NASA Technical Memorandum 4503

Preliminary Flight Results of a Fly-by-Throttle Emergency Flight Control System on an F-15 Airplane

June 1993

Frank W. Burcham, Jr., Trindel A. Maine, C. Gordon Fullerton,and Edward A. Wells

NASA Technical Memorandum 4503

Frank W. Burcham, Jr., Trindel A. Maine, C. Gordon Fullerton

Dryden Flight Research FacilityEdwards, California

Edward A. Wells

Mc Donnell Douglas Aerospace CompanySt. Louis, Missouri

National Aeronautics and Space Administration

Office of Management

Scientific and Technical Information Program

1993

Preliminary Flight Results of a Fly-by-Throttle Emergency Flight Control System on an F-15 Airplane

PRELIMINARY FLIGHT RESULTS OF A FLY-BY-THROTTLE EMERGENCY FLIGHT CONTROL SYSTEM

ON AN F-15 AIRPLANE

Frank W. Burcham Jr.*

Trindel A. Maine**

C. Gordon Fullerton† NASA Dryden Flight Research Facility

P.O. Box 273Edwards, California 93523-0273

Edward A. Wells‡

McDonnell Douglas Aerospace CompanySt. Louis, Missouri

Abstract

A multi-engine aircraft, with some or all of the flightcontrol system inoperative, may use engine thrust forcontrol. NASA Dryden has conducted a study of thecapability and techniques for this emergency flight con-trol method for the F-15 airplane. With an augmentedcontrol system, engine thrust, along with appropriatefeedback parameters, is used to control flightpath andbank angle. Extensive simulation studies have beenfollowed by flight tests. This paper discusses the princi-ples of throttles-only control, the F-15 airplane, theaugmented system, and the flight results including ac-tual landings with throttles-only control.

Nomenclature

CAS control augmentation system

CG center of gravity

DEEC digital electronic engine control

EMD engine model derivative

HIDEC Highly Integrated Digital Electronic Control

*Chief, Propulsion and Performance Branch. AIAA Asso-ciate Fellow.**Senior Aerospace Analyst. †Aerospace Research Pilot. ‡Senior Engineer.

Copyright 1993 by the American Institute of Aeronauticsand Astronautics, Inc. No copyright is asserted in the UnitedStates under Title 17, U.S. Code. The U.S. Government has aroyalty-free license to exercise all rights under the copyrightclaimed herein for Governmental purposes. All other rightsare reserved by the copyright owner.

HUD heads up display

PCA propulsion controlled aircraft

V airspeed, kts

α angle of attack, deg

γ flightpath angle, deg

φ bank angle, deg

Introduction

In an emergency situation, throttles can be used toaugment or replace aircraft flight control systems. Air-craft flight control systems are extremely reliable be-cause of the multiple control surfaces, hydraulic sys-tems, sensors, control computers, and control cablesused to achieve high levels of control system redun-dancy and reliability. However, there are rare occa-sions when potentially disastrous flight control systemfailures do occur. This is particularly true for mili-tary airplanes operating in a hostile environment. Atsuch times, any other form of flight control, includingpropulsion, would be welcome.

Some aircraft with multiple engines may be con-trolled to a rudimentary degree with the throttles.The use of differential thrust induces yaw and thenormal dihedral effect results in roll. Many trans-port airplanes exhibit nose-up pitching moments fromthrust that may be useful for pitch control. In addi-tion, most airplanes have positive speed stability (ifspeed is increased, the airplane will climb, and viceversa). Airplanes with total hydraulic system failureshave been flown for substantial periods with only en-gines for control.1 The following are examples of lossof hydraulic power:

• A B-747 aircraft lost its entire hydraulic systembecause of a pressure bulkhead failure. It wasflown for almost an hour using throttle control,but the crew were forced to learn by trial and er-ror, and the plane eventually hit a mountain.

• An uncontained engine failure on a DC-10 causedthe loss of all hydraulics. The crew used throttlesfor control under extremely difficult circumstancesand were able to execute an emergency crash land-ing at an airport, and many lives were saved.

