportional ~o the app~ied force and the roll Tate gyro output signal is proportional to rate of roll, t he force that must be • applied to the control column handgrip to produce a manoeuvre of a ~iven roll rate will be constant 9 except at high and low airspeeds when the force to produce a given rate of rol l is increased . As described above, force exerted on the handgrip in a left or right motion relative to the neutral position will produce a~ 1 leron movement in proportion to the applied force, and the ai eron movement will ce~e .when the integrator output is balanced by the feed- back signal from the parallel servo. Meanwhile the. resultant roll rate will have produced a roll. rate gyro signal to nullify the signal from the transducer, and the sy s tem will remain balanced until a new command signal appears. If the force exerted on the handgrip is reduced to zero, the signal fro m the roll rate gyro is no _longer opposed by a signal from the transducer amplifier and ther efore acts as a new command signal. Being of equal s trength but of opposite phase to the original command s i gnal 9 its appearance at the integrator input results in the cancelling of the original signal there and 9 consequently, in the cessation of integrator output. The feed-back signal from the parallel servo thereupon finds itself unopposed and returns the parall el servo and therefore the ailerons and the control column to a position where the roll rate is zero. A slight time lag occurs between the receipt of a signal by the integratol' and the subsequent output by the integrator of a signal to the servo amplifier. While t his del ay has a useful effect in preventing the transmission to the control col umn via the parallel servo of short-term corrective movements of the ailerons (see sub paras .3 ,2. 16 and .3,2.17), it would also have t he undesirable effect of delaying t he r esponse of the ailerons to command signals but for the action of the differential servo s. Command signals whi ch are f ed t o the integrator are also fed to two differential servos connected into the aileron control linkage. These el ect ro- hydraulic unit s operate the ailerons immediately in response to the command signals . Aileron movement is then taken over by the parallel servo as output ~ecomes avail able from the integrator, As the command signal cir cu i t to the differential servos does not incorporate an int egrator the differential servos return to their original position as soon as the command signal is nullified by t he action of tbe roll rate gyro. ~-----------------'S~EC~R~ET~-------- 3.2.12 3.2.13 3.2.14 3.2. 15 3.2.16 3.2.17 •• • 70/MAINT 13/1 (continued) This allows them to return to their main task of damping uncontro l led movements about the roll axis as described in sub paras. 3.2.14 through 3.2.18. The operation of the command circuits under command signals from the AFCS in the automatic mode of operation is identical with that in the normal mode . In the automatic mode the command signal from the AFCS substitutes for the signal from the transducer. Selecting the l anding gear down introduces certain changes in the command circuit. One effect of this action is to prevent the roll rate gyro signal and the airspeed compensation from acting upon the command si gnal. In this condition, the command signal is supplied directly to the parallel servo amplifier. Consequently j movement of the ailerons is directly proportional to control column force regardless of airspeed. Sensing Circuit In addition to its function in the command circuit, the rol l rate gyro functions as a sensing device to detect uncontrolled movement about the roll axis. Since the gyro output signal is proportional to rate of roll rather than to aircraft displ acement, comparatively greater control of the ailerons can be achieved automatically before the displacement become s significant. The output signal from the gyro i s supplied to the circuits of the parallel and differenti al servos to correc t uncontrolled movement and maintain stable flight. Signals from the rate gyro resulting from movement exceeding 4 cps about the axis are defined as short term signals and effect the operation onl y of the differential servos. The short term signal s are effectively blocked out of the parallel servos circuit due t o the operating time lag of the parallel servo integrator. The differential servos are positioned continuously by the short term signals to suit stability requirements. Corrective movements of the differential servos will not effect movement of the control cplumn. Long term signal s resulting from movemen~ of les ~ than 4 cps about the axis are suppl ied to the diffcrenti~l servos and t he parall el servo. The corrective movement of the parallel servo wil l effect a movement of control column_ via the mechanical linkage of the aileron quadrru:it• _Th~s action ensures that the balance of the command circuit is maintained • PAGE 15 3.2.18 3.2019 3.2.20 302. 21 3,2023 SECRET 70/MAINT 13/1 When a corrective movement i s initiated and the aircraft commences to return to its original attitude the signal produced by the gyro will be of the opposite'phase 0 In t~e case of short term movement, the differential servos will. ~hen be move~ ~n the opposite direction, positioning. the ailerons to limit the correction. In the case of long ~erm movements, the signal cancels the output of the integrator, thus permitting the servo feed-back signals to operate through the amplifier and return the ailerons to their original position. Two trim circuits are incorpor ated in t he r oll axis damper channel, via a roll rate trim circuit and a pressure trim circuit. When the channel is in the pre-enga ge condition, the emergency trim actuator is operative in place of the roll rate trim. Roll Rate Trim Circuit When the roll axis damper channel is engaged the roll rate trim circuit is used to trim the aircraft by deflecting the ailerons. The circuit is controlled by the trim switch on the control column handgrip. Operating the trim switch drives a motor-positionea potentiometer incorporated in the trim unit. The voltage picked off this potentiometer is supplied to the summing point of signals from the r ate gyro and signals from the force transducer or the AFCS. The trim signal operates in t he command circuits in t he same way ad does a command signal. When t he trim switch is operated, the movement of the aircraft about the roll axis produces a si gnal from the r oll rate gyro. This signal nullifies the trim signal with the same results as when a command signal is nµllified. Until the trim switch is again operated, transduc er si gnals will add to or subtract from the standing signal from the trim potentiometer to effect further movement of t he ailerons. When the roll axis channel is disengaged, the trim switch will govern a trim unit instead of the roll rate trim circuit. The trim unit serves to trim the ailerons via movement of the aileron quadrant. The unit comprises an electrical actuator and a spring. One end of the unit is attached to the aircraft structure. The other end is attached to the aileron quadrant. When the trim switch is ope~ated the trim actuator increases or decreases the loading on the spring so effecting movement of the aileron quadrant. PAGE 16 ; _:--____________ ____ _:S::ECR::::.:ET.:_ _______ _ .3.2.24 3, 225 3.2.27 3. 