NASA B737 Flight Test Results of the Total Energy Control System

NASA CR-178285

NASA B737 Flight Test Results of the Total Energy Control System

(MASA-CR-7782€5) N A S A B737 B E S U L I S OF TBE T O T A L SICE86I (Eoeing Commercial Airplane a v a i l : & T I S HC A O 6 f H F A01

Final Report

P f I G f t T TESZ CCBTEit2L SYSZEM CC-) 103 p

CSCL O l C

Kevin R. Bruce

Seattle, Washington @e@@+- - AiF@aReCMaQaRjr

Prepared for NASA TSRV

January 1987 CONTRACT NAS1-17509

63/0 8

#87-26918

t

ffnclas 0090708

https://ntrs.nasa.gov/search.jsp?R=19870017485 2018-02-14T18:19:26+00:00Z

1. Report No. NASA CR-178285

2. Government Accession No. 3. Recipient's Catalog No.

Boeing Commercial Airplane Company PO. Box 3707 Seattle, WA 98124

4. Title and Subtitle NASA B737 Flight Test Results of the Total Energy Control System

NAS1-17905

5. Report Date

6. Performing Organization Code

July 1987

7. Author@)

9. Performing Organization Name and Address

Kevin R. Bruce ,/'

I

15. Supplemmtary Note8 NASA Technical Monitor, J. R. Kelly Boeing Technical Supervision, G. R. Flynn and Dr. D. Cannon Boeing Contract Manager, W. C. Botts

8. Performing Organization Report No.

10. Work Unit No.

18. Abstract

12. Sponsoring Agency Name and A d d m National Aeronautics and Space Administration Langley Research Center Hampton, VA 23665

The Total Energy Control System (TECS) is an integrated autopiloffautothrottle developed by BCAC that was test flown on NASA Langley's Transport System Research Vehicle (i.e., a highly modified Boeing B737). This system was developed using principles of total energy in which the total kinetic and potential energy of the airplane was controlled by the throttles, and the energy distribution controlled by the elevator.

TECS integrates all the control functions of a conventional pitch autopilot and autothrottle into a single generalized control concept. This integration provides decoupled flightpath and maneuver control, as well as a coordinated throttle response for all maneuvers. A mode hierarchy was established to preclude exceeding airplane safety and performance limits.

13. Type of Report and Period Covered

14. Sponsoring Agency Code

The flight test of TECS took place as a series of five flights over a 33-week period during September 1985 at NASA Langley. Most of the original flight test plan was completed within the first three flights with the system not exhibiting any instabilities or design problems that required any gain adjustment during flight.

19. Security Classif. (of this report) 20. Security Classif. (of this page)

Unclassified Unclassified

17. Key Words (Suggested by Author@)) Flight Control System Integrated AutopiloffAutothrottle Total Energy Control System

21. No. of Pages 22. Price

18. Distribution Statement Unclassified-Unlimited

I I I FORM GCQ5 For sale by the National Technical Information Service. Springfield, Virginia 22161

TABLE OF CONTENTS Page

1.0 Summary .............................................................. 1

2.0 Introduction ............................................................ 2

3.0 Symbols and Abbreviations ............................................... 3

4.0 Basic Concepts of TECS ................................................... 5

5.0 Architectural Features of the System ....................................... 9

6.0 Flight Test Results ....................................................... 11

6.1 Flightpath Angle Mode ............................................... 11 6.2 Altitude Select and Hold Mode ........................................ 12 6.3 CASMach Mode .................................................... 12 6.4 Velocity Vector Control Wheel Steering Mode ............................ 13 6.5 Underspeed Protection ............................................... 14 6.6 Energy Exchange Maneuvers .......................................... 14 6.7 Approach and &Around ............................................. 15

7.0 Pilot Reaction to Flying the TSRV With TECS ............................... 17

8.0 Conclusions ............................................................. 18

9.0 References .............................................................. 19

10.0 Appendixes ............................................................. 95

1.0 SUMMARY

The Total Energy Control System (TECS) is an integrated autopilot/autothrottle developed by BCAC that was test flown on NASA Langley’s Transport System Research Vehicle (Le., a highly modified Boeing B737). This system was developed using principles of total energy in which the total kinetic and potential energy of the airplane was controlled by the throttles, and the energy distribution controlled by the elevator. TECS integrates all the control functions of a conventional pitch autopilot and autothrottle into a single generalized control concept. This integration provides decoupled flightpath and maneuver control, as well as a coordinated throttle response for all maneuvers. A mode hierarchy was established to preclude exceeding airplane safety and performance limits. The flight test of TECS took place as a series of five flights over a 33-week period during September 1985 at NASA Langley. Most of the original flight test plan was completed within the first three flights with the system not exhibiting any instabilities or design problems that required any gain adjustment during flight.

1

2.0 INTRODUCTION

In 1979 NASA funded Boeing to begin the conceptual development of a fully integrated automatic flightpath and speed control system. The work was carried out under NASA contracts NAS1-14880 (1979-1980) and NAS1-16300 (1980-1981). Detailed design and simulator implementation was carried out under Boeing IR&D funding from 1979-1982. The outcome of this work was the Total Energy Control System (TECS). Following successful detailed simulator development of TECS at Boeing, NASA awarded a contract (NAS1-17509) in 1983 for the flight test of TECS on NASA Langley's Transport Systems Research Vehicle (TSRV), a highly modified Boeing

Flight test of TECS took place in September 1985 at NASA Langley in a series of five flights over a 3-week period. Most of the original flight test plan was completed in the first three flights. The final two flights were demonstration flights of TECS.

B737-100.

This document discusses:

The basic concept of TECS The architectural features of the system used for the flight test program The detailed flight test results

2

3.0 SYMBOLS AND ABBREVIATIONS

3.1 SYMBOLS

............... D Drag E ED, . . . . . . . . . . . . . Energy rate distribution error Es .............. Specific total energy rate EsE ............. Total energy rate error g ................ Acceleration due to gravity h ................ Altitude h . . . . . . . . . . . . . . . Altitude rate KEI, KEP, Km, K p Gain constants K G E ~ ........... Switched gain s . . . . . . . . . . . . . . . . Laplace operator TREQ . . . . . . . . . . . . Thrust required V ............... Airspeed VSTUL .......... Stall speed 9 ............... Filtered airspeed V ............... Rate of change of airspeed V C ~ .......... Maximum acceleration command V C ~ ........... Minimum acceleration command VE .............. Rate of change of airspeed error V- ........... Maximum acceleration

y . . . . . . . . . . . . . . . . . Flightpath angle yc ............... Flightpath angle command YE ............... Flightpath angle error 6Ec .............. Change in elevator command

.............. Change in throttle command 7~ ............... Time constant of energy rate loop

............... Total energy of system

w ............... Airplane weight

. TD . . . . . . . . . . . . . . . Time constant of energy rate distribution loop

3.2 ABBREVIATIONS

AFD . . . . . . . . . . . . . Aft flight deck AFDTH .......... Aft (flight deck) throttle (deg) ALPHA .......... Angle of attack (deg) ALT HOLD ....... Altitude hold mode ALT SEL ......... Select altitude mode ALTCOR . . . . . . . . . Altitude (barometric) (R) ALTSUM . . . . . . . . . Altitude command (R) ALW ............ Angle of attack (wing) (deg)

3

AOA ............. Angle of attack AZ .............. Vertical acceleration (fils2) CAS ............. Calibrated airspeed (kn) CMP . . . . . . . . . . . . Control mode panel DAS ............. Data acquisition system DATAC .......... Digital Autonomous. Terminal

Access Communication System

DME ............ Distance measuring equipment EADI ............ Electronic attitude display indicator

EPRA ........... Engine pressure ratio (average of engines)

DECMD .......... Elevator command (deg)

EPR . . . . . . . . . . . . . Engine pressure ratio

FLAP ............ Flap handle position (deg) FMS ............. Flight management system FPA ............. Flightpath angle FPASUM ......... Flightpath angle command (degrees) FPA SEL . . . . . . . . . Flightpath angle select mode GA .............. Go-around mode GAMEPS ......... Flightpath error (rad) GAMMAI ........ Flightpath angle (rad) GS .............. Glideslope capture and tracking mode HDOT ........... Height rate (Ws) HOR PATH LASSUM KGEPS KVEPS . . . . . . . . . . Switched gain (nominally unity) for speed to elevator path LRU . . . . . . . . . . . . . Line replaceable unit NCDU ........... Navigation control and display unit ND .............. Navigation display PFD ............. Primary flight display PITCH ........... Pitch angle (deg)

....... Horizontal path mode ......... Calibrated airspeed command (kn)

.......... Switched gain (nominally unity) for flightpath to elevator path

PMC . . . . . . . . . . . . Panel mounted controller Q . . . . . . . . . . . . . . . Pitch rate (deg) RFD . . . . . . . . . . . . . Research flight deck STAB . . . . . . . . . . . . Stabilizer position (deg) TECS . . . . . . . . . . . . Total Energy Control System TKA SEL . . . . . . . . Track angle select mode TSRV ............ Transport System Research Vehicle VEL-CWS . . . . . . . . Velocity vector control wheel steering mode VTER . . . . . . . . . . . True airspeed error (Ws) WEIGHT . . . . . . . . . Weight of airplane (lb)

4

4.0 BASIC CONCEPTS OF TECS

The basic concepts of TECS are discussed in References 1 through 3. However, a review of the design philosophy and theoretical concept is presented in this section. The work of developing an integrated autopilot/autothrottle was originally initiated to solve the problems identified with conventional uncoupled autopilots and autothrottles:

1. Cross-coupling errors in speed and altitude occur when maneuvering due to the design of autopilots and autothrottles as single inputlsingle output control systems. For example, a speed change can not be accomplished by only a change in throttle setting, but must be accompanied by an elevator retrim if altitude is to be maintained. Conversely, an FPA change cannot be achieved by an elevator deflection, but must be coordinated with a change in throttle setting. Autopilot, autothrottle, and flight management systems (FMS) control laws have developed over a long period of time that has led to duplication of function in the autopilot and FMS computer.

2.

These problems led to a general design philosophy for TECS: 1.

2.

Design the system as a multi-input/multi-output system.

Design with a generalized inner loop structure and design the outer loop functions to interface with the common inner loop, thus minimizing software duplication.

Provide underspeed and overspeed protection for all modes. 3. This philosophy integrates the conventional pitch and speed control functions into a single control system, which facilitates the replacement of the autopilot and autothrottle found in current airplanes by a single autoflight line replaceable unit (LRU).

