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RWAvionics & Communications Collins

FCS-80( ) Flight Control System installation manual

Collins General Aviation Division

Collins Divisions Cedar Rapids, Iowa, 52498 (319) 395-1000 Cable COLLINRAD Cedar Rapids

Rockwell International

February 21, 1990

TO: HOLDERS OF COLLINS FCS-80( ) FLIGHT CONTROL SYSTEM INSTRUCTION BOOK (5230766515)

SIXTH EDITION HIGHLIGHTS The attached instruction book (installation manual) completely replaces the existing book. The book was revised to incorporate the following equipments 562C-8F/8G Yaw Damper Computer and SVL-80 Series Actuator. All descriptions, diagrams, and tables have been updated to reflect the new equipments. The installation section has new mating connector pin assignment diagrams and updated interconnect diagrams and outline and mounting diagrams. Changes to this manual have been implemented in a manner that preserves the information needed for servicing earlier model equipment. Revised areas of text are identified with black bars in the margin. Service bulletins (SBs) and service information letters (SILs) are not supplied with new and revised publications. A listing of all SBs/SILs issued to date is in the General Aviation Equipment Service Bulletin/Information Letter Index (523-0766944). Remove the SBs/SILs from your old book and add them to the bulletins section of your new book. Discard the remainder of the old instruction book.

PUBLICATIONS DEPARTMENT

1/2

@523-0766515-00611A)

6th Edition, 21 February 1990

RWAvionics & Communications Collins

FCS-80() Flight Control System installation manual This publication includes: Description FCS-80() APA-80() APC-80() APP-80() ASU-80() FGC-80() FGP-80()/81() NAC-80/LAC-80 SVO-80() 562C-8F/8G SVL-80 Installation FCS-80() Operation FCS-80()

Collins General Aviation Division Rockwell Collins, Inc. Cedar Rapids, Iowa 52498 Printed in the United States of America © 1990 Rockwell Collins, Inc.

(FCS-80_IM_FEB_21/90)

523-0766516 523-0766517 523-0766518 523-0766519 523-0770790 523-0766520 523-0766521 523-0766522 523-0766523 523-0775962 523-0776022 523-0766524 523-0766525

WARNING INFORMATION SUBJECT TO EXPORT CONTROL LAWS This document may contain information subject to the International Traffic in Arms Regulation (ITAR) or the Export Administration Regulation (EAR) of 1979 which may not be exported, released, or disclosed to foreign nationals inside or outside of the United States without first obtaining an export license. A violation of the ITAR or EAR may be subject to a penalty of up to 10 years imprisonment and a fine of up to $1,000,000 under 22 U.S.C.2778 of the Arms Export Control Act of 1976 or section 2410 of the Export Administration Act of 1979. Include this notice with any reproduced portion of this document.

CAUTION The material in this publication is subject to change. Before attempting any maintenance operation on the equipment covered in this publication, verify that you have complete and up-to-date publications by referring to the applicable Publications and Service Bulletin Indexes.

SOFTWARE COPYRIGHT NOTICE © 1990 Rockwell Collins, Inc. All Software resident in this equipment is protected by copyright.

We welcome your comments concerning this publication. Although every effort has been made to keep it free of errors, some may occur. When reporting a specific problem, please describe it briefly and include the publication part number, the paragraph or figure number, and the page number. Send your comments to:

Publications Department MS 106-124 Collins General Aviation Division Rockwell Collins, Inc. Cedar Rapids, Iowa 52498 or by Internet E-Mail to: [email protected]

z523-0766516-006118 6th Edition, 21 February 1990

ûú7l_ed_Yi#95 48

2 +50 >95 48

2 +50 >95 48

2 +50 >95 48

g Milliseconds

6 6 11 3

6 6 11 3

6 6 11 3

6 6 11 3

g Milliseconds

6 15 11 1

6 15 11 1

6 15 11 1

6 15 11 1

Cat J

Cat O or P

Cat S

Cat O or N

5 to 2000 0.020 3.0

5 to 55 0.010 (0.10 for P) 1.5

5 to 2000 0.036 10

5 to 55 0.010 (0.10 for N) 1.5

1.0

55 to 2000 0.25 1.0

1.0

55 to 2000 0.25 (1.0 for N) 1.0

Altitude Exposure Operation Overpressure Decompression

Feet

High temperature Exposure Operation Short term Continuous Operation

ºC

30 minutes

Low temperature Exposure Continuous operation

ºC

Temperature variation

ºC

Cycles 2. Humidity Cycles Temperature Level Time

SERVO ACTUATORS

ELECTRONICS

ºC Percent Hours

3. Shock Operational Positions Level Time Cycles Crash safety Positions Level Time Cycles 4. Vibration Frequency Level Acceleratoin, max

Hz Inches da g

Frequency Acceleration, constant Time per axis

Hz g Hours

Revised 21 February 1990

1-8

description 523-0766516 Table 1-5. Environmental Requirements of RTCA DO-138. TECHNICAL STANDARD ORDER-C9c FLIGHT CONTROL SYSTEM QUALIFICATION LEVELS TEST

UNITS CONTROLS

SERVO ACTUATORS

ELECTRONICS

SENSORS

5. Power input High variation Voltage Voltage Frequency

Ac Dc Hz

127 30.8 420

127 30.8 420

127 30.8 420

127 30.8 420

Low variation Voltage Voltage Frequency

Ac Dc Hz

103 25.2 380

103 25.2 380

103 25.2 380

103 25.2 380

Low voltage Voltage Voltage Frequency

Ac Dc Hz

100 22.4 400

100 22.4 400

100 22.4 400

100 22.4 400

NA

NA

Cat E

NA

6. Explosion Gasoline/air ratio Temperature Max test altitude 7. Icing Altitude Temperature Humidity Cycles

13:1 +71

ºC Feet

40,000 NA

NA

Feet ºC Percent

NA 40,000 –55 to +32 95 5

8. Dielectric strength Voltage, operating Time

Multiplied by Seconds

5 5

5 5

5 5

5 5

Conducted voltage transient test

Para 10

Para 10

Para 10

Para 10

Audio frequency conducted susceptibility

Para 11

Para 11

Para 11

Para 11

Audio frequency magnetic field susceptibility

Para 12

Para 12

Para 12

Para 12

Radio frequency susceptibility

Para 13

Para 13

Para 13

Para 13

9. Electromagnetic interference

Revised 21 February 1990

1-9

description 523-0766516 Table 1-6. Environmental Requirements of RTCA DO-160. TECHNICAL STANDARD ORDER-C9c FLIGHT CONTROL SYSTEM QUALIFICATION LEVELS TEST

UNITS CONTROLS

1. Temperature/altitude Altitude operation High temp operation Low temp operation High temp exposure Low temp exposure Decompression Overpressure

Feet ºC ºC ºC ºC Feet Feet

2. Humidity cycles Temperature Level Time

ºC Percent Hours

3. Vibration a. Frequency b. Level c. Acceleration, max

Hz Inches da g

SERVO ACTUATORS

ELECTRONICS

SENSORS

Cat A2

Cat D2

Cat D2

Cat D2

15,000 +70 –15 +85 –55 40,000 –15,000

50,000 +70 –55 +85 –55 NA NA

50,000 +70 –55 +85 –55 NA NA

50,000 +70 –55 +85 –55 NA NA

Cat A

Cat A

Cat A

Cat A

50 95 48

50 95 48

50 95 48

50 95 48

Cat K

Cat R

Cat O

Cat O

5 to 2000 0.010 0.25

a. 5 to 55 b. 0.010 c. 1.5

a. 5 to 57 b. 0.030 c. NA

a. 5 to 55 b. 0.010 c. 0.5

a. 55 to 2000 b. NA c. 0.5

a. 57 to 350 b. NA c. 5.0

a. 55 to 2000 b. NA c. 0.5

a. 350 to 2000 b. 0.0008 c. NA a. 500 to 2000 b. NA c. 10.0 Time per axis

Hours

4. Explosion Gasoline/air ratio Temperature Max test altitude

1.0

1.0

1.0

1.0

Cat X

Cat X

Cat E

Cat X

13:1 71 50,000

ºC Feet

5. Waterproofness

Cat X

Cat X

Cat X

Cat X

6. Hydraulic fluid

Cat X

Cat X

Cat X

Cat X

7. Sand and dust

Cat X

Cat X

Cat X

Cat X

Revised 21 February 1990

1-10

description 523-0766516 Table 1-6. Environmental Requirements of RTCA DO-160. TECHNICAL STANDARD ORDER-C9c FLIGHT CONTROL SYSTEM QUALIFICATION LEVELS TEST