2.28 19 A 70/ MAINT 13 /1 Pressure Trim Circuit The pressur? t rim cir cuit is operative automatically when the roll ~is channel.of ~he damper system is disengaged. The f~nction of the_c~rcuit is to simulate command signals rel ative to the position of the ailerons. These signals are supplied to the input of the integrator. By this means , when the system is re-engaged, the output of the integrator wi l l correspond with the position of the control column and will ensure smooth re- engagement. The operation of t he circuit is such that during emergency mode operation , the normal feed- back signal produced through movement of t he output rod of the parallel servo will conti~ue to be applied to the summing point preceding t he servo amplifier . This signal will eff ect t he operation of the amplifier and in turn the transfer valve of the parallel servo. Note t hat in the emergency mode of operation, action by the t r ansfer valve does not influence the output rod, which is not sub j ected to hydraulic pressure during emergency mode operation, The operation of the t ransfer valve results in a pressure differential between the two hydraulic metered supply lines, This action moves a pressure sensing ram which is connected to the core of a trim transformer, The signal from the trim transformer is supplied to the integrator which provides an output signal to null the l'eed,..back signal. Due to the inherent time lag of the integrator a portion of the trim transformer s i gnal by=passes the integrator to speed control response , See also Dis= engage Circuits sub paras 3, 2,28 and 2 , 3.29 . Disengage Circuits The roll axis channel disengage cir cuits operate automatically if the roll rate exceeds 159 degrees per second. The circuits are controll€d-by a roll rate disengage accel erometer which at the predetermined roll rate limit compl~tes ~upp~y circuits to the differential servos centering circuit and the parallel servo disengage circuit. Note t hat the pow~r suppl y to the damper system components is uninterrupted in the emergency mode. The roll axis disengage circuit wil l also oper at e if the yaw axis monitor oper at es due to excessive skid, side slip~ or transverse accel eration. Oper ation of the disengage circuits r evert s the ch~nn~l to the · · · 11 • t t he R-P AXIS indicator pre-engage condition and i umina es light on t he master warning panel. In the pre~ engage · • · t is operative and the condition, the pressure trim circui . roll rate tr i m actuation switch is transferred t o the aileron t rim circuit. PAGE 17 ~--------------- ---_ _ _ SE_C_R_,E_T _________ _ 70/MAINT 13/1 3.2.29 3.2.30 When the channel is in the pre-engage condition the output of the force t ransducer and the roll rate gyro is routed to the trim unit to ensure smooth re-engagement. A trim circuit amplifier, fitted in the circuit immediately preceding the trim unit, discriminates upon the phasing of the applied signals to effect the proper direction of rotation of the motor-driven potentiometer of the trim unit. The :effect of this provision is such that any unbalanced signals appearing in the command circuit initiate the oper ation of the trim unit and are balanced out by the resulting output signal from the trim unit. This ensures that the channel will be re-engaged into a balanced circuit and, as described in sub paras 3.2.24 and 3.2.25 in agreement with the prevailing position of the ailerons. But for this provision, an unbalanced signal , resulting, for example, from f ailure of the gyro to nullify a roll rate signal, would produce unexpected and undesirable movement of the ailerons. Note that if an unbalanced signal is being nulled by the trim u~it when re-engagement occurs, the effect is equivalent to introducing an out- of- trim condition on the axis. This is due to the trim signal standing at the summing point of signals from the force transducer and normal accelerometer. In this condition, to regain level flight with the control column in neutral, the trim actuation switch must be operated. PAGE 18 .,, ?5 c,,. 3: :J: t C > 3: .,, m ?' ;,,:, 0 ,- ,- PILOT'S ~ BUTTON \ I --- ROLL RATE ~ ~:~: I \ \ ~ ACTUATOR : STEP r----0-----o 1 0 ;,,:, C 0 == z TRIM ay a fJ MONITOR ROLL RATE DISENGAGE TO PARALLEL SERVO DISENGAGE CKT. TO DIFFERENTIAL SERVO CENTERING CKT. PRESSURE TRIM LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO > •I PARALLEL SERVO ~ 'Ill:::: ~ c::i ~ ~ ~ c::i ~ ... " ., .. ... ' ., ., ., ... .., ..., ~ ' ;i: ' ..., ..., .. Tg ... ~ 'I ~ X I t C )> ~ ,, m ?' ;a 0 r- r- ► X pi m z c:, ► Q .!" Q m ► ;a C 0 :J: z ROLL RATE STICK FORCE PILOT'S BUTTON - • - •I TRIM I I •Cx'l STEP I I 0--a,---~ ay ll./3 MONITOR ROLL RATE DISENGAGE TO PARALLEL SERVO DISENGAGE CKT. ~ 00 ~ I ~ '-" I ... TO DIFFERENTIAL SERVO CENTERING CKT. LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO PARALLEL >------, SERVO ~ -.:::: :::.:, c:::, ~ :::.:, :::.:, c:::, ~ .. "S ,. .. .. .... ,. .,, ,. .. .., .... ~ .... t .... .... .,, .. T§ -n p 'I ~ :I: I °' .I>, 0 > ~ .,, m ?O ;,a 0 ,... ,... > ~ .!" m z a > a !" a m > ;,a 0 0 ,£ z ROLL RATE STICK FORCE PILOT•~ BUTTON - • - • I TRIM I I • O---~ ay II. f3 MONITOR ROLL RATE DISENGAGE TO PARALLEL SERVO DISENGAGE CKT. ~ m = I ~ = ' I-' TO DIFFERENTIAL SERVO CENTERING CKT. LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO '>----PARALLEL SERVO ~ -.:::: :a::i c:::, ~ :a::i :a::i c:::, ~ .. "' "' .. .. ~ .., .. .. .., .., ~ --t -.., ..., .. Tg ..., l5 co ~ ::c 0- .i. 0 > ~ .,, m JO ;,a 0 ,.. ,.. > >< p; m z C) > C) !" C) m > ;,a C .,, ROLL RATE STICK FORCE PILOT•~ BUltUN :..·-~ •I TRIM 1-------- I STEP I cY---~ ay a f3 MONITOR ~-TO PARALLEL SERVO DISENGAGE CKT ROLL RATE DISENGAGE l-------00--4 ....___. TO DIFFERENTIAL SERVO CENTERING CKT LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO "' .... ' .... 0 = ' .... ) .,1 PARALLEL ' ' SERVO :a:.. ~ :i::i c::, :a:.. :i::i :i::i c::, ~ .. "" ., "' ... ' ., " ., ... .., .... ~ ' ~ .... .., .. T§ ... p '° 3: ::c t C > 3: .,, m ?O ~ 0 r r- > X p; n 0 3: .,, :;; PILOT'S _a BUTTON ROLL RATE STICK FORCE V, ::c 0 :e z ~ TRIM ........,. _____ _ :e :::j ::c ~ m ~ ~~r~ I \ STEP I I 0----o----~ '.: ACTUATOR z m z C'I )> C'I m )> z C C'I m )> ~ C .,, PRE TRIM aytJJ MONITOR ROLL RATE DISENGAGE PRE 0 DN ENG TO PARALLEL SERVO . DISENGAGE CKT. TO DIFFERENTIAL SERVO CENTERING C.KT. LEFT DIFFERENTIAL SERVO RIGHT DIFFERENTIAL SERVO ~ 00 ' 6 = ' ,._, PARALLEL~ SERVO :i:,., ~ ~ ~ :i:,., ~ ~ ~ ~ ,. "S ., .. ,. .... .. .., .. ,. .., ..., ... .... :l:: .... ..., .,, t::, Ta ... 1749 A SECRET 70/MAINT 13/1 3.3 Yaw Axis (l<,igure lb) General 3.3.1 3.3.2 3.3.3 3.3.5 J.J.6 The_yaw axis ch~nel of the damper system corrects yawing· motiontand provides for co- ordinated rudder movement in manoeuvres l!l's , neo:e.ssary, 0 The ch~nel inco: por~tes two independent systems. One system_ is operative in the normal mode of damper sys tem. operation. The other system is operative only ih the event of th: normal system disengaging, i. e . in emergency mode operation. Except that a circuit is deleted in the emergency mode, both systems are identical in operation. No independent command signals are supplied to the yaw axis channel in either the normal mode system or the emergency mode system. The motivating signals are provided by sensing devices which operate automatically in particular attitudes of the aircraft . These signals are in most . ' instances, consolidated at various summing points and the resultant supplied to a servo amplifier. Signals applied to the amplifier are amplified to a usable level and the phasing discriminated against a reference. The output of the amplifier initiates the operation of a rudder dual differential ;servo -val ve which in turn operates the rudder hydraulic jack control valve to