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NASA Technical Memorandum 4745

Development and Flight Test

of an Augmented Thrust-Only

Flight Control System on an

MD- 11 Transport Airplane

Frank W. Burcham, Jr., Trindel A. Maine,

John J. Burken, and Drew Pappas

July 1996

NASA Technical Memorandum 4745

Development and Flight Test of

an Augmented Thrust-Only

Flight Control System on an

MD- 11 Transport Airplane

Frank W. Burcham, Jr., Trindel A. Maine, and John J. Burken

Dryden Flight Research Center

Edwards, California

Drew Pappas

McDonnell Douglas Aerospace

Long Beach, California

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1996

DEVELOPMENT AND FLIGHT TEST OF AN AUGMENTED

THRUST-ONLY FLIGHT CONTROL SYSTEM ON AN

MD-11 TRANSPORT AIRPLANE

Frank W. Burcham Jr.,'

Trindel A. Maine*, John J. Burken*

NASA Dryden Flight Research Center

Edwards, California

Drew Pappas*

McDonnell Douglas Aerospace

Long Beach, California

Abstract

An emergency flight control system using only engine AFS

thrust, called Propulsion-Controlled Aircraft (PCA), has AGL

been developed and flight tested on an MD- 11 airplane.

In this thrust-only control system, pilot flightpath and c.g.

track commands and aircraft feedback parameters are

used to control the throttles. The PCA system was LSAS

installed on the MD-ll airplane using software EPR

modifications to existing computers. Flight test results

show that the PCA system can be used to fly to an FADEC

airport and safely land a transport airplane with an FCC

inoperative flight control system. In up-and-away

operation, the PCA system served as an acceptable FCP

autopilot capable of extended flight over a range of FDS

speeds and altitudes. The PCA approaches, go-arounds, FPA

and three landings without the use of any normal flight

controls have been demonstrated, including instrument GW

landing system-coupled hands-off landings. The PCA HDG

operation was used to recover from an upset condition.

In addition, PCA was tested at altitude with all three IAS

hydraulic systems turned off. This paper reviews the ILS

principles of throttles-only flight control; describes the

MD- 11 airplane and systems; and discusses PCA system MCDU

development, operation, flight testing, and pilot ND

comments. PCA

PFD

TRK

"Chief, Propulsion Branch. Associate Fellow AIAA V/S

tAerospace Engineer

"_PCAProgram Manager

Copyright © 1996 by the American Institute of Aeronautics and

Astronautics. Inc. No copyright is asserted in the United States under

Title 17,U.S. Code. The U.S. Government has a royalty-free licenseto

exercise all rights under the copyright claimed herein for

Governmental purposes. All other rights are reserved by the copyright

owner.

Nomenclature

autoflight system

above ground level, ft

center of gravity, percent of mean aerodynamic

chord (fig. 1)

longitudinal stability augmentation system

engine pressure ratio

full-authority digital engine control

flight control computer

flight control panel on cockpit glareshield

flight deck simulator

flightpath angle, deg

gross weight, lb

heading

indicated airspeed, kn

instrument landing system

multifunction control and display unit

navigation display

propulsion-controlled aircraft

primary flight display

track angle, deg

vertical speed, ft/min

Introduction

In the last 20 years, pilots of B-747, B-52 (The

Boeing Company, Seattle, Washington); L-1011

(Lockheed Martin, Palmdale, California); C-5A

(Lockheed Martin, Marietta, Georgia); and DC-10
(McDonnell DouglasAerospace(MDA),LongBeach,

California)aircrafthaveexperiencedmajor flight

controlsystemfailuresandhavehadtousethrottlesfor

emergencyflightcontrol.In themajorityof cases,a

crashresulted;theB-747,DC-10,andC-5Acrashes

claimedover1200lives,l

Toinvestigatetheuseof enginethrustforemergency

flight control,the NationalAeronauticsandSpace

Administration,Dryden Flight ResearchCenter

(DFRC),Edwards, California,conductedflight,ground

simulator,andanalytical studies.Oneobjectivewasto

determinethedegreeof controlavailablewithmanual

manipulationof enginethrottlesfor variousclassesof

airplanes.Testsin simulationhaveincludedB-720,

B-747,B-727,MD-11,C-402,C-17,F-18,andF-15

airplanes.In addition,B-747,B-777,MD-11,T-39,

Lear24,F-18,F-15,T-38,andPA-30airplaneswere

flighttested.Thepilotsuseddifferential throttlecontrol

to generatesideslip,whichthroughthedihedraleffect

resultsinroll.Symmetricthrottleinputswerealsoused

tocontrolflightpath. Thisobjectivehasshownsufficient

controlcapabilityfor all testedairplanesto maintain

grosscontrol;flightpathand trackanglemaybe

controlledto withina fewdegrees. Forall airplanes

tested,it isexceedinglydifficulttomakeasaferunway

landingusingmanualthrottles-onlycontrol 2becauseof

problemsin controllingthephugoidanddutchroll

modes,slow engineresponse,and weakcontrol

moments.

Toprovidesafelandingcapability,DFRCengineers

andpilotsconceivedanddevelopedasystemthatuses

only augmentedenginethrust,calledPropulsionControlledAircraft(PCA).A PCAsystemusespilot

flightpathinputs and airplanesensorfeedback

parametersto provideappropriateenginethrust

commandsfor emergencyflight control. This

augmentedsystemwasfirst evaluatedon a B-720

transport airplanesimulation. 3

Later,simulationstudiesand flight testswere

conductedonanF-15airplaneto investigatethrottlesonly controPand to investigatethe in-flight

performanceof anaugmentedPCAsystem. _ThePCA

systemflighthardwareandsoftwareweredeveloped

and implementedby MDA. Flighttestingincluded

landingsusingPCAcontrol.-_ThePCAtechnologywas

alsosuccessfullyevaluatedusinga simulationof a

conceptual megatransport. 6Anothermajorsimulation

studyhasbeenconductedat NASAAmesResearch

Center(ARC),MoffettField,California,usingthe

advancedconceptsflightsimulatc:.Over,100 simulated

PCA landings have been flown by more than 20 pilots]

and a PCA system has been evaluated on a B-747

simulation.

As a follow-on to the F-15 PCA research, MDA and

DFRC developed and tested a concept demonstration

PCA system for the MD-11 transport airplane. An

intermediate objective was to demonstrate controlled

up-and-away flight over a flight envelope from 150 to

250kn below 15,000 ft. Another was to make low

approaches to a runway that could have been continued

to a survivable landing. The goal was to make actual

PCA landings. To make the PCA demonstration apply to

the large number of transports with two wing-mounted

engines, the MD-11 PCA system primarily used only

the wing engines. The PCA system used only software

changes to existing digital systems on the MD-I 1. In

over 30 hr of testing, the PCA system exceeded the

objectives, served as a very acceptable autopilot, and

performed landings without using any flight controls. _

Later tests studied PCA operation over the full flight

envelope, in upset conditions, with the hydraulics turned

off, and coupled to an instrument landing system (ILS)

for hands-off landings. Sixteen pilots flew PCA

demonstration flights. Analysis of the control system

design and performance is given in reference 9. The

longitudinal control details are given in reference 10.

Details of the hardware architecture, software

development, and testing are given in reference 11.

