NASA Technical Memorandum 4746

Flight Test of a Propulsion-Based

Emergency Control System on the

MD-11 Airplane With Emphasis

on the Lateral Axis

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

John Feather, Steven Goldthorpe and Jeffrey A. Kahler

July 1996

NASA Technical Memorandum 4746

Flight Test of a Propulsion-Based

Emergency Control System on the

MD- 11 Airplane With Emphasis

on the Lateral Axis

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

Dryden Flight Research Center

Edwards, California

John Feather and Steven Goldthorpe

McDonnell Douglas Aerospace

Long Beach, California

Jeffrey A. Kahler

Honeywell, Inc.

Phoenix, Arizona

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1996

FLIGHT TEST OF A PROPULSION-BASED EMERGENCY CONTROL

SYSTEM ON THE MD-11 AIRPLANE WITH EMPHASIS

ON THE LATERAL AXIS

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

NASA Dryden Flight Research Center

Edwards, California

John Feather* and Steven Goldthorpe*

McDonnell Douglas Aerospace

Long Beach, California

Jeffrey A. Kahler*

Honeywell, Inc.

Phoenix, Arizona

Abstract Nomenclature

A large, civilian, multiengine transport MD-II

airplane control system was recently modified to

perform as an emergency backup controller using engine

thrust only. The emergency backup system, referred to

as the propulsion-controlled aircraft (PCA) system,

would be used if a major primary flight control system

fails. To allow for longitudinal and lateral-directional

control, the PCA system requires at least two engines

and is implemented through software modifications.

A flight-test program was conducted to evaluate the

PCA system high-altitude flying characteristics and to

demonstrate its capacity to perform safe landings. The

cruise flight conditions, several low approaches and one

landing without any aerodynamic flight control surface

movement, were demonstrated. This paper presents

results that show satisfactory performance of the PCA

system in the longitudinal axis. Test results indicate that

the lateral-directional axis of the system performed

well at high altitude but was sluggish and prone to

thermal upsets during landing approaches. Flight-test

experiences and test techniques are also discussed with

emphasis on the lateral-directional axis because of the

difficulties encountered in flight test.

*AIAA Member, Aerospace Engineer

*Aerospace Engineer

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 license to

exercise all rights under the copyright claimed herein for

Governmental purposes. All other rights are reserved by the copyright

owner.

Clls

c.g.

EPR

FADEC

FCC

H

Klleft

Klrigm

Klat

Klc

Klead

Kp

Kphf

gphid

Kpr

Kthe

Kve

gver

Kvi

KWO

KCAS

roll caused by the dihedral term

center of gravity

engine pressure ratio

full-authority digital engine control

computers

flight control computer

altitude, ft

left lateral engine gain

right lateral engine gain

lateral feed-forward gain, lb/deg

heading error feed-forward gain, deg

command prefiiter gain

longitudinal damping gain, deg

bank angle rate feedback gain, deg/sec

bank angle rate feedback gain, deg/sec

turn coordination feedback gain, deg/sec

pitch attitude (0) gain, deg

longitudinal proportional gain, deg

longitudinal proportional gain, lb/deg

longitudinal integrator gain, deg

wash out gain, deg

knots calibrated airspeed

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American Institute of Aeronautics and Astronautics
LSAS

rn.a.c.

P

PCA

PCALAT

IS

V

Verr

W

O_

Y

7cmd

"_err

9

0

¢

_ cmd

longitudinal stability augmentation system

mean aerodynamic chord, percent

roll rate, deg/sec

propulsion-controlled aircraft

lateral PCA engine command, lb

yaw rate, deg/sec

Laplace transform

airspeed, kn

velocity error

takeoff gross weight, lb

angle of attack, deg

flightpath angle, deg

flightpath angle command, deg

flightpath angle error, deg

pitch attitude angle, deg

pitch attitude angle rate, deg/sec

bank angle, deg

roll attitude, deg

heading angle, deg

heading angle command, deg

Introduction

Aircraft flight control systems are designed with

extensive redundancy to ensure a low probability of

failure. However during recent years, several aircraft

have experienced major flight control system failures,

leaving engine thrust as the only usable control effector.