cause the deflection of the rudder in the direction appropriate to the phase of the amplifier unit input. A self balancing or feed-back circuit is introduced from the rudder dual differential servo t o t he servo amplifier. Movement of the dual differential servo results in feed- back signal proportional to such movement, appearing a t a summing point preceding the amplifier input. When this signal nulls the input signal to the servo amplifier, the servo circuit is balanced and further movement of the rudder does not take place. Upon the input signal to the amplifier being cant:.elledg the feed-back signal operates the amplifier causing the rudder to . be returned to its original position. Signals are supplied by the sensing devices in the norma~ mode system in respect of yaw rate, t ransverse acceleration, and aileron position differential by, respectively, a yaw rate gyro a transverse accelerometer, and a transducer. A signal ~hich is the product of pitch r ate and aileron ncsition is also provided. All signals are compensated for airspeed by air data transducer components. PAGE 19 3.3.7 3.3.8 3.3.9 3.3 .10 SECREI' 70/MAINT 13/1 The ya~ rate and ai~eron position signals are supervised ~ an integra~or uni~ whose action is similar to that of t he pitch rate axi~ sensing circuit integrator in that it has no effect on s~gnals of short duration but nullifies steady signals. The integrator also supervises partially the · transverse accelerometer signals when the landing ge · selected down. ar is Yaw Rate The ~aw rat.0 -signal-prbdu.ced by the yaw rate gyro is of a magnit~de and_phase appropriate to the rate and direction of yawi ng motion. Since the gyro signal is proportional to rate of yaw r at her than to displacement of the aircraft co m~aratively quick corrective action by the rudder can be' ~chieve~ before the displacement becomes significant . ITw in straight an~ leve~ f li ght or during turning movements, the_yaw :at: signa~ is t he result of transient yawing motion , it is applied through the various .summing points to the servo amplifier to effect a corrective mo vement of the r udder . As the aircraft returns to its original attitude, yaw rate signals of the opposite phase are produced which cause the deflection of the rudder in the opposite direction t o prevent overshoot. Steady state yaw r ate signals will cause an initial deflection of the rudder but, due to the action of the integrator the rudder will be returned to its original position by the feedoaiback signal. Aileron Position An aileron position signal is produced when the ailerons are mo ved to commence a turn or to maintain stability. The position of the ailerons is t ransl ated by a transducer into an electr ical signal , the magnitude and phase of which is dependent on the degree and direction of aileron movement. Note that, as the transducer signal is hi-passed by the filter integrator, the low rate simultaneous up deflection of t he ailer ons at high altitude will not result in a signal from t he transducer . See also sub . para. 3. 2.2. The t ransducer signal i s applied through the various .:summing points to t he servo amplifier to effect the deflection of the rudder in the direction appr opriate t o turn the ai rcr aft and overcome adverse yawing moment resulting from aileron movement. If the aileron position transducer signal is the resul~ of a co mmanded continuous rolling manoeuvre or of an applied ai l eron trim the signal will be, in effect, a s teady ~tate signal . Ther efore, as wi t h a steady s t ate yaw rate signal , the integrat or act ion will cause th~ ~u~der to be :eturned to its original position after the initial deflection or short term co - ordination. .,PAGE 20 -------------------~SE~CR~ET~------ .3 • .3. 10 .3 • .3.11 .3 • .3.12 .3 • .3.1.3 .3 . .3 . 14 70/MAINT 1.3/1 (continued) T~ansducer signals r:sult~ng ~rom the positioning of the ailerons by the se~sing circuit will continually position the rud~er to provide co-ordination as required by the . corrective ~o~ements of the ailerons. As the signals normally originated by the sensing circuit are of relatively short duration 9 they will not be nulled due to the operating time l ag of the integratoro. Transverse Acceleration Sl~p_or skid in any flight attitude resul ts in a signal originated by the transverse accelerometer being supplied to the servo amplifier to produce a corrective movement of the r·udder . As the sides lip motion is eliminated by the applied correction, the .ouj:.put of the transverse accelerometer correspondingly decreases so permitting the servo feed-back signal to decrease the deflection of t he rudder • When the landing gear is selected down !/ certain circuit changes are eff ected to by-pass the major port ion of the transverse accelerometer signal to t he summing point of the yaw rate and aileron position signals. This action subiects the accelerometer signal t o supervision by the integrator which nullifies the signal. By t his means, the pilot can introduce sides lip without opposition from the yaw axis channel as the yaw rate signals resulting from intentional sideslip will also be attenuated by t he integrator. The portion of the accelerometer signal which is not by-passed tends to maintain stability in the axis. Ailerons Position - Pitch Rate The simultaneous command of a large rate of roll and degree of pitch would introduce an adverse yawing_m~ment. Th? signal which is the product of aileron position ~d pitch rate is utilized in the yaw axis to counteract this condition. The magnitude of this signal is directly proportional to the aileron position commanded and degree of pitch rate. The signal supplements the short term rudder co-ordinat~on . signal so increasing the deflection of the rudder in rel~tion to the !I attitudes commanded . As this signal is not. supervised by the integrator rudder deflection is applied until the command signals are cancelled. 'ii" th the t • of the combined aileron position and pitch • excep ion . "d t·cal rate signal input, the emergency mode sy stem is i enfi th with that of the normal mode sy stem. The changeover rom e normal mode to the emergency mode is effected by t~~ pu~~ switch on the control column handgrip or th~ ope~a io~. a rudder monitor which initiates the operation o cer ain PAGE 21 18 A ~-----------------~SE~CR~ET~----- ?0/MAINT 13/1 3. 3.14 3.3.15 3.3 ,16 3.3.17 3,3,18 ( continued) electrical c~rcuits of the rudder dual differential servo an~ re~ults in the_illumination of the roll and/or pitch axis disengage indicator light on the master warning panel The operation of the emergency yaw damper =stem is · di t' d b th • 11 • t. f t -., in ea e .Y. e i umina ion o he yaw emergency damper operative indicator ~ also located on the master warning panel. Note that when the rudder monitor is operated 9 the pitch axis and roll axis damper circuits are al~o disengaged, See the Pi tch Axis and the Roll Axis Disengage Circuits. Rudder Monitor The rudder monitor circuit includes a sideslip switch i.e. the beta vane in the nose boom and two transverse accelerometers which are located 13 feet and 40 feet . ' respectively 9 forward of the aircraft's centre of gravity. The sideslip switch is set to open and interrupt the normal mode sy st em supply if the sideslip angle exceeds ten degrees. This function is primarily for low airspeeds in which condition a small acceleration may correspond to a large angle of sideslip. The action of the accelerometers is such that one detects direction of sideslip and the other yawing acceleration . If yawing ~cceleration is sensed which tends to reduce sideslip the changeover circuit will operate at a transverse acceleration value of approximately :± o.6 11g 11 at the centre of gravity. Conversely 9 if the yawing acceleration is such that it tends to increase sideslip the changeover circuit will operate at the relatively lower transverse acceleration values of += 00 2 ng 11 • This configuration is necessary due to the fact that in co- or9inated manoeuvres rel.ati vely large values of transverse and yawing accelerations are required for which the monitor must make provision. Rudder Feel . Trim and Hinge Moment Limitation A trim circuit is provided to permit the rudder to be trimmed to compensate for steady state yawing motion resulting from asymmetric flight or other cause~. The circuit is controlled by a trim switch located rn_the front cockpit and consists of an actuator fitted as an in t egr~l component of the rudder rear quadrant assem?ly. 0pe~ating the trim switch results in the actuator moving the linkage of the rudd er hydraulic jack control valve. The ~o~em~nt of the rudder by the trim circuit effects the positioning of the rudder bar. PAGE 22 ' ________________ _______ SE_C_R_EI' _______ _ 70/MAINT 13/1 3 . .3.19 1749 A In addition to the trim actuator, the rudder rear quadrant assembly incor porates a rudder feel and hinge moment limitation system. This system functions automatically and its action is to decrease the rudder movement for a given force on the rudder bar in proportion to increasing dynamic pressures. This is done to counteract the increased sensitivity of the rudder at high dynamic pressures. PAGE 2.3 .,, !5 -0 ~ ::c t 0 > 3: -a m '!" -< > == > X .;; YAW RATE AILERON POSITION ..,1,---- ; LATERAL ~ ACCEL. z '------ I" PITCH ~q--~ RATE I AILERON \" a I POSITION YAW \r I RATE • t; ,-------, z ..,.. AILERON ffi POSITION :E '----.... LATERAL ACCEL. ay or Say q I AMPL. AMPL. CONTROL VALVE NORMAL EMERGENCY DIFFERENTIAL SERVO RUDDER MONITOR ~ = ' ~ ' >-" 1) ay =~.&g P=:t10° OR 2)TWO AC CELEROMETER CONFIGURATION :t:,,, ~ ~ c:::i :t:,,, ~ ~ c:::i ~ .. "' ., .. .. ' ., ... ., ... .,, .... .... ' ii: ' .... .,, .. T§ ~ 1749 ,4 -------------------- ~S~EC~R~ET~------••'• 70/MAINT 13/1 GROUND TEST EQUIPMENT 4,1 Introduction 4,1.1 The Ground Test Equipment described in this report is the equipment that will be used for the flight test phase of the Damper System program. The equipment has been designed with consideration for its ultimate employment by the RCAF with possible modification and simplification as found required during the flight test phase of the Damper System. 4,1,2 Two types of equipment are used, the first line equipment for primary inspection and the second line equipment that is used to service faulty Damper components. 4, 1.3 The first line test equipment consists of the 'following units: (a) UG 6004A=l Damper Test Set (b) UG 6005A-l Damper Auxiliary Test Set (c) MB-1 Pitot-Static Test Set (d) UG 6006A-l "G" Limiter Test Set This equipment will provide fov the following: (a) (b) ( C) ( d) Quantitative "go- no-go II t ests on the Damper Amplifier Calibrator, Hydraulic System, and Control Surface Linkage. Quantitative tests on the Pitot-Static System including sensors and tr ansducers. Qualitative tests on the rate gyros and accelerometers . Loc ation of faulty components. 4.1 . 4 The second line test equipment consi s ts of the following units: UG 60Q3A- l Damper Test Stand UG 6002A- l Rate Table UG 6004A- l D~mper Test Set UG 6005A- l Dampec Auxiliary Test Set MB- 1, Pitot-Static Test Set ( a) ( b) ( C) (d) ( e) (f) ( g) UG 6006A-l 11G11 Limiter Test Set . Servo Hydraulic Test Stand (Requi:ement_for this piece of Test Equipment is being investigated) This equipment will provide for the fo l lowing: (a) Calibration (b) Quantitative tests on gyros and accelerometers . (c) Vertification of faulty components , (d) Pre- installation tests of a whol e Damper Sy stem PAGE 24 ------------------- ~S::'.'.EC~RE::'.:T:__ ____ _ 70/MAINT 13/1 4,2 UG 6004A- l DAMPER TEST SEI' 4,2,1 Description 4,2. 1.1 The UG 6004A-l Damper Test Set performs "go-no- go" tests on the Damper Amplifier Calibrator, hydraulics, ru_id control_ sur~ace linkage. When used in conjunction with the MB=l Pitot- Static Test Set it permits quantitative tests to be performed on the air data sensors and transducers as well. 4.2.1.2 The UG 6004A-l Damper Test Set consists of a source of step and pulsed 400 cycle voltage, Damper Amplifi er Calibrator Test Selectors A & Band a bridge circuit for measuring control surface deflection. When t he Test Selectors A & Bare r otated to the desired positi on a signal of the proper amplitude and duration is appl ied to the portion of the Damper bridge selected. A vo l tage is simultaneously applied to the"qc' scheduler set ting the system gain t o t he desired value. This will cause the control surface t o deflect to some pre-determi ned angle which will be measured by the bridge circuit. One arm of the bridge circuit is a surface position potentiometer whose wiper is mechanically connected to the contro l surface and whose output voltage is therefore proportional t o the surface deflection. The other arm of the bridge is a voltage divider whose output is a voltage corresponding to the desired surface defl ection. The voltages from the two arms of the bridge are applied to the SURFACE INDI CATOR which indicates the difference between these voltages. 4.2.1.3 Selected re s i s tors in series with the SURFACE INDICATOR change the sensitivity such that the coloured area at the centre of the indicator marked "Good" alweys represents t he allowable limits of system performance . 4.2.2 Operation The UG 6004A-l Damper Test Set will be of the portabl e 11fly - away' t ype and capable of being carried by one man . It will have the approximate dimensions: Length Width Height Weight - 20 i nches - 18 inches - 8 inches - 40 l bs. The operation of the UG 6004A- l Damper Test Set will be explained wi t h reference to Figure 11 . -:-----------------~S~EC~R.filEr~----- 4.2.2.4 70/MAINT 13/1 The sensors are disconnected by removing two connector fro m t ~e Damper Amplifier Calibrator and connecting s in t heir place two connectors from t he Damper Test Set. As the TEST SELECTORS A and Bare rotated pre-selected vol~ages ar~ applied to each of the Dampe; Amplifi~r Calibrator inputs in turn causing control surface deflection. The output of the surface position potentiometer is compared with t he voltage fr om the voltage divider and the difference appears on the SURFACE I NDI CATOR as a 11Good II or 11Bad" indication. Circuits whose input signals are high-passed are checked in two steps: first as a gain, and secondly as a high pass signal. In order to test the gain, the output of the high-pass circuit must be zeroed. This is done by allowing the high- pass unit to run until its output is zero when no input is applied to the bridge. The input to the high~pass amplifier is then gr ounded. The gain of the circuit is then checked as described above and if this is satisfactory, the input signal is maintained and the second step undertaken. The high pass amplifier is connected to the Damper circuit for one second causing its mo tor to run applying a voltage to t he Damper that cancels a portion of the original step input si gnal. The amount of cancellation will depend upon the gain of the high- pass amplifier and will appear as a r eduction in the surface deflection. A new bucking voltage will be appli ed to the SURFACE I NDICATOR allowing the operation of the high-pass