This paper describes the MD-I1 airplane, principles

of throttles-only flight control, and throttles-only control

modes. The PCA system implementation, operation,

development, and test are discussed. In addition, control

laws, pilot comments, flight test results, and

recommended future PCA applications are presented.

Airplane Description

The MD-II airplane is a large, long-range, threeengine, wide-body transport. This 202 ft long airplane

has a wingspan of 170 ft and a maximum takeoff gross

weight of 618,000 lb (fig. 1). The test airplane, ship 560,

was configured for flight test and had no interior

furnishings.

Flight Control Systems

The MD-11 airplane has a mechanical flight control

system with irreversible hydraulically powered

actuators. Three systems provide hydraulic power.

Essential control functions may be maintained on

any one of these three systems. Dual elevators provide

pitch control. A hydraulic jackscrew-actuated horizontal
Jii

12° _ /__ in.

_ _ 59ft 2 In

26 ft 10 in. __ _

9_-__Mean aerodynamic chord

25% c.g.--_ ","7 ////-/57 ft 9 in.

202 ft "'I

960203

(a) Three-view drawing.

(b) Photograph.

Figure 1. The MD- 11 airplane.

EC 95-43247-4

stabilizer provides pitch trim. Inboard and outboard

ailerons supplemented by wing spoilers provide roll

control. Dual rudders provide yaw control. Hydraulic

actuated leading- and trailing-edge flaps are provided. In

case of hydraulic system failure, the stabilizer would

effectively lock, the flaps would drift up, and other

surfaces would float.

The standard MD-11 airplane is equipped with an

advanced cockpit. Figure2(a) shows the electronic

displays on the main panel, glareshield control panel,

and automated systems management located on the

overhead panel. The mechanical flight control system of

an MD-I 1 airplane has longitudinal stability augmented

by a longitudinal stability augmentation system (LSAS)
and a dual yaw damper. The autoflight system (AFS)

includes autopilot, autothrottle and speed control, stall

warning, turn coordination, flap limiting, windshear

detection, and other features. A flight management

system integrates the autopilot navigation, and autoland

functions. The crew communicates with the flight

management system using the multifunction control and

display units (MCDU) located on the center pedestal.

Figure 2(b) shows the autoflight system crew

interface and display. The glareshield flight control

panel (FCP) includes a thumbwheel for selecting

flightpath angle (FPA) or vertical speed (V/S). A

heading or track knob is used to select the desired

heading or track. The primary flight display (PFD)

shows the selected and actual flightpath, heading, or

track. These data are shown on the navigation display

(ND) and on the PFD. A speed selection knob normally

commands the autothrottle speed and Mach number

hold.

E_agia_

The MD-I 1 airplane tested was powered by three

high-bypass ratio Pratt & Whitney (East Hartford,

Connecticut) 4460 turbofan engines in the 60,000-1b

thrust class. Two engines are mounted in underwing

pods, 10ft below and 26 ft 10in. outboard of the

nominal center of gravity (c.g.). The third engine is

located at the base of the vertical tail 20 ft above the

vertical c.g. and inclined 2.5 ° nose-up. These engines

have full-authority digital engine control (FADEC)

systems. The crew normally controls the engines with

electronic throttles which command a power setting

based on engine pressure ratio (EPR). The EPR varies

from just below 1.0 at idle to approximately 1.6 at

maximum power; thus, each 0.1 EPR is approximately

10,000 lb of thrust.

The FADEC systems normally accept small

(approximately + 5 percent) EPR trim commands from

the flight management system to closely maintain

engine limits or thrust settings and eliminate the need

for throttle stagger to match engines to a given EPR. As

is typical for high-bypass turbofans, thrust response near

idle power is initially slow. Once thrust reaches

approximately 20 percent, the thrust response improves

dramatically. In the midthrust range, thrust has a time

constant of approximately 0.5 sec. Thrust decay also

shows similar nonlinear effects.

Fuel System

The MD-11 airplane is equipped with fuel tanks

located in the wings, center fuselage, and horizontal tail.

Maximum fuel quantity is 259,000 lb. Fuel may be

and display unit

(a) Cockpit.

Figure 2. The MD-I 1 cockpit and autoflight system.

4
Heeding/track knob

Speed ,mbwheel

AFS override

peddle switches

Autoflight button

Flight control panel

- Roll window )itch window

Flightpath

command

bar

Flightpath

command

Primary flight display

Navigation display

(b) Autoflight system crew interface and display.

Figure 2. Concluded.

960205
transferred to the tail tank, which can move the c.g. aft

up to 10 percent from nominal. Fuel may be dumped

down to a total of 40,000 lb.

Weight and Center of Gravity

With a full payload of 122,700 lb, the MD-I 1 airplane

has zero-fuel weight of 400,000 lb. Maximum gross

weight (GW) is 618,0001b, and maximum landing

weight is 430,000 lb. The longitudinal c.g. is given in

percent mean aerodynamic chord (fig. 1). At cruise

altitude, the fuel management system transfers fuel to

the tail tank that maintains the c.g. at an airline

selectable aft limit of approximately 32 percent. For

takeoff and landing, the c.g. is normally approximately

20 to 24 percent. The forward limit is 12 percent.

Landing Gear

The MD-11 airplane is equipped with a steerable

nosewheel and three main landing gears. The gear may

be lowered with an alternate system independent of the

hydraulic system. Limited braking may be obtained

with inoperative hydraulic systems using brake

accumulators.

In_trumcn_atiqn

The test MD-11 airplane was equipped with a large

data acquisition system that can display and record

several thousand parameters from the digital data buses.

This system included airdata, engine, and inertial

parameters as well as flight control and PCA systems

parameters. A real-time plotting and hardcopy capability

was also available. Cameras were installed in the

cockpit to record the pilot's out-the-window view.

Equipment was also installed to record communications

and pilot comments.

MD- 11 Simulation

The MD-I 1 flight deck simulator (FDS) used for the

majority of the MD-11 PCA research was a highfidelity, fixed-based, piloted simulator with a great deal

of actual flight hardware in the cockpit, systems, and

electronics. This simulation included a projection digital

display system capable of displaying Long Beach,

Edwards, and other Southern California visual scenes.

Scenes from Yuma, Arizona, can also be displayed. In

addition, an off-line version of the full MD-11

simulation was hosted on an IBM mainframe computer.

This nonlinear simulation was used for control system

development and evaluation.

For flight control system and engine control system

integration testing, the MD-II bench simulation was

used. This simulation allowed actual PCA software in

Flight Control Computers (FCC) and one FADEC to

operate through the data buses and used a database

similar to the FDS. Pilot inputs could be simulated with

a joystick interface.

Prilaciples of Thr0ttles-Only Flight Control

This section describes the principles of throttles-only

flight control. Examples are for the MD- I 1 airplane.

-Lal.e, ml=I_,r, li_z_

Differential thrust generates sideslip which, through

the dihedral effect, results in the airplane rolling to a

desired bank angle which results in a turn and change in

aircraft heading. Figure 3 shows an open-loop throttle

step response for the MD- 11 airplane at 220 kn. The 10°

throttle split results in approximately 20,0001b of

differential thrust and a roll rate averaging 1.5 deg/sec.

Note that the EPR lags the throttle by approximately

1 sec, and the roll rate lags the yaw rate. A lightly

damped dutch roll mode is excited by this throttle step.

Full differential thrust for the MD- 11 airplane at a speed

of 150kn yields a peak roll rate of approximately

8 deg/sec.