In other cases hydraulic loss caused partial loss offlight controls:

• A C-5A cargo airplane had a major structural fail-ure that caused loss of all hydraulics to the tail.This airplane was flown for 1/2 hr with the throt-tles, but on a landing attempt, the airplane hit theground short of the runway, broke up, and allaboard were killed in the resulting fire.

• B-52 airplanes have experienced two cases of totalloss of hydraulic power to the rudder and eleva-tor. Thrust and wing spoilers were used for pitchcontrol. In one case, the crew tried to land theairplane, and hit very hard, breaking off the nosesection on impact. The rest of the airplane wasdestroyed by fire, but the entire crew survived. Inthe second instance, procedures developed as a re-sult of the first accident were used. The B-52 hada hard landing, but was repairable.

The NASA Dryden Flight Research Facility has beenconducting flight, ground simulator, and analyticalstudies to investigate the use of the propulsion systemfor emergency flight control. One objective, determin-ing the degree of control power available for variousclasses of airplanes, has shown a surprising amountof control capability for many airplanes. The secondobjective was to provide awareness of and techniquesfor manual throttles-only control.1 Airplanes studiedto date include the B-720, MD-11, F-15, B-727, T-38,Learjet, and B-747. The third objective is to investi-gate possible control modes that could be developedfor future airplanes.2

NASA Dryden and McDonnell Douglas AerospaceCompany (MDA, St. Louis, MO) developed an aug-mented control system for the F-15 which uses feedbackto provide throttle commands for emergency flight con-trol. An initial flight evaluation of this propulsion con-trolled aircraft (PCA) system has recently been flown.Comparisons of flight and simulation results of the F-15airplane flown with manual throttles-only control aregiven in ref. 3.

This paper reviews the principles of throttles-onlycontrol, the design of the PCA system, and preliminary

results of the first flight evaluation of a PCA system,including landings without the use of flight controls.Also presented is how the PCA system performs atconditions beyond the design envelope.

Principles of Throttles-Only Control

The principles of throttles-only flight control, pre-sented in refs. 1 and 3, will be reviewed here, usingexamples for the F-15 airplane.

Roll: Differential thrust generates sideslip, which,through dihedral effect, results in roll rate. Roll rate iscontrolled to establish a bank angle which results in aturn and change in aircraft heading. Full differentialthrust for the F-15 yields a roll rate of about 12 to15 deg/sec.

Pitch: Pitch control due to throttle changes is morecomplex. There are several effects that occur on theF-15.

1. Flightpath angle change due to speed stability. Allstable airplanes, including the F-15, exhibit pos-itive speed stability. For a short time (approxi-mately 15 sec), a thrust increase will cause a speedincrease, which will cause a lift increase whichcauses a pitch rate increase, and a climb (if al-lowed to continue, this effect will be oscillatory,see phugoid, no. 4). The degree of change to theflightpath angle is proportional to the differencebetween the initial trim airspeed and the currentairspeed, hence, the flightpath angle tends to in-crease as speed increases.

2. Pitching moment due to thrust line offset. If theengine thrust line does not pass through the centerof gravity (CG), there will be a pitching momentintroduced by thrust change. For many transportaircraft, the thrust line is below the CG, and in-creasing thrust results in a desirable nose-up pitch-ing moment, the magnitude being a linear functionof the thrust change. This is the desirable geom-etry for throttles-only control, because a thrustchange immediately starts the nose in the samedirection as that needed for the long-term flight-path angle change. The effect is more a functionof change in thrust than of change in speed, andoccurs near the time of the thrust increase. Forthe F-15, the thrust line passes within ±1 in. ofthe vertical CG, depending on fuel quantity, andthis effect is small.

3. Flightpath angle change due to the vertical compo-nent of thrust. If the thrust line is inclined to theflightpath, as is commonly the case, an increase inthrust will cause a direct increase in vertical ve-locity, i.e., rate of climb, and a resulting increase

2

in flightpath angle. For a given aircraft configu-ration, this effect will increase as angle of attackincreases (i.e., as speed decreases).

For the F-15, the combined effects of the en-gine thrust produce a nose-up pitching responseof about 2.5 deg/sec for a throttle step from trimto intermediate power on both engines.