UNITS CONTROLS

ELECTRONICS

SERVO ACTUATORS

SENSORS

8. Fungus

Cat X

Cat X

Cat X

Cat X

9. Salt spray

Cat X

Cat X

Cat X

Cat X

10. Magnetic effect

Cat A

Cat A

Cat A

Cat A

1º 0.3 to 1.0

1º 0.3 to 1.0

1º 0.3 to 1.0

1º 0.3 to 1.0

Cat B

Cat B

Cat B

Cat B

Allowable deflection Distance

Degrees Meters

11. Power input High variation Ac voltage Frequency Dc voltage

Volts Hz Volts

122 420 30.3

122 420 30.3

122 420 30.3

122 420 30.3

Low variation Ac voltage Frequency Dc

Volts Hz Volts

104 380 24.8

104 380 24.8

104 380 24.8

104 380 24.8

Emergency low Ac voltage Frequency Dc voltage

Volts Hz Volts

104 360/440 20

104 360/440 20

104 360/440 20

104 360/440 20

Cat B

Cat B

Cat B

Cat B

+78 –22 +48 NA 230 peak

+78 –22 +48 NA 230 peak

+78 –22 +48 NA 230 peak

+78 –22 +48 NA 230 peak

Cat B

Cat B

Cat B

Cat B

200 to 1000 0.56 1000 to 15,000 1.4

200 to 1000 0.56 1000 to 15,000 1.4

200 to 1000 0.56 1000 to 15,000 1.4

200 to 1000 0.56 1000 to 15,000 1.4

Cat A

Cat A

Cat A

Cat A

Induced signal susceptibility

Para 19.0

Para 19.0

Para 19.0

Para 19.0

Rf susceptibility

Para 20.0

Para 20.0

Para 20.0

Para 20.0

Spurious emissions

Para 21.0

Para 21.0

Para 21.0

Para 21.0

12. Voltage spikes Dc intermittent + Dc intermittent – Dc repetitive + Dc repetitive – Ac repetitive

Volts

V rms

13. Audio susceptibility A.1 Frequency B.1 Induced voltage A.2 Frequency B.2 Induced voltage

Hz V rms Hz V rms

14. Electromagnetic interference

Revised 21 February 1990

1-11

description 523-0766516

Table 1-7. Weight and Power Requirements. WEIGHT PER UNIT (APPROX) TYPE NO

DESCRIPTION

115 V AC, VA kg

APA-80/80G APA-80A APA-80C/80M APA-80F APC-80/80M APC-80A/F APP-80/80M APP-80A ASU-80 FGC-80A/B/M FGC-80F, G FGP-80/80M FGP-81/81A FGP-81M NAC-80/LAC-80 SVO-80( ) 334C-6B 334D-6A 351B-6( ) 351B-7( ) 329B-8Y 331A-9G 332D-11( ) 345A-7B 562C-8F 590A-3K 614E-22D

POWER

Autopilot amplifier Autopilot amplifier Autopilot amplifier Autopilot amplifier Autopilot computer Autopilot computer Autopilot panel Autopilot panel Avionics switching unit Flight guidance computer Flight guidance computer Flight guidance panel Flight guidance panel Flight guidance panel Accelerometer Primary servo Primary servo (yaw) Trim servo Primary servo mount Primary servo mount Attitude director indicator Horizontal situation indicator Vertical reference Rate-of-turn sensor Yaw damper computer Air data control Remote heading/course selector

2.45 2.05 2.45 2.05 2.40 2.40 0.68 0.68 0.68 2.72 2.77 0.45 0.23 0.27 0.32 2.59 2.59 1.63 1.13 1.22 3.63 3.4 3.3 1.0 1.36 2.13 0.86

lb 5.4 4.5 5.4 4.5 5.3 5.3 1.5 1.5 1.5 6.0 6.1 1.0 0.5 0.6 0.7 5.7 5.7 3.6 2.5 2.7 8.0 7.5 7.2 2.2 3.0 4.7 1.9

MAX 24 24 24 24 25 25 NA NA NA 29 29 NA NA NA NA

NA NA NA 40 30 63 46 14 45 7

NOM 24 24 24 24 25 25 NA NA NA 29 29 NA NA NA NA

28 V DC, W MAX

130 130 130 130 12 12 40 40 25 5 5 NA NA NA NA (From APA-80) (From 562C-8F) NA 28 NA NA NA NA 40 NA 30 NA 37 NA 20 NA 14 60 30 6 7 NA

NOM *42 *42 *42 *42 12 12 40 40 25 5 5 NA NA NA NA

18 NA NA NA NA NA NA 20 6 NA

W LIGHTS NA NA NA NA NA NA **5 **5 NA NA NA **5 **5 **5 NA NA NA NA NA NA 8 4 NA NA NA NA 6

*Servos at quiescent current **Does not include panel annunciation

1.5 AIRCRAFT CONTROL AND SYSTEM PERFORMANCES The autopilot system controls the aircraft over the whole flight profile including cruise, climb, descent, and category I and II approaches. The autopilot is capable of being engaged throughout the entire flight envelope and during normal aircraft attitudes. The autopilot engagement is independent of the yaw damper.

Revised 21 February 1990

The autopilot provides transient-free operation and mode switching. The system is synchronized to eliminate engage and disengage transients. The surface and control wheel activity is kept to the minimum necessary to provide good tracking and stability and has the frequency and amplitude of the control surface movement limited to a level acceptable to the aircraft crew and passengers. Performance data for the basic system is detailed in Table 1-8.

1-12

description 523-0766516 Table 1-8. APS-80 Performance Data. MODE Autopilot manual pitch command

Autopilot manual roll command

HDG

PARAMETER

VALUE

Command rate

Programmed

Pitch-command limit

17 or 25º

Pitch hold

±0.25º smooth air ±1º moderate turbulence with respect to the gyro reference

Roll command limit

32º

Roll rate limit

5º /second max

Roll-angle limit

27º ±3º

Roll-rate limit

5º /second

Roll acceleration limit

Provided by a 1-second low-pass circuit

Accuracy

±0.5º smooth air ±1º moderate turbulence with respect to heading datum

VOR

Greater than 30 nmi Beam-intercept angle limit

±90º

Roll-angle limit

27º ±3º

Roll-rate limit

5º/second

Roll acceleration limit

Provided by a 1-second low-pass circuit

VOR ON-COURSE SUBMODE

APPR LOC

Roll-angle limit

12.5º

Crosswind correction

±45º

Beam-hold accuracy

15 µA maximum averaged over a period of 10 minutes with no crosswind

LOC CAPTURE SUBMODE

Greater than 10 nmi

Beam-intercept angle

±90º

Beam overshoots

1 maximum of less than ±25 µA when not turn limited

Roll-angle limit

27º ±3º

Roll-rate limit

5º/second

LOC ON-COURSE SUBMODE

(Cont)

Roll-angle limit

12.5º

Crosswind correction limit

±45º

Localizer beam tracking

Cat H limits (±25 µA)

Revised 21 February 1990

1-13

description 523-0766516 Table 1-8. APS-80 Performance Data. MODE

PARAMETER

VALUE

GS SUBMODE

GA

GS beam intercept limit

±20 to 50 µA, below 2500-ft radio altitude

GS beam overshoot

±75 µA

Pitch command limit

20º ±2º

GS beam tracking

Cat II limits (±75 µA)

Fixed pitch-up command

0 to 15º pitch-up, adjustable

Wings level command ALT

VS

IAS

MACH

Altitude-hold engage range

–1000 to +55,000 ft

Altitude-engage first overshoot

Approximately 10% of climb or descent ft/min at less than 2000 ft/min

Altitude-hold accuracy

±50 ft maximum deviation at sea level or 0.2% of altitude, whichever is greater, nonmaneuvering

Maneuvering without aircraft configuration changes

±75 feet or ±0.3%, whichever is greater, of barometric altitude error measured at the input to the flight computer under normal aircraft operating conditions

Maneuvering with minor aircraft configuration changes

±125 ft or ±0.5% of altitude, whichever is greater

Change of altitude during maximum roll angle

±75 ft or ±0.5% of altitude, whichever is greater

Altitude-hold range limits

–500 to +500 ft from engaged altitude

Pitch-command limit

20º ±2º

VS-hold engage range

0 to ±8000 ft/min

VS-hold accuracy

±250 ft/min or 10%, whichever is greater, of the selected VS ±500 ft/min with large power change

Pitch-command limit

20º ±2º

IAS-hold engage range

100 to 500 knots

IAS-hold accuracy

±5 knots nonmaneuvering, ±10 knots in turn, during turbulence, or with reasonable thrust change

Pitch-command limit

20º ±2º

MACH-hold engage range

0.3 to 0.99 Mach

MACH-hold accuracy

0.01 Mach nonmaneuvering, ±0.02 Mach in turn, during turbulence, or with reasonable thrust change

Pitch-command limit

20º ±2º

Revised 21 February 1990

1-14

description 523-0766516 Table 1-8. APS-80 Performance Data. MODE ALT PRESELECT

1.6

PARAMETER

VALUE

Preselect capture range

0 to 50,000 ft

VS capture range

8000 ft/min with Collins ADS-80( ); 4000 ft/min with Collins 590A-3( )

Capture g's

0.3 g increment maximum

so that a vertical mode cannot be selected unless a lateral mode is selected except when the autopilot is engaged.