The basic concept of TECS is to control the total energy of the airplane. The total energy of the system can be expressed as the sum of the potential and kinetic energy.

Where: W = airplane weight h = altitude g = acceleration due to gravity V = airspeed

The specific energy rate is given by:

V i . E s = h + - 8

5

Normalizing by velocity, then:

Es h v v - - - - + - = - y + - v v g g

Where y = flightpath angle @PA>

(3)

From the equations of motion along the flightpath, the thrust required to maneuver is:

Treq = W ( 7 + f ) + D (4)

Where D = airplane drag

Assuming that drag variation is slow, equations (3) and (4) show that the engine thrust required to maneuver is proportional to the specific energy rate of the system. Alter- nately, it can be stated that the throttles control the rate at which energy can be added to or deleted from the system.

In response to speed or flightpath changes then, a control law can be developed that uses the throttles to drive the total energy rate error to zero.

Where

E S E V E - = Y E + 7 V

This control law uses proportional plus integral control to reduce the total energy error to zero with a first order time constant, 7~ = K p / Km. However, achieving a speed maneuver without flightpath perturbation, or vice versa, requires coordinated elevator and thrust inputs. An energy rate distribution error EDE can still exist (e.g., too high an FPA and too low an acceleration). Correction of the energy rate distribution error can be accomplished by feeding back the difference of the acceleration error term v E / g and the FPA error YE.

Using proportional plus integral control, the elevator control is:

Where K,,, KEI = elevator proportional and integral gains respectively.

This control law calls for the use of the elevator to redistribute the energy rate error ED^ equally between FPA and acceleration. The response has a first order time constant y~ = KEP/KEI. This concept is shown in Figure 1.

6

To ensure coordinated response to both speed and FPA changes, both the total energy rate error and the energy rate distribution error must go to zero simultaneously. This requires that ideally the dynamic response of equations (5) and (7) should be identical (i.e., q = TD and K n = KEI).

In addition, the thrust and pitch response are matched by designing engine and pitch inner loops to minimize variations due to the engine or aerodynamics.

The engine loop is designed to produce the net thrust at all flight conditions. This is achieved by converting the net thrust command produced by the control law to a command relating to the engine pressure ratio (EPR), and then closing a feedback loop around the engine using this variable. Software limiting is provided to ensure that neither throttle or EPR limits are exceeded.

The short period pitch dynamics is stabilized in a conventional manner by the feedback of pitch and pitch rate. The gains were selected to match the thrust dynamics. The variable elevator effectiveness is compensated for by gain scheduling the elevator command as a function of dynamic pressure.

An important aspect of the overall TECS design is the utilization of a common inner loop for each mode of the autopilot (Figure 2.). By generating a common flightpath error signal (YE) irrespective of which autopilot mode is engaged, software duplication is minimized and system response is consistent. 'Ib implement the FPA mode, for example, the YE signal is generated by differencing an FPA command (yc) from the control mode panel with a y signal computed from height rate (%) and filtered velocity @).

For control of altitude and glideslope, outer loop control is added to generate the normalized yc signals and provide exponential capture of the desired altitude or glideslope. All inputs, except for the velocity vector control wheel steering mode (VEL-CWS) and go-around mode (GA), are rate limited to provided 0.lg maximum normal acceleration during maneuvers.

The CAS and Mach speed modes both use outloop control, similar to the altitude and glideslope modes, to provide exponential capture of the parameter. An acceleration command (VC) drives the inner loops.

The control modes are divided into speed or path priority so that, when throttle limiting occurs, either speed or flightpath has priority and control of that parameter is maintained. The other parameter goes open loop and is not controlled. The mode logic of TECS is set up so that FPA, altitude, and GA mode have speed priority, while glideslope and VEL-CWS mode have path priority.

Control of the system, while throttle limiting, is an important aspect of TECS that required significant development on a nonlinear simulator to achieve an implementation with minimum software that provides accurate control and is consistent with the speedlpath priority for each mode. This implementation is discussed in more detail in a later section that covers the flight test results.

The result of the TECS design is a fully integrated system that has predictable consistent performance in all modes. Duplication of software has been minimized by maintaining a common inner loop structure.

7

5.0 ARCHITECTURAL FEATURES OF THE SYSTEM

n e TECS flight tests were condxcted on NASA Langley's Transport Systems Re- search Vehicle (TSRV) that is a highly modified Boeing 737-100 aircraft (Figure 3) designed to investigate advanced navigation, guidance, control, and display concepts applicable to the emerging National Airspace System. In this aircraft, the entire ex- perimental flight management system, i.e., all navigation, guidance, and control functions, and primary pilot displays are under software control and can be reprogrammed to suit the requirements of a particular experiment.

The "XCS flight tests employed the full-up test configuration wherein the aircraft is flown from a Research Flight Deck (RFD) mounted in the cabin of the aircraft and referred to in this report as the Aft Flight Deck (AFD). The AFD features programmable electronic Primary Flight Displays (PFDs), Navigation Displays (NDs), a Navi- gation Control and Display Unit (NCDU), a glare-shield mounted Control Mode Panel (CMP), and Panel-Mounted Controllers (PMSs) that take the place of conventional column and wheel controls. The CMP (Figure 4) was modified for these tests by replac- ing selected baseline switch legends with ones corresponding to the TECS modes.

Figure 5 is a simplified block diagram showing the arrangement of the principal components of the research system. The system is built around two Norden digital flight computers that are militarized versions of the general purpose PDP 11/70 computer. Both computers are interfaced to the AFD and an extensive array of sensors by a global data bus known as the Digital Autonomous Terminal Access Communication (DATAC) system. This installation is the frrst practical application of DATAC, a high- speed (1-MHz) multitransmitter/receiver data bus that only requires a single twisted wire pair to link all components on the bus. In the TSRV, two DATAC buses are actu- ally used, one for navigation, guidance, and control functions, and the other for the programmable display system. A data acquisition system (DAS) is also coupled to the DATAC bus and is set up to record approximately 540 data channels at 20 samples per second with selected channels at 100 samples per second.

The TECS algorithms are programmed into the Norden Computers via floppy disk and replace the pitch control and throttle control algorithms of the baseline software. The system,

1. 2. 3. 4. 5.

as flown on the TSRV airplane, consists of the following longitudinal modes:

Velocity vector control wheel steering (VEL-C WS). Glide slope capture and tracking (GS). Flightpath angle select @PA SEL). Altitude select and hold ( U T S, ALT H). Go Around (GA).

The system longitudinal modes can be flown in either CAS or Mach speed modes.

~ B E C ~ ~ W G PAW BLANK W 8 T FUJWD

9

The TECS software is integrated with the baseline lateral modes to give default mode pairing. Engagement of a longitudinal mode, for example, would result in the engagement of a corresponding mode in the lateral axis, or vice versa (Figure 6). On “power up” of the TECS system, the system is default in VEL-CWS and CAS modes. Engage- ment of FPA SEL, ALT H, ALT s, or GA modes gives Track Angle Select (TKA SEL) in the lateral axis.

The lateral VEL-C WS mode incorporates roll rate commandaltitude hold for maneuvering flight and track angle hold for nonmaneuvering flight.

In integrating TECS with the baseline, certain selected navigation, guidance, and lateral control functions are retained. These functions include dual DME, aided inertial navigation, and a Horizontal Path (HOR PATH) guidance mode in which the flight- paths can be selected by the pilot from a prestored data base or can be constructed in real-time using the NCDU. The baseline vertical path and time control modes were not enabled during these tests.

10

6.0 FLIGHT TEST RESULTS

This section discusses in detail the results of the flight test series. A summary of the figures and test conditions is given in Appendix 1. The tasks have been grouped ac- cording to mode of operation and do not necessarily follow the order in the test plan (Appendix 2). References 4 and 5 provide a summary of test results and highlight the unique aspect of the system operation.

6.1 FLIGHTPATH ANGLE MODE

Figures 7 through 13 show the performance of the system for FPA dialed in through the CMP while the airplane holds constant speed. Figures 7 through 10 show the performance for a mid-altitude mid-speed condition and Figure 13 shows a low-altitude low-speed condition.

Figure 7 shows the linear response of the system to a change in FPA of f 2.5 deg. The response shows no transient overshoot. Speed error was generally held within 1.5 kn although peak error during the maneuver was 2.5 kn and the maximum vertical acceleration change was about 3 Ws2.

Figures 9 and 10 show the effect of throttle limiting during a large FPA change. As discussed in Section 4, each mode of the autopilot is prioritized to control either speed or path when throttle limiting occurs. In FPA mode the strategy is simple, speed control has priority. Hence, on dialing in a large FPA change, the throttles will limit. The system will now stop controlling FPA, the throttle integrator is limited and the crossfeed between FPA and elevator is cut (i.e., the gain KCEm (Figure 2) is set to zero). The system controls speed through elevator and with the throttle at the forward limit, any energy not required to maintain speed is available to increase the FPA and, hence, the climb rate.

Figures 10a and 10b provide a good example of the throttle limiting case. In a steady state climb the airplane can maintain about a 6 deg FPA at 200 kn, 10,000 ft, which corresponds to a climb rate of about 40 ft. Maximum speed deviation during this maneuver is about 2 kn. Figure 10b shows the gain KGEm going to zero during the throttle limiting condition.

Figures 11 and 12 show the performance at high speed high altitude. In this condition the system cannot achieve the +2.5 deg command FPA and can only climb at 0.5 deg and sustain the 300-kn velocity. Throttle limiting occurs at 47 deg. In the -2.5 deg FPA case no limiting occurs and the system holds speeds within 3 kn. Figure 12 shows the result of a large negative FPA change in which the throttle limits at 8 deg. The system holds speed within a maximum error of 3 kn and the maximum FPA descent is at -3 deg. The results of tests 2.1.3 and 2.1.5 were not plotted as they were identical to test 2.1.1, as shown in Figure 11.

Performance of the system at low speed, low altitude is shown in Figure 13. It can be seen that the system limits in a climb of about 2 deg. Speed was held in limiting

11

situations within 3 kn, and normally within 2 kn. Tests 3.1.3 and 3.1.5 are similar to 3.1.1 and are not shown.