In some of these emergency situations, the engines were

used "open-loop" to maintain control of the flightpath

and bank angle of the airplane. Perhaps, the best known

use of manual throttles-only control occurred in

July 1989 on United Airlines flight 232. At cruise

conditions, a DC-10 (McDonnell Douglas Aerospace

(MDA), Long Beach, California) suffered an

uncontained tail engine failure that caused the loss of all

hydraulics. After the failure, the airplane trimmed at

approximately 210 kn with a significant yaw angle

because of damage to the center engine nacelle. Under

extremely difficult circumstances, the crew used wing

engine throttles for control and was able to crash land at

the Sioux City, Iowa, airport, and over one-half of the

people onboard were saved. _ In the majority of cases

surveyed, major flight control system failures have

resulted in crashes with a total of over t200 fatalities. 2

The challenge was to create a sufficient degree of

control through thrust modulation to control and safely

land an airplane with severely damaged or inoperative

control surfaces. The objective of the propulsioncontrolled aircraft (PCA) emergency backup system is

to command the engines to maneuver the airplane to a

safe landing without moving the normal aerodynamic

control surfaces. This system requires that the airplane

have at least two engines, preferably two wing engines,

functioning. In addition, it is assumed that the normal

control surfaces are not locked in a hardover position

where aerodynamic forces would exceed the moments

resulting from the engine's thrust.

To investigate the PCA concept, the NASA Dryden

Flight Research Center, Edwards, California, performed

nonlinear and linear analytical studies and conducted

several flight-test programs. Gross control can be

obtained by using the throttles in the open-loop sense

(manual throttles-only), but making a safe runway

landing is exceedingly difficult because of low phugoid

and dutch roll damping coupled with the high pilot work

load near the ground. 2-6 To improve performance and

reduce the pilot work load, the PCA program was

developed.

This paper concentrates on the difficulties

encountered during flight test of the lateral-directional

controller and the solutions found. Comparisons of

linear simulation models to flight-test results and

control law design processes are presented. In addition,

an overview of the longitudinal controller is presented.

Test Vehicle Description

Figure 1 shows the MD-11 airplane, a large, longrange, three-engine, wide-body transport. This 202 ft

long airplane has a wing span of 170 ft and a maximum

takeoff gross weight (W) of 618,000 lb.

Flight Control Systems

The MD-11 airplane has a mechanical control system

with stability augmentation provided by the flight

control computers (FCC) (Honeywell, Inc., Phoenix,

Arizona) and with irreversible hydraulically powered

actuators. Three independent systems provide hydraulic

power for intended fail-safe capability. Essential control

functions can be maintained by any one of the three

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American Institute of Aeronautics and Astronautics
__170 ft 6 in.

t ° _ in.

//_

9"_-ft _Mean aerodynamic chord 7 in.

lo ft "_ /'//-/_

125% c'g'-_Oft ',"7 ///-/s? fi, in.

. 202 ft N

960203

Figure 1. The MD- 11 airplane.

systems. Dual elevators on each horizontal stabilizer

provide pitch control, and pitch trim is provided by a

moveable horizontal stabilizer. Roll control is provided

by inboard and outboard ailerons supplemented by wing

spoilers. Yaw control is provided by a dual rudder

mounted on a single vertical stabilizer.

The basic control design for the lateral dynamics is

called the yaw damper and longitudinal stability

augmentation system (LSAS) for the pitch dynamics.

Hydraulic devices provide the force through control

cables from the control column that moves the

aerodynamic surfaces commanded from FCC. The FCC

operates at 20 samples/sec.