Pitch control caused by throttle changes is more

complex than lateral--directional control. Several effects

occur. These effects include flightpath angle changes

caused by speed stability; pitching moment resulting

from thrust line offset, and flightpath angle changes

because of vertical component thrust and phugoid

oscillation.

Flightpath Angle Change Caused by Speed Stability

The majority of airplanes, including the MD-11,

exhibit positive speed stability. Over a short period

(approximately 10 sec), a thrust increase will cause a

speed increase. In turn, the speed increase causes a lift

increase. When the lift becomes greater than the weight,

the airplane will climb. The long-term effect is

oscillatory. (See Phugoid subsection.)

Pitching Moment Resulting from Thrust Line Offset

If the engine thrust line does not pass through the

vertical c.g., a pitching moment will be introduced by

thrust change. For many transport aircraft, the thrust line

is below the c.g., and increasing thrust results in a

desirable nose-up pitching moment. For the MD-II

airplane, if the three engines are used equally, the
Bank

angle,

deg

r Differential throttle step input

30

20 ....................

10 ....................

0

oll rate i

Roll and

yaw rate,

deg/sec

,.................. p .....

#,1"

. . -f

J

.5

0

Angle of

sideslip, - .5

deg

- 1.0

Throttle

angle,

deg

Engine

pressure

ratio

1.5

60 .................

5O

1.4

1.3

1.2

1.1

s Left

.... ._- Right

- i

(3 10 20

Time, sec

960206

Figure 3. The MD-11 flight response to differential throttle step inputs, PCA off, 220 kn, flaps up, gear down, yaw

dampers off, and center engine idle.
resultant thrust line is near the vertical c.g., and this

effect is small. The wing engines are 10 ft below the

vertical c.g. If only the wing engines are used, the noseup pitching moment is significant. The center engine of

the MD-11 is 20 ft above the vertical c.g. and causes a

strong nosedown pitching moment with thrust increase.

Flightpath Angle Change Because of the Vertical

Component of Thrust

If the thrust line is inclined to the flightpath, as is

commonly the case, an increase in thrust will increase

the vertical component of thrust, which will cause a

vertical acceleration and a resulting increase in

flightpath angle. For a given airplane configuration, this

effect will increase as angle of attack increases.

For the MD-11 airplane, the combined short-term

effect of a thrust increase is to produce a nose-up

flightpath response. Figure 4 shows a time history of the

step throttle increase of the wing engines at 220 kn.

Thrust responds within approximately 1 sec. Pitch

attitude and the resulting angle of attack increases

approximately 0.3 °, and airspeed increases for the first

10 sec.

When the wing-mounted engines are used for control,

pitching moment because of thrust offset is the strongest

component. In addition, a throttle advance increases

pitch attitude and angle of attack such that the long-term

effect is a climb at reduced trim airspeed. The converse

is also true. A reduction in wing engine thrust causes a

descent at increased speed.

The phugoid, or longitudinal long period oscillation

of an airplane, may be excited by a pitch, thrust,

or velocity change. For the MD- 11 airplane, the phugoid

is lightly damped. Figure 5 shows a flight example.

This phugoid oscillation, with the same airplane

configuration and flight conditions as those shown in

figure 4, was excited by a pullup which results in a

lightly damped oscillatory climb. Note that the average

trim airspeed is reduced from 220 to 210 kn, and the

average angle of attack is increased and is out of phase

with the airspeed. With throttles fixed, a slight

oscillation in EPR occurred because of the changing

flight conditions. Properly sized and timed throttle

inputs can be used to damp unwanted phugoid

oscillations?

Relative Position of Inlet to Exhaust Nozzle

The relative positions of the inlet and the exhaust

nozzle of each engine may have an important effect for

throttles-only flight control. The ram drag vector acts

through the centroid of the inlet area, along the

flightpath, and thus rotates with respect to the airplane

geometric reference system as angles of attack and

sideslip change. The gross thrust vector usually acts

along the engine centerline and, thus, maintains its

relationship to the airplane geometric reference system.

Ram drag can be a significant percentage of gross thrust,

particularly at low power settings where it may

approach the magnitude of the gross thrust.

In the pitch axis, having the inlet located above the

engine centerline is beneficial. An increase in throttle,

which will increase ram drag and gross thrust, will

result in a nose-up moment. This is the case for the B-2

airplane and for the center engines of the B-727 and

L-1011 aircraft. If the inlet is located below the engine

centerline, an increase in thrust causes an undesirable

nosedown moment. The F-16 and F-18 aircraft are

examples of such a configuration. Podded engines

typically have the inlet and nozzle closely aligned, and a

neutral effect results.

In the yaw axis, the principles are similar. The

desirable geometry has the engine nozzles outboard of

the inlets, so an increase in thrust would result in a

favorable yawing moment. Unfortunately, this nozzle

placement is not the case for many fighter airplanes that

have the inlets outboard of the engines. For the MD- 11

airplane, with podded engines, these inlet and nozzle

effects are small.

Trim Speed Control

Once the normal flight control surfaces of an airplane

are locked at a given position, the trim airspeed of the

majority of airplanes is only slightly affected by engine

thrust. In general, the speed will need to be reduced to

an acceptable landing speed. This reduction requires

developing nose-up pitching moments. Methods include

moving the c.g. aft, lowering flaps, increasing the thrust

of low-mounted engines, decreasing the thrust of highmounted engines, or burning off or dumping fuel.

Extending the landing gear often decreases trim speed

because it requires an increase in engine thrust.

On the MD-I 1 airplane, several ways to control trim

speed exist. Flaps and stabilizer trim require hydraulic

power which may not always be available. The center

engine may be used as a moment-generating device to

change angle of attack and, hence, to control the speed.

Increasing center engine thrust has a strong nosedown

pitch effect and increases the trim speed. Starting with

all throttles equal, increasing thrust on the wing engines
225

Throttle Increase

I

Airspeed,

kn

22O

215

210

4

3

Fllghtpath 2

angle,

deg 1

0

-1

I

I ' i

Rate of

climb,

ft/sec

3O

2O

10

0

-10

I

t

............. j .............. , .............. .......... : .............

6.0

Angle of

attack,

deg

I

1.4 I -- - ------ -- -- _- Left

1.3.............. I.............. :.............. i.............. _.............

L_Ri_g

1.2 I

Engine

pressure

ratio

Throttle

angle,

deg

I

7O I

t SI i ._'Right

_o................. .-:...:._L.....! ............. i............

.... i '. i

50

0 10 20 3O 40

Time, sec 960207

Figure 4. Flight response of MD-11 airplane to step throttle increase, PCA off, LSAS off, flaps up, gear down, and

center engine idle.
Altitude,

ft

Flightpeth

angle,

deg

Airspeed,

kn

Angle

of attack,

deg

13,500

13,000

12,500

12,000

4

2

0

-2

220

215

210

205

200

5.5

5.0

4.5

4.0

Phugoid initiated

by pullup

........... Angle

of attack

Measured

" ',.................................. --:--- Faired "

Engine

pressure

ratio

1.3

1.2

1.1

1.0

........... ! .................... _- Right ...............

0 25 50 75 100 125

Time, sec

960208

Figure 5. Time history of MD- 11 phugoid, LSAS off, gear down, flaps up, light turbulence, fixed throttles, center

engine idle, initiated by an elevator pull-up and release.