6.2 AIZI'ITUDE SELECT AND HOLD MODE

Figures 14 to 20 show the system performance for altitude changes at the three flight conditions discussed in Section 6.1. The linear performance at mid-altitude and mid- speed is shown in Figure 14. The system shows excellent transient response and speed error is less than 2 kn. Throttle limiting does occur during the descent at approximately 92 sec into the run but only for about 2 sec. The performance for large altitude changes is shown in Figures 15 through 17. In Figure 15 an altitude increase of 2,500 R is selected. In this case speed error is shown to increase up to 5 kn, about 20 sec after throttle limiting occurred. In the TECS design a feedforward was used to minimize speed error during a long climb. In simulation studies the speed loop feedforward gain KVT was set to 1.5, this provided a compromise over the whole flight regime. However, in fight test it was shown that increasing this gain KVT to 3, reduced the maximum speed error to about 3 kn. Simulation studies should be carried out to check the flight test results and optimize the gain value.

The low-altitude low-speed performance is shown in Figure 19. In test 3.2.3 (MI the pilot dialed up 2,500 R, the airplane reaches a steady-state climb rate of about 600 ft per minute, and the pilot changes command from 2,500 R to 1,000 R and the system captures that new altitude.

6.3 CAS/MACH MODE

The performance of the system to speed changes is shown in Figures 20 through 22 for mid-speed mid-altitude condition. For these cases, speed changes of up to 100 kn are performed. Altitude error is less than 20 R in all cases. The system has a smooth response and no excessive throttle activity is noticed. In Figure 22, which shows a f 100 kn speed change, throttle limiting occurs at 48 and 9 deg. Speed steps are also carried out at high altitude (Figures 23 through 26) and low altitude (Figure 27). Un- fortunately, the airplane had conservative throttle limiting at high altitude. The actual throttle limits are set at lower values than the software limits in the TECS program and, hence, the control did not reconfigure for this limiting situation. It can be seen in Figure 23 that the new speed target is captured successfully with no overshoot, however, a large altitude deviation occurs (in the order of 100 ft). This test is repeated in flight R461 at time 14:23:44. On that occurrence, an altitude error in the region of 65 R occurs. The details of the physical throttle limiting are not apparent in the baseline documentation and show a difference between the actual airplane and simulation. This test case has previously been successfully demonstrated on the NASA simulator.

Mach changes are shown in Figures 25 and 26. Unfortunately, the same problem with throttle limiting occurs in these cases as occurred earlier in the CAS changes. In Fig- ure 25 the pilot dialed in two incremental changes in Mach number of 0.25 Mach, and followed with a decrease in Mach number of 0.05 Mach. Altitude error increases in the

12

region of throttle limiting that occurs around 1,000 sec into the flight, otherwise altitude error is small throughout these runs. A glitch on the Mach number signal shows this signal going to zero at about 1,030 sec. The true airspeed error signal gives a clearer picture of the error going to zero with time as no glitch appears on that signal. Figure 26 shows the effect of the software limiting for large variations in Mach number. Altitude error increases up to 150 ft during that run. Unfortunately, again a glitch exists on the Mach number and it is shown going to zero at about 1,170 sec. Reference to true airspeed error signal (VTER) shows that the system captures the new velocity exponentially with no overshoot. Performance of the system at low altitude is shown in Figure 27. Examination of test 3.3.3, which starts at 215 sec into the run, shows that the speed dows not reach the commanded airspeed dialed in on the control mode panel and a stand-off of about 20 kn exists. This is caused by the pilot dialing in an airspeed above the flap placard for flaps 40 deg. The system increases speed to approximately 166 kn and holds that speed until the command is dialed down below the flap placard speed as can be seen at approximately 312 sec. Altitude error during that maneuver is shown to be about 50 ft.

Test 2.7.2 (Figure 28) is an example of the automatic Mach to CAS (and vice versa) switching. The airplane is executing a descent maneuver and holding constant Mach number. Just below 20,000 R the system automatically switches to holding a constant CAS of 330 kn. The maneuver is executed with about 0.lg acceleration.

6.4 VELOCITY VECTOR CONTROL WHEEL STEERING MODE

The Velocity Vector Control Wheel Steering Mode (VEL-CWS) is a manual flying mode with control augmentation that significantly reduces pilot workload when flying a path defined in inertial space. In this mode a column deflection commands a rate of change of FPA. Integration of the column command gives the FPA command ('yc) to which the system attempts to control. The performance of the VEL-CWS mode is shown in Figures 29 through 36. The response of the system at mid-speed mid-altitude conditions is shown in Figures 29 through 33. The response at high-altitude is shown in Figures 34 through 36.

The response to small column inputs is shown in Figure 29. In this case the pilot has put in a small impulse response on the column to get a gamma command of 2112 deg. The airplane climbs at 700 Wmin, and airspeed changes by approximately 2 kn. Throt- tle response is smooth and coordinated with no excessive activity. In test 1.5.3 (Figure 30) the pilot has put in a command of +5 deg FPA and the airplane is just on the verge of limiting the throttles. It is a good example of the airplane climbing out at maximum rate without losing airspeed.

In test 1.5.4 (Figure 31) the pilot has commanded -5 deg gamma, and the throttle has gone to the idle limit. In this situation the limit logic of TECS in the VEL-CWS mode means that path control has priority. Speed control is given up and the gain KmB, shown in Figure 2, is set to zero. The crossfeed gain KGEB is maintained. In this situation the speed error builds up as the path control is maintained. This can be seen as speed increases from the 200 kn that is commanded, up to 215 kn. If the pilot maintains the descent, the speed increases until overspeed protection engages and holds speed at a maximum.

Test 1.5.5 (Figure 32) is an example of a large y input in which the throttles do not limit and speed is maintained within about 3 kn. Figure 33 shows a large column input that causes throttle limiting. In this case FPA has priority and the command is reached. However, speed bleeds off rapidly and the underspeed protection mode switches in. FPA then decreases despite additional column inputs until minimum speed is reached. This corresponds to about 178 kn for flaps zero. A steady-state gamma of about 11.5 deg is reached. A negative column input decreases gamma and the airplane returns to the initial commanded speed of 200 kn.

Figures 34 through 36 show the VEL-CWS response at high altitude (20,000 R). In all these cases throttle limiting occurs. The system behaves as predicted in these limiting conditions and shows a smooth gamma response in each case.

6.5 UNDERSPEED PROTECTION

Underspeed is a priority control mode that switches in automatically irrespective of whether the mode engaged has speed or path priority. In the underspeed protection mode, the actual wing angle-of-attack is compared with the reference alpha. The reference alpha value corresponds to 1.3VSTALL for each flap setting. The error signal is converted to a normalized acceleration command WCm) that is computed continu- ously. The mode engages underspeed protection automatically (Figure 2) when:

Examples of the operation of the underspeed protection for angle-of-attack (AOA limiting) are shown in Figures 37 through 39. Figure 37 illustrates the situation of dialing the speed command to a very low value. The airplane slows down and when Vcm equals V, the mode switches in a transient-free manner and controls to the reference alpha. At flaps 0 deg this corresponds to about 178 kn. As the flaps were extended to flaps 40 deg, the speed decreases as the system controls to the new reference. Altitude deviation during the maneuver was negligible.

Figure 38 shows the situation of dialing speed to a large value and slowly retracting the flaps from 40 to 0 deg. Figure 39 shows a nonstandard situation in which speed has been dialed down and the flaps rapidly extended to 40 deg. The maneuver was accomplished safely with little deviation although considerable pitchdown was experienced in the airplane.

6.6 ENERGY EXCHANGE MANEUVERS

In Figures 40 and 41 the change in kinetic energy has been set equal to the change in potential energy. This is not a normal maneuver for a pilot, but it does demonstrate clearly the coordination between the throttle and elevator, with the TECS system, for a dual maneuver that changes both speed and altitude. In test 1.7.1 (Figure 40) the airplane loses 600 R of altitude and gains 20 kn in airspeed. The energy trade is accomplished by using the elevator. Aft flight deck throttle motion is only 0.5 deg

14

through the whole maneuver. Figure 41 shows a modified test 1.7.1. In this case the pilot changes speed by 20 kn and increases and then decreases altitude by 500 R. The change in potential and kinetic energy is not exactly equal and there is a small throttle movement of about 3 deg.

Test 1.7.3 (Figure 42) is an excellent example of a descent and speed decrease executed as a double maneuver. The airplane is cruising at 15,000 ft, 300 kn, gear and flaps up, the pilot simultaneously dials speed and altitude down. In this situation, prior to the throttles limiting, the autopilot treats the maneuver in a linear fashion and starts descending and decelerating. When the throttle limits out it is impossible to control speed and path, and the system has priority set to control speed and allow path to go open loop. In addition, a strategy is adopted that during descents 100 percent of the available energy is put into deceleration. For descents the energy is negative and it is the drag that can be split into either maintaining path and losing speed or vice versa. Once throttles limit, 100% of the drag is put into decelerating the airplane. The airplane levels out and slows down to 250 kn. Once this desired speed is captured, the airplane pitches over and descends at a rate commensurate with the drag configuration and maintaining airspeed.

During a simultaneous descent and decrease in speed that causes the throttle to limit, 100% of the available energy rate is put into decelerating the airplane. However, during a climbout situation (Figure 43) when the pilot has dialed in a large speed and altitude change, the limit logic computes the total energy available and splits the energy so that 50% is used for climb and 50% is used for acceleration. This is achieved by setting the Vc signal path limiter (Figure 2) to 50% of total available energy (Le., 0.5 V + -yJ. This ratio is readily adjustable by changing one gain in the software if it is determined by pilot evaluation that an alternative ratio is desirable and that priority should be given to achieving speed or altitude capture fwst. In the example, the airplane is climbing to 10,000 R altitude and increasing speed to 270 kn. The energy is split evenly between climbing and accelerating. The acceleration is approximately 2.5 Rlsz, which in energy terms corresponds to a y of about 0.08 Once the reference speed of 270 kn is captured, the climb rate doubles from about 35 ft to a maximum of 70 R. This can be seen to occur 48 sec into the run.

6.7 APPROACH AND GO-AROUND

The full-time underspeed protection, discussed in Section 6.5, can be used during approach to reduce the pilot workload. Figure 44 (tests 3.6.1 and 3.6.2) shows an approach and go-around situation. The airplane approaches the glideslope in altitude hold mode at 3,000 ft, 150 kn, flaps 15 deg, gear down. At glideslope intersection, shown by the glideslope error signal reducing to zero, the flaps are extended rapidly from 15 to 40 deg. The airplane overshot the glideslope by about 30 ft and it is considered that this overshoot was caused by a soRware problem that caused premature switching to glideslope capture prior to the control law logic switch point. On capture of the glideslope the command speed is dialed from 150 kn to a low value. Speed reduces to the value commensurate to the flap setting. For safety reasons the command speed is dialed up to 120 kn prior to go-around (220 sec into the run).