4. Phugoid. The phugoid is the longitudinal long-period oscillation of an airplane. It is a motionin which kinetic and potential energy (speed andaltitude) are traded. The phugoid oscillation isexcited by a pitch, or velocity change, and willhave a period of approximately one minute, andmay or may not damp naturally. Properly sizedand timed throttle inputs can be used to dampunwanted phugoid oscillations.1

Speed Control: Once the flight control surfaces of anairplane are locked at a given position, the trim air-speed of most airplanes is only slightly affected by en-gine thrust. Retrimming to a different speed may beachieved by other techniques, such as variable stabi-lizer control, CG control, lowering of landing gear, andflaps, etc. In general, the speed must be reduced to anacceptable landing speed; this implies developing nose-up pitching moments. Methods for doing this includemoving the CG aft, lowering the flaps, and extendingthe landing gear. For the F-15, moving the air inlets tothe full-up emergency position reduces the trim speedby at least 20 kts, and lowering the flaps lowers thetrim speed by at least 30 kts.

Trim speed is also affected by changes in weight. Asweight is reduced (such as by burning fuel), the liftremains constant, so the airplane tends to climb. Tomaintain level flight, the throttle setting must be re-duced, which reduces speed. On the F-15, this effectreduces trim speed by approximately 1 kt every 2 min.

Stability: The flight controls-failed stability of an air-plane is also an important consideration for throttles-only control. Large transport airplanes typically havegood basic static stability. Yaw dampers may be usedfor increasing the dutch roll mode stability, but goodpitch, roll, and yaw static stability is usually builtin. This stability remains if the flight control systemshould be lost. For fighter airplanes, the airframe mayhave lower levels of static stability, with adequate sta-bility being achieved with mechanical and/or electronicstability augmentation. Thus in the case of flight con-trol system failure in a fighter, the basic stability maybe considerably reduced, and the control requirementsfor a PCA system will be more difficult. (The previ-ous comments do not apply to the long-term phugoidstability which will likely be a problem for both fighterand transport aircraft).

Airplane Description

The F-15 airplane (Fig. 1) is a high-performanceair superiority fighter airplane with a maximum Machcapability of 2.5. It has a high wing with 45° of sweepand twin vertical tails. The two afterburning turbofanengines are mounted close together (4.25 ft apart at thenozzles) in the aft fuselage. Air inlets for the enginesare located on the fuselage sides, ahead of the wings.

Engines

The NASA F-15 is powered by F100 engine modelderivative (EMD) engines, designated PW1128 bythe engine manufacturer. These engines feature a3-stage fan and a 9-stage compressor, each driven bya 2-stage turbine. A mixed flow augmentor exhauststhrough a variable-area convergent–divergent nozzle.The PW1128 is a derivative of the F100-PW-220 en-gine, and features an improved fan, higher turbine tem-perature capability, and a 15-segment augmentor.

The digital electronic engine control (DEEC) systemcontrols the F100 engine. Closed-loop control of enginepressure ratio and airflow is provided at intermediatepower and above. At lower power, fan rpm is controlledas a function of throttle angle. At low power settingswith the landing gear extended, the nozzle opens toreduce thrust. The DEEC transmits engine parame-ters in digital format to the data bus, and also receivesinputs for throttle commands on the data bus.

Because of the development nature of the PW1128engines used in the NASA F-15, the DEEC softwarehas some nonproduction effects, one of which is a slowdecay of thrust at low power settings. An engine modeldeveloped by MDA accurately represents the dynamicresponse of the PW1128 engines at the low-speed–low-altitude condition.

Inlets

The F-15 is equipped with variable-geometry 2-dimensional external compression horizontal ramp in-lets. Since these inlets are well forward and outboard ofthe aircraft CG, pitching, rolling, and yawing momentsare developed by the inlet aerodynamics as engine air-flow changes. Although these forces and moments aresmall in conventional flight, they become significantwhen the flight controls are locked. If hydraulic pres-sure to the actuators is lost, the inlet ramps both driftto a full-up position. This was the condition used for allF-15 PCA tests.