FUNCTIONAL DESCRIPTION OF MODES

The following paragraphs describe the functional operation of the system modes. Specific system configurations may result in operational differences. These differences will be explained as they occur. Refer to the pilot's guide for precise operational descriptions.

b. 1/2-Bank Limit Selection of 1/2 BANK causes the normal bank limit in HDG, VOR LOC, or a NAV mode to be reduced to one-half its usual value. This function is interlocked so that selection of APPR or capture of the localizer will clear it.

1.6.1 Computed Modes Selection of computed modes is from the FGP-80 (Figure 1-3) and from various switches on the aircraft control wheel. Annunciation is on the instrument panel. The FGP-80 mode buttons are push on-push off momentary pushbuttons. Pushing a button once causes that mode to latch and automatically clears any existing noncompatible modes. A lighted gullwing indicator is located above each mode button and will illuminate if all conditions for the mode are satisfied. The gull wing is a 2-section indicator. When a single FGP-80 is used for dual flight computer control, the left half of the indicator indicates the condition of the left system and the right half for the right system. A second push of the mode button will drop out the mode and the gull-wing indicator is turned off. Following is a description of each of the system modes. a. Basic Mode The basic mode is when no mode is selected in the lateral or vertical channel. The ADI command bars will be driven out of view. When the autopilot is engaged, this will be the autopilot manual mode allowing the pilot to manually control the autopilot using the turn knob or pitch wheel on the autopilot panel. The basic mode also occurs when initial power is applied or when system power is interrupted. Interlocks are provided

Revised 21 February 1990

Note An alternative to the manual selection of 1/2-bank limit is automatic switching using an altitude switch in the air data system. This switch is set to operate at high altitude. With this configuration, the 1/2 BANK button would not be included on the FGP. c.

Heading Mode Heading mode is selected by pushing the HDG button. The heading angle is preselected by positioning the heading marker on the HSI to the desired heading. To provide smooth and accurate control of the aircraft in heading mode under all flight conditions, nonlinear programming is provided. Standard gains are used for small heading changes. When a large heading error is selected, the gain is increased using true airspeed for programming the gain. This higher gain is used until a small heading error is again present, at which time the gain is reduced back to the standard gain.

d. VOR LOC Mode VOR- or LOC-only mode is selected by pushing the VOR LOC button. The frequency tuned on the navigation receiver determines whether VOR or

1-15

description 523-0766516

Figure 1-3. FGP-80( )/81( ) Flight Guidance Panel

Revised 21 February 1990

1-16

description 523-0766516 When a LOC frequency is tuned, the VOR LOC mode selects LOC operation without glideslope. The operation is identical to VOR operation except the computation gains are modified and the gains are programmed with radio altitude. The localizer inbound course is set with the HSI course arrow. The intercept angle is set using the heading marker. If radio altitude is not available due to not having a radio altimeter installed or due to a radio altimeter failure, the gains are step programmed with middle marker.

LOC computation will be used. Initial selection arms the mode until a capture is sensed. In arm, the system automatically selects preselected heading and actuates the HDG and ARM annunciation. The course intercept angle is selected by moving the heading marker on the HSI to the desired heading. The system flies heading until conditions for a capture are present. The capture is an adaptive type that compensates for course intercept angle and distance from the station. Also, capture is forced immediately if VOR LOC mode is selected with a small radio deviation value. During capture, the amount of bank angle is kept to the minimum required to roll out on the center line of the radio beam. After acquiring the beam, the system will track it with adequate crab angle to compensate for the crosswind, and annunciation will change to capture. The system has the capability of having the VOR deviation linearized with DME distance. During this operation, the deviation bar on the HSI presents linear rather than angular deviation. Logic is provided which causes the system to revert back to angular deviation when the DME is invalid or as a function of external logic such as DME hold or a pilot override if the DME is not collocated with the VOR station. The linearization function is provided as an interconnect option and may not be activated for a particular system. LIN DEV (linear deviation) is annunciated when provided in system interconnect wiring. When the aircraft approaches the VOR cone of confusion, the radio rate is monitored, and when it is above a prescribed level, the system will maintain selected VOR heading with crosswind correction applied. This is essentially the longterm heading reference present at the time the cone was entered. DR (dead reckoning) will be annunciated. The course may be changed to a new outbound course while in the cone. The system will command an equal heading change retaining the crosswind correction present at cone entry. The system will come out of dead reckoning approximately 2 minutes after cone entry or less than a minute after a VOR FROM indication is sensed and radio rate is below an acceptable value. The system will then resume tracking outbound on the VOR course. After capture, the course cut angle is limited to 45 degrees.

Revised 21 February 1990

Back localizer operation is automatically selected in VOR LOC when a localizer frequency is tuned on the navigation receiver and the aircraft heading is greater than 105 degrees from the inbound localizer course heading. To use, the inbound (front) course is set with the HSI course arrow and the intercept angle is controlled by the heading marker. Gains are programmed using radio altitude. The ADI expanded scale LOC deviation pointer and the HSI deviation bar will have correct display phasing. Vertical control of the aircraft is maintained using altitude hold, vertical speed hold, or pitch hold. A B/LOC annunciation is provided when a back localizer mode voltage is present in the system. e.

APPR Mode APPR mode is selected for any type of approach: localizer/glideslope, VOR, RNAV, etc. The navigation receiver frequency and external navigation switching determines which type of approach will be used. The lateral operation is essentially the same as in VOR LOC mode except the gains are slightly tighter to give increased precision during the approach. When APPR is selected, the system is automatically switched to selected HDG. In the same manner as VOR LOC, sensing of a capture condition will cause the system to turn onto the inbound course and to track it. Appropriate ARM and CAP annunciation will be displayed during this transition. Vertical guidance is computed from the glideslope signal when a localizer frequency is selected. Glideslope capture is independent from localizer capture. During a typical approach, glideslope will normally capture after the localizer. The capture can be from either above or below the beam. Prior to capture, the system may be flown in any of the other vertical modes. When capture is sensed, the vertical mode being used is automatically switched off. A bias signal of the

1-17

description 523-0766516 when the desired flight conditions exist. The hold reference value is established and the system will compute commands to maintain this reference value. Interlocks are provided to ensure that only one mode can be selected at a time and that a lateral mode has been selected. (A lateral mode selection is not required if the autopilot is engaged). Annunciation is available for each hold mode, although a particular installation may not have VS hold, IAS hold, or MACH hold annunciated. Also, ALT hold annunciation may be combined with ALT CAPT annunciation.

correct phase, pitch-down when below the beam and pitch-up when above the beam, is injected into the computation to start the aircraft down the glidepath. Glideslope gain is programmed with radio altitude. The gain is reduced as altitude decreases. If a fault in the radio altimeter occurs causing the data to be invalid, a constant glideslope gain is used until middle marker where it is reduced. Appropriate GS ARM and GS CAP annunciation will be displayed. For both lateral and vertical modes, the system can be armed when a flag is present but interlocks are provided to ensure that capture is inhibited when the respective flag is showing; that is, a lateral channel capture can occur only when a NAV valid is present and a glideslope capture when a GS valid is present. In B/LOC mode, glideslope capture is inhibited. f.

h. PITCH Hold Mode When no vertical mode is selected and a lateral mode is selected, the system is in PITCH hold mode. It will hold the pitch attitude that was present when PITCH hold was selected. (When the autopilot is engaged, this is the autopilot manual elevator mode and the pitch hold computation will synchronize to the autopilot commands). PITCH hold annunciation is provided. Vertical sync mode may be used to modify the PITCH hold reference.