15

The TECS system at present does not include a flare control law; therefore this series of tests go-around is initiated at 300 ft. On go-around the control law commands a large FPA. This gives large elevator and throttle commands that cause the airplane to pitch up and climb at the maximum rate attainable while maintaining speed. As discussed earlier, during climbout the energy is split between gaining altitude and increasing speed. By coincidence, in this example the airplane reaches the commanded altitude of 3,000 ft and the new command speed of 225 kn simultaneously.

Elevator activity increases significantly during this run. The TECS system includes a complementary filter in which the time constant varies as a function of altitude. The effect is to improve path-tracking at lower altitudes and improve the system performance to windshear response. However, this has an effect of increasing elevator activity.

Test 3.6.3, 3.6.4 (Figure 45) shows high-speed capture of the glideslope from altitude hold mode. At 105 sec into the run, speed was dialed down to 100 kn, the system decreases speed and holds at approximately 180 kn as the underspeed protection mode switches in. in this run, the same software problem that was previously discussed caused premature switching to glideslope capture. This caused the airplane to pitch up and attempt to capture the glideslope early. Once the glideslope has been captured, the flaps are dropped rapidly to 40 deg and the speed decreases from 180 kn to approximately 120 kn. At 300 ft go-around is initiated, speed is dialed up to 200 kn and the airplane climbs out to 3,000 ft altitude. Figure 46, test 3.6.5, shows the end of the climbout and capture of the commanded altitude of 2,000 ft. This capture is done with negligible overshoot and smooth throttle response.

Figures 47 and 48 show tests 3.6.6 through 3.6.7. This case is very similar to tests 3.6.1 through 3.6.3 except that it shows a low-speed glideslope capture with flaps 40 deg and gear down, prior to glideslope interception.

Test 3.6.9 (Figure 49) shows the ability of the TECS system to capture the glideslope from above. In this case the airplane approached the glideslope flying straight and level, flaps set at 40 deg, gear down. Speed was dialed down to below 1.3 VREF at the start of the run and the airplane was flying in the AOA mode during the whole approach. The airplane flew through the glideslope at about two dots above the glideslope. Approximately 75 sec into the run, the FPA was dialed down to 4 . 5 deg. The airplane approaches the glideslope from above and at about 157 sec into the run the airplane captures the glideslope. The results show an excellent transition and smooth capture.

Glideslope capture and tracking using the VEL-CWS mode is shown in Figure 50 (tests 3.7.1 and 3.7.2). The airplane is flying straight and level, flaps 15 deg, CAS 150 kn. Speed is stabilized at about 150 kn at 58 sec into the run. At 66 sec the pilot pushes the column forward using the VEL-CWS mode and captures the glideslope. Pathtracking during the run is excellent and the pilot workload required to maintain good glideslope tracking is small. Go-around maneuver is initiated at about 400 ft by the pilot pulling a large column input and simultaneously dialing speed command to 180 kn, and retracting the flaps.

16

7.0 PILOT REACTION TO FLYING THE TSRV WITH TECS

Pilot reaction to flying the TSRV with TECS was very favorable. Explanation of the TECS concept and philosophy to pilots was very important so that the operation of TECS was predictable and consistent in flight.

A major difference from a conventional system is the full-time autothrottle. Tradition- ally, autothrottles have been heavily 'kiticized for excessive activity and unpredictable or counterproductive motion in maneuvers. TECS solves that problem and produces an autothrottle that is predictable and minimizes unnecessary throttle activity, both in situations of simultaneous altitude and speed changes and when flying the VEL-CWS mode. The pilot does not need to switch off the autothrottle and fly manual throttles to achieve satisfactory flying qualities.

With TECS, the mode logic is simple and the mode hierarchy is straightforward so the behavior is predictable in throttle limiting conditions. In all events, limiting prevents stall and overspeed irrespective of the mode or combination of modes.

The strategy adopted during simultaneous speed and altitude changes of splitting energy (50%/50% during climb, 100% into deceleration during descent) is flexible and, although considered the preferred system, can readily be changed by adjusting one gain.

The VEL-CWS mode is the preferred method of flying the airplane during climb-out and approach. It allows the pilot, as opposed to the autopilot, to fly the airplane with great precision and yet greatly reduces the workload. Interception and tracking of the glideslope, even in turbulence, is a simple task.

17

8.0 CONCLUSIONS

The Total Energy Control System was successfully flight-tested on the NASA Langley TSRV (B757) in a total of five flights. The system did not exhibit any instabilities or design problems that required gain adjustment in flight. No major problems were encountered during the tests.

The success of the flight tests validates the extensive use that was made of analysis, simulation, and hot bench checkout to thoroughly develop, check out, and verify the system design prior to flight tests. No system tuning was necessary in flight.

The integrated autopilot/autothrottle received favorable pilot reaction. The full-time autothrottle, inherent in the design, was predictable during maneuvers and did not exhibit unnecessary throttle activity.

Performance in all modes was comparable to simulation results. Path tracking was excellent, altitude deviation during large speed changes was less than 20 R. Speed tracking during maneuvers was generally less than 2 kn although a peak error of 5 kn was noted on certain large altitude changes.

The velocity vector control wheel steering (VEL-CWS) mode received very favorable pilot reaction because of its consistent performance over the flight regime, the predictable and responsive throttle, and the reduction in workload that the mode allows when carrying out precision tracking tasks.

TECS has provided NASA with a state-of-the-art integrated autopilot/ autothrottle suitable for use with future NASA experiments such as 4D navigation.

18

9.0 REFERENCES

1. Lambregts, A. A., “Vertical Fiightpath and Speed Control Autopilot Design Using Total Energy Principles,” AIAA Paper 83-2239CPY August 1983.

2. Lambregts, A. A., “Operational Aspects of the Integrated Vertical Flight- path and Speed Control System,” SAE Paper 831420, October 1983.

3. Lambregts, A. A., “Integrated System Design for Flight and Propulsion Control Using Total Energy Principles,” AIAA paper 83-2561, October 1983.

4. Bruce, K. R., Kelly, J. R., and Person, L. H. Jr., “NASA B?3? Flight Test Results of the Total Energy Control System,” AIAA Guidance, Navigation and Control Conference, AIAA 86-2143 CP, August 1986.

5. Kelly, J. R., Person, L. H. Jr., and Bruce, K. R., “Flight %sting TECS - The Total Energy Control System, Aerospace Technology Conference, Long Beach, California,” SAE Paper 861803, October 1986.

19

Figure 1. Total Energy Control System Concept

r&EWLNG PAGE BLANK NOT FILMED

21

t

w +

-0

B

I 4 C .e E

s Q E

c

E -

22

1

Low I Level

Flight Control

Units

Figure 3. NASA Langley TSR V Showing Internal Arrangements

CONTROL MODE PANEL

UMP

TEST

Figure 4. TSRVflECS Control Mode Panel

23

k v i a a t i o n and 1 Flighi Control Sensors

b

RFD Controls

L I

Flight Control

CMP and NCDU

Y Norden Display No. 1 System

Norden

c. m d

n Data Acquistion System

4/ Strip Charts

1 I U

I uian

U

Flight Control Engineer Terminal

r

Flight Management Engineer Terminal

Figure 5. Research System Architecture

.

24

Pitch and Speed r - I - - 7

I

VEL GS FPA ALT ALT cws LOC SEL S H

1

Figure 6. TECS Control Modes

25

5 FPASUM

-5

0.088 GAMMA1

-0.088

Az

CAS

HDOT

DECMD

AFDTH

30

20

205

195

CAS 200 kn Flaps 0 Gear Up

Alt 10,000 ft

Wt 86,000 Ib

Flight R455:

Time 15:32:19 133552

GAMEPS

25 50 75 100 125 150 175 200 Time (sec)

~ _ _ _ _ _ _ _

Figure 7. Test 1.1.1-1.1.2; FPA Mode +/-2.5 deg

26

5 FPASUM

-5

GAMMA1

Az

CAS

HDOT

DECMD

AFDTH

GAMEPS

0.088

0.088

30

20 . 205

195

50

0

0

-0.1 . 0 10 - 20 30 40 50 60 70 80

Time (sec)

Figure 8. Test 1.1.3; FPA Mode +5 deg

CAS 200 kn

Flaps 0

Gear Up

Alt 10,000 ft

Flight R455:

Time 1 540: 15 1541 :37

27

0 FPASU M

-1 0

0 GAMMA1

-0.1756

FPNCAS Mode

Flaps 0

Gear Up

Flight R455:

Time 15:43:05

GAMEPS

200 IASSUM

CAS 1 95

ALPHA

Az

AFDTH

FLAP

DECMD

10

5

35

25

40

0

0

.A

0

.A - : 0 12.5 25 37.5 50 62.5 75 87.5 100

Time (sec)

Figure 9 (a). Test 1.1.4; FPA Mode -5 deg

28

PITCH 10

0

Q

STAB

FPNCAS Mode

Flaps 0

Gear Up

Flight R455:

Time 15:43:05

7

5

EPRA 10

-1 0

KGEPS

KVDEPS

DECMD

90,m

0.5

-0.5

1

0

--T .r

0 12.5 25 37.5 50- 62.5 75 87.5 100 Time (sec)

Figure 9 (b). Test 1.1.4; FPA Mode -5 deg

29

10 FPASUM

-5

0.1746 GAMMA1

-0.0873

Az

IASSUM

CAS

HDOT

GAMEPS

AFDTH

DECMD

30

20

205

195

.1

-0.1

40

0

I I I I I / I I 1 1 I 1 , . I I

0 I - ! I I i l l I I I I I I ! I I 2

-2 0 7.5 15 22.5 30 37.5 45 52.5 60

Time (sec)

CAS 200

Flaps 0

Gear Up

Alt 5000

kn

ft Flight R455:

Time 154335 1547: 10

Figure 10 (a). Test 1.1.5; FPA Mode + 15 deg

30

8 ALPHA

6

1 KGEPS

STAB

( 1 r I

EPRA

5.4

5

86,100 WEIGHT

85,900

1 KVDEPS

0.9998

DECMD

0 15 30 45 60 75 90 105 120 Time (Sec)