Flight Control System

The F-15 has a mechanical flight control system aug-mented by a high-authority electronic control augmen-tation system (CAS). Hydraulic power is required forall flight control surfaces. The NASA F-15 airplane

3

is equipped with a digital electronic flight control sys-tem (DEFCS) which replaced the standard F-15 analogelectronic CAS. In the flight control mode (CAS off,with the mechanical system pitch and roll ratio changemechanisms set in the “emergency” fixed position) theflight control system surfaces remain stationary as longas the stick and rudder pedals are not moved.

Instrumentation

The F-15 airplane was instrumented to measure theparameters required for the throttles-only flights. Alltypical engine and airplane parameters were measured.Data from individual sensors and data from the digitalcontrol system data buses (each engine and the digi-tal flight control system) were recorded on an onboardpulse code modulation system. Data were telemeteredto the ground for real-time recording, analysis, anddisplay.

The F-15 has a heads-up display (HUD) which pro-vides flight information such as airspeed and altitude. Avelocity vector symbol displays the precise flightpathrelative to the ground. A HUD video camera was pro-vided and the signal, along with the pilot’s microphone(hot mike) was also telemetered to the ground.

Propulsion Controlled Aircraft System

The PCA system features on the F-15 are shown inFig. 2. Much of the equipment used by the PCA systemwas previously installed on the NASA F-15 for otherintegrated control research as part of the Highly In-tegrated Digital Electronic Control (HIDEC) system,4

and included the digital flight control computer, thegeneral-purpose research digital computer, the F100EMD engines with DEECs, the cockpit HUD and con-trol system input–output, interface equipment to allowthese systems to communicate, the “emergency” flightcontrol mode, and the data system and tape recorder.The PCA system was implemented by adding onlythe attitude command thumbwheel controllers in thecockpit.

Figure 3 shows the F-15 HIDEC airplane cockpit,the PCA equipment, the thumbwheel controllers, theHUD, the navigation control interface (NCI), and theswitches and control panels from the PCA and HIDECsystems.

The PCA system was designed for a limited-envelopefirst flight evaluation of augmented throttles-only con-trol. It was designed to function at airspeeds between150 and 190 kts at altitudes below 10,000 ft. It wasassumed that the airplane would be trimmed to the de-sired test conditions prior to PCA system engagement.

Figure 4 is a block diagram of the PCA system struc-ture. All of this equipment except the thumbwheelcontroller panel was previously installed. The various

avionics and PCA units communicate with each othervia digital data buses. The logic for the PCA controllaws resides in the general-purpose research computer,in FORTRAN code. Digital inputs are received fromthe digital flight control system, the inertial navigationsystem (INS), the airdata computer, the digital enginecontrols, and from the pilot’s pitch and roll thumb-wheels. The PCA system sends throttle commandsto the internal DEEC throttle command logic withoutdriving the throttles in the cockpit. No commands aresent to the inlets during PCA operation.

Figure 5 is a block diagram of the PCA control laws.In the pitch axis, pilot thumbwheel command for flight-path angle is compared to the sensed flightpath an-gle, with flightpath angle rate and velocity also avail-able as feedbacks to assist in phugoid damping. Col-lective (equal) thrust commands are sent to both en-gines to obtain the commanded flightpath. The thumb-wheel flightpath command is displayed to the pilot onthe HUD with the box shown in Fig. 3.

In the roll axis, the pilot bank angle command iscompared with yaw rate, roll rate, and bank angle; dif-ferential thrust commands are issued to both enginesto obtain the commanded bank angle.

The pitch and roll axis control laws were devel-oped by MDA and Dryden using linear models, non-linear simulations, and finally in full nonlinear pilotedsimulations.

Variable gains, filters, multipliers, and gain sched-ules can be selected by the pilot, and are available atmost points within the PCA software. These featuresprovide a great deal of flexibility for testing.

Numerous automatic features were installed to dis-engage the PCA system in case of malfunction, if thepredefined limits were exceeded, or if the pilot movedthe stick or throttles.