to be read as a "Good" or "Bad II deflection on the SURFACE INDICATOR. Integrators are tested by running them for one secon~ with no other signals being applied to the Damper bridge. Satisfactory operation of the integrator will result in a "Good II indication on the SURFACE INDICATOR. Schedulers are tested automatically during the series of tests described above. The input t o the scheduler must be varied for each input signal in order that the resulting surface deflection remain constant. On the UG 6on1A _l Damper Test Set Additional controls ~ are as follows: (a) DIRECTION SWITCH= Tests on t he Damper are_nor~ally erformed with the control surface defle~ting in ~ne direction only. If the operator desire~ to check h t f the control surface deflection he t e symme ry o N s · t h t may do this by switching the DIRECTIO wi c o the "negativen position . PAGE 26 , -------------------~SECR~E::_T ____ _ _ ,_ 4.2.2.6 70/MAINT 13/1 (continued) This reverses the phase of the input signal causing the control surface to deflect in the opposite direction The bucking voltage applied to the SURFACE INDICATOR 0 . is also reversed, so that the control surface deviation indication is still -valid. ( b) ZERO LEFT & JilGH'r TRIM CONTROLS - These two controls can be operated when the TEST SELECTORS A and B are selected to t he SL (Streamline) position. They permit the technician to zero his SURFACE INDICATOR even if the control surface is not perfectly streamlined. · Two controls are used because the two elevator s and two ailerons operate independently a:nd it is not likely that their motion will track exactly. (c) (d) TEST SELECTOR A - This control applies a signal to each of the inputs to the pitch and roll axis of the Damper bridge as it is rotated through successive positions. Each position of the switch is marked by a symbol representing the particular sensor that this input replaces. These same symbols will appear on all schematic and block diagrams furnished with the test equipment giving the technician a clear reference to that portion of the Damper bridge he is testing. • TEST SELECTOR B - This control operates in exactly ·the . same manner as TEST SELECTOR A. No inputs wil l be supplied to yaw axis, however, unless TEST SELECTOR A is in the TO B position for TEST SELECTORS A & Bare interlocked to prevent simultaneous inputs to the two axes, When TEST SELECTOR B. is selected to the AFCS position it will be possibl e to use the Damper Test Set with the AFCS Test Set to test t he AFCS equipment . TEST SELECTOR A & B positions are as follows: PAGE 27 ~-----------------~SEC~RE~T _____ _ ELEVATOR AILIBON 70/MAINT 13/1 A B 1. Streamline SL 1. Streamline 2. Stick Force css SL 2. Pitch Rate & 3. Normal Accel n: Ail. Pos. 4, Pitch Rate q RUDpER. 3. Yaw Rate q&a r 5, Hi pass qh NORMAL 4. Lat. Accel 6. Pitch Disengage fD a 5. Ail. Pos. .la 7. Streamline SL 6. Hi-Pass HP 8. Integrator INTEG 0 { 7. Streamline SL 0 Streamline SL RUDDER 8 0 Yaw Rate 10. Stick Force r css "EMERG. 9. Lat. Accel. 11. Roll Rate ay p o. Ail. Pas. ba 12. Roll Disengage RD 1. Hi-Pass HP 13. Streamline SL 12. Pitch=:J--GEAR DOWN 14, Integrator INTEG. 13. Roll . . 15. TO B 14. qc Normal qcNCR 15. Altitude Alt 16. AFCS AFCS (e) ELEVATOR/AILERON Switch - This switch is used in conjunction with the TEST SELECTOR A to ensure that the motion of the l eft and right elevators or ailerons are symmetrical. Normally , these surfaces will be tested with the ELEVATOR/AILERON switch in the 11Left 11 position but it should be selected to the "Right" position for one test as a check on the symmetry of the surface deflection. Symmetrical surface deflection will ,appear as equal and opposite defl ections on the SURFACE INDICATOR, (f) EXTERNAL/INTERNAL Switch~ This switch is included as part of the UG 6004A=l Damper Test Set to simplify wiring and to reduce the number of interconnecting wires between this unit and the UG 6005A-l Auxiliary Test Set. When the EXTERNAL/INTERNAL switch is selected to the INTERNAL position, and the TEST SELECTCRS A & Bare rotated, the correct values of input voltages to the Damper bridge, the•q c' schedulers, and the bucking voltage to the SURFACE INDICATOR are automatically selected. This reduces the Damper Test procedure t o one of rotating a series of switches and observing the deflection of an indicator. When the EXTERNAL/INTERNAL switch is select ~d to the EXTERNAL position the input voltage and the ''q c" inputs are transferred to the Auxiliary Test Set, where the 'magnitude of these signals may be set to any desired value, (g) INTEGRATOR/HIGH=PASS Switch - This switch is to be pressed whenever the TEST SELECTORS A ·& B are in the HP or INTEG position.. It st arts the timer motor connecting these units to the circuit for one second. · ~ SECRET -~:-------~- 1749 A 70/MAINT lJ/1 4,2.2.6 (continued) (h) A-C POWER Switch - This switch applies A-C power to the Damper Test Set. (i) DI-C POWER SwitcR - This switch applies D-C power t o the Damper Test Set. (j) D-C ADJUST POTENTIOMETER - This control permits the technician to set the D-C voltage to the VOLTAGE ADJUST position of the SURFAC~INDICATOR when t he D-C PRESS TO ADJUST switch is actuated, , (k) D- C PRESS TO ADJUST Switch - This switch shoul d be pressed after the D-C is turned on and before the t est begins. It allows the D-C ADJUST potent iomet er to be used to adjust the D-C voltage entering the test set . When the voltage is set to the D-C VOLTAGE ADJUST line of the SURFACE INDICATOR the indicator sensitivity will be correct. JYKD A! KCKAFT LIMITED SECRET AVRO AR ROW► -------.:;__~ \ .... ·.-· -· ~·· ===""'"==-"'-'==--'-'· ·~···"':·: ... •. : ... ::::": ... '--: ~·:::::=::::-::::·• •. {; ======(7]=====- ZERO LEFT TRIM Q DIRECTION ZERO RIGHT TRtM ELEVATOR,,AILERON SURFACEJNOICATOR ~ HP•IIITEG \EMERGEHC)' ,/\. SL ray Ru., GEAR / ·11p \ / ~a 11 / D AITTO~ <>IOO DEFLECTION M " qcAOJllSTPSF 16 18 20 (f """PIJl e "'"'""'' © ,,~,\,l,/1~{',. ~ .. ,! {2 9\ :~_: / I~ ~ D-UG6005A-I / I \ TEST SET-DAMPER, AUXILIARY J~ ~MANUAL START HONEYWELL CONTROLS LIMITED \JIii © TORONTO, CANADA ~ , FIG. 12 UG 6005A -l DAMPER AUXILIARY TEST SET --:------------------___.::SEC~RE~T:___ ________ ~- ~,.,,. . 70/MAINT 13 /1 4,4 MB-1 PIT0T - STATIC TEST SET 4,4.1 De s£!:i.J2.:li2!l (Reference Figure 13) 4.4.1.1 The MB-1 Pitot-Static Tett Set will be used to test the normal and emergency pitot-static sy stems. 4,4.1.2 The Test Set consists of the following equipmentg (a) (b) (c) (d) (e) (f) ( g) (h) 1 hand operated vacuum pump 1 hand operated pressure pump 5 needle-type control valves 1 1 1 1 1 vacuum gauge 0-to 809 000 ft. altimeter machmeter (to roach. 2.5) thermometer circular slide rule type computer. 4.4.1.3 The Test Set is a portable unit enclosed in a metal carrying case which has the following dimensions: Length Width Height Weight - 18 inches - 10.5 inches - 6 inches - 23 lbs. JYRD A/I/Cl/AFT LIMITED FIG. 13 MB-1 PITOT STATIC TEST SET SECRET AVNO ARNOW► -~--------------~S~EC~REI'~------ 70/MAINT 13/1 4,5 UG 6006A-l 11G11 LIMITER TEST SET 4,5,l Description 4, 5. 1. 1 4.5. 1.2 4,5.2 Operation The UG 6006A-l "G" Limiter Test S t • e will permit an independent check to be made of the "G" Limiter ca libration. Th 11G11 L' • t e imi er Test Set is a self- contained unit and will be _stored in the UG 6004A~l Damper Test Set carrying case ~ it will have the following dimensions: Length Width Height Weight = 5 inches - 4 inches - 6 inches - J lbs, 4. 5.;2.1 It will be necessary to disconnect the accelerometers and stick force signals from the 11G11 Limiter Calibrator to test the 11G11 Limiter. 