ALT SEL Mode Altitude preselect mode causes the system to capture the altitude that is set on the altitude alert panel. Any other vertical mode may be selected prior to the system sensing the altitude capture point. Before reaching the capture point, ALT ARM will be annunciated. At capture, other vertical modes will be switched off and a command will be generated to level the aircraft onto the selected altitude. The capture point is adaptive and will move toward or away from the selected altitude as a function of the vertical speed present at capture. After capture is sensed, ALT annunciation is provided. If the altitude set knobs on the alert panel are moved during the capture phase, the vertical computation will drop back to pitch hold mode. When the altitude error is small, the system senses a track condition and starts tracking the preset altitude level. After reaching a track condition, the system will automatically provide commands to compensate for altitude errors resulting from airspeed changes, resetting of barometric setting on the altimeter, or resetting the altitude level on the altitude alert panel. Selection of another vertical mode after capture will clear the ALT SEL mode.

g. ALT, VS, IAS, and MACH Hold Modes Each of the air data hold modes operates identically. The appropriate mode button is depressed

Revised 21 February 1990

i.

Vertical Sync Mode Vertical sync mode is controlled by the sync button on the aircraft control wheel. As the name implies, the reference for the specific vertical mode selected (excluding APPR), ALT hold, VS hold, IAS hold, MACH hold, and PITCH hold, is synchronized to the existing flight conditions at the time the vertical sync button is released. Vertical sync mode does not clear any existing mode.

j.

GA Mode Go-around mode is selected by depressing the GA button. GA may be selected at anytime and will cause the autopilot to disengage. A fixed pitch-up, go-around command (the amount of pitch is a function of aircraft type) and a wingslevel command are computed. After becoming established in the go-around attitude, the autopilot may be reengaged. Since the heading marker is not active in GA or APPR mode, the missed approach heading may be preset after selection of either. GA is cleared by selecting any new mode. The lateral mode will continue to be wings level until a new lateral mode is selected. A GA annunciation is provided until GA mode is cleared.

1-18

description 523-0766516 command bars are driven by the autopilot except in APPR mode.

k. Special Modes Specific installations may be customized to utilize the blank mode buttons remaining on the controller. Generally, a spare lateral button would be used for advanced navigation systems such as INS, RNAV, vlf, etc. Operation would be similar to VOR mode. The spare vertical button operation would depend upon the type of navigation system used. Logic on this button allows it to be armed similar to ALT SEL mode and could, therefore, be used for capture like ALT SEL. 1.6.2 Autopilot Modes Control of the autopilot is through the APP-80( ) Autopilot Panel (Figure 1-4) and from various switches on the aircraft control wheel. The APP-80( ) provides engagement control, AP XFR and TURB switching, and the controls for manual pilot inputs. When engaged, the autopilot has three basic modes in each channel: guidance mode, manual mode, and autopilot sync mode. a. Lateral Guidance Mode Anytime a lateral mode is selected on the flight guidance panel, the autopilot aileron channel uses the computed lateral guidance command from the flight guidance system. In this mode, the ADI

b. Vertical Guidance Mode Anytime a vertical mode is selected on the flight guidance panel, the autopilot elevator channel uses the computed vertical guidance command from the flight guidance system. c.

Manual Aileron Mode When there is no lateral mode selected on the flight guidance panel, the pilot may use the TURN knob on the APP-80( ) to manually control the aileron channel. The TURN knob may be of a rate type (APP-80) or a position type (APP-80A). The rate TURN knob is identified by noting that the knob is spring loaded to a center detent position. The aircraft roll rate commanded is proportional to the displacement of the TURN knob. Movement of the knob out of detent, when a lateral mode is selected and the autopilot is engaged, will automatically clear any lateral mode selected except APPR. When roll attitude is below 5 degrees and knob is in detent, heading will be held. If roll is above 5 degrees with knob in detent, the existing roll attitude will be held. The autopilot may be engaged at any reasonable roll attitude. Below 5 degrees, the system will go to

Figure 1-4. App-80( ) Autopilot Panel

Revised 21 February 1990

1-19

description 523-0766516 computations are not affected by AP sync but since the vertical sync mode and AP sync mode use a common control, the sync button, the vertical guidance computations will simultaneously synchronize when the autopilot does. When the sync button is released, the autopilot will return to lateral and/or vertical guidance modes. In manual aileron mode and/or manual elevator mode, AP sync effectively operates like the TURN knob and pitch wheel except the pilot manually positions the aircraft, using the aircraft controls, to a new attitude reference.

heading hold; between 5 and 32 degrees, the system will hold existing roll angle at time of engagement; and above 32 degrees, the system will roll back 32 degrees. The position TURN knob is identified by noting that when the knob is released, it will stay in any position between the two knob stops. The bank angle commanded is proportional to knob displacement. The lateral mode clear and heading/roll hold logic are the same as the rate knob. The TURN knob will not be active when autopilot is initially engaged, when all lateral modes are cleared, when a lateral mode is unselected, after AP SYNC operation, or when AP XFR is operated. The TURN knob will become active when moved into and out of detent. d. Manual Elevator Mode When there is no vertical mode selected on the flight guidance panel, the pilot may use the pitch wheel on the APP-80( ) to manually control the elevator channel. The pitch wheel control is a rate type control, spring loaded to a center detent position. Movement of the wheel from detent provides a pitch rate command proportional to the amount of wheel displacement. The signal is programmed with true airspeed over the whole flight regime that effectively results in equal wheel displacements at two different flight conditions producing approximately the same g loading during the pitching maneuver. If a vertical mode is selected on the flight guidance panel, then movement of the pitch wheel will clear the mode, except in APPR. In the detent position, the autopilot will hold the pitch attitude present when the wheel went to the detent. The reference is changed by moving the wheel to get the desired pitch rate and releasing it back to the detent when the desired pitch attitude is attained. Pitch hold annunciation is available. The autopilot may be engaged in any reasonable pitch attitude. e.

Autopilot Sync Mode AP sync mode is controlled using the aircraft control wheel sync button. AP sync allows the pilot to manually move the aircraft to a new reference without disengaging the autopilot. Depressing the sync button causes both the aileron and elevator channels of the autopilot to synchronize. In flight guidance mode, lateral flight guidance

Revised 21 February 1990

f.

TURB Mode Depressing the TURB mode button on the APP80( ) softens the autopilot gains to reduce aircraft control surface movement in turbulent conditions. The TURB button is an alternate action button and a second push will turn the TURB mode off. An interlock disables the TURB mode when APPR mode is selected. An indicator above the button illuminates when in TURB mode.

g. AP XFR Autopilot transfer allows the pilot to select which flight guidance system is used to provide guidance commands to the autopilot. In the normal position, the guidance commands are from the pilot's FGS. Pressing the AP XFR button transfers the APS signal and monitor inputs to the copilot's system. An indicator above the AP XFR button illuminates when transferred to the copilot's system. AP XFR is an alternate action button and a second push will transfer the system back to the pilot. Annunciation is provided indicating a transferred condition. The AP XFR function is not installed in systems that do not have a copilot's flight guidance computer. In this type of system, the AP XFR button will be missing. 1.6.3 Autopilot/Yaw Damper Engaging The autopilot and yaw damper are engaged using the engage levers on the APP-80( ). The interlocks between the two levers are dependent upon specific aircraft requirements. Two basic configurations are provided: (1) a mechanical and electrical interlock that allows the yaw damper to be engaged independently from the autopilot, but requireing the yaw damper to be engaged anytime the autopilot is; and (2) the autopilot and yaw damper engagement are totally independent. (In the first configuration, a mechanical interlock simultaneously picks up the

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description 523-0766516 yaw damper engage lever anytime the autopilot lever is engaged.) Each time the autopilot is engaged, a self-test sequence is activated that automatically tests safety critical portions of the autopilot. The autopilot engage lever must be held in the up position for approximately 1 second to allow the test sequence to be completed. If the self-test sequence is completed successfully and all interlocks are correct, the system will engage. If a monitored fault exists in the system, the equipment is improperly configured or the aircraft is improperly configured, and the autopilot will not engage. The system will disengage automatically or may be manually disengaged by the pilot. A summary of nonengage and disengage requirements are shown in Table 1-9. Autopilot engage (Figure 1-18) and yaw damper engage (Figure 1-19) configurations vary with specific aircraft installations. Reference the appropriate aircraft SIL for correct configuration control. Refer to Figure 1-20 for a logic diagram of the autopilot and yaw damper engage requirements.

function monitors the difference between the pilot's and copilot's sensors; that is, compasses, vertical references, localizer receivers, and glideslope receivers. The limit deviation sensing function monitors the absolute values of the pilot's localizer deviation and glideslope deviation. Compass and attitude (pitch and roll) data are compared continuously and are independent of modes except for changing sensitivity levels. Localizer and glideslope comparators are active when the receivers are tuned to a localizer frequency. When either a comparator or a limit sensor exceeds the prescribed levels, an annunciator for that specific parameter will be illuminated. If the parameter is a comparator type, a MASTER WARN annunciator will also light. The MASTER WARN can be reset and will remain off until a different parameter exceeds the prescribed limits. The individual annunciators will automatically latch when the signal is outside the prescribed trip level and will remain latched until it is reset while inside the trip level limit.