The MD-11 airplane is equipped with a flight

management system which integrates autopilot,

navigation, and autoland functions. The autopilot

control includes a thumbwheel for commanding

flightpath angle, Ycma, and a heading knob for

commanding the desired heading, Wcma"

Three Pratt & Whitney (Palm Beach, Florida)

PW4460 high-bypass ratio turbofan engines in the

60,000-1b thrust class power the MD-11 airplane. Two

engines are mounted in underwing pods. The third

engine is located at the base of the vertical stabilizer.

Each engine has a full-authority digital engine control

(FADEC) system in which the software was modified

for the PCA program. The modification allows for an

increased range of engine pressure ratio (EPR)

commands to be sent from the FCC. The range was

increased from 5 percent of the trimmed EPR to

approximately 0.9 to 1.50 EPR. The wing engines are

121 in. below the nominal vertical center of gravity

(c.g.), and the tail engine is 240 in. above the vertical

c.g. with its thrust axis inclined 2.5 ° (nozzle pointing

down). The crew controls the engines with electronic

throttles which command a power setting based on EPR.

As is typical for high-bypass turbofan engines, thrust

response at low power settings is initially slow. Once

thrust levels exceed 20 percent, the engine response

improves dramatically. An "approach idle" setting when

the flaps are extended beyond 27 ° maintains the idle

revolutions per minute at a sufficiently high level so that

the 8 sec from idle to full power requirement can be

met. A cruise idle or minimum idle setting can require

as much as 12 sec to go from idle to full power. 2 If PCA

were engaged with minimum idle setting, a pilotinduced oscillation could occur because of the large

time lags; therefore, another modification to the FADEC

system included setting the approach idle when the PCA

system was engaged.

PCA Control System Design

The design of the longitudinal and lateral-directional

control laws assumes that the normal control surfaces

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arenotfunctioningandarenotin ahardoverposition.

Giventhe limitedenginebandwidth,PCAcontrol

requiresarelativelystableopen-loopplantor slightly

unstablepoles.As a transportairplane,theMD-11

easilymeetsthiscriterion.Ofthecasesinvestigated, the

leaststableopen-looplateral--directional poleswerefor

aflightconditionof 195kn/0.0° flaps.In addition,the

leaststableopen-looplongitudinal eigenvalues werefor

a flightconditionof 145kn/28° flaps. The MD-11

engine bandwidth is limited to approximately 2 rad/sec.

The PCA system uses engine thrust modulation driven

by a closed-loop controller to increase the damping and

allow the pilot to land safely. In the longitudinal axis,

the phugoid mode needed improvements. In the

lateral--directional axis, the dutch roll poles needed

enhancement.

The initial PCA system was designed to have minimal

impact on existing hardware and software. The resulting

system only required software changes. The flight

control panel was used for the pilot input paths. The

heading control knob was used for the

lateral--directional controller, and the flightpath angle

thumbwheel was used for the longitudinal axis.

Collective thrust commands to the wing engines

provided pitch or longitudinal control, and differential

thrust commands provided lateral--directional control.

The PCA system was designed with the flexibility to

change the control gains in flight (within safety limits),

allowing for improved performance or robustness.

However for safety reasons, the controller architecture

could not be modified other than zeroing out selected

feedback paths.

Nonlinear simulators were heavily used to adjust the

initial gains determined from classical linear design.

These control laws were developed by a team from

McDonnell Douglas Aerospace; Honeywell, Inc.; Pratt

& Whitney; and NASA.

Longitudinal Control

The longitudinal control law commands the flightpath

angle and augments phugoid damping (fig. 2). The

feedback signals selected were pitch attitude, 0; pitch

attitude rate, 0; velocity error, Verr; and flightpath

angle, y. Velocity feedback was not used initially, but it

was added later for improved longitudinal control.

These signals were already available to the primary

stability augmentation system. Hightpath angle error,

"Yerr' was passed through a proportional plus limited

integral compensator to maintain an acceptable tracking

task for the pilots. Washed out pitch attitude and attitude

rate were summed after the integrator for improved

phugoid damping. Reference 7 gives detailed

information regarding the longitudinal controller.