Flight R455:

Time 154535

Figure 10 (b). Test 1.1.5; FPA Mode + 15 deg

31

5 FPASUM

-5

Flight R457:

0.0873 Time 16:42:05 GAMMA1

-0.0873

GAMEPS

IASSUM

CAS

ALPHA

Az

AFDTH

FLAP

DECMD

0

-0.41

300

290

1

-1

35

25

35

25

40

0

60 120 180 240 300 360 420 480 Time (sec)

Figure 11 (a). Test 2.1. 1, 2.1.2; FPA Mode +/-2.5 deg

32

PITCH

Q

STAB

EPRA

1

-1

3.6

3.2

2

1

81,500 WEIGHT

80,500

1 KGEPS

0

1 KVDEPS

0.9998

0 DECMD

___ _ _ .~

0 60 120 180- 240 300 360 420- 480 4

Time (sec)

Figure 11 (b). Test 2.1.1, 2.1.2; FPA Mode +/-2.5 deg

33

Flight R457:

Time 16:42:05

5 FPASUM

-5

0.0873 GAMMA1

-0.0873

GAMEPS

IASSUM

CAS

ALPHA

Az

AFDTH

FLAP

DECMD

305

295

1

-1

35

25

2

-2 0 15 30 45 60 75 90 105 120

Time (sec)

Flight R457:

Time 1651 :45

Figure 12 (a). Test 2.1.4; FPA Mode -5 deg

34

PITCH

Q

STAB

EPRA

WEIGH T

0

4

Flight R457:

Time 1651 :45 0

-1 e

3.6

3.2

1.5

0.5

1 KGEPS

0

1 KVDEPS

0.9998

DECMD 2

-0 15 30 45 60 75 90 105 120 Time (sec)

Figure 12 (b). Test 2.1.4; FPA Mode -5 deg

35

FPASUM

GAMMAD

IASSUM

CAS

GAMEPS

PITCH

ALPHA

Az

AFDTH

FLAP

DECMD

0

-1 0

135

125

0

-0.1

0

10

10

0

35

25

35

25

40

5

-5 0 50 100 150 200 250 300 350 400

Time (sec)

FPA Mode

Altitude 5000 ft

Flaps 40

Gear Down

Flight R456

Time 18:49: 10

Figure 13 (a). Test 3.1.1, 3.1.2, 3.1.4; FPA Mode +/-2.5 deg, -5.0 deg

36

0 PITCH

-1 0

Q

STAB

EPRA

2

-2

7.5

6.5

87,000 WEIGHT

86,000

KGEPS

0.9996 KVDEPS

0.9994

DECMD 5

-5 - 0 50 100 150 200 250 300 350 400

Time (sec)

FPA Mode

Altitude 5000 ft

Flaps 40

Gear Down

Flight R456

Time 18:49: 10

Figure 13 (6). Test 3.1.1, 3.1.2, 3.1.4; FPA Mode +/-2.5 deg, -5.0 deg

37

10,500 ALTSUM ALTCOR

10,000

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

200

180

50

-50

40

0

2

-2 . ~~

0 15 30 45 60 is 90 105 120 Time (sec)

Flight R455:

Time 15:49: 19

Figure 14 (a). Test 1.2.1 and 1.2.2; Alt +300 ft, -300 ft

38

STAB

Q

KGEPS

KVDEPS

EPRA

VTER 10

-1 0

WElGt i 85,900

T

85,700 0 15 30 45 60 75 90 105 120

Time (sec)

Flight R455:

Time 1549: 19

Figure 14 (b). Test 1.2.1 and 1.2.2; Alt +300 ft, -300 ft

39

13,000 ALTSUM

ALTCOR 10,000

IASSUM

CAS

Az

200

180

35

25

PITCH 10

0

HDOT

ALW

AFDTH

DECMD

50

-50

40

0

160 175 190 205 220 235 250 265 280 Time (sec)

Flight R455:

Time 15:54:00

Figure 15 (a). Test 1.2.3; Alt +2,500 ft

40

STAB

Q

5

0

Flight R455: 2 Time 15:54:00

-2

1 KGEPS

0

1 KVDEPS

0000

EPRA

VTER 10

-10

85,900 WEIGHT

85,700 160 175 190 205 220 235 250 265 280

' Time (sec)

Figure 15 (b). Test 1.2.3; Alt +2,500 f t

41

16,000 ALTSUM

ALTCOR 12,000

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

200

190

35

25

15

0

8

6

40

0

0

-4 0 20 40 60 80 100 120 140 160

Time (sec)

Alt/CAS

Mode

M idAlti tude

Flaps 0

Gear Up

Flight R457:

Time 16:02:55

Figure 16 (a). Test 1.2.3M; Altitude Change + 4,000 f?

42

STAB

Q

KGEPS

2

-2

1 KVDEPS

aaan

EPRA

VTER

..;I""" .

I .

10

-in

85,600 WEIGHT

85,200 80 100 120 140 160 -- 0 20 40 60

Time CGc,

Figure 16 (b). Test 1.2.3M; Altitude Change +4,000 ft

Alt/CAS Mode

Mid-Altitude

Flaps 0

Gear Up

Flight R457:

Time 16:02:55

13,000 ALTSUM

ALTCOR 10,000

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

2001

1801

4

0,

0

-40

7

5

15

5

0

50 75 100 125 150 175 200 4 0 25

Time (sec)

Flight R455:

Time 155:55:28 .

Figure 1 7 (a). 1.2.4; Alt -2,500 ft

44

STAB

Q

Flight R455:

Time 1555528 2

-2

1 KGEPS

0

1 KVDEPS

.9999

EPRA

VTER

1.4

1

5

-5

86,500 WEIGHT

75 100 125 150 175 200 85,500

- 0 25 50 Time (sec)

Figure 17 (b). Test 1.2.4; A/t -2,500 ft

45

27,000 ALTSUM

ALTCOR 23.000

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

305

295

30

20

5

3

3

2

40

0

2

-2 0 35 70 105 140 175 210 245 280

Time (sec)

Flight R457:

Time 16:56:03

Figure 18 (a). Test 2.2.1-2.2.3; Alt + 300, -300, +2,500, -2,500 ft

46

STAB

Q

KGEPS

KVDEPS

4

3

0.5

-0.5

1

0.9998

EPRA

.5

0 VTER

1.5

-20

80,500 WEIGHT

79,500 0 35 70 105 140 , 175 210 245 280

Time (sec)

Alt Mode

Altitude 25,000 ft

Flaps 0

Gear Up

Speed 300 kn Flight R457: Time 16:56:03

Figure 18 (b). Test 2.2.1-2.2.3; Alt + 300, -300, +2,500, -2,500 ft

47

ALTSUM

ALTCOR

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

6000

2000

135

125

35

25

10

-1 0

0

-40

10

0

40

0

5

-5

5

-5 - 0 55 110 165 220 275 330 385 440

Time (sec)

Alt Mode

M id-At ti tude

Flaps 40 Gear Down

Flight R456: Time 19:02:00

Figure 19 (a). Test 3.2.1-3.2.4; Altitude Changes 300 f t , 2,500 f t

48

STAB

Q 0

-4

Alt Mode

M id-AI ti tude

Flaps 40 Gear Down

Flight R456: 1 Time 19:02:00 KGEPS

0

0.9996 KVDEPS

.9994

EPRA

VTER 0

-20

86,000 WEIGHT

84,000 0 55 110 165 220 275 330 385 440

Time (sec)

Figure 19 (b). Est 3.2.1-3.2.4; Altitude Changes 300 ft, 2,500 ft

49

270 IASSUM

CAS 230

10,000 ALTSUM

ALTCoR 9,950

Az

PITCH

32

28

5

-5

ALPHA

50 VTER

-50

AFDTH 30

10

DECMD

0 25 - 50 75 100 125 150 175 200 Time (sec)

Figure 20 (a). Test 1.3.1 and 1.3.2; CAS +20 -20

AltlCAS

Mid-Altitude

Flaps 40

Gear Up

Flight R455:

Time 16:01:17

50

Q

- . . _ . . . = _ . .

1

_ _ . - n

EPRA

I I ..

I 1 . - . . . . . .. , . . . , . .,

.- -. . . -_ . . .. - . . - - . - . ~. . . .___.- . . ., ..

1 THRSTC

KGEPS

KVDEPS

0

-5

1

-1

0,000

0

10,000 WEIGHT

70,000

VTER 1 00

-1 00

1 DECMD

75 ioo- 125 150 175 200 -1 0 25 50

Time (sec)

Flight R455:

Time 16:Ol: 17

Figure 20 (b). Test 1.3.1 and 1.3.2; CAS +20 -20

51

300 IASSUM

CAS 230

10,000 ALTSUM

ALTCoR 9,950

Az

PITCH

ALPHA

KGEPS

AFDTH

DECMD

32

28

5

-5 I

. . - . . . . . . . . . - . . . . . . . . . .

4 ...... -- .. \

......... . . . . . . . . . .

30

0

2

-2 I 200 225 250 275 300 325 350 375 400

Time (sec)


52

Flight R455:

Time 16:09:00

Q

1 KVDEPS

EPRA

- _ . . . . . . . ... - . _ _ .. - ., . . 4b

. . . . . . . . . . . . . . . . . . 0

.- . . . . . . . . . . . . . . . _. . . - . . . . . . . . . . .

1

-1

10,000 THRSTC

0

1 KGEPS

0

100,000 WEIGHT

70,000

1 DECMD

-1 200 225 250 275 300 325 350 375

Time (sec)


Flight R455:

Time 16:09:00

53

IASSUM

CAS

300

100

CASIAlt

Hold

Mode

Flaps 0

Gear Up

AI t

10,000 ft

Flight R455:

ALTCOR

PITCH

Q

AFDTH

DECMD

4 \- I I

. . . . . . . . . . . . . . . . . . . . . . . . . . ... ... . . . . . . . . . . . . . . . . . . . - . . . - . .