4.5.2.2 Signals simulating the outputs of the accelerometers will be applied to the 11G11 Limiter calibrator . A step input of pre-determined amplitude will be fed into the stick input. This signal when high-passed in the 11G11 Limiter Calibrator will apply a pul se to the "G 11 Limiter bridge causing the disconnect relay to operate . 4.5.2.3 Additional tests will apply simulated accelerometer sisnals to the 11G11 Limiter bridge to cause the di sconnect relays to operate. Both the normal and emergency limi ters will be checked for positive and nega tive 11G1" s. Panel lights on the 11G11 Limiter Test Set will indicate proper operation of the "G" Limiter. PAGE 33 ~ SECRET ~:-------~~-- 70 /MAI NT 13/1 4,6 UG 6003A-l DAMPER TEST STAND 4,6 0 1 Descriptio!}, (Reference Figure 14) 4.6.1.1 4,6.1.2 4.6.1.4 4,6.2 Operation 4.6 . 2.2 The UG 6003A- l Damper Test Stand is a combination bench and transit case. Damper components are mounted on brackets which form part of the stand and are electrically interconnected by means of cables provided with the Test Stand, ~pace is provided to mount the UG 6004A-l Damper Test Set l' and t he UG 6005A= 1 Damper Auxiliary Test Set on a shelf above the Damper components. Space is also provided to mount the MB-1 Pitot Static Test Set, The UG 6003A-l Damper Test Stand is of the "fly-away" type. The overall stand dimensions and weights without Damper components are approximately as follows: Closed for transportation Length - 48 inches Width = 34 inches Height = 36 inches Weight = 200 lbs. Open Length - 77 inches Width = 34 i nches Height = 60 inches Weight - 200 lbs. The power required to operate the Test Stand will ~e 27.5 volts d=c with an input current of 10 amps an 115 volts ~ three phase, 400 cps at 5 amps. The .Tes.t., Stand will be used for/a;!~:;~:f1~~~~~nents~ confirming faulty components . ~ p It will allow tests of a complete Damper Sys em. ts to be isolated~ gains to be adjusted ~ faulty comp~ne~ and the cause of the fault determine• f d back potentiometers are Manually adjustable ee . d t. ve ick-off connected provided to replace the in uc i pt saturation of the to the servo unit. This wi~_l prev:~ep input will be 'f' In pract~.ce l' a servo ampli ier. t nd the feedback applied to the Damper Sys em_aimum deflection of the • dJ'usted for min t d potentiometer a . d' ators which are loca e SERVO CONTROL CURRENT_i~ ic Test Set. A surface on the UG 6005A-l Auxiliary d to the feed back position potentiometer gange PAGE 34 _____________________ ___ S_ECR....::...::E::..T _____ _ 70/MAINT 13/1 4.6.2 . 2 (continued) potentiometer will provide the necessary voltage for operating the SURFACE INDICATOR on the UG 6004A=l Damper Test Set . ::!! ~ i'. C G> 0- 0 0 .., !'.: C ► 3: .,, m :,g .... Cl .... V, .... ► z C f UG6004A-1 DAMPER TEST SET UG6005 A -l :41ii II DAMPER A UXILIARY TEST SET 'G ' LIMITER CALIBRATOR ACCELEROMETERS I \W J ~~11~ ~ ~ I. JUNCTION BOX ~ ~ ~-[rJ- - -.lf:11 _J - l,_--i71JJ IL1] . . . ,. q- I/ ,\ i)J \ 7 /: ~ • fL...4:_.___~ ~ -~ ;; ., s S SS ~ mm,RFA)', ! , ~~ ~ '!a 0 0 0 I IA~ ffi !'/':"' ~ · I \\\ ~~~~RARfllff 't~ -L__ ___ 'L:11 I LJ/'--- . ~7• ,,en~• 0 ·:: DAMPER AMPLIFIER CALIBRATOR - --------=::_~ ~-:---~~ :_. __ _,,--~~ - ~ ~ -- ~~ -= _:;-~ ~ '--~ ~~--~ -; ~ - ~ ;:::===-__ - CONTROL COLUMN ,/ / \ "No,oN mmo• HYDRA ULIC SIMULATOR ::t:,,. 'oi:::: ~ ~ ::t:,,. ~ ~ ~ ~ ' ... ,., "" ' ~ ... .. ... -., .... ~ .... l::: .... .... "' .. Tg llit:-----------------~SEC~RE~T------~ · ' 70/MAINT 13/1 4, 7 UG 6002A- l RATE TABLE 4o7.l Description (Reference Figure 15) 4o7ol.2 4o7ol.3 4,,7.2 0J2eration The UG 6002A=l Rate Table is a Genisco Model 181 Rate Table mounted with an operating console in an air- transportable case t hat also serves as a stando The operating console contains all the swi t ches and meters necessary to perfor m quantitative t ests on the rate gyros and accelerometerso The gyro tests include linearity j sensi ti vi ty and maxi11tum turning rate as well as the measurement of spi n motor currents, The accelerometers can be tested for sensitivity , linearity and rangeo Those accelerometers that perform switching operations can be tested to ensure that their operation is within the allowable limits of performance, Provision is made to che ck t he cont i nuity of the heaters in those gyros and accelerometers so equipped, The Rate Table approximate dirnensions and weight are as follows: Closed for transportiom Length - 48 inches Width = 48 inches Height - 36 inches Weight - 400 lbso Opern = 48 inches Length Width = 48 inches Height - 56 inches Weight = 400 lbso (Reference Figure 15) b t t d bv comparing their The accelerometers can e es e J . . b d by mounting two linearities, Thi s can e ~ne d connecting accelerometers on the rotatinG_beamtants buck each t hem in such a manner t~at the~r o_u pu t he difference other• A meter is provided to _mea::\ithi n some in t he outputs and if the outiuts eters are considered predetermined levels the acce erom serviceableo er ate a switch at some Those accelerometers that op t d by rotating t hem at t • can be tes e preset accelera i on . . st above and below h t elerations JU a speed such ta ace_ h O eration, or non these l i mi ts are obtained,_T el i~it switch will be an operation of the accelera ti~n . f the accelerometer, indication of t he serviceability o AY/10 AIIICRAFT LIMITED RATE-OF-TURN TABLE AVHO SECRET ARNOW► ACCELEROMETER MOUNTING BRACKETS ~ -:-- t ~~~Lftl . • . ~~~- - . . . <~ 'El .... ~'.,~ ,~, I ff;,' , I ' I i FIG. 15 D-UG6002A-I RATE TABLE -:---------~SECR~ET--- 7O/MAINT 13/1 5, FIRST LINE TESTING 5, 1 Test Equipment Required (a) Hydraulic Test ·Machine .Trailer (b) A-C Ground Power Unit (c) UG 6004A-l Damper Test Set (d) UG 6OO5A-l Damper Auxiliary Test Set (e) MB-1 Pitot - Static Test Set (d) UG 6OO6A- l 11G" Limiter Test. Set 5,2 Preparation for Test 5.2.1 Rotate all switches on the UG 6OO4A-1 Damper Test Set to the left, turn all power switches OFF and centre all controls. Turn t he EXTERNAL-INTERNAL switch to INTERNAL. This readies the UG 6OO4A-l Test Set for checking the Damper. 5,2,2 With reference to Figure 15 connect the UG 6OO4A- l Damper Test Set to the aircraft as follows: (a) Remove connector Rl O87/5 from the Damper Amplifier Calibrator and connect in its place P29 from the Damper Test Set. (b) Remove connector Rl087/8 , from the Damper Amplifier Calibrator and connect in its place P3O from the Damper Test Set. (c) Remove the Damper Test Shorting Plug No. 2 and connect in its place P31 from t he Damper Test Set. This connects t he UG 6OO4A- l Damper Test Sett lho .the;-:D.amp.er and surface position potentiometers permitting "go-no-go" checks to be made on the Damper Amplifier Calibrator , hydraulic system and control surface linkage. 5,2 ,3 Connect the A-C Ground Power Supply Unit and Hydraulic Test Machine Trailer to the a i r ·craft. 5,3 Test Procedure 5. 3.1 5,3 . 2 5.3,3 Select t he MASTER ELECTRIC switch , located in the pilot' s cockpi t t o the ON position. Select the PO\·lER switch , located on t he Damper Funq_~i-?_n Selector Panel in t he pilot 8s cockpit to the ON posi tion. Carry out the followin g on the Damper Test Set. • h t the ON position (a) Select the A-C and D-C POWER switc es ? . ted and check that the POWER ON lights are illumina • PAGE 37 DAMPER AMPLIFIER CALIBRATOR .., ?i 0: SHORTING PLUG NO 2 (DAMPER TEST) n 0 z z m n .... z G') .... m "' .... m 0 C ~ m z .... .... 