10
and decreasing thrust on the tail engine to idle thrust

reduces speed by 20 to 30 kn. Lowering the landing gear

with the alternate gear extension system reduces speed

an additional 15 to 17 kn because of the increased thrust

required for the wing engines to maintain level flight.

This approach leaves the wells open and the landing

gear doors exposed.

Trim speed is also affected by changes in weight. As

weight is reduced (for example, by burning or dumping

fuel), the lift remains constant, so the airplane tends to

climb. It is assumed that the c.g. remains constant. To

maintain level flight, the throttle setting must be reduced

to reduce speed until lift and weight are again in

balance. On the MD-11 airplane, flying at low speed,

this effect reduces trim speed by approximately 1 kn for

every 3000 ib of fuel consumed (approximately every

12 min).

Figure 6 shows the MD-11 trim airspeed as a function

of stabilizer setting and c.g. position. With these and

other data over a range of GW obtained from the

nonlinear simulation, determining the stabilizer-fixed

speed range available as a function of GW and c.g. is

possible. Fuel transfer either to or from the tail tank or

between the main tanks provides speed change of up to

50 kn. Note that at the forward c.g., such as 12 percent,

the airplane has strong speed stability (steep slope of

stabilizer as a function of speed); whereas at aft c.g.,

such as 34 percent, speed stability is reduced by at least

50 percent.

Speed Effects on Propulsive Control Power

The propulsive forces (differential thrust for lateral

control and collective thrust for flightpath control) tend

to be relatively independent of speed. On the other hand,

the aerodynamic restoring forces that resist the

propulsive forces are proportional to the dynamic

pressure, which is a function of speed squared. This

relationship results in the propulsive control power

being approximately inversely proportional to the

square of the speed. 2

Surface Float with Hydraulics T_rned Off

When the hydraulic system fails, a surface will float to

the zero hinge moment condition. For the MD- 11 hinge

geometry, this position is essentially the trail position.

Simulator studies on the MD- 11 airplane indicate that a

total hydraulic failure would cause the ailerons to float

with the trailing-edge up. The amount depends on

speed, thus reducing lift and increasing trim airspeed.

Rudder float would have a negligible effect on trim

speed but would reduce directional stability. Elevators

are usually trimmed to nearly zero force; hence, elevator

float would have a small effect. The stabilizer is

trimmed with a jackscrew actuator which, in case of

Trailing 2 Center of

edge gravity,

down 1 -- "...... "...... _..... _..... _ .................... ".................... percent

0 ......... ' 34

_ _ - _..... - - -. 30

Stabilizer

......... ....... _._. .... ' ..... : ..

,," , . J

-4 j_.._.7..' ...... :...;_I. .... ..'_< ' _ .......................

" i i 77" 7 _/'J '

-5 .............. _........ -L:

:/ : / /

-6 ....... ./i ...... ,Z..... i..... /

Trailing /"/ •V I

edge - 8

up 140 160 180 200 220 240 260 280 300 320 340 360

Calibrated airspeed,kn 96o2o9

Figure 6. The MD-11 trim stabilizer position for a range of speed and c.g. positions, 560,000-1b GW, gear down, flaps

up, 5000-ft altitude.

11

American Instituteof Aeronautics and Astronautics
hydraulic system failure, remains fixed because of

friction.

Throttles-Only Control Modes

This section describes the two throttles-only control

modes: manual throttles only and PCA control. These

modes were tested in simulation and in-flight.

Manu_l ThrottlcPOnly Control

For these tests, the crew turned off the LSAS, yaw

dampers, speed protect system, and fuel transfer system;

trimmed the airplane; released the flight controls; and

used only the throttles for flight control. In this mode,

the airplane behaved much like it would with a total

flight control system failure. Then, pilots manually

moved the throttles for control.

Propulsion-Controlled Aircraft Control Logic

In the PCA system, closed-loop control of engine

thrust is provided to satisfy pilot-commanded flightpath

and ground track. Figure 7 shows a simplified block

diagram of the PCA control laws. For the two-engine

PCA system, the center engine is used manually by the

pilot as a low-frequency speed trimmer. The lateral axis

shows only the track mode, but a bank angle control

mode was also available.

In the lateral axis, differential thrust is controlled to

provide precise track angle control (fig. 7(a)). The pilot's

autopilot track knob command (selectable to integer

degrees) is compared with the sensed track angle,

generating a track error. This error is converted into a

bank angle command and compared to the sensed bank

angle, generating a bank angle error with the maximum

bank angle limited to + 20 °. There is bank angle, yaw

rate, and roll rate feedback for stabilization and dutch

roll damping; differential thrust commands are issued to

the wing engines to obtain the commanded track.

Normal MD-11 bank angle limits could be selected on

the heading or track knob, up to a limit of_+20 °. The

pilot's track command is displayed by a bug on the

compass and a vertical bar on the PFD (fig. 8) as well as

on the ND cursor.

Longitudinal control is designed to provide precise

flightpath control. In the pitch axis (fig. 7(b)), the pilot

input through the autopilot thumbwheel command for

flightpath angle is limited, compared to the sensed

flightpath angle, and limited again. This command is

selectable to a tenth of a degree. A low-gain integral

path is provided in pitch to eliminate steady-state errors

and pitch angle rate feedback to assist in phugoid

damping. Collective (equal) thrust commands are sent to

the wing engines to obtain the commanded flightpath.

The thumbwheel flightpath command selected on the

FCP is displayed to the pilot on the PFD using the

existing flightpath command bar (fig. 8). The PCA

flightpath commands were limited from 10 ° to-10 ° and

to errors to + 3 °.

Logic is provided to prevent the wing engines from

being driven to very low power settings where the

Commanded

heading/track t.. . i_ I Lelt I

angle -_ Commanded /'m.egra. t°r II _._,. I I enoine I-_

\ bank angle _ I'1switc .nlng N-'_ r] I oain I

\ II ,oglc I1--_--II I- I

Pilot \ Bank \ I' ' +,.L

input / _+_ _ \ ÷_| + ,,_ Ia°"/pitchl I_terall +

)-e- 1 / _ _-_ __r{ _ priority t-I thrust _ )

Track _ W I "Z I "-11 l°gic II gain I _(

knob - ? Command - ? _ ' I I - T

I ,,,,,,t I Sensed bank angle _ , I

I I Sensed bank rate _ Feedback I ]

I ] Sensed yaw rate o_ network ] F --]

• IRi tl

Sen.d Sensed t_2 J

heading/track bank

angle angle

(a) Lateral control system (track mode only].

Figure 7. The MD- 11 PCA block diagram (center engine speed control loop omitted).

Left engine

roll thrust

command

Right engine

roll thrust

command

960210

12
Commanded

flightpath

angle -_

Flightpath Command

angle limit

thumbwheel E

Sensed

FPA

Ilnte( "atorI _ F-- e_

lil ,it I-I _ I-t G, _ L

Io llC I _

+ r

FPA _ + f _ IRoll/pltch I _ +/,.._

_'_ _1 Gain _ )--_ priority M Gain

Error ' ' ""/ I logic I _ _-

limit Sensed pitch angle O-_ Feedback L

Sensed pitch rate _ network / m

(b) Longitudinal control system.

Figure 7. Concluded.