F-15 Simulations

High-fidelity simulations of the PCA system in theF-15 airplane were available at NASA Dryden and atMDA. These simulations included nonlinear aerody-namics, control systems, and nonlinear engine modelsas well as the PCA logic. Pilot-in-the-loop simulatorswere used for closed-loop pilot evaluations. Batch ver-sions were used for open-loop system response and todevelop, evaluate, and test the PCA software. TheMDA simulators included a high-fidelity visual systemprojected onto a dome, whereas the Dryden simulationused a smaller monitor; both were adequate for thePCA evaluation. Linear models of the PCA systemwere also developed at MDA and at Dryden for controlsystem development and analysis. MDA tests includeda hardware-in-the-loop piloted simulation in which the

4

actual flight software and computers were part of thesimulation.

Test Conditions and Procedures

The F-15 PCA system was tested primarily in 2 con-figurations; 150 kts with flaps down, and 170 ktswith flaps up. Test altitudes ranged from 2300 ft (10 ftabove the runway) to 15,000 ft. The pilot config-ured the airplane with the CAS off, and pitch and rollratios and inlets in the “emergency” position, whichis the position that would occur if hydraulic pressurewere lost. The landing gear was lowered hydraulically,although it could have been lowered with the emer-gency pneumatic extension system. The flaps were low-ered with the electric actuation system installed on theNASA F-15.

The pilot trimmed the airplane to the desired testcondition with the stick, engaged the PCA system us-ing the “couple” button on the right throttle, and op-erated the thumbwheels with no inputs to the stick andthrottles. The flight controls remained active, but notused, as a safety feature. In some cases, the systemwas engaged 70 kts above its original design envelope.

Test techniques were developed to assess thethrottles-only control capability of the F-15 airplaneand simulation. Open-loop tests, including small con-trolled throttle steps were flown, and control capabilitywas compared to the simulation.

Small step commands in pitch and roll duringlevel flight were made when the augmented systemtests were initially conducted. Once these tests werecomplete, combinations of pitch and roll commandswere tested, and finally, simulated approaches weremade. Manual control techniques were also used forcomparison.

Results and Discussion

The F-15 PCA system was evaluated during a seriesof flights. The initial tests consisted of engaging thePCA system in level flight and observing the systemoperation. Later tests included response to step inputsand approaches to the runway.

Step Inputs

Numerous step thumbwheel command inputs havebeen made to flightpath and bank angle axes at varyingweights, airspeeds, and gain combinations. These stepinputs allow detailed postflight comparisons of actualflight performance with simulation predictions, and be-tween differing flight control configurations tested. Fig-ure 6 shows a response to a small negative flightpathangle command at 150 kts, with the flaps down. Theinitial throttle decrease is followed by throttle modu-lation to achieve the desired flightpath with minimum

overshoot. The average fan speed, a good approxima-tion of thrust, is also shown in Fig. 6. Approximately11 sec is required to achieve the 1.8°-decrease in flight-path angle. A comparison of the nonlinear simulationat this condition shows a slightly slower response, butreasonably good agreement with the flight data.

Roll response to a full roll step command is shownin Figs. 7(a) and 7(b). Roll control was initially poorbecause of low roll rate, as shown in Fig. 7(a), with28 sec required to achieve the commanded bank angle.Only a small differential throttle command was gener-ated by the control laws. This low roll rate was dic-tated by results from the hardware-in-the-loop simula-tion, in which higher gains caused a limit cycle oscilla-tion. Extensive flight evaluations were then conductedto improve roll performance. After several iterations,the roll response was greatly improved by changes ingains, yaw rate filtering, and adding bank angle feed-back as shown in Fig. 7(b), with the commanded bankangle being reached within 6 sec. A significant degreeof differential thrust was commanded in this test. Noevidence of the limit cycle oscillation was seen in theflight tests. Again, comparison to the nonlinear simu-lation prediction for this condition is reasonably good.

Runway Approaches

The PCA system was typically engaged on the down-wind leg of approaches to the Edwards runway. Turnswere made to the base leg, and onto final approachabout 5 miles out. Figure 8 shows the command andactual flightpath (glide slope) and bank angle valuesfor a low approach and PCA go-around at 150 kts. En-gine throttle settings, height above the ground, and air-speed are also shown. This approach showed good con-trol with very light turbulence. Flightpath was main-tained within approximately 1° of command until thego-around was initiated. Most of the throttle motionis differential to maintain the commanded bank angle.Bank angle lags pilot inputs by approximately 3 sec.At 100 ft above the ground, as planned, the pilot ini-tiated a go-around by moving the flightpath commandup to command a climb. The system response wasconsidered adequate by the pilot.