1.6.4 Autopilot Trim

1.6.6 Annunciation

The autopilot automatically trims the elevator when it is engaged. When the elevator is out of trim by a prescribed amount, an out-of-trim annunciator will illuminate. Should a fault occur in the autopilot trim system, a dual monitor will detect it and indicate a trim fail condition to the pilot. In some systems the autopilot will be disengaged automatically rather than showing a trim fail annunciation. Should trim runaway occur, monitors provide trim power interrupt.

The annunciation used with the FCS-80 varies considerably depending upon the system configuration and the specific customer requirements. Table 1-10 is a listing of annunciation showing the recommended list, the comparator annunciation, and an optional list.

Provisions are made in the system to install trim indicators for each autopilot channel. These will provide a direct reading of the out-of-trim condition of each channel. In some systems, an annunciation may be provided to indicate long-term aileron mistrim.

1.7 DETAILED DESCRIPTION OF MODES The FCS-80 Flight Control System is composed of four general subsystems: FIS-85 Flight Instrument System, FGS-80 Flight Guidance System, APS-80 Autopilot System, YDS-80 Yaw Damper System, and associated sensors. Following is a simplified description of the FGS-80 and APS-80. A functional system diagram of the FCS-80( ) system is shown in Figure 1-21.

1.6.5 Comparator and Limit Monitoring Installations for use under category II weather conditions will use the comparator and limit monitoring system designed into the APS-80. The two types of functions provided are signal comparison and magnitude limit deviation sensing. The signal comparison

Revised 21 February 1990

1.7.1 FGS-80 Guidance Computation The following paragraphs describe in detail the specifics of the guidance computation. The discussion is function oriented rather than equipment oriented.

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description 523-0766516 Table 1-9. Nonengage and Disengagement Requirements With Options SYSTEM

Collins Standard

RESULT

MANUAL DISENGAGE

Both autopilot and yaw damper disengage. AP DISENG annunciator flashes. Reset annunciation with disconnect button, GA button, or trim switch.

Press emergency disconnect button. Move AP and YD disengage levers down.

AUTOMATIC DISENGAGE Loss of yaw damper valid

FAILURE TO ENGAGE

Yaw damper incorrectly configured Aircraft wiring incorrect Open disengage button

Only autopilot disengages and AP DISENG annunciator flashes. Reset with emergency disconnect button, GA button, or manual electric trim switch.

Move AP disengage lever down.

Select GA mode. Move manual electric trim switch.

Any of the automatic disengage criteria Aircraft wiring incorrectly configured

Fault in servo loop circuits Loss of gyro valid Loss of system power

APA-80 incorrectly configured APC-80 incorrectly configured

Fault in engage circuits

Failure to pass AP selftest sequence

Option 1

Same as above except AP DISENG annunciator is steady and not reset.

Same as standard

Same as standard

Same as standard

Option 1A

Same as option 1 except AP DISENG annunciator is reset with emergency disconnect button, GA button, or trim switch.

Same as standard

Same as standard

Same as standard

Option 2

Same as standard

Same as standard

Same as standard except manual electric trim switch is not active with AP engaged plus AP trim failure disengages autopilot.

Same as standard

Option 3

Same as standard except yaw damper and autopilot are mechanically and electrically independent.

Same as standard except AP lever controls only autopilot and YD lever controls only yaw damper. Emergency disconnect disengages both AP and YD.

Same as standard except yaw damper monitor is not interlocked into AP engage circuits.

Same as standard

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description 523-0766516 Table 1-10. FCS-80 System Annunciators ANNUCIATOR

CONDITION TO ACTIVATE

COLOR

RECOMMENDED HDG

Selection of HDG mode or navigation arm

Green

D/R

Computations in heading memory

Amber

LIN DEV

Linear deviation being displayed

White

V/L ARM

Selection of VOR LOC mode and prior to radio beam capture

Amber

V/L CAP

Selection of VOR LOC mode and after radio beam capture

Green

B/LOC

System in back localizer mode

White

ALT SEL

Selection of altitude preselect mode

Amber

ALT

Selection of altitude hold mode or after preelect capture

Green

GS ARM

Selection of APPR mode prior to glideslope beam capture

Amber

GS CAP

Selection of APPR mode after glideslope beam capture

Green

PITCH

Selection of pitch hold mode

Green

GA

Selection of go-around mode

Green

AP XFR

Autopilot commands from right flight guidance computer

Green

EL OUT OF TRIM

Elevator servo holding torque for a period of time

Amber

AP TRIM FAIL (NO. 1) AP TRIM FAIL (NO. 2)

Dual monitoring when elevator trim system fails or runs away when AP is engaged

Red

AP DISENG

Autopilot disengaged (flashing)

Red

MASTER WARN

Compass, roll, pitch, GS, or LOC difference has been exceeded. Latches. Has reset capability.

Red

LOC LIMIT

When the aircraft deviates from the localizer center line by a preset amount

Amber

GS LIMIT

When the aircraft deviates from the glidescope center line by a preset amount

Amber

ATT

When two pitch or two roll attitude signals disagree by a present amount

Amber

COMP

When two compass systems disagree by a preset amount

Amber

LOC

When two localizer deviation signals disagree by a preset amount (nominal 30 mV dc)

Amber

GS

When two glideslope deviation signals disagree by a preset amount (nominal 40 mV dc)

Amber

COMPARATOR (FOR CAT II SYSTEM)

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description 523-0766516 Table 1-10. FCS-80 System Annunciators ANNUCIATOR

CONDITION TO ACTIVATE

COLOR

OPTIONAL VOR LOC

Selection of VOR LOC mode

Green

NAV

Selection of NAV (or specific name of other NAV mode) mode

Green

NAV ARM

Selection of NAV mode (or specific navigation mode name ) and prior to lateral capture

Amber

NAV CAP

Selection of NAV mode (or specific navigation mode name) and after lateral capture

Green

APPR

Selection of APPR mode

Green

IAS

Selection of IAS hold mode

Green

VS

Selection of VS hold mode

Green

MACH

Selection of MACH hold mode

Green

TURB

Selection of autopilot turbulence mode

Green

AP BANK HOLD

Autopilot in bank hold when using turn knob

Green

AP HDG HOLD

Autopilot in heading hold when using turn knob

Green

AIL OUT OF TRIM

Aileron servo holding torque for a period of time

Amber

YD DISENG

Yaw damper disengaged

Amber

YD FAIL

Yaw damper failed

Red

ALT ARM

After altitude preselect mode and prior to capture altitude

Amber

ALT TRK

After altitude preselect mode and when vertical speed is less than 100 ft/min

Green

ALT CAPT

After altitude preselect mode and after capture altitude is achieved

White

Pilot interface to the system is through the flight guidance panel and the flight instrument controls. 1.7.1.1 Heading Mode With heading mode selected (Figure 1-5), a command is computed to capture and track the heading selected using the heading marker on the horizontal situation indicator. The selected heading error signal is the angular difference between the compass heading and the selected heading. The heading error is programmed with true airspeed to give proper gain throughout the flight regime. The programmed signal is mixed with an autopilot filter compensation

Revised 21 February 1990

signal and passed through a bank limit. The resulting bank command signal is used by the autopilot lateral channel as a command input and is also mixed with roll latitude data to make a composite lateral guidance signal for display on the attitude director indicator command bars. 1.7.1.2 VOR Mode Automatic capture and tracking of a selected VOR course with crosswind correction is provided by the FCS-80 (Figure 1-6). When VOR LOC mode is selected and the receiver is tuned to a VOR frequency, the guidance computation switches to the arm submode,

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description 523-0766516

Figure 1-5. HDG Mode, Block Diagram

Figure 1-6. VOR Mode, Block Diagram

which uses selected heading for guidance until capture conditions occur. Captures may be made up to 90 degrees from the selected inbound VOR course. A capture/track sensor monitors radio deviation and bank command levels. Depending upon the course intercept angle, distance from the station, and speed of the aircraft, a turn command will be computed that will command a specific minimum bank angle to turn onto the course.