0 )

0

-2

40

0

5

0 50 100 150 200 250 300 350 400 -5

Time (sec)

Time 16:14:38 - 16:20:00

Figure 22. Test 1.3.5, 1.3.6; CAS Change +lo0 kn, -700 kn

54

320 IASSUM

CAS 280

24,000 ALTSUM

ALTCOR _ _ _ ~

23.800

CAS Mode

Speed 300 kn

Alt 25,000 ft

Gear Up

Flaps 0

35 Flight R457: Az Time 17:01:05

25

PITCH

ALPHA

VTER 0

-200

40 AFDTH

0

DECMD 0

-4 0 19 38 57 76 95 114 133 152

Time (sec)


0 Q

-1

2 EPRA

1

10,000 THRSTC

0

1 KGEPS

n

1 KVDEPS

0.9998

79,800 WEIGHT

79,400

VTER

DECMD

- 0 19 38 57 76 9 5 114 133 152

Time (sec)

CAS Mode

Speed 300 kn Alt 25,000 ft

Gear Up

Flaps Up

Flight R457:

Time 17:01:05


56

300 IASSUM

CAS 220

20,000 ALTSUM

ALTCOR 19,800

Az

PITCH

A I ni I A n u nn

VTER

AFDTH

DECMD

5

4

11

120

-1 20

40

n V '

2

- - -_ - -2 700 725 750 775 800 825 850 875 900

CAS Mode

Speed 250 kn

Alt 25,000 ft

Gear Up

Flaps 0

Flight R461:

Time 14:28:12

Time (sec)

Figure 24 (a). Test 2.3.3-2.3.4; CAS +50 -50

57

0 Q

-1

CAS Mode

1.5 Speed 250 kn

0.5 Gear Up

EPRA Alt 25,000 ft

Flaps 0

10,000 Flight R461:

Time 14:28:12 THRSTC

-1 0,000

1 KGEPS

KVDEPS

95,000 WEIGHT

85,000

VTER

DECMD

200

-200

200

. 10

-1 0 700 725 750 775 800 825 850 875 900

Time (sec)

Figure 24 (b). Test 2.3.3-2.3.4; CAS +50 -50

58

MACSUM

MACH

0.7

0.6

IASSUM

CAS

325

300

Mach Mode

Mach 0.65

Altitude 20,000 ft

Flight R461:

Time 14:37:31

PITCH

ALPHA

VTER

AFDTH

DECMD

5

-5

50

-50

40

0

10

-10 . - 915 940 965 -990 1,015 1,040 1,065 1,090 1,1.15

Time (Sec)

Figure 25 (a). Test 2.4.1-2.4.2; Mach +0.05 -0.05

59

0 Q

EPRA

-2

1.5

0.5

Mach Mode

Mach 0.65 Altitude 20,000 ft


10,000 THRSTC

-1 0,000

1 KGEPS

0

KVDEPS 1

-1

. . . . . . .. . . . i . .

. . . . . . . . . . . . . . . . . . . . . .

-1

95,000 WEIGHT

85,000

VTER 200

-200

200

-200

10 DECMD

-1 0 915 940 965- 990 1,065 1,015 1,040 1,090 1,115

Time (sec)

Figure 25 (b). Test 2.4.1-2.4.2; Mach + 0.05 -0.05

60

MACSUM

MACH

IASSUM

CAS

Mach Mode

Mach 0.65

Altitude 20,000 ft

Flight R461:

340

300

Time 14:41:20

20,000 ALTCOR

19,800

PITCH

ALPHA

VTER

AFDTH

DECMD

5

5

1

-1

100

-1 00

40

0

10

-1 0 1,120 1,170 1,220 1,270 1,320 1,370 1,420 1,470 1,520

Time (sec)

Figure 26 (a). Test 2.4.3-2.4.4; Mach +O. 75 -0.75

61

Q 0

Mach Mode

1.5 Mach 0.65 EPRA Altitude 20,000 ft

0.5 Flight R461: Time 14:41:20

10,000 THRSTC

-1 0,000'

- . . . . . . ...... _ _ -.- __ ~ . . . . . ~ . . . . . . . KGEPS

... -. .. -. - .. - . .. . . . . _ _ . . . ._ .. - . . . . . . . - . _. . . . . . . . . . . -- ... - - . . . .

01

KVDEPS

WEIGHT t " ' J I t 1

I I I 1 . 1 I d 85,000

VTER t I I t I 1 1 - 1 I C I I I I I i I I

DECMD

1,120 1,170 1,220 1,270 1,320 1,370 1,420 1,470 1,520 Time (sec)

Figure 26 (b). Test 2.4.3-2.4.4; Mach +O. 15 -0.15

62

IASSUM

CAS

200

100

CAS Mode

Altitude 5,000 ft

Flight R456:

Time 19:33:00

5,050 ALTSUM

ALTCoR 4,950

Az

PITCH

ALPHA

VTER

AFDTH

DECMD

35

25

0

-1 0

0

-200

40

20

5

-5 0 50 100 150 200 250 300 350 400

Time (sec)

Figure 27. Test 3.3.1-3.3.4; CAS +20, -20, +50, -50 kn

63

320 IASSUM

CAS 280

ALTSUM

ALTCOR

Az

PITCH

ALPHA

VTER

AFDTH

DECMD

20,000

0

40

30

-5

2

0

0

-20

20

1

-1 0 40 80 120 160 200 240 280 320

Time (Sec)

Mac h/C AS

Altitude 20,000 ft Gear Up

Flaps 0

Flight R461:

Time 155535

Figure 28. Test 2.7.2; Mach to CAS Switch

64

DCOL 0

-3 -

V-CWSICAS Mode

GAMMA1 0 Flaps 0

Gear Up

-0.1 Alt 10,000 ft

Flight R455:

Time 18:22:00 10,500 ALTCOR

9,500

CAS 200

196

Q

30 AFDTH

10

0 DECMD

-4 0 40 80 120 160 200 240 280 320

Time (sec)

Figure 29. Test 1.5.1, 1.5.2; Wocity Wctor-CWS

65

DCOL

GAMMA1

ALTCOR

CAS

Az

Q

AFDTH

DECMD

0.1

-0.1

200

150

30

10

5

-5 0 15 30 45 60 75 90 105 120

Time (sec)

V-CWSICAS Mode

Flaps 0

Gear Up

Alt 10,000 ft

Flight R455:

Time 18:29:30

Figure 30. Test 1.5.3; klocity kctor-CWS

66

DCOL

GAMMA1

ALTCOR

CAS

Q

AFDTH

0

-2

0

-0.1

2,000

0,000

200

1 90

DECMD

80 100 120 140 160 0 20 40 60 Time (sec)

V-CWSIC AS

Mode

Flaps 0

Gear Up

Alt 10,000 ft

Flight R455:

Time 18:29:30

Figure 3 1. Test 1.5.4; klocity kct0r-C WS

67

DCOL 2

-2

GAMMA1

ALTCOR

CAS

Az

Q

AFDTH

DECMD

0.1

-0.1

12,000

8,000

205

1 95

2

-2

5

-5 - 0 20 40 60 80 100 120 140 160

Time (sec)

V-CWSICAS

Mode

Flaps 0

Gear Up

Alt 10,000 ft


Figure 32. Test 1.5.5; Velocity Vector-CWS

68

DCOL

GAMMA1

V-CWSiCAS

Mode

Flaps 0

Gear Up

Alt 10,000 ft

12,000 Wt 90,000 Ib

Flight R455: ALTCOR

8,000 Time 18:37:00

CAS 200

1 50

Az

Q

AFDTH

40

20

0

-4

40

0

40

0

5 DECMD

-5 0 20 40 60 80 100 120 140 160

Time (sec)

Figure 33. Test 1.5.6; klocity kctor-CWS

69

1 DCOL

-1

0.5 GAMMA1

-0.5

21,000 ALTCOR

20,000

CAS

Az

Q

AFDTH

DECMD

300

280

35

25 I

0

-1

50

30

0 12 24 36 48 60 72 84 96 Time (sec)

V-CWSIC AS

Mode

Flaps 0

Gear Up

Alt 20,000 ft


Figure 34. Test 2.5.7; V-CWS +2.5 deg

70

.

DCOL 0

4 - 1

V-CWS Mode

-0.02 Flaps 0 GAMMA1

Gear Up

-0.06 Alt 20,000 ft

Flight R461:

21,000 Time 14:53:43 ALTCOR

20,000

CAS

Az

Q

AFDTH

DECMD

302

300

30

26

0.2

-0.2

30

10

1

-1 0 3.5 7 10.5 14 17.5 21 24.5 28

Time (sec)

Figure 35. Test 2.5.2; VCWS -2.5 deg

71

DCOL

GAMMA1

1

-1

0

-1

V-CWS Mode

Flaps 0

Gear Up

Alt 20,000 ft

Flight R461:

20,000 Time 14:54:20 ALTCOR

16,000

CAS

Az

Q

AFDTH

DECMD

320

?nn

35

3c

0

-1 . .

12

R

0

-2 0 7 14 21 28 35 42 49 56

Time (sec)

Figure 36. Est 2.5.4; V-CWS -5 deg

72

ALFREF

ALW

AONAlt Mode

10,000 Alt 10,000 ft

Flight R455:

Time 18:47:00-19:09:00

ALTCOR

9,900

CAS

FLAP

Az

PITCH

AFDTH

200

1 00

30

0

35

25

10

0

40

0

40

0

DECMD

-200 275 350 425 500 575 650 725 800 Time (sec)

Figure 3 7 (a). Test 7.6.1-8; A OA Protection

73

5 STAB

-5

EPRA

AOAIAlt

Mode

Alt 10,000 ft

Flight R455:

Time 18:47:00-19:09:00

THRSTC

WEIGHT

VTER

Q

ALPHA

DECMD

90,000

00000

4

-50

-1 50

0

-2

200 275 350 425 500 575 650 725 800 Time (sec)

Figure 37 (b). Test 1.6.1-8; AOA Protection

74

10 ALFREF

ALW 0

AONAlt

10,000 Mode

Alt 10,000 ft

9,900 Flight R455:

ALTCOR

CAS

FLAP

Az

PITCH

AFDTH

DECMD

Time 18:47:00-19:09:00

200

100

30

0

35

25

10

0

40

0

5

-5 - . - 7

800 goo i o o o i , ioo i,ioO 1,300 1,400 1,500 1,600 Time (sec)

Figure 38 (a). Test 1.6.9-1 5; AOA Protection

75

STAB

EPRA

5

-5

AONAlt Mode

2 Alt 10,000 ft

Flight R455:

1 Time 18:47:00-19:09:00

10,000 THRSTC

0

90,000 WEIGHT

00000

VTER

Q

-1

450

50

0

-2

0

10 ALPHA

5

5 DECMD

-5 800 900 1,000 1,100 1,200 1,300 1,400 1,500 1,600

Time (sec)

Figure 38 (b). Test 1.6.9- 15; AOA Protection

76

ALFREF

ALW

ALTCOR

CAS

FLAP

Az

PITCH

AFDTH

10

n

4,500

4,400

AONAlt Mode

Flight R456:

Time 15:52:00

35

96

5 DECMD

-5 0 15 30 45 60 75 90 105 120

Time (sec)

Figure 39. Test 3.5.1; AOA Protection

77

ALTSU M

ALTCOR

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

10,500

9,500

AlffCAS Mode

280 Mid-Altitude

Flaps 0

240 Gear Up

Flight R457:

35 Time 15:50:35

25

2

-2

24

0 5 10 15 20 25 30 35 40 Time (sec)

Figure 40. Test 1.7.1; Energy Change (KE=PE)

78

10,500 ALTSUM

ALTCoR 9.500

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

260

220

Alt/CAS Mode

M id-AI ti tude

Flaps 0

Gear Up

Flight R457:

35 Time 15:50:35

25

10

0

50

xn

5.5

4.5

22

18

1

I . . l t ' ! ' : I ! ! ' ! ! ! ' ! ' ! ' 0 15 30 45 60 75 90 105 120

Time (sec)

Figure 4 1. Test 1. 7.1; Energy Change (KE= PE)

79

ALTSUM

ALTCOR

15,000

10,000

IASSUM

CAS

A2

PITCH

HDOT

ALW

AFDTH

DECMD

300

240

35

3c GV .