0 l> ;;; n ~ =!l ~ MAGNETIC AMPLIFIER (HINGE MOMENT LIMITER) RATE GYRO PACKAGE Rl087/ 6 P7B 7A P6 UG6004A-J DAMPER TEST SET ~ 'Ill::: :ii:. c:::i ~ :ii:. :ii:. c:::i ~ .. "' .. .. .. ... .. ... .. .. .,, ._, .... ... t ... ._, .,, .. T§ .... ~:------------ ~S~ECRE~T _ __ _ 70/MAINT 13/1 (b) Press the D=C PRESS TO ADJUST push- switch and adjust the D=C ADJUST control until the .pointer of the SURFACE INDIOATOR is at the D- C VOLTAGE ADJUST position. (c) Enga ge the Damper by pressing the DAMPER ENGAGE push=switch located on the Damper Test Set . (d) Ensure that TEST SELECTOR A is selected to the first "SL" posiJion. With the Damper Test Set as above a zero signal is applied to the pitch axis bridge which should ;esult in zero surface deflection producing a zero reading on the SURFACE INDICATOR. Should tpe control surface not be streamlined a non zero reading will be read on the SURFACE I NDICATOR. This indicates residual signal in the damper bridgep faulty mechanical linkage or a faulty servo control valve operation. If the surface is only slightly off the streamline pos.ition it is quite likely that the trouble lies in faulty mechanical linkage and the procedure in sub para 5. J.4 should be followed. 5.J.4 If the SURFACE INDICATOR does not read zero adjust the ZERO LEFT TRIM control unit it does. The control index should not meed to be rotated more than __ 0 •in order to zero the SURFACE INDICATOR. 5.J. 5 Select the ELEVATOR/AILERON switch to the ''R. 11 position and .adjust the ZERO RIGHT TRIM control until the SURFACE INDICATOR reads zero. Select the ELEVATOR/AILERON switch to the 11L11 position. This test ensures t hat both the left and right elevators are in t he streamline position or that the streamlining has been corrected for electrically i n the t est set. !'!QTil;g- All tests but one on the pitch axis damper will be performed by energizing the differential servos only. In additionp the tests are performed on the "gear up" confi guration so that the fader potentiometer to the differential servo bridge is at its full gain end 9 and the parallel servo fader potentiometer is at its zero gain end . Furthermore~ the output of the integrator in the parallel servo circuit is fed-back to the input of the integrat or forming a hi gh pass circuit to prevent the integrator from drifting. These precautions prevent the tests on the diff erential servos from being confused by such factors as improper operation of the integrator and . altitude scheduler or by integrate~ dri ft. The deflection of the control surface for ail · tests on the different i~l servos will be approximately five degrees. The operation of the integrator and parallel servo wi ll be tested at the "PD II position of the TEST SELECTOR A and the fader circuits will be tested at the PITCH GEAR DOWN position of the TEST SELECTOR B. SECRET 7O/MAINT 13/1 5.3.6 , Select the TEST SELECTOR A to the 11CSS 11 position. A signal voltage simulating that ~rom t he stick f orce transducer is injected into the pitch axis bridge causing the control surface to deflect. A second voltage is simultaneously applied to the D=DG 33A unit causing it to run to some predetermined p~sition9 automatically adjusting the gain of the Damper circuit. The combination of input voltage p manual gain and automatic gain determines the control surface deflectio~. The signal voltage and scheduling voltage are set to produce a deflect ion of approximately 5°. 5.3.7 A potentiometer attached to t he control surface produces an output voltage proportional to the surface deflection and. this voltage is cotapared with a fixed voltage in the Test Set and displayed on the SURFACE INDICATOR, If the differemce between these two voltages is within tolerance the needle of the SURFACE I NDICATOR will deflect within the GOOD portion of the scale, 5,3.8 Select the ELEVATOR/AILERON switch to the "R 11 position 9 the SURFACE I NDICATOR should read GOOD. Return the ELEVATOR/AIIERON switch to the ''L" position. This checks that the motion of the left elevator is equal to that of the r i ght elevator. 5. 3.9 Select the DIRECTION switch from the 111' 11 position to the 11 =11 position. This causes the elevators to deflect in the opposit e (down) direction. A GOOD reading should be obtained on the SURFACE INDICATOR indicating that the motion of the control surface is symmetrical. Return the DIRECTION switch to the n+n 5.3.11 position. Select the TEST SELECTOR A to the 11n 11 position. The SURFACE INDI CATOR should read GOOD 0 I t should be noted that it is no longer necessary to rotate the ELEVATOR/AILERON switch or the DIRECTION switch to check that the symmetry of the elevator de flect ion is correct since this symmetry has already been establ ished on theuCSS11 position of the TEST SELECTOR A. Select the TEST SELECTOR A to the 11J 11 position. A GOOD indication on the SURFACE INDICATOR should be obtained, Select the , TEST SELECTOR A to the "_q h11 position and ?res~ the HP=INTEGRATOR push=swi tcho When the TEST SELECTOR A is. first selected to the "q h n position , the SURFACE I NDICATOR will deflect off the scale but will gradually come back to the GOOD portion of the scale within _ seconds af~er t~e HP=I NTEGRATOR push=switch is pressedo A GOO~ reading w~ll. indicate that the high pass port i on of the ?itch rate circuit is operating properly 0 During all the previous tests (e~cept 5 3 11· above) the high pass unit has been connected to the Damper b;idge in such a manner that no signal will reach the parall el servo amplifiero PAGE 39 ~--------------__ ____:::SE~C~REI':'..:._ ________ _ Ii! 1749 A 7O/MAINT 13/1 5o3ol2 (continued ) Whe~ the T~ST SE~EX::TO~ A is selected to the "q h" position a signal simulating pitch rate is applied to the bridge causing the elevators to deflect 9 (the high pass unit is grounded to· prevent drift ) and the SURFACE INDICATOR to deflect off the scaleo When the HP=INTEGRATOR switch is pressed the ground is removed and the hi gh pass unit is connected to the bridge for ___ seconds causing the high pass unit to operate and "washing out " a portion of the input signal 9 there-by- reducing the amount of elevator deflectiono If the operation of the high pass circuit is correctD the needle of the SURFACE INDICATCR will return to the GOO_D portion of the scale 0 This test actuallj" checksthe time constant of the high pass unit 0 5, 3o 13 Select the TEST SELECTOR A to the "PD" posi tiono The SURJ:"ACE INDICATOR should read approximately 5°. Hold down the PITCH DISENGAGE push- switch , The SURFACE INDICATOR reading should now decrease to zero. If the SURFACE INDI CATOR does not return to zero but remains fixedj a faulty differential servo is indicated. If the surface deflection increases, the parallel servo is not disengagedo When a faulty differential servo is indicated it will be necessary to repeat step 503,13 but with the ELEVATOR/ AILERON switch in the "R II position to determine if the fault is common to both differential servoso 1:ill.'