Revised 21 February 1990

VOR deviation is programmed with DME distance to linearize the signal. This signal is monitored to sense the presence of the cone of confusion and then routed through a beam noise filter where it is mixed with a course datum error damping signal. A second course datum signal feed is mixed with the signal before and after a course cut limit circuit. This in effect cancels course datum and ensures that a maximum command will not exceed a 45degree course intercept angle.

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description 523-0766516 After the system switches to track submode, the course datum signal is washed out to remove any steady state course error that would cause the system to parallel the VOR course when a crosswind is present. The system will crab the aircraft into the wind by the correct amount to track the center of the VOR course.

(XTD) to sense capture. Crosswind correction is normally disabled in this mode since it is provided by the navigation system. At capture, the computation switches to the roll command signal that is routed through the bank limit and mixed with roll attitude data similar to heading mode.

The composite bank command signal is run through the bank limit and mixed with roll attitude data similar to heading mode. When a cone of confusion is sensed, the memorized course angle with crosswind correction is used to provide guidance commands until the aircraft has passed through the cone. The system then reverts back to VOR track mode and will continue to fly outbound on the VOR course. Outbound course changes may be made while over the cone without rearming the system.

1.7.1.4 Approach Mode (Lateral)

1.7.1.3 NAV Mode (INS, RNAV, Etc) Selection of NAV mode enables the computation to capture and track a selected navigation course from an external navigation system. A diagram with simplified switching and logic is shown in Figure 1-7. Captures are made using a preselected heading, track angle error (TAE), and crosstrack deviation

Lateral approach mode for localizer is diagrammed in Figure 1-8 The gains and filter time constants are selected to accommodate the specific approach geometry for the navigation system used and are changed as required to differentiate between VOR, LOC, BLOC, and RNAV approaches. LOC deviation is programmed with radio altitude to allow higher gains during all of the approach without jeopardizing stability. As with VOR, selected heading is used to determine the intercept angle. When capture is sensed, the aircraft is guided onto the LOC inbound course. The programmed deviation signal is filtered and mixed with damping signals derived from course datum and lagged roll attitude to increase tracking accuracy and response close in on the approach. As with VOR crosswind correction, course cut limiting and bank limiting functions are provided. The localizer computation also has a roll washout feed that

Figure 1-7. NAV Mode, Block Diagram

Revised 21 February 1990

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description 523-0766516

Figure 1-8. Approach Mode (Lateral), Block Diagram

eliminates long-term roll signals of up to 3 degrees of roll. This permits an aileron mistrim or gyro installation leveling error equivalent to 3 degrees of roll without any resultant steady-state errors. As much as 15 to 20 microamperes of beam error can occur if this feed were removed. Back localizer computation is similar to the front course except GS capture is defeated and radio data is reversed to give correct display and computation phasing. 1.7.1.5 Approach Mode (Vertical) When very tight control of the glideslope beam tracking is required, the common method of using glideslope deviation as a position signal and pitch attitude as a damping signal is not adequate. One of the difficulties in obtaining the desired performance is a result of the position gain increasing as the aircraft approaches the transmitter due to the glideslope beam angular geometry. When pitch attitude and ground speed change, this results in deviation from the glideslope beam. Generally a steady-state pitch attitude other than 0 degree is required to produce a vertical descent value

Revised 21 February 1990

to track the glidepath. This steady-state attitude is removed by a pitch washout circuit or by limited forward integration of the glideslope deviation signal. The time constants must be selected with care to maintain adequate stability and tracking accuracy. Use of normal acceleration to derive a damping signal and programming of glideslope deviation with radio altitude allows the use of much less pitch attitude derived damping and increased radio gains. This improves gust response and increases tracking accuracy and stability while making the system less sensitive to airspeed changes, pitch trim changes, aircraft configuration changes, and power changes. The glideslope is more critical that the localizer to these parameters. Figure 1-9 shows a block diagram of the glideslope computation. Glideslope deviation is programmed with radio altitude from 1000 to 100 feet. The gain is then reduced at a greater rate and is zero at 50 feet. If a flare computation is provided, it would provide commands from 50 feet to the runway. The programmed glideslope signal is low pass filtered to remove undesirable noise from the basic glideslope position data. A capture bias signal is inserted at this point to provide an initial command (either up or down

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description 523-0766516

Figure 1-9. Approach Mode (Vertical), Block Diagram.

depending on whether the capture is from above or below the beam) to start the aircraft tracking down the glidepath. The output of a forward integrator is also introduced at this point to remove long-term glideslope position errors due to any bias output from the damping signal circuits. The damping signal is derived from vertical acceleration and a limited amount of pitch. This signal is corrected for g errors in turns by mixing with processed roll data. Accelerometer installation leveling errors and biases are removed by a long-term washout. The resulting vertical acceleration signal represents the proper damping signal after it is programmed with true airspeed and mixed with derived vertical rate. The composite acceleration damping and glideslope signal passes through the pitch limit to form a pitch command signal prior to mixing with pitch attitude, producing a vertical guidance signal for display on the ADI command bars and for autopilot commands. The pitch signal goes through a washout to remove tracking errors when a pitch attitude is present. 1.7.1.6 Air Data Hold Modes A diagram of the air data hold modes is shown in Figure 1-10. The modes include altitude hold (ALT),

Revised 21 February 1990

vertical speed hold (VS), indicated airspeed hold (IAS) and Mach hold (MACH). The air data sensor synchronizes until one of these modes is selected. At that point, the sensor is locked at that reference level, and an error signal proportional to the deviation from this reference is generated. The noise is filtered from the signal and, in the case of VS and ALT, programmed with true airspeed. The specific air data mode error signal (only one vertical mode can be selected at a time) is mixed with an acceleration-derived damping signal identical to that used for glideslope. The pitch washout and pitch limit is also the same as glideslope. 1.7.1.7 Altitude Preselect Mode Altitude preselect mode is basically altitude hold mode with a preselect/capture capability as shown in Figure 1-11. The altitude preselect computation is performed by utilizing an altitude command signal derived from altitude error, which is the difference between barocorrected altitude from the altimeter and the pilot preset altitude from the preselected panel. The system will automatically capture and track the chosen altitude on the preselect panel from any vertical mode.

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description 523-0766516

Figure 1-10. Air Data Hold Modes, Block Diagram

Figure 1-11. ALT Preset Mode, Block Diagram

Revised 21 February 1990

1-29

description 523-0766516 The capture is performed by passing the preselect error signal through a filter to remove noise and then to a capture/track sensor. When the correct combination of error and vertical speed occurs, the system senses capture and injects a capture profile command into the channel. As the aircraft approaches the selected altitude, a track condition is sensed and the computation is switched to altitude track. The altitude track reference is then stored in a memory circuit that compares this stored reference and the actual altitude. The memory output is altitude track error and is then put through a noise filter, appropriately programmed, and mixed with the acceleration damping signal and the washed out pitch signal. There is a specific difference between basic altitude hold mode and the altitude preselect track condition. Once altitude hold is selected, the altitude reference does not change. But in altitude preselect track, the preselect error from the preselect panel is continually monitored, programmed with altitude, and applied to a slew rate deriver circuit. The resulting signal updates (moves) the altitude error memory reference slowly. Therefore, on a longterm basis, any changes in airspeed, baroset, etc, which causes the aircraft to move from the selected altitude, will cause a correction error to occur so that the altimeter reading and the preselected altitude setting will agree since the aircraft will be commanded to move to this new altitude reference. 1.7.2 APS-80, Autopilot Computation The autopilot provides automatic control of the aircraft through the aileron and elevator control surfaces. The specifics of this computation are discussed in the following paragraphs. 1.7.2.1 Automatic Control -- Aileron Channel The aileron channel has two basic modes of operation: a manual command mode and a guidance mode. The manual command mode allows the pilot to control the autopilot using the turn knob. This knob is a rate control; that is, it is spring loaded to a center detent position and the amount of roll rate commanded is proportional to the knob displacement from the center detent position. (A position-type turn knob is available as an option.) The operation with the turn knob is dependent upon the bank angle at which the knob is returned to the center detent. When the bank angle is greater than 5 degrees, the system will hold that angle. When the angle is less than 5 degrees, it will hold heading. The autopilot