4

2

20

0 40 80- 120 160 200 240 280 320 Time (sec)

Alt and CAS Mode

Flaps 0

Gear Up

Flight R455: Time

Figure 42 (a). Test 1.7.3; Altitude Change -5,000 ft, Speed Change -50 kn

80

STAB

Q 2

-2

Alt AND CAS Mode

Flaps 0

Gear Up

Flight R455:

Time 1

KGEPS

0

1 KVDEPS

.9999

EPRA

VTER

1.5

0.5

0

-200

87,000 WEIGHT

86,800 0 40 80 120 160 - 200 240 280 320

Time (sec)

Figure 42 (6). Test 1.7.3; Altitude Change -5,000 ft, Speed Change -50 kn

81

10,000

ALTCOR 0

ALTSUM

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

300

200

AltICAS Mode

Mid-Altitude

Flaps 0

Gear Up

Flight R459:

35 Time 19:17:45

25

10

6

60

20

5

3

50

45

DECMD

"0 12 24 36 48 60 72 84 96 Time (sec)

Figure 43 (a). Climbout

82

KGEPS

0.9996 KVDEPS

0.9994

AlffCAS Mode

M id-AI ti t ude

Flaps 0

Gear Up

Flight R459: Time 1 9: 17:45 100

VTER

-1 00

FLAP 0

0.0002

HDOT

0.15 GAMMA1

0.05

Q

.

0

0 12 24 36 48 60 72 84 96 Time (sec)

Figure 43 (b). Climbout

83

ALTSUM

ALTCOR

IASSUM

CAS

ALFREF

ALW

FLAP

HDOT

GSEFT

AFDTH

DECMD

2000

0

200

0

10

0

40

0

5

0 60 120 180 240 300 360 420 480 -5

Time (sec)

Figure 44. Test 3.6.1, 3.6.2; Approach and Go-Around

Alt/Glides lope/Go-Around


84

ALTSUM

ALTCOR

IASSUM

CAS

Az

4000

0

400

0

40

20

10 PITCH

10

HDOT

ALW

50

-50

10

0

DECMD 0

-20 0 50 100 150 200 250 300 350 400

Time (sec)

Figure 45 (a). Test 3.6.3R, 3.6.4R; Capture Glideslope

AWGS Mode

Altitude 3,000 ft

Flaps 0

Gear 0

Flight R456:

Time 17:43:58

STAB

Q

KGEPS

KVDEPS

EPRA

VTER

10

0

AWGS Mode

2 Altitude 3,000 ft

Flaps 0

-2 Gear 0

Flight R456:

1 Time 17:43:58

0

1

0

-400

94,000 WEIGHT

93.000 0 50 100 150 200 250 300 350 400

Time (sec)

Figure 45 (b). Test 3.6.3R; Capture Glideslope

86

ALTSUM

ALTCOR

IASSUM

CAS

Az

PITCH

HDOT

ALW

AFDTH

DECMD

2000

1800

220

200

40

20

10

0

0

-1 00

5

-5 0 12.5 25 37.5 50 62.5 75 87.5 100

Time (sec)

Alt Sec It Hold

Flight R456:

Time 17:52:00

Figure 46. Test 3.6.5; Alt Select to Alt Hold

87

ALTSUM

ALTCOR

IASSUM

CAS

ALFREF

ALW

Az

GAMEPS

GSEFT

AFDTH

FLAP

DECMD

2000

0

Alt/Glideslope

Flight R456:

Time 18:08:04

10

0

0.2

_n 3 -V.& .

500

rnn

40

10

5

-6 - 0 40 80 120 160 200 240 280 320

Time (sec)

Figure 47 (a). Est 3.6.6 (R); Glideslope Capture

88

2,000 ALTSUM

ALTCOR

PITCH

Q

EPRA

AWGlideslope

Flight R456:

Time 18:08:04

5

-5

5

-5

91,600 WEIGHT

91,200

KGEPS

KVDEPS

0 40 80 120 160 200 240 280 320 Time (sec)

Figure 47 (b). Test 3.6.6 (R); Glideslope Capture

89

2,000 ALTSUM

ALTCOR

IASSUM

CAS

0.0873 GAMMA1

-0.0873

ALFREF

ALW

HDOT

FLAP

AFDTH

DECMD

10

0

20

-20

40

0

0 20 40 60 80 100 120 140 160 Time (sec)

Figure 48. Test 3.6.7; Go-Around and Climbout

Glideslope/Go-Around/Alt

Flight R456:

Time 18:Ol:lO

ALTSUM

ALTCOR

IASSUM

CAS

FPASUM

GAMMA1

GSEFT

4000

0

1 30

100

5

-5

0

-0.1

10 ALFREF

ALW 0

AFDTH

FLAP

DECMD

0 30 60 90 120 150 180 210 240 Time (sec)

f PNGlideslope

flight R456:

Time 19:28:15

Figure 49 (a). Test 3.6.9 (R); Glideslope Capture From FPA Mode

91

ALTSUM

ALTCOR

GAMEPS

Az

PITCH

HDOT

EPRA

4000

0

FPNGlideslope

0.1 Flight R456:

Time 19:28:15

-0.1

35

25

5

-5

0

-40

1.4

1

89,500 WEIGHT

88,500

0 DECMD

l I i l 1 I / i t q 1 1 1 -10 -.

0 30 60 90 120 150 180 210 240 Time (sec)

Figure 49 (b). Test 3.6.9 (R); Glideslope Capture From FPA Mode

0.5 DCOL

-0.5

Flight R456:

Time 18:38:40 0.0873 GAMMA1

-0.0873

2000 ALTCOR

0

180 IASSUM

CAS 140

10 ALFREF

ALW 0

0.05 GAMEPS

-0.05

AFDTH

FLAP

40

0

5 DECMD

-5 -. __ 0 45 90 135 180 225 270 315 360

Time (sec)

Figure 50 (a). Test 3.7.1, 3.7.2; V-CWS Approach

93

0.0873 GAMMA1

-0.0873

PITCH

Q

EPRA

WEIGH

Az


0

-4

88,000 T

87,000

35

25

20,000 THRSTC

0

5 DECMD

0 45 90 135 180 225 270 315 360 Time (sec)

Figure 50 (b). Test 3.7.1, 3.7.2; V-C WS Approach

94

Appendix 1: Summary of Figures and Test Conditions

Figure Test Flight Conditions Alt (f t) Speed (kn) Remarks

6.1 FPA MODE

7(a) (b) 1.1 .l-1.1.2

8 1.1.3

9(a) (b) 1.1.4

10(a) (b) 1.1.5

1 l(a) (b) 2.1.1- 2.1.2

6.3 CAS/MACH MODE

2O(a) (b) 1.3.1 - 1.3.2

2 1 (a) (b) 1.3.3 - 1.3.4

22 1.3.5 - 1.3.6

23(a) (b) 2.3.1 - 2.3.2

24W (b) 2.3.3 - 2.3.4

25b) (b) 2.4.1 - 2.4.2

10000

10000

10000

10000

25000

25000

so00

5000

10000

10000

10000

10000

25000

5000

10000

10000

10000

24000

24000

24000

200

200

200

200

300

300

130

130

200

200

200

200

300

130

250

250

250

300

300

300

A Y i 2 + O

A Y +So

A Y -5'

A Y +1S0

A Y i2+O2.1.3 and 2.1.5 not plotted response as 2.1.1

A Y - S o

A Y i2+O3.1.3 and 3.1.5 not plotted response as 3.1.1 '

A Y -5'

A H i 300ft

A H +25OOft

A H +4OOOft

A H -2500ft -

A H i300, i2500f t

A H 2300, 22500ft

A V 220

A V f 5 0

A V i 1 0 0

A V 220

A V +SO

AM i.05

95

Figure Test Flight Conditions Alt (ft) Speed (kn) Remarks

26(a) (b) 2.4.3, 2.4.4 24000

27 3.3.1, 3.3.4 5000

28 2.7.2 20000

6.4 V-CWS MODE

29 30 31 32 33 34 35 36

1.5.1, 1.5.2 10000 1.5.3 10000 1.5.4 10000 1.5.5 10000 1.5.6 10000 2.5.1 20000 2.5.2 20000 2.5.4 20000

6.5 UNDERSPEED PROTECTION MODE

37b) (b) 1.6.1, 1.6.8 10000

38(a) (b) 1.6.9, 1.6.15 10000

39 3.5.1 10000

6.6 ENERGY EXCHANGE MANEUVERS

40 1.7.1 10500

41 1.7.1 (M) 10000

6.7 APPROACH AND GO-AROUND

44 3.6.1, 3.6.2 3000

45(d (b) 3.6.3, 3.6.4 3000

46 3.6.5 3000

300

150

330

200 200 200 200 200 300 300 300

200

200

200

250

250

300

300

150

150

150

48 3.6.7

AMf.15

AV f 20, t: 50

MACH to CAS Switch

A i 2 4 - O

A Y +So

A Y +loo A Y +15O A Y + 2 S 0

A Y -5 '

A Y -.So A Y -5'

Flaps: Oo - 40° (Slow Transition) Flaps: 40° - Oo (Slow Transltion) Flaps: Rapid Change

- 600ft, + 2Okts

- SOOft, + 2Okts

Simultaneous -5OOOft , -50kts

Climb Out + SOOOft + 75kts

ALT *GS *GA

High Speed Glideslope Capture

ALT SEL * ALT Hold a t End of Climb Out

Low Speed Glideslope Capture (Flaps 40, Gear Down Prior to Capture)

Go Around and Climb Out

GS Capture from above

V-CWS Approach

.