Lateral-Directional Control

Lateral-directional control is obtained by using

differential throttle inputs to generate yaw, resulting in

Pilot

flightpath

command Left engine

commend,Ib

Tail engine

command,Ib

Right engine

commend,Ib

P,tc,..,tu0.

senior

Velocityerror

[]

Figure 2. Longitudinal MD-I l PCA control law.

Flightpath

angle

960310a

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American Institute of Aeronautics and Astronautics
roll caused by the dihedral term, Ct_. The

lateral--directional control law tracks heading angle

commands, Udcmd (fig. 3). The feedback signals

selected to improve the dutch roll damping and closedloop performance consisted of bank angle, _; bank

angle rate, +; yaw rate, r; and heading angle, W. Bank

angle rate was included for dutch roll damping, and yaw

rate and bank angle were included for turn coordination.

The resulting output from the PCA lateral-directional

controller (PCALAT) was then used to modulate the

thrust of each engine.

Linear Analysis

In the longitudinal and lateral-directional axes,

classical methods were initially used to design the

controllers. The classical single-input-single-output

methods included root locus and frequency analysis

methods, such as Bode or Nyquist techniques. Classical

linear design was used for acceptable rise time and

damping characteristics. This design was also evaluated

in a nonlinear simulation where the gains were adjusted

to increase damping or performance.

From linear analysis, the longitudinal flight

conditions were H= 12,000 ft, V= 175 kn, flaps = 28 °,

slats = extend, c.g. = 23 percent m.a.c., W= 360,000 lb,

and gear down. The phugoid damping increased by a

factor of 12 when the PCA system closed the loop. The

stability margins for flightpath loop were acceptable for

the stability criteria of 6 dB and 45 °.

From linear analysis, the flight conditions were H =

12,000 ft, V= 175 kn, flaps = 28 °, slats = extend, c.g. =

23 percent m.a.c., W = 360,000 lb, gear down and

default gains. The dutch roll damping increased by a

factor of 5.2 with the closed-loop system allowing for

good airplane response characteristics. However, the

default gains used in the initial flight-test phase

produced a sluggish response of the airplane near the

ground and required a different set of gains to improve

the roll rate, p.

Simulation

Flight control system design and analysis for aircraft

relies on mathematical models of the vehicle dynamics.

These models are brought together to form a linear or

nonlinear simulation for design and evaluation. The

development of the PCA control algorithms used a sixdegree-of-freedom nonlinear simulation program and

linearized state-space models. In addition, a fixed-base,

piloted, flight deck simulator was used. This simulator

had an option to run hardware-in-the-loop which

included FCC and FADEC.

For linear analysis and simulation, the engines were

modeled as a first-order Laplace transform, s, shown in

Pilot

_l/cmd

Limit 5

_-_-] in

Left eng" e

___ .__ PCALAT [ _ command, lb

command, Ib

[_]

< __ Bank angle rate

Bank angle, _

Heading,

Figure 3. Lateral-directional control law. Note that shaded boxes (Klc, Kla t, Kpr, Kphid)

during flight test.

960311a

were modified gains

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equation 1, with a rate limit of one-half the engine thrust

output in pounds per second (eq. 2).

1

eng(s) - (1)

(0.5s + 1)

Engrate = trimthrust/2 (lb/sec) (2)

Evaluation of this simple model is shown in the Results

and Discussion section through time history matching

of flight and simulation aircraft angular positions and

rates.

Software Implementation

The FCC provides a host of functions, including

autopilot, autothrottle, and navigation. These computers

also provide a flight management system. The PCA

logic interfaced with existing sensor signals and sent

commands to the engine FADEC over a 429 data bus.

Pratt & Whitney modified the engine controllers to

accept a full-authority EPR command from the FCC

which ranged from 0.9 to approximately 1.5. The PCA

system incorporated a safety disengage capability which

was accomplished by the pilot through moving the

throttle lever or by pressing a cockpit FCC switch.