.!]g = Re ":lengage the Damper by- pressing the ENGAGE push-switch:, Return the ELEVATOR/AILERON switch to the 111 11 position. Select the TEST SELECTOR A to the 11SL 11 position and check that the SURFACE INDICATOR is at zero deviation for both left and right elevat orso This test ensures the proper streamlining of the control surface before checking the pi tch axis integrator. Select the TEST SELECTOR A to the 11INTEG" position and press the HP=INTEGRATOR push=switch. The SURFACE INDICATOR needle will first go off the scale but should return within the GOOD portion within _ second if the integr~tor is ope:ating properly 0 This test checks both the operation of the integrator and of the parallel servo j for i n this case no signal is applied to the differential servo; the control surface motion being carried by- the parallel servoo The integrator is tested by applying a signal of known amplitude and d~ation wh~n the HP=INTEGRATOR push switch is pressed 9 causing the uni~ to run and produce a pre~determined known surface deflectiono Select the TEST SELECTOR A to the AILERON "SL" position and check that the left and right ailerons are streamlined by means of the ELEVATOR/AILERON swi t ch as des,eribed in step 5 3 7 for the elevator 0 Return the ELEVATOR/AILERON switch t~ the llL" posi tion0 Select the TEST SELECTOR A to t~e '.'CSS 11 position and ensure the SURFACE INDICATOR deflects within the GOOD portion of the scaleo ----------------~S~ECR~l!."l'~-------.. '"'-·- 5.3.17 5.3 .18 5.3.19 5.3.20 5.3.21 5.3.22 5.3 . 23 70/MAINT 13/1 Select the DIRECTION switch t th it 11 • • ELEVATOR/AILERON . o e , .-:- position and the . switch to the "R II position A GOOD reading should be obtained on the SURFACE INDICATOR indicati~g ~~t control surface motion is symmetrical. Test 5.3.16 5.3.17 are necessary to ensure proper operation of both the left and right aileron servo sy stems as each is separately powered. Select the ~EST. SELECTOR, A to the 11P11 position and ensure that the deflection is GOOD. It is no longer necessary to rotate the J)I~ECTION switch and the ELEVATOR/AILERON switch to ensure that aileron_ position i s symmetrical since this has already be~checked in 5.3.16 and 5.3.17 and only those portions of the bridge that are common to both surfaces are now being tested. Select the TEST SELECTOR A to the 11RD 11 posi.tion 0 The SURFACE INDICATOR should read approximately 5°. Press the ROLL DISENGAGE push=swit ch to test-the proper operation of the differential servo disengage relay. If the ·relay is oper ating properly the differential servo should become disconne cted and return to a streamline positi on due to the satisfactory operation of the centering spring in the servo unit. Select the TEST SELECTOR A to the 11 SL 11 position. This removes all signal inputs to the roll axis bridge and ensures that t here are no spurious signals being appl ied to the servo amplifier. It is necessary to ensure streamlining at this point preparatory to checking the integrator. Select the TEST SELECTOR A to the 11INTEG 11 position. The SURFACE INDICATOR should deflect off scale as a voltaee applied to the SURFACE INDICATOR representing the final deflection will not be reached ·until ____ seconds after the HP=INTEGRATOR push= switch is pressed. If the integrator is operating satisfactorily the control surface should slowly return to its predetermined position which will be indicated by a GOOD indication on the SURFACE INDICATOR. Select the TEST SELECTOR A t o the 11TO B 11 position. This automatically transfers all functions from the TE~T. SELECTOR A to the TEST SELECTOR B. If the Test Set was or~grna~ly t P roperly the TEST SELECTOR B should be at its first se up 1 . • • t by "SL" position. Check that the stream ining is correc - · that t he SURFACE INDICATOR reads zero. If not1 ensuring .1 · t d rotate the ZERO LEFT TRIM control unti i oes . U I/ • Select the TEST SELECTOR B to the ~ ""·' ~ ~ > 2; t:I > Cf.l Cf.l 8 H ~ t 'cl 'cl trj 8 H :>< H I '~.~if.I hp-;.ff:.,~ --=--►, Cf.l ~ t< AiResearch 525328 Release trimming during , .... emergency operation Buffalo Instr. 25101-A30E- Indicate side slip angle. Co. 1-Bl Buffalo Instru. Provide side slip and angle Company of attack signal to Damper sv s tem and Indicator Day stron Ltd. 1889 Type 8 Indicate control surface response Humphrey FTOl-0101-1 Provide electrical command !2! 0 0 signal to AMP. CAL~ to I\) Stick force Garrett 47648 Amplifiers signal from•q•· transducer for'q"actuator - AVRO PART DESCRIPTION NO. GYR0-Yaw Rate (Normal) 7-1362-231 GYR0 - Yaw Rate (Emerg) 7-1362-231 GYR0 - Pitch Rate 7-1362-234 GYR0 - Roll Rate 7-1362-235 t:l ~ ~ Parallel Servo - Aileron 7-3262-167 :,,- :z: t:l Parallel Servo - Elevator 7-3262-165 :,,- C/l C/l 0 (") H Differential Servo - Left 7-3260-11 :,,- 1--'3 Aileron trj t:l ~ Cl Differential Servo - Right 7- 3260-11 Aileron ,H '"d i 1--'3 Differential Servo - Left 7-3260-11 Elevator t"-1 H C/l Differential Servo - Right 7-3260-11 '"d 1--'3 Elevator :t> ~ Dual Differential Servo- 7-3283-7 --..J 1\.) Rudder Damper Circuit Breaker 7-1282-259 Panel V - ----__;. MANUFACTU RER'S NAME PART NO. FUNCTION M.H. D-GG72A- l Supplies angular displacement rate si1mals to AMP-CAL M.H. D-GG72A-l Supplies angular displacement rate signals to AMP-CAL M.H. D-GG72A-3 Supplies angular displac ement rate signals to AMP-CAL M,H. D-GG72A-2 Supplies angular displacement ' rate sirmals to AMP-CAL M.H. MG-55M-2 Provides command signal to aileron control valve M.H. MG-55N~2 Provides command signal to elevator control valve M.H. MG-51D-l Provides damping motion signal to left aileron control valve. M.H. MG-51D-l Provides damping motion signal to right aileron control valve. M.H. MG-51D-l Provides damping motion signal to left elevator control valve. M.H. MG-51D-l Provides damping motion signal to right elevator control valve M.H. D-MG62B-l Provides damping and turn co-ordination signals to rudder control valve. Avro Aircraft Provides wiring and equipment protection. ~ '"d ~ t:l H ~ H :z: 0 C/l t p M --..J l:,J I t:J ~ -i3 H :z: c:, ;,> :z: t:J ::,. en ~ 0 H ;,> >--3 t%J t:J t%J C) c::: H 'cl ~ :z: >--3 t-1 H (/) >--3 ~ • • AVRO PART DESCRIPTION NO. Potentiometer - Left 7-1162-173 Elevator Differential Servo position Potentiometer - Right 7-1162-173 Elevator Differential Servo Position Potentiometer - Left Aile ror 7-1164-11 Position Potentiometer - Right 7-1164-11 Aileron Position Potentiometer - Rudder 7-1100-21 Position Potentiometer - Left 7-1100-21 Elevator Position Potentiometer - Right 7-1100-21 Elevator Position Differential Transformer - 7-1100-23 Left Aileron Position Differential Transformer - 7-ll00-23 Right Aileron Position .i~ MANUFACTURER 1 S NAME PART NO. FUNCTION Garrett Provides signals to Damper "G" Limiter Garrett Provides signals to Damper 11G11 Limiter Helipot Corp. RPl 7-0101-1 Provides L. Aileron position to indicator. Provides L. Aileron position for ground Tests Helipot Corp. RPl 7-0101-1 Provies R. aileron position to indicator. Provides R. aileron position for ground Test. ;,> 'cl 'cl ~ t:J H I>< H Helipot Corp. SX-1274 Provides rudder position to I (/) indicator. Provides rudder position for ground test M g ~ Helipot Corp. SX-1274 Provides L. Elevator position siimal to indicator --..J Helipot Corp. SX-1274 Provides R. elevator position si=al to indicator ~ ~ H . ~ GL Colli1ts LMI'-1241X Provides L Aileron position I-' signal to Damper during normal < and emergency damper operation 1--' I G.L. Collim_s - LMI'-1241X Provides R. Aileron position signal to Damper during normal :z: 0 !,.. and emergency damper operation ► 'u ~ § H z Q ► § ► en en 0 0 H ~ ~ § H i ~ ~ ~ - • DESCRI PTION Linear llqlf Actuator = Rudder Trim Actuator - Rudder Potentiometer = Elevator Parallel Servo Position AVRO PART! MANUFACTURER 1 S I NO. NAME I PART N0 0 7-1583-61 AiRe·search 2985- 2 7- 1583-63 AiResearch 39886 7- 1358-13 M.H. D- LG-16K-l FUNCTION Provides displacement signal proportional to ''qc• to rudder feel unit. Provides trimming to -.rudder Supplies stick position signal to 11G" Limiter ► 'u 'u ~ t:l H >< H en @ ~ --J -z ~ H z >-3 z 1,..... ? ~ \.1l ,.....

Arrow maintenance instructions: damper system Report No. 70/MAINT 13/1

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