LeR engine

pitchthrust

command

_.ht'_ Right engine

!he _ pitch thrust

command

960211

Flight control panel

®

®

®

®

Number 2 AFS override paddle down to enable PCA

(in FCC number 1)

Autoflight pushed to engage PCA, see empty

Speed window, AP1

Heading or track button pushed to select track for

lateral mode, turn knob to select, pull to engage,

shown commanding 224 ° track, track command

shown as the vertical bars.

V/S/FPA pushed to select the FPA vertical mode,

to engage, move FPA thumbwheel command.

Shown at - 2.8 °.

Primary flight display 960212

Figure 8. Flight control panel and primary flight display showing PCA flightpath and track mode selection.

13
engine response is very slow. In addition, integrator

wind-up protection is provided. The priority logic also

prevents saturation of pitch or roll commands.

Variable gains, filters, multipliers, gain schedules, and

modes (selectable by the flight test engineer using the

MCDU) were available at the majority of points within

the PCA software, providing a great deal of flexibility

for testing. More detailed design and analysis of the

PCA control laws are discussed in reference 9. The

speed control loop and augmented longitudinal control

loop that uses the center engine is discussed in

reference 10.

Implementation

The PCA logic resides in one of the two FCC. The

FCC provides a host of functions, including autopilot,

autothrottle, navigation, and flight management system.

Honeywell (Phoenix, Arizona) developed the code for

the control laws, interfaced them to existing sensor

signals, and sent commands to the engine FADEC

computers over an ARINC 429 data bus.

Pratt & Whitney modified the FADEC to accept a

full-authority EPR command instead of the normal

+ 5-percent EPR commands. For this initial flight test, a

PCA disengage capability in case the throttles were

moved was also incorporated. This capability instantly

gives the pilots normal throttle control, if needed. The

FCC 1 contained the PCA software; FCC2 was

unmodified. The PCA system implementation is

discussed in detail in reference 11.

Operation

The PCA system is activated by selecting FCC 1 and

disabling FCC 2 by lowering the AFS override paddle

switch, and pushing the autofligbt button on the FCP

(fig. 8). The PCA engagement is indicated by the AP 1

(autopilot 1) indication and the empty box on the PFD

speed window. Longitudinal control is commanded

through the FPA thumbwheel. Lateral control is effected

through the track knob. In the track mode, turning the

knob selects the track, and pulling the knob sends the

command to the PCA software. The PCA system may

be disconnected by several methods. Each is indicated

by a red flashing box on the PFD and an aurally

annunciated autopilot warning.

Development

The PCA system was developed over a period of

3 years, beginning with manual throttles-only tests in

the FDS. Although very difficult at first, a rapid learning

curve developed. With some experience and a very high

workload, gross control could be achieved. That is, bank

angle, heading, and flightpath could be controlled within

a few degrees. Because of the low-mounted wing

engines, a pilot using only the wing engines could

simply increase thrust slightly to increase flightpath and

reduce thrust slightly to decrease flightpath. Phugoid

damping could also be achieved using this technique;

for example, add thrust during the descent and reduce

thrust during the climb. The center engine could be used

as a trimming device to control speed. Increasing center

engine thrust has a strong nosedown pitch effect and

increases the trim speed.

Differential thrust was effective in inducing sideslip,

which resulted in the aircraft rolling. At speeds in excess

of 170 kn, the dutch roll mode was adequately damped.

In addition, bank angle could be controlled reasonably

well. Below 170 kn, dutch roll damping decreased, and

bank angle control was difficult. At the lower speeds,

using the throttles to attempt to damp the dutch roll was

quite difficult. Although gross control was certainly

possible and improved with practice; a manual throttlesonly landing on a runway was very difficult.

Based on the success of the F-15 PCA development

and the gross control capability of manual throttlesonly-control in the FDS, MDA performed a study of a

PCA system for the MD-I 1 airplane. A linear model

based on the FDS was developed and used to design a

PCA control law. This control law worked well on the

linear simulation. When tried on the FDS, however,

performance was poor. The problems were primarily

caused by the nonlinear engine response characteristics.

Later, a nonlinear model was developed for control law

analysis and development. Initially, the center engine

was used along with the wing engines; however, to

simplify the initial control design, only the wing engines

were used.

The PCA system design and development studies

were conducted in 1993. The prime contractor, MDA,

established subcontracts with Honeywell, Pratt &

Whitney, and General Electric (Evendale, Ohio). The

preliminary design was completed in fall 1994. A test

airplane with Pratt & Whitney engines was selected, and

the final design was complete in March 1995. MDA

conducted integration tests in the bench simulation to

verify proper communication over the data buses as well

as proper engagement, annunciation, operation, and

disengagement, jl A ground test on the MD-I 1 airplane

verified end-to-end operation, including verification that

PCA commands could be sent to the FADEC and that

the engines would respond properly.

14
Flight Tests

The PCA flight test program began by assessing the

operational characteristics at 10,000ft in smooth air.

The evaluation consisted of turning off the LSAS and

yaw dampers, trimming the airplane, releasing the

controls, and engaging the PCA system. Two

configurations were tested: a clean condition at

approximately 220 kn and a takeoff configuration with

28 ° flaps and gear extended at approximately 175 kn.

At altitude, pitch and track control were good. The

pilots were impressed that the PCA system performed

almost as well as the normal autopilot. This system

controlled flightpath to a few tenths of a degree. With a

level flightpath command, the system also held altitude

to within + 20 ft.

Figure9 shows a flight response to a series of

flightpath steps and a comparison to the FDS. The pilot

first selected a -2 ° flightpath step. Both engine EPR

decreased sharply because of the flightpath error, then

almost immediately began to increase because of the

pitch rate feedback. The 2° command was reached in

approximately 7sec, then overshot approximately

25 percent. Angle of attack followed EPR closely, and

airspeed varied inversely as it did in the open-loop

throttle steps. These FDS data show more angle-ofattack change and less flightpath overshoot than the

flight data.

Lateral control was evaluated at 10,000 ft with step

inputs in track angle command. Track was controlled to

within 1°. In addition, track captures showed no

overshoot, but initial response was slow. Next, the PCA

system was tested at 5000 ft. Then, the test included

approaches to the runway in turbulent air.

Pitch control remained very good, track control was

sluggish. On approaches it was difficult to anticipate the

sluggish track response. Some gain changes improved

the initial lateral control response. Data in reference 9

show the original and improved track response. The

time for a 5° track change was reduced from 22 to

17 sec; maximum bank angle increased from 4 ° to 6°.

Figure 10 shows a time history of a large track

command with the improved control gains and with the

gear down and flaps at 28 °. The pilot commanded a right

80 ° turn; the engine differential thrust response resulted

in a maximum roll rate of 3.5 deg/sec. The 20 ° bank

limit was reached in approximately 10 sec. Flightpath

angle dropped approximately 0.5 ° but corrected back

with a loss of only 30 ft. Once stabilized in the 20 °

bank, the EPR increased from 1.20 to 1.25. Airspeed

increased approximately 7kn, and turn rate was

approximately 4 deg/sec. In the rollout, flightpath was

again well controlled.

With the improved track response, approaches to the

runway were made. After establishing the airplane

configuration with the gear down, a takeoff flap setting

of 28 °, and the stabilizer trimmed for an approach speed

of approximately 170kn, the flight controls were

released and not touched, thus simulating a flight

control failure. The PCA system was activated. The

pilots usually set the PCA flightpath command to the

desired glideslope and then spent the majority of their

attention making track corrections to achieve and

maintain runway alignment. In the summer turbulence

at Yuma, Arizona, on their first few approaches, judging

the inputs required to maintain runway alignment was

difficult.