Revised 21 February 1990

can be engaged in any reasonable attitude and will hold heading or bank according to the preceding discussion. In guidance mode, the autopilot will follow precisely the commands computed in the flight guidance computer. Figure 1-12 is a block diagram of the aileron channel. Switching and logic are provided internal to the system for accepting the manual command and guidance (bank command) inputs. With no guidance mode selected, the turn knob is active. When a mode is selected on the flight guidance panel, the bank command input is routed to the autopilot. The bank command input is from either the left or right guidance system depending upon the position of the AP XFR switch on the autopilot panel. The command signal is acceleration and rate limited to ensure that response to step and transient inputs is smooth to give passenger ride comfort. This smoothed signal is mixed with bank producing an aileron steering signal. Heading hold error is computed with heading data from the compass system. Power normalization is used to remove false rates due to power transients. A synchronizer follows the heading until heading hold mode logic stops the synchronizer providing the heading hold reference. The error between actual heading and the reference is then mixed with derived heading rate to provide damping to the heading hold error signal. The heading rate signal is also used to compensate for the lag in the signal smoothing filters by mixing it ahead of the flight guidance bank limit. The bank hold synchronizer is similar in operation to the heading hold circuit in that it synchronizes to the actual bank until bank hold logic locks the synchronizer to produce a command error between the reference bank angle and actual bank angle. The command signal resulting from mixing bank with either manual commands, guidance commands, heading hold error, or bank hold error is then mixed with the programmed derived bank rate damping signal. This composite command signal is gain adjusted and programmed prior to being applied to the servo amplifier inputs. An autopilot synchronizing submode is selected by depressing the aircraft control wheel SYNC button. When this occurs, the entire aileron channel synchronizes by sampling the differential voltage across the primary servo motor and applying the voltage as positive feedback to the input of the channel. This

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description 523-0766516

Figure 1-12. Aileron Channel, Block Diagram

Revised 21 February 1990

1-31

description 523-0766516 gives the pilot a supervisory override function, allowing the aircraft to be manually moved without disengaging the system. 1.7.2.2 Automatic Control -- Elevator Channel The elevator channel, like the aileron, has two basic modes of operation: a manual command mode and a guidance mode. In the manual command mode, the input is controlled by moving the pitch wheel on the autopilot panel. This control is also a rate control spring loaded to a center detent. The system holds the pitch attitude when the wheel is returned to the detent. In the guidance mode, the autopilot will follow the vertical guidance signal from the flight guidance computer. The manual command mode is operational when no vertical mode is selected on the flight guidance panel. Selection of a vertical mode results in the elevator channel accepting the guidance signal.

Figure 1-13 is a diagram of the elevator channel. As with the aileron channel, the guidance input from the left or right guidance computation is a function of the AP XFR switching. Switching transients are eliminated with special synchronization and switching circuits in the vertical guidance signal path. A mode switch selects either the vertical guidance or pitch wheel signal. The pitch hold circuit synchronizes to the pitch attitude until pitch hold logic is present. At that time the pitch reference is stored on the synchronizer and compared with pitch attitude to produce the pitch hold error signal. As with the aileron channel, elevator motor differential voltage is mixed into the channel in AP SYNC mode to synchronize the channel when the pilot manually moves the aircraft with the autopilot engaged. The command signal is summed with power normalized derived pitch rate for damping the elevator channel and an up-elevator command signal. The composite command signal is gain adjusted and

Figure 1-13. Elevator Channel, Block Diagram

Revised 21 February 1990

1-32

description 523-0766516 programmed before being applied to the elevator servo amplifier inputs. Up elevator is derived from roll data, which is absoluted (to always produce an up command), shaped, gain adjusted, and programmed.

annunciator to indicate to the pilot that the elevator has not been trimmed adequately to relieve the primary servo torque. A threshold and time delay ensure that the mistrim annunciation is present only when predetermined values are exceeded. The input to the threshold and gain circuits controls the trim drive.

1.7.2.3 Automatic Control -- Elevator Trim The elevator trim for the autopilot is composed of the trim drive circuits and independent dual trim monitors. The differential voltage across the elevator primary servo is proportional to the torque being applied to the elevator surface. The trim system automatically trims the elevator trim to reduce this long-term torque to zero. Motor voltage (Figure 1-14) is simultaneously applied to trim threshold and gain circuits and to out-of-trim monitor. The out-of-trim monitor drives a mistrim

The threshold and time delay keep the system from continually trimming at low torque levels. A reference generator is used to generate a triangular waveshape for the trim gain and duty cycle functions. The trim gain circuits take the elevator servo motor error voltage and the triangular wave and produce a pulse output proportional to the gain setting and the voltage value. The duty cycle control is an adjustment that controls the rate at which pulses are

Figure 1-14. Elevator Trim, Block Diagram

Revised 21 February 1990

1-33

description 523-0766516 applied to the trim circuit outputs. The trim interface circuits combine the delayed motor error voltage, the gain control signal, and the duty cycle control to produce interlocked trim up and trim down drive signals to control the trim power amplifiers that interface into the aircraft trim system. The interlocks are arranged so that an arm voltage and its corresponding trim pulse will occur together. Should a condition occur that would result in an arm voltage and the opposite trim pulse occurring simultaneously, the trim system will shut down, stopping the trim. The dual trim monitors are driven with elevator motor voltages that are electrically isolated from each other. The design is such that no single fault will cause both monitors to fail together. The trim generator output signal phase is compared with the phase of the motor voltage to determine that the system is trimming in the correct direction. A phase

difference indicates a trim runaway. A separate circuit determines loss of trim by monitoring that the trim is running anytime motor voltage is present. The logic outputs of these monitors are combined using OR logic with a maximum duty cycle monitor to produce dual trim failure annunciation logic signals. These are displayed to the pilot to warn him of any trim faults. Monitors provide automatic trim power interrupt should trim runaway occur. 1.7.2.4 Automatic Control -- Servo Amplifiers Identical servo amplifiers are used for both the aileron and elevator channels. These servo amplifiers are designed to provide fail passive operation. The fail passiveness of the servo is achieved by a dual amplifier design with appropriate monitoring as shown in Figure 1-15. During most system operations, commands to the servo do not require servo forces up to the maximum

Figure 1-15. Fail Passive Servo, Block Diagram

Revised 21 February 1990

1-34

description 523-0766516 torque value. The servo command for the aileron or elevator is connected to servo command A and servo command B inputs. Since the servo amplifier is a dual amplifier driving the top and bottom of a single servo motor, the command inputs to the two halves of the amplifier must be opposite phase to cause a differential drive voltage to occur across the motor. Therefore, servo command B is shown inverted. The signal then goes through output switches that are normally closed unless a predetermined g level in the pitch channel or a roll/roll rate level in the roll channel is sensed. At that time, the switches would open removing the commands from the amplifier input. The servo commands are mixed with servo position and routed through voters and torque limiters before being amplified to drive the servo motor. The motor is mechanically connected to the dc rate generator to supply feedback to the amplifier. Dual rate signals from the rate generator are independently integrated to obtain servo position feedback to be summed with the command signals to close the servo loops. Rate damping is used to shape the servo response characteristics. The amplifier circuit is designed to operate around a 14-V dc level. Under normal idling conditions, when no input is present, there is no motor current, and both sides of the motor are floating at the 14-V dc reference level. This is accomplished by applying a 14-volt common mode reference and using internal feedback to drive the output up to this reference. Since the input commands to the amplifier are of opposite polarities, a differential voltage is created across the motor with the circuit voltage gain equaling 20. It should be emphasized that when the amplifier and motor are operating normally, the voltage on each side of the motor will change by the same magnitude with a common input, one up and one down from the 14-V reference. This characteristic provides a point to monitor circuit operation by comparing the motor center voltage against an external reference. Torque limiting is accomplished by precise control of current into the motor. Two sets of torque limiters are provided, one set on each side of the motor. Either set will control the maximum current through the motor, and hence, will control maximum torque available from the motor. In essence, there are two types of torques being controlled by a torque limit set. One of these is a fast response low torque limit that is adequate for usual aircraft maneuvering. The

Revised 21 February 1990

second limit is a torque rate limit that allows additional torque to be applied at a slow rate. This allows large configuration changes to be controlled by the higher torque available with the rate limited torque without exceeding any certification requirements. The voters are used to reduce the effects of tracking errors and tolerances in the servo loops. The voters also provide instantaneous hardover protection. These are midvalue voters that choose the middle voltage when comparing the two command inputs and ground. The input signals to the voters are compared to detect large discrepancies at the inputs which would indicate that a fault condition is present. Should this occur, the autopilot would automatically disengage. The response limiters and the input comparator are critical to ensuring that the fail passive servo is protected from upstream faults. Therefore, these functions are tested each time the autopilot is engaged. The system will engage only if the test is passed. 1.7.3 Stability Augmentation System (SAS) A self-contained yaw damper is used to provide the yaw stability augmentation function and turn coordination. The servo amplifier/servo motor is similar in operation to that used in the aileron and elevator channels. It is a dual amplifier driving a dc motor with integrated rate being used for position feedback. Since a usual requirement for an SAS system is that the yaw channel must be able to be trim without disengaging, a torque washout circuit relieves the torque at the servo motor in a short period of time. A block diagram of the yaw damper is shown in Figure 1-16. Yaw rate is processed to provide appropriate filtering for the yaw damping function. This yaw damping signal is summed with an appropriately programmed and filtered roll signal for turn coordination. The composite damping and coordination signal is adjusted with the channel gain to match the requirements of a specific aircraft and is programmed with indicated airspeed to provide optimum control over the whole flight regime. The yaw damper is available to drive either a standard rotary servo typically used in a parallel damping system or a linear actuator used in series damping systems.