In a test to evaluate PCA response in ground ef-fect, the pilot flew with PCA control to within 10 ftof the runway. The pilot decoupled the system at thispoint as planned, and made only a minimal stick in-put in the remaining 2 sec until touchdown. Fig-ures 9(a) and 9(b) show a time history of this approach.Weather conditions included a 5-kt tailwind and verylight turbulence, with occasional small upsets causedby thermals. Figure 9(a) shows 83 sec of the ap-proach. Flightpath command varied between –1 and–2° for most of the approach, and flightpath was main-tained within 0.5° of the command, except when mild

5

thermal activity caused a pitchup at 23 sec and againat 60 sec. Bank angle commands were generally smalland bank angle was maintained, considering the 3-seclag, within 3 °. At 70 sec, the pilot increased the flight-path command to initiate a landing flare.

The last 6 sec prior to touchdown are shown inFig. 9(b). Flightpath gradually was reduced and bankangle remained small. As the F-15 entered ground ef-fect at about 15 ft above the ground, an increase inflightpath angle caused by increased wing lift was seen.At 10 ft, the pilot disengaged the PCA system andmade a small aft stick input (seen in the stabilizer po-sition data) to reduce the sink rate. At about 5 ft, anose-down pitch is seen in the pitch rate and angle ofattack data, because of the effect of ground effect onthe fixed horizontal tail. This reduction in angle of at-tack, which reduces lift, causes the flightpath to remainconstant for the last second. Touchdown rate of sinkwas about 4.5 ft/sec. The pilot made a larger aft stickinput at touchdown to control de-rotation. The pilotconsidered the system performance to be good on thisapproach.

A time history of the last 56 sec of the first PCAlanding is shown in Fig. 10. The conditions for thislanding included an 8-kt wind down the runway, andalmost no turbulence. The pilot reduced the flightpathcommand to 1° at an altitude of 200 ft, and to 0.4 ° at80 ft, resulting in a very shallow final approach. At analtitude of 20 ft, 6 sec before touchdown the groundeffect begins. With no flight control input to counterthe ground effect, the nose pitched down to –1.8 ° attouchdown, at which point the pilot made an aft stickinput to cushion the landing. The PCA system addedthrust and increased airspeed by 4 kts trying to counterthe pitchdown. Bank angle control and lineup was goodthroughout the final approach. A small correction tothe right was made just before touchdown. The bankangle at touchdown was –1 ° and the touchdown wasapproximately 8 ft to the left of the runway centerline.Following this landing, another approach was made. Inthis case, the control tower requested a 360 °-turn forspacing 6 miles from the runway. The pilot made thisturn and then continued the approach and landed, allwith PCA control.

Engagement at Unusual Attitude

Another test was to engage the PCA system after theairplane was maneuvered to unusual attitudes, such asmight occur with an actual loss of flight controls. Al-though the PCA system was not specifically designedto handle such conditions, simulation studies indicatedthat it could safely recover the F-15 from a range ofupsets. The most severe test (Fig. 11) was initiated at250 kts at 15,000 ft, with a 22°-nose down and 78°-banksituation. The pilot moved the inlets to the emergency

position and engaged the PCA system. The PCA sys-tem increased the right engine thrust to intermediatepower; the wings were rolled level within 15 sec, thepullout reached 3 g and 320 kts, with a loss of alti-tude of 3000 ft. Following the pullout, the airplaneentered a climb. With no pilot action and a zero bankcommand, the airspeed would have decayed to approx-imately 100 kts; in this case the pilot terminated thetest at an airspeed of 150 kts.

Hydraulic System Failure Simulation

Tests were also conducted to determine the trimspeed variations after a simulated hydraulic failure.Starting from 260 kts and level flight, the CAS wasturned off and the inlets were switched to emergencyas would occur with loss of hydraulic pressure. ThePCA system was engaged, and the new trim speed was200 kts. The flaps were then lowered electrically, andthe trim speed was reduced to 160 kts. Landing gearextension caused no change in trim speed. From thiscondition, fuel could be burned off to achieve a 150-ktapproach speed.