These features provided pilots with normal throttle and

conventional control surface response if needed.

Flight Test Maneuvers

A series of evaluation maneuvers were flown at a

condition of 12,000 ft, 175 KCAS, gear down, flaps

extended to 28 °, slats extended, c.g. at 23 percent m.a.c.

and 360,000 lb. The pilot stabilized the airplane at this

flight condition with the PCA system turned on and

completed a series of step inputs. During PCA flight-test

operations, the hydraulic system was powered with the

control surfaces fixed.

To mcrease confidence in the system, low approaches

were performed in a graduated series of decreasing

altitudes until a landing was accomplished. Note that the

flaps were set at 28 ° (takeoff flap position) to obtain low

landing speeds. Other flight conditions were flown, such

as 0.0 ° flaps and a range of c.g. positions (23 to

31 percent m.a.c.), altitudes, and speeds. 2 Low

approaches to 50 ft above the ground were flown with

0.0 ° flaps, gear down, and airspeed of approximately

195 kn. These cases were not allowed to touch down

because of programmatic decisions and airplane rental

agreements. Although the 0.0 ° flap approach speeds

would have been pushing the upper limitations of a

normal MD-11 landing (204-kn tire speed), during an

emergency these conditions would be acceptable. The

PCA flying qualities with the flaps at 0.0 ° were wellbehaved. No noticeable stability or performance

degradation occurred.

Results and Dis¢¢ssion

This section describes the linear simulation to flight

result comparison for the longitudinal and lateraldirectional axes. In addition, results are presented for

the improved lateral-directional controller and one

MD- 11 PCA landing.

Comparison of Simulation to Flight Evaluation

Figure 4 shows a longitudinal axis comparison of

flight and linear simulator results for a series of

flightpath angle step inputs at a flight condition of

175 kn, 12,000 ft, 28 ° flaps, and gear down. There is an

overshoot of the flightpath response compared to

command of approximately 30 percent for both the

simulator and flight-test results. This overshoot did not

concern the pilots even at low altitudes. In general, the

linear model represents the flight dynamics reasonably

well and is adequate for control design. Reference 7

provides additional information regarding the different

longitudinal modes flown.

Figure 5 shows a lateral--directional axis comparison

of flight and linear simulation results for a heading

angle command, step input using the default gains at the

previous flight condition. The time history traces of the

simulator model matched the flight data reasonably

well. However, the pilots rated the lateral-directional

response poor near the ground, and modifications to

improve the sluggish response were needed.

The piloted simulations did indicate some

lateral--directional challenges but not to the extent of the

flight-test evaluations. Matching a pilot's work load, or

gain, in a simulator to flight is a difficult task, especially

during a turbulent day with a low bandwidth control

system. The level of turbulence has a large impact on the

PCA controller performance and pilot ratings,

especially in the roll axis. These approaches and landing

occurred on a hot August day with light turbulence.

Although the winds were light (approximately 9-kn

head wind), high thermal disturbances caused roll

upsets.

Lateral-Dircqtional Axis Modification

The major controls challenge of the MD-II PCA

system was to improve the lateral-directional axis

response without a new control law release. The pilots

6

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Flightpath

angle,

deg

0

-1

-2

-3

.5

Pitch

attitude

rate, 0

deg/sec

4.0

Angle

of attack,

deg

Velocity,

ft/sec

-- -- Flight

.... Simulator s

--- -- y command .-.v_

- i \ ,,',4,,

' '} ._ ,1."

•

I I I

1.0 --

I I I I I I

3.5

3.0

305 --

300

295

290

285

28O

0

sss '_ _

d f

• t I

I I I I I I

10 20 30 40 50 60

Time, sec

7O

960312

Figure 4. Comparison of flight and linear simulation flightpath angle step response.

7

Amencan Institute of Aeronautics and Astronautics