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description 523-0766516

Figure 1-16. Yaw Damper, Block Diagram

1.8 SYSTEM SAFETY CONSIDERATIONS The FCS-80 system provides additional capability in features, increased control authority, and increased performance, while maintaining equal or improved safety levels over previous flight control systems. The system has been thoroughly analyzed and tested to ensure that the probability of an undetected failure resulting in a malfunction is acceptable. 1.8.1 Servo Failure Protection The servo system is fail passive to eliminate the primary servo loop as a source of malfunctions due to internal failures in the pitch and roll axes of the flight control system. Adequate internal adjustments are provided to allow torque and torque rate levels to be set to limit the effective control surface deflection. The servo loops are monitored to disengage the autopilot when a safety critical failure occurs.

Revised 21 February 1990

There are basically three levels of safety in the autopilot servo motor/mount for each control axis to protect against jam failures. The first level of safety is the capability to back drive the servo motor with manual control column force. If the servo is engaged, the force required to back drive the motors is a back torque, which just exceeds the torque limiter input to the motor. If not engaged, the servo motor can be back driven with much less manual control force. Should a motor or gearing jam occur, which has been shown to be extremely remote, manual override is accomplished by manual control input to overcome the slip clutch. This level of safety is extremely reliable. The slip clutch is designed to prevent foreign objects or matter from reaching the slip clutch area. The slip clutch is the last device coupling the servo to the capstan and thence to the primary control cables. A third override capability is provided by the ability to overcome the engage clutch at a level above the slip clutch.

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description 523-0766516 1.8.2 Command Malfunction Control Hardovers can occur due to faults in sensors and computation upstream from the servo loop. Command input malfunctions are controlled by voters, performance sensing, and cutout switching. In the elevator channel, acceleration is sensed and the command inputs are opened anytime the g level exceeds preset values. A similar condition occurs in the aileron channel except a combination of roll and roll rate is used to open the aileron command inputs. When these inputs are opened, the control surface is driven toward a streamlined state. This action provides the pilot adequate time to respond while none of the certification limits are exceeded. 1.8.3 Servoed Command and Attitude Displays Servoed displays provide increased torque to withstand vibration and reduce sticking. Servos also allow position monitoring so incorrect displays can be flagged. 1.8.4 Status and Warning Annunciation Interlocks, monitors, and warnings give an immediate indication to the crew when a major problem exists. This may be in the form of flags on the instruments, disengagement of the system, or a warning indication. Interlocks and monitors on the modes ensure that incompatible or unreliable mode conditions do not occur. System annunciation both for system mode condition and warning is positive and integrated into the system functions. The positive annunciation is such that all modes, monitors, and interlocks are correct before the annunciation is activated. The annunciation outputs are capable of controlling dual lamps on both sides of the cockpit. 1.8.5 Self-Test In the APS-80, critical safety functions are tested each time the autopilot is engaged. The functions tested include the aircraft performance limit circuits, monitors, accelerometers, and engage/ disengage circuits. The system must pass the test or it will not engage. 1.8.6 Comparator Warning The comparator warning circuit (Figure 1-17) compares the operation of the ILS receivers, the vertical gyros/instruments, and the compass systems.

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The warning outputs, when displayed, greatly reduce the heavy workload of the pilot during an instrument approach by performing full time comparison of the dual flight director systems. If a significant difference exists between the subsystems being compared, a warning indicator lamp lights to indicate an out-of-tolerance condition. Individual warning lamps light to indicate which system is unreliable. Master warning lamps light coincidentally with any subsystem warning except LOC and GS LIMIT. The master warning may be reset independently of the subsystem warning lamps to allow the master warning to operate when a second subsystem comparison is out of tolerance. For comparison of heading, bank, and pitch directly at the output displays, the comparator system employs bootstrap differential resolver (3¸ to 2¸) synchros on each of the instrument mechanism shafts carrying the displayed information to be compared. By originating comparison signals at these points, the entire systems, including display mechanisms, are compared. A differential resolver in the pilot's instrument is energized as a transmitter, producing a 3-wire synchro signal describing the position of the first display. The 3-wire signal then is transmitted to the similarly located differential resolver in the copilot's instrument. The second differential resolver is connected as a control transformer, and its null winding output is a direct measure of the angular discrepancy between instrument displays. In addition, the orthogonal (maximum) output winding of the second unit produces a known constant (for small errors) voltage, which provides a monitor for the presence of energizing power and for assuring circuit integrity, and power normalization to make the trip points independent of line voltage variations. The following are basic advantages of the differential resolver comparison system: a. Only one voltage source is required, and variations in it produce only second-order effects on comparison output. b. The comparison is performed directly in terms of the angular accuracy of a synchro system, which is a specified, controlled, and measurable characteristic. c. The maximum voltage output is available to distinguish between true null output and power or circuit failure (a necessity in nulling comparison).

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description 523-0766516 d. A minimum of added circuits and components is required, and an output directly related to discrepancy is obtained. e. The output has a known constant relationship to the compared quantities and a constant sensitivity for all shaft positions. As an example, the compass channel monitors the sine output of a differential resolver. The sine output is a signal proportional to the angular difference between the heading displayed on two course indicators. When this difference signal becomes greater than a preset limit, a dc warning voltage is provided to light an external warning lamp, indicating a compass out-of-tolerance condition. The cosine output is monitored to ensure that power is applied to the resolver and to take out line voltage variations. Bank and pitch information has a similar comparison. The threshold for the compass comparator is modified by attitude information to allow a larger disagreement if the aircraft is not in straight and level flight. The glideslope and localizer channels utilize the comparison technique to assess subsystem information quality. The two outputs of a dual navigation system are accepted as inputs to the comparison circuit. The difference between these two signals is amplified

Revised 21 February 1990

and applied to a level sensor. When this amplified difference signal becomes larger than a preset limit, a dc warning signal is provided to light an external warning lamp, indicating an out-oftolerance error in the information. The integrity of the comparator circuit is obtained by the use of amplifier biasing, ac tracer-type level detectors, and a self-test function. The level sensors used for comparing the heading roll, pitch, localizer, and glideslope signals are actually the fail-safe tracer-type comparator like the ones used to monitor the servo-amplifier operation. Failures in this type of comparator are selfmonitoring. In addition, the input amplifiers and signal conditioning chains are biased-up to provide a common mode signal to the level sensors and thus failures in these circuits are also self-monitoring. Therefore, the only failures that would not be selfmonitored would be the output driver stages and the annunciator lamp itself. This circuitry and lamps will be tested when a TEST button is pressed. Additional monitoring of ILS approach performance is provided by glideslope and localizer deviation limit sensors, which drive warning annunciators whenever specified deviations are exceeded.

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description 523-0766516

Figure 1-17. Computer Warning, Block Diagram

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description 523-0766516

Figure 1-18. FCS-80( ) Autopilot Engage and Configuration Control, Interconnect Diagram

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Figure 1-19. FCS-80( ) Yaw Damper Engage and Configuration Control, Interconnect Diagram

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Figure 1-20. FCS-80( ) Autopilot and Yaw Damper Engage Requirements, Logic Diagram

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Figure 1-21. FCS-80( ) Flight Control System, Functional System Diagram

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z523-0766517-006118 6th Edition, 21 February 1990

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