Concluding Remarks

The first flight evaluation of an augmented propul-sion controlled aircraft system on the F-15 airplane hasbeen conducted. An augmented throttles-only feed-back control system has been shown to provide stableoperation to step inputs and acceptable operation forlanding approaches and actual landings. The systemhas also been tested at conditions beyond its designenvelope, including engagement at unusual attitudesand at speeds 100 kts above approach speeds.

References

1Burcham, Frank W., Jr., and C. Gordon Fullerton,Controlling Crippled Aircraft–With Throttles, NASATM-104238, 1991.

2Gilyard, Glenn B., Joseph L. Conley, Jeanette L.Le, and Frank W. Burcham, Jr., “A Simulation Eval-uation of a Four-Engine Jet Transport Using EngineThrust Modulation for Flightpath Control,” AIAA-91-2223, June 1991.

3Burcham, Frank W., Jr., Trindel A Maine, andThomas Wolf, Flight Testing and Simulation of an F-15Airplane Using Throttles for Flight Control, NASATM-104255, 1992.

4Stewart, James F., Frank W. Burcham, Jr., andDonald G. Gatlin, Flight-Determined Benefits of Inte-grated Flight-Propulsion Control Systems, NASA TM-4393, 1992.

6

Fig. 1. NASA F-15 HIDEC flight research aircraft.

Fig. 2. PCA features on the F-15 airplane.

18.67 ft

63.75 ft

920241

42.83 ft

4.25 ft

F-15

Digital flight control computer

Digital electronic engine control (DEEC)

F100 EMD engines

HIDEC

Digital interface

HUD

General purpose research computer

"Emergency" flight control mode in which surfaces won't move

Thumbwheel panel

Cockpit input/ output and switches

Data system and recorder

930096

7

Fig. 3. F-15 PCA cockpit configuration.

8

Fig. 4. PCA hardware block diagram.

Fig. 5. Simplified block diagram of the F-15 PCA logic.

F-15 horizontal situation indicator

Digital electronic

engine control

SerialF-15

attitude direction indicator

Modified flight control

computer pitch

Air inlet controller

Data recorder

Data link transmitter

F-15 inertial

navigation set

HIDEC control panel

F-15 actuators

Flight control sensors

F-15 airdata

computer

Inlet actuators

DEFCS modified flight

control computer roll-yaw

NCI panel select PCA

modes

Interface unit

Vehicle management

system computer

Air inlet control sensors

Heads-up display

HUD

Instrumen- tation

system

Central computer

General purpose

computer

H009 bus

Software modified for PCA Added for PCA 930097

1553 bus

Thumbwheel control panel

Flightpath angle

command� Flightpath angle rate�

Calibrated airspeed�

+�

+�

Collective throttle command�

Flightpath angle thumbwheel command�

+�

– Weight�

Mil�

idle�

+�

–

Bank angle –

Bank angle command�

Gain�

Flightpath angle

Gain�

Gain�

Gain� Filter� Filter�

Gain� Filter�

+�V� Filter�

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Differential thrust command�

Lead-lag filter�Weight�

+�

+�Yaw rate

Roll rate sin(α)�

cos(α)�

Filter�

Filter�

V�

930098

9

Fig. 6. Response of the F-15 PCA system to a flightpath angle step from 0 to –1.8°.

47

45

44

43

42

Averagepower

lever angle,deg

7700

7600

7500

7400

Thumbwheel command

Simulation

0

–.5

–1.0

–1.5

–2.0

–2.5

Averagefan speed,

rev/min

Flightpathangle,

deg

0 10 20 30 40

Time, sec

46

Measured

930099

Measured

Measured

10

(a) Initial bank angle control logic.

(b) Improved bank angle control logic.

Fig. 7. Effect of a step bank angle command on bank angle, 150 kts, flaps down.

Bankangle,

deg

25

15

5

–5

Differentialthrust,

lb

1000

500

0

–5000

Time, sec

5 10 15 20 25 30 35

930100

Thumbwheel command Simulation

Measured

Measured

Bankangle,

deg

30

20

10

0

Differentialthrust,

lb

4000

–10000

Time, sec

5 10 15 20 30 35

930101

3000

2000

1000

0

Measured

SimulationThumbwheel

command

25

11

Fig. 8. Time history of a PCA approach and go-around.