The MD-11 airplane was flown to Edwards AFB,

California, for initial PCA landings where there is a

15,000 ft long and 300 ft wide runway. Three PCA low

approaches and go-arounds were flown. Continuous

light turbulence existed, and occasional upsets from

thermals occurred. However, PCA performance was

judged adequate to proceed to PCA landings. On the

first intended landing initial lineup and flightpath

control was good (fig. 11). Based on simulation

experience, the pilot selected a flightpath of-1 ° at 140 ft

above ground level (AGL). The flightpath overshot.

Then at 30 ft AGL, the sink rate was increasing to

8 ft/sec, so the safety pilot, as briefed, made a small

nose-up elevator input, then allowed the airplane to

touch down under PCA control. The touchdown was

25 ft left of the runway centerline and 5000 ft from the

threshold at a sink rate of 4.5 ft/sec. The MD-II

airplane was stopped using reverse thrust and brakes,

but spoilers or nosewheel steering were not used.

The second landing used a slightly different flightpath

control technique. 9 The PCA pilot made small track

changes to maintain runway lineup and set the flightpath

command at -1.9 ° for the initial part of the approach.

Airspeed was 175 kn. At 200 ft AGL, based on the

experience with the first landing, the pilot shallowed the

flightpath to -i ° and at lOOft to -0.5 °. The airplane

touched down smoothly on the centerline at a 4 ft/sec

sink rate 3000 ft from the threshold with no inputs from

the safety pilot. The thermal updraft upset the airplane

bank angle to 9° at 100 ft AGL; the PCA track mode

corrected without pilot input. The airplane was stopped

using reverse thrust and light braking, but no flight

control inputs were required.

15
Flightpath

angle,

deg

2

1

0

-1

-2

-3

-- Flight

...... Commend

Simulation

............................... :.... i"""11 :......... i"_......... :

i i , , i

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

V ' 'i '

180 . _-. :

s ,J _--

.... I- _

Airspeed, 170

kn

160

,'_ f-,,,. 5.O

. ,/ ,_ l _ •

I : i / i\/..M \.

Angle l l [" "_ ' A " I" I_ ..LJ_ f 1 --v-'-.--_,.

4.0 --- I.. t- ........ '..r ._....... " ..... o,...c,, ..... i..........

3.5 ..... _ .............................................

3,0 '

1.5

Left engine

pressure 1.3

ratlo

1.1

.................... _ ................ _''_"t _'_- J' .,,;.,_ " t ;/ \ : "-_ - ;-

__" " _f :'".: - [ "S< -_: .... V[_z....... i :........

20 40 60 80 1O0 120 140

Time, sec 960213

Figure 9. The PCA response to thumbwheel flightpath step commands, gear down, flaps 28% flight and flight deck

simulator, 10,000-ft altitude, and center engine idle.

16
Track,

deg

s°°° I _

Altitude, - : : " ' :_ " I

4800

1

Flightpath

angle, 0 ....

deg

-1

Track command from 80°to 160 °

Flight data

Measured

--- Command

175

Airspeed, 170

kn

165

20

Bank 10

angle,

deg

2

Roll rate,

deg/sec

0

-2

1.4

Engine

pressure 1.3

ratio

1.2

1.1

i i "" i i

_!----4 ' i

0 20 40 60 80 1 O0

Time, sec 96o214

Figure 10. Time history of MD-l 1 PCA track change of 80 °, gear down, 28 ° flaps, and center engine idle.

17
Flightpath - 1

angle,

deg - 2

-3

226

Magnetic 225

track,

deg 224

223

Bank

angle, 0

deg

-5

6

Angle

of attack, 4

deg

0

5

Elevator

position, 0

deg

-5

1.3

1.2

1.1

1.0

.8

Engine

pressure

ratio

Flight data

Measured

.... Command

Touchdown, 4.5 ft/sec

25 ft left of centerllne

............. : ............. , ............. , " " 1_-- - ....

' ' i J ,

h

ill

Safety pilot

__ elevator input

Thrust reversing

Left

.... Right

m m Center

0 25 50 75 I00

Time, sec

125

960215

Figure 11. Time history of PCA landing, 28 ° flaps, and 180 kn.

18
From the two landings in light turbulence, it was

observed that PCA generally controlled track and pitch

to within + 0.5 ° of command (disregarding the 1° bias in

the track command). The EPR variations were normally

approximately +0.1; a 0.4 EPR differential thrust was

used to correct for the thermal upset. Ground effect was

similar to that seen in the simulator.

The only significant problem encountered in PCA

testing to this point was the sluggish and difficult to

predict lateral control on approaches in turbulence.

Pilots found that three or four approaches were required

before adequate lineups were consistently achieved.

Figure 12 shows the fourth approach of a PCA pilot,

flown with a 1.5° rudder offset to test the PCA system

robustness. The approach was flown at Yuma, Arizona,

and the pilot achieved runway lineup just inside 5 miles.

Continuous light-to-moderate turbulence occurred

because of thermal activity, but the pilot was still able to

maintain adequate runway alignment and glideslope

down to 200 ft AGL. At this point, the planned PCA goaround was initiated. Go-around performance was good.

Approximately 60 ft were lost before achieving a

positive rate of climb. Note the + 2-kn airspeed changes

from the turbulence, the increased average airspeed of

almost 10 kn on the glideslope, and the extensive EPR

activity to maintain control in the turbulence.

Envelope Expansion

Following successful completion of the PCA

landings, it was decided to test PCA operation beyond

the original design envelope. Figure 13 summarizes the

MD-11 design and flight test envelope. Based on

simulation results, PCA performance was expected to

degrade as altitude increased because of reduced engine

response and degrading dutch roll characteristics. The

PCA was tested at middle and aft c.g., at altitudes from

200 to 30,000 ft mean sea level, at speeds from 160 to

360 kn, with a 1.5° rudder offset, with a 7° aileron

offset, and with the hydraulics turned off. Approaches

were flown with a takeoff flaps setting of 28 °, with slatsonly, and without flaps or slats. The center engine was

used for speed control. The PCA engagement in mild

upset conditions was tested.

The third FADEC was installed on the center engine,

providing a low-frequency speed control. It worked very

well; trim speed could be changed 30 kn above and

below the initial trim speed. Pitch control was

maintained within 1°. _°

The PCA system tests were conducted at 15,000,

20,000, and 30,000 ft. Contrary to predictions from the

simulation, operation was good; pitch control was

essentially unchanged at these altitudes. Lateral control

was stable as well.

For aft c.g., at 30,000 ft, the c.g. was shifted from 24

to 32 percent, and step response tests were conducted.

Response was good. The PCA system was tested at

maximum airspeeds of 360 kn and Mach 0.83 with no

degradation in performance.

The ability to recover from upsets was tested. With

the PCA system turned off, the airplane was upset using

the normal controls. When the desired conditions were

achieved, the pilot released the controls and activated

the PCA system. All recoveries were successful; the

most severe recovery occurred from a 45 ° bank and a 7°

dive. When engaged, the PCA logic tries to hold the

commanded flightpath and track at the time of

engagement. In the upsets, the track deviates several

degrees during the recovery, so after flightpath has been

stabilized, the airplane returned to the reference track

command.