96

Appendix 2: Flight Test Conditions

Test No.

1.1 .l

1.1.2

1.1.3

1.1.4

1 .?.5

1.2.1

1.2.2

1.2.3

1.2.4

1.3.1

1 A 2

1.3.3

1.3.4

1.3.5

1.3.6

1.4.1

1.4.2

1.4-3

1.1.4

1.5.1

1 s . 2

1.5.3

WOOE

~~~

FPA

ALT

CAS

CAS

MACH

V-CWS

Initial Conditions ~~

SPEED [AS ( k t r )

200

200

250

200

300

0.5/271

200

ALT (ft)

10000

10000

10000

10000

10000

10000

10000

- GEAR

- UP

UP

UP

~~

FLAPS

~~

Parameter

Ay t 2.5

Ay - 2.5

Ay t 5.0

AT - 5.0

Ay t 15.0

Ad t 300 f t

AH - 300 f t

AH t 2500 f t

AH - 2500 f t

AV t 20 k t

AV - 20 k t

AV t 50 k t

AV - 50 k t

AV t 100 k t

AV - 100 k t

AI! t 0.05

AM - 0.05

AM 0.15

AI! - 0.15

Ay t 2.5

AT -2.5

Ay + 5.0

Description of Test Condi!im

Rrrponrr t o smll Jnd l r r g r FPA coarrrndr

T h r o t t l r r roach a f t l i m i t . TECS glvrr UP con t ro l l i ng FPA but u in t r l n r contro l o f r p r r d

T h r o t t l r r r tach forwJrd lldt, TECS g l v r r up con t ro l l i ng FPA but Dv ln t r l n r control o f r p r r d

Chock rp r rd ortor. maximum cllnb r a t r

Rrsponsr t o small and largo vrloclty (CAS) camrvnds

Obrrrvr nuxlmun helght e r ro r during speed chrnges

Rrrponse t o sm11 and large Mach c o m n d s

Examlnr V - M pr r fo rmncr u l t h r w l l rnd large column Inputs. Exanlnc g r n r r ~ l V-CWS per formncr I n smooth r l r and turbulencr (If tVJ l lab le)

97

Description of Test Condition

For small FPA changes. system controls FPA and speed

for la rger pos l t l ve comunds, the t h r o t t l e s w l l l l l m l t ; the system w l l l continue t o contro l t o FPA and speed v i11 decrerso. AOA protect lon w l l l prevent s t a l l . Mr in ta ln c m n d f o r s u f f i c i e n t tln t o d m n r t r r t e AOA protect ion.

Dia l down speed. Exmlne AOA p ro tec t l on a t f l a p sett lngs. 0 + 40.. Extend gear a t Flaps IS..

Zero n r t enrrgy exchange ( r r t o o f KE change equals tho r r t o of PE change) Exocuto r s double maneuver

Executr as doublo maneuver

Exomto as double maneuver

\.

ORfGlNAL PAGE IS OF KlOR Q U A W

Test No.

2.1.1

2.1.2

2.1.3

2.1.4

2.2.1

2.2.2

2.2.3

2.2.4

2.3.1

2 .3 .2

2.3.3

2.3.4

2.4.1

2.4.2

2.4.3

2.4.4

2.5.1

2.5.2

2.g.3

2.5.b

2.6.1

MODE

FPA

ALT

CAS

MACH

v-cws

vMAX

Initial Conditions

SPEED [AS (k ts)

325

325

300

250

M 0.65

300

CAS 300

ALT (ft)

25000

25000

25000

10000

25000

20000

- GEAR

- UP

UP

UP

UP

UP

-

FLAPS Parameter

by t 2.5.

Ay - 2.5.

by t 5.0.

Ay - 5.0'

AH t 300 f t

AH - 300 f t

AH t 2500

AH - 2500

AV t 20

AV - 20

AV t 50

AV - 50

AH .OS

w -.as w - .IS AM -.IS

Ay t 2 .5

AY - 2.5

Ay t 5.0

AY - 5.0

t A V 100 k t s


Rrsponsr t o small and largo FPA coanrnds

T h r o t t l r s rrach fONJrd l ia i t . TECS g iv rs up c o n t r o l l i n g FPA bu t maintains contro l o f sprrd T h r o t t l r r reach a f t l i a i t , TECS g iv rs up con t ro l l i ng FPA but maintains contro l o f sprrd

Rrsponrr t o smll and largo a l t l t u d r comnrnd

Chock sprmd o r r o t , l u x i a u n climb rat.

Obsrrvr nvximua h r i g h t rrror during sprrd changrs

Ex rm ln r V - C H prrforlarncr with small and largo c o l u m inputs. E r ra in r gen r r r l V-cwf performance i n smooth a i r and turbul rncr I f condit ions perai t . commands, tho t h r o t t l r s w i l l l i m i t ; tho system w i l l continuo t o contro l t o FPA and speed w i l l incrrasr. V,,

pro t r c t l on v i11 p r r v r n t ovrrsprrd.

For 1JrgO. n r g r t i v r

Oarcrnd from 25000 f t and chock oprrat lon of VW nod..

&st No.

2.7.1

2.7.2

3.1.1

3.1.2

3.1.3

3.1.4

3.1.5.

3.2.1

3.2.2

3.2.3

3.2.4

3.3.1

3.3.2

3.3.3

3.3.4

3.3.5

3.3.b

MOO€

A u t w t l c AS + MACH Sultchlng

Autonra t 1 c IACH + CAS Sui tc h l ng

FPA

ALT

CAS

Initial Conditions

SPEEO AS (kts)

330 ktt

M 0.7

'ref +

10 kts

'ret +

10 kts

"ref ' 10 kts

- ALT

( ft)

1 8000

-

20000

5000

5000

5000

- GEAR

- UP

UP

OOUN

DOWN

-

f LAPS

0

'

0

bo

40

Parameter

by + 2.5

AT - 2.5 by + 5.0

AT - 5.0

Ay + 15

AH + 300 ft

AH - 300 ft AH + 2500 ft

AH - 2500 ft AV + 20 kt

AV - 20 kt AV + 50 kt

AV - 50 kt AV + lo0 kt

AV - 100 kt


1. Fly CAS. ALT HOLD mode rt le000 ft.

2. Dial in 20000 ft and select ALT SEL.

3. Climb to 20000 ft. 4. CAS + MACH Autowtlc

sultching occurs at 330 kts. 0.7H and rpprox. 19000 ft.

1. Fly MACH. ALT H O U m o d e at 20000 ft.

2. Dial In 18000 ft and select ALT SEL.

3. Descend to 18000 ft. 4. Observe MACH 4 CAS

autwrtlc sultchlng.

Response to small and large FPA coararnds

Throttles rrach aft llolt, TECS gives up controlllng f P A but nrlntrins control of speed Throttles reach forward llrlt, TECS glves up controlling FPA but ulntalns control of speed

Response to snull and large ALT c-nds.

Check speed error, u x l w a climb rate

Response to small and largo CAS comnrnds.

Observe maximum height error durlng speed changes.

100

Test No.

3.4.1

3.4.2

3.4.3

3.4.4

3.5.1

3.5.2

3.6.1

3.6.2

3.6.3

3.6.4

3.6.5

3.6.6

3.6.7

3.6.8

MODE

MACH

AOA

AOA

ALT + &S HOLD

6.A.

ALT + 6S

6.A.

ALT SEL + ALT HOLD

ALT HOLD,-, 6s

6.A.

F PA

Initial Conditions

SPEED IAS (kts)

0.5/277

175

125

1 SO

1.3 vrei

250

l.' 'rot

''3 'rrf

',e3 'ret . .

' e 3 'ret

'e3 'ref

10000

5000

5000

3000

500

3000

500

-1 000

2000

500

1000

- GEAR

-

UP

QW

1

1

0

1

OWN

1

1

1

-

FLAPS

0

40

15

40

0

40

40

40

40

40

Parameter

an t 0.05

AH - 0.05

AM t 0.1s

AH - 0.15

SELECT ALTITUDE 2000 ft

FPA tS*


0111 doun speed. Examine AOA protcctlon during rapld flap rxtenslon (0 + 40.). Extend gear at flaps 15..

Loavo spred comnand at 125 kts. Examinr ADA protrction during rapid flap rrtraction ( 4 0 * 4 * ) . Rotract grar at flaps IS*.

Intrrcrpt glldrslopr from ATL HOLD moda at 150 kt. 3000 ft. Continuo drscent on glldrslope and extend flaps to 40.. Drcrrasr sprrd to 1.3 Vrrf.

Engagr 6.A. at 500 ft and clld out to 3000 ft. Autwtlcally engagr ALT HOLD at 3000 ft. Incrrasr speed to 250 kts during climb out. . Captutr glldrslope. Dial d o m sprrd to- 100 kts. Extend flaps uhlle flying dobm glldrslopr.

Engagr go around at 500 ft.

Engagr A l l SELECT and fly airplanr to approach poi 1 t lon.

Capturr glldrslopr at 1.3 Vrrf and fly d m to 500 ft altitudr.

At 500 ft altitudr rngagr go-around mode, cllnb out to 1000 ft.

Engagr FPA climb out to 3500 ft and fly airplrnr to approach position apprpximately 10 nn from runway threshold.

101

Test No.

3.6.9

3.1.1

3.1.2

MODE

FPA * G S

v-cus

V-CYS

Initial Conditions

S P E E D IAS (kts)

l s 3 'rat

150

150

ALT (ft)

3500 (-lorn range from thrrrh- old)

3000

500

- GEAR

1

1

1

FLAPS

40

15

15

102

Parameter

FPA -4.


Approach and crpturr g l id r r lopr from rbovr a t -4. FPA.

Approach and f l y doun glidoslopr using V-GUS modo.

&around a t 500 f t a l t i t u d r using V-CYS -0.

NASA B737 Flight Test Results of the Total Energy Control System

Documents

Transcript of NASA B737 Flight Test Results of the Total Energy Control System