0

200

400

600

146

148

150

152

154

2

0

5

0 20 40 60 80

Go-aroundinitiated

0

–2

–5

–1050

45

40

35

Time, sec930102

Radaraltitude,

ft

Airspeed,kts

Flightpathangle,deg

Bankangle,deg

Throttleangle,deg

800

Thumbwheel command

Measured Thumbwheel command

Measured

Left

Right

12

(a) 83 sec of landing approach.

Fig. 9. Time history of F-15 PCA landing approach, flaps down.

Time, sec

Airspeed,kts

180

170

160

150

1

0

–1

–2

–3

Flightpathangle,

deg

10

5

0

–5

–10

Bankangle,deg

0 10 20 30 40 50 60 70 80 9020

40

60

80

Powerleverangle,deg

PCA disengage Touchdown

0

200

400

600

Radarheightof main

gear,ft

930103

Left

Thumbwheelcommand Measured

Thumbwheelcommand

Measured

Right

13

(b) Last 6 sec of landing approach.

Fig. 9. Concluded.

Time, sec

Angle of attack,

deg

10

8

6

4

1

0

–1

–2

Flightpath angle,

deg

Stabilizer position,

deg

78 80 82 84 86 887

9 x 103

Pitch rate, deg/sec

20

30

40

50

Radar height of main

gear, ft

930104

10

0

1

–1

–3

4

0

–4

8Fan speed,

rev/min

PCA disengaged Touchdown Struts compressed

Measured

Right

Thumbwheel flightpath command

Left

14

Fig. 10. Time history of PCA landing.

300

200

100

0

Radar altitude,

ft

50

40

30

Power lever angle, deg

0 10 20 30 40 50 60Time, sec

2

–2

–6

Average stabilator position,

deg

160

155

150

145

Airspeed, kts

1

0

–1

–2

Flightpath angle, deg

10

5

0

–5

–10

Bank angle,

degThumbwheel

command

Thumbwheel command

Measured

Measured

Touchdown

Right

Left

930222

15

Fig. 11. Time history of F-15 PCA engagement at unusual attitude.

Time, sec

Airspeed, kts

300

250

200

150

–40

Flightpath angle,

deg

0

–50

Bank angle, deg

0 5 10 15 20 25 30 35

20

40

60

80

Power lever angle, deg

15

16 x 103

Altitude, ft

930105

PCA engaged Pilot takeover

100

50

–20

0

20

40

100

350

14

13

12

Mil

Idle

Right

Left

16

REPORT DOCUMENTATION PAGE

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NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18298-102

Preliminary Flight Results of a Fly-by-Throttle Emergency FlightControl System on an F-15 Airplane

RTOP 533-02-34

Frank W. Burcham, Jr., Trindel A. Maine, C. Gordon Fullerton, andEdward A. Wells

NASA Dryden Flight Research FacilityP.O. Box 273Edwards, California 93523-0273

H-1911

National Aeronautics and Space AdministrationWashington, DC 20546-0001 NASA TM-4503

A multi-engine aircraft, with some or all of the flight control system inoperative, may use engine thrust for control.NASA Dryden has conducted a study of the capability and techniques for this emergency flight control method forthe F-15 airplane. With an augmented control system, engine thrust, along with appropriate feedback parameters, isused to control flightpath and bank angle. Extensive simulation studies have been followed by flight tests. This paperdiscusses the principles of throttles-only control, the F-15 airplane, the augmented system, and the flight resultsincluding actual landings with throttles-only control.

F-15 airplane, flight control, flight-propulsion control integration, propulsioncontrol, throttles-only control

AO2

16

Unclassified Unclassified Unclassified Unlimited

June 1993 Technical Memorandum

Available from the NASA Center for AeroSpace Information, 800 Elkridge Landing Road, Linthicum Heights, MD 21090; (301)621-0390

Presented as AIAA 93-1820 at the 29th AIAA/SAE/ASME Joint Propulsion Conference, Monterey, CA,June 28–30, 1993.

Unclassified—UnlimitedSubject Category 08