Figure 14 shows a time history of tests in which all

three hydraulic systems were turned off. Before the time

shown, all hydraulic systems were turned off for several

minutes. The outboard ailerons had floated to 12°, and

the inboard ailerons had floated to 5° . Meanwhile, the

elevators had floated near zero. Contrary to simulation

results, the floating surfaces resulted in a nose-up pitch

and caused a reduced trim speed. It was planned to

lower the landing gear, which was expected to lower the

trim speed further.

Because airspeed was already near the minimum

speed for flaps-up flight, the hydraulic system 3 was

turned on as shown. The ailerons returned to the trim

position, causing a pitchdown. Then, the stabilizer was

retrimmed to an increased speed position. When the

system 3 hydraulics were again turned off, the ailerons

again foated up, and the pitchup again occurred. The

PCA system reduced the throttles to idle to reestablish

commanded flightpath, and speed stabilized at 212 kn.

After a small track change, the landing gear was

lowered using the alternate gear extension system. With

the gear doors down and the wheel wells remaining

open, speed decreased to 195 kn. Note that lowering the

gear moved the aileron and elevator slightly. The

hydraulic pressure was zero, but of course there was still

fluid in the system.

The PCA maintained flightpath well in this transient.

Track and pitch control at this point were normal and

identical to that with the hydraulics systems turned on.

19
Altitude,

ft AGL

Airspeed,

kn

Flightpath

angle,

deg

190

185

180

175

I ............... ]

Flight data

Measured

Command

2

0

-2

-4

Magnetic

track,

deg

40

3O

2O

10

0

........ jI

!

!.... i ................ i.......

P

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

20

Bank

angle, 0

deg

-20

I_A ........... i_ii ...... ------ Right

! i !/ _- Center

1.3 i ''.'_ .... _: -'i ....

Engine _ __}]_

pressure

ratio 1.1

.9

0 50 100 150

Time, sec 960216

Figure 12. The MD-11 PCA system approach and go-around with 1.5 ° rudder offset, gear down, 28 ° flaps, continuous

light-to-moderate turbulence and thermal upsets. This was the pilot's fourth PCA approach.

20
30 x 103 24 and 32 percent c.g.

PCA flight

test envelope

20 360 kn

Altitude,

ft

PCA

h_

10

:s2 and 3

turned off

0 I I

100 200 300 400

Airspeed, kn

_,PCA landings PCA approaches with no flaps, slats only, and rudder offset

• PCA ILS-coupled autolands Demonstration to 16 pilots; airlines, DoD, FAA, end industry

960217

Figure 13. The MD-11 PCA design and flight test envelope.

21
Retrim

12

Surface 8

position, 4

deg 0

-4

19

Altitude,

ft xl0- 3 17

15

4

Flightpath 2

angle, 0

deg - 2

-4

100

Magnetic 360

track,

deg 260

160

20

10

0

Bank

angle,

deg

Airspeed,

kn

Engine

pressure

ratio

. 1 and 2 hydraulics off 4

.... 3hy(J ............................... Main landing Cenlt_-i-[dter an ng....... 3-h-d'y

gear extension gear extended on

Aileron

-- Outboard

........ Inboard

Elevator

......... Outboard

........ Inboard

--- Stabilizer

120 ° turn

Flight data

-- Measured

Command

240

220

200

180

1.5

1.3

-- 212

Speed chang e

200 400 600 800

Time, sec

960218

Figure 14. Time history of MD- 11 PCA operation with hydraulic systems turned off.

Left

.... Right

m__ Center

22
A simulated landing approach was made using a -2.5 °

glideslope with a small track change. A large track

change was commanded at 560sec. Bank angle

stabilized at its 20 ° limit, and flightpath was stepped

back to level flight. The center main landing gear was

then extended, resulting in another small change in

surface positions. The PCA control remained normal.

Hydraulics systems were then turned back on still under

PCA control.

The PCA operation as an autopilot was sufficiently

good that at times, PCA remained engaged for 1 hr or

more. Overall in the 30 hr flight program, over 25 hr of

PCA operation were logged. No unplanned PCA

disengagements occurred.

Simulation-To-Flight Comparison

The flight test results generally compared well with

the flight deck simulation. The pitch control in-flight is

somewhat less damped than in the simulator, possibly

because the vertical c.g. in the test airplane is lower than

in the simulation. The somewhat difficult lateral control

effects on approach were well-modeled. Anticipated

lateral control instability at high altitudes, based on

simulation, were not found in-flight. The trim speed

change caused by turning off the hydraulics systems was

the right magnitude but the wrong sign.

Pilot Comments

Pilot comments were recorded on the cockpit video

immediately after each approach. The project pilot

reported that pitch control was excellent and that all of

the workload was in the lateral task. In turbulent air,

anticipating the lead needed for track changes was

difficult. In addition, getting established on centerline

well out and using only small inputs was necessary.

After the PCA landings, the project pilot was really

impressed that control was so good and commented that

not only a survivable landing but a normal landing could

be made using PCA control. Other pilots also found that

the lateral control was sufficiently sluggish that

considerable practice was needed to learn how to

maintain runway lineup.

Instrument Landing System-Coupled Approach

It was obvious that the lateral control was sufficiently

sluggish that a pilot could not be assured of making a

successful approach on the first try with a significant

level of turbulence. An ILS-coupled PCA mode had

been checked out on the PCA simulation at DFRC and

ARC 7 and had worked very well. As a result, the ILScoupled mode was incorporated into the MD-II PCA

system. This modification was accomplished by

Honeywell. Existing MD-11 ILS-coupled logic and

displays were used as much as possible. Because this

modification was for concept demonstration only, no

attempt was to be made to optimize the ILS-coupled

system. Requirements were to capture the ILS localizer

first within a 30 ° intercept. The glide slope was to be

captured from below. The glideslope to be flown was

adjustable from 0.5 to 1.5 dots below the glideslope.

(One dot is 0.35°.) A simple two-step autoflare

capability was also added. All ILS approaches were

successful, including two actual hands-off landings.

Figure 15 shows a time history of an ILS-coupled

landing. Before the time shown, the pilot had selected a

track to intercept the localizer in level flight

approximately 12 miles from the runway and pressed

the approach-and-landing (appr/iand) button (fig. 8) to

arm the ILS capture. At localizer capture, the system

rolled the MD-11 airplane onto the localizer with one

overshoot and maintained a slight descent rate until

glideslope capture at 10miles out. The glideslope

capture resulted in an overshoot. Once established on

the ILS, localizer tracking was excellent, and glideslope

tracking was adequate. The preprogrammed flares at

130 and 30 ft, arrested the sink rate and resulted in a

touchdown 1 ft from the centerline, at 5 ft/sec sink rate,

and at a point 1200 ft beyond the threshold. After

touchdown, the PCA system was deactivated by moving

the throttles to idle. The nose lowered, and the

nosewheel touched down smoothly. The pilot used

differential braking to slow and maintain directional

control. Another ILS-coupled landing was made with a

touchdown sink rate of less than 2 ft/sec and on the

centerline. 9

These ILS-coupled approaches and landings were

made under nearly ideal weather conditions, but

simulator and computer analysis showed that safe

landings could be made in turbulence levels up to

moderate conditions. The ability of the coupled system

to make small thrust changes to correct a deviation

immediately provided a major improvement in

capability over pilot-in-the-loop control as well as a

huge reduction in pilot workload.

D¢monstration Flights

The PCA system was demonstrated to 16pilots

representing airlines, U.S. Air Force, U.S. Navy,

industry, and Federal Aviation Administration. After a

simulator session, each pilot flew manual throttles-only

control, then flew with the PCA system using the FCP

knobs. Each also made an approach to a virtual runway

23
Radar

altitude,

ft AGL

GIIdeslope

deviation,

deg

Flightpath

angle,

deg

Airspeed,

kn

Locallzer

deviation,

dots

Magnetic

track,

deg

Bank

angle,

deg

Engine

pressure

ratio

Localizer capture

3O0O

2O00

1000

0

.75

.50

.25

0

Touchdown

q ..................... __

Flightdata

Glidealope

2

1 Dots

0

Ground track

Glideelope capture

or

-2

-4

170

160

150

1

0

1

2

240

220

200

180

20

10

0

-10

1.3

1.2

1.1

1.0

.9

AP,_ __ ^ ^^_ _ '..... ^. -_^;__

0 I O0 200 300

Time, sec

960219

Left

.... Right

_- Center

Figure 15. Time history of PCA ILS-coupled approach and hands-off landing, 28 ° flaps.

24
100 ft AGL. The majority of pilots used the ILS-coupled

mode. A few used PCA flightpath and track control. In

the very smooth air, even the PCA approaches using the

FCP knobs were successful. All pilots were positively

impressed with the PCA system.

Future Applications

Based on the success of these MD-11 airplane and

previous F-15 PCA flight tests, a PCA system appears to

provide an acceptable backup flight control system

capable of safe landings independent of hydraulic

power. The addition of a slow electric actuator for

trimming the stabilizer for trim speed control would

eliminate the need to plan and use c.g. and weight

control for controlling speed.

Concluding Remarks

A Propulsion-Controlled Aircraft (PCA) system using

closed-loop thrust for control has been designed,

developed, and tested on an MD-11 airplane. This

system was implemented with software changes to

existing flight and engine control computers and uses

the autopilot knobs on the glareshield flight control

panel. Flight tests demonstrated that PCA can

successfully provide control adequate for up-and-away

flight and for runway landings. The PCA instrument

landing system-coupled mode was used to make handsoff landings. The PCA system was tested over a wide

flight envelope and found to perform somewhat better

than predicted. Flight-to-simulation comparisons were

generally reasonably good. Sixteen demonstration pilots

flew the PCA system, and pilot comments were very

favorable.

References

_Frank W. Burcham, Jr., Trindel Maine, C. Gordon

Fullerton, and L. Dean Webb, Development and Flight

Evaluation of an Emergency Digital Flight Control

System Using Only Engine Thrust on an F-15

Airplane, NASA TP-3627, 1996.

2Burcham, Frank W., Jr., and C. Gordon Fullerton,

Controlling Crippled Aircraft- With Throttles, NASA

TM- 104238, 1991.

3Gilyard, Glenn B., Joseph L. Conley, Jeanette L. Le,

and Frank W. Burcham, Jr., "A Simulation Evaluation of

a Four-Engine Jet Transport Using Engine Thrust

Modulation for Flightpath Control," AIAA 91-2223,

June 1991. (See also NASA TM-4324, 1991.)

4Burcham, Frank W., Jr., Trindel Maine, and Thomas

Wolf, Flight Testing and Simulation of an F-15 Airplane

Using Throttles for Flight Control, NASA TM-104255,

1992.

5Propulsion Controlled Aircraft, NASA TM- 104278,

1995. (See also internet address: http://www.dfrc.nasa.

gov/dryden/html.)

6Gerren, Donna S., Design, Analysis and Control of a

Large Transport Aircraft Using Selective Engine Thrust

as a Back-up for the Primary Flight Control, NASA

CR-186035, 1995.

7Bull, John, et al., Piloted Simulation Tests of

Propulsion Control as Backup for Loss of Primar).'

Flight Controls for a Mid-Size Jet Transport, NASA

TM-110374, 1995.

8Dornheim, Michael A., "Throttles Land "Disabled"

Jet", Aviation Week and Space Technolog3, Sept. 4,

1995, p.p. 26-27.

9Burken, J., J. Feather, S. Goldthorpe, J. Kahler,

T. Maine, and EW. Burcham, "Flight Test of a

Propulsion-Based Emergency Control System on the

MD-11 Airplane," AIAA 96-3919, July 1996.

l°Burken, John, et al., "Longitudinal Emergency

Control System Using Thrust Modulation (MD-11),"

AIAA 96-3062, July 1996.

l_Feather, J.B. and F.W. Burcham, Jr., Development

and Flight Test of a Propulsion Controlled Aircraft

(PCA) System and Application to the MD- 11 Airplane.

MDC 96-K0202, McDonnell Douglas Aerospace, Long

Beach, California, March 1996.

25
REPORT DOCUMENTATION PAGE For,.Approved OMB No. 0704-01 88

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT'tYPE AND DATES COVERED

July 1996 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Development and Flight Test of an Augmented Thrust-Only

Flight Control System on an MD- 11 Transport Airplane

6. AUTHOR(S)

Frank W. Burcham, Jr., Trindel A. Maine, John J. Burken,

and Drew Pappas

7.PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9.SPONSORING/MONOTORING AGENCYNAME(S) ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU 505-68-10

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2107

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-4745

11. SUPPLEMENTARY NOTES

Presented as AIAA 96-3742, AIAA Guidance, Navigation, and Control Conference, San Diego, California, July 29-3 l, 1996.

EW. Burcham, Jr., T.A. Maine, and J.J. Burken, NASA Dryden Flight Research Center, Edwards. California. Pappas,

McDonnell Douglas Aerospace, Long Beach, California.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 08

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

An emergency flight control system using only engine thrust, called Propulsion-Controlled Aircraft (PCA), has

been developed and flight tested on an MD-11 airplane. In this thrust-only control system, pilot flightpath and

track commands and aircraft feedback parameters are used to control the throttles. The PCA system was

installed on the MD-11 airplane using software modifications to existing computers. Flight test results show

that the PCA system can be used to fly to an airport and safely land a transport airplane with an inoperative

flight control system. In up-and-away operation, the PCA system served as an acceptable autopilot capable of

extended flight over a range of speeds and altitudes. The PCA approaches, go-arounds, and three landings

without the use of any normal flight controls have been demonstrated, including instrument landing systemcoupled hands-off landings. The PCA operation was used to recover from an upset condition. In addition, PCA

was tested at altitude with all three hydraulic systems turned off. This paper reviews the principles of throttlesonly flight control; describes the MD-11 airplane and systems; and discusses PCA system development,

operation, flight testing, and pilot comments.

14. SUBJECTTERMS

Emergency flight control; MD-11 airplane; Propulsion-controlled aircraft;

Propulsive control; Throttle-only control

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF REPORT OFTHIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500 Available from the NASA Center for AeroSpace Information, 800 Elkridge Landing Road,

Linthicum Hetghts, MD 21090; (301)621-0390

15. NUMBER OF PAGES

28

16. PRICE CODE

AO3

20. LIMITATION OF ABSTRACT

Unlimited

Standard Form 298 (Rev. 2-89)

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298-102

National Aeronautics and

Space Administration

Code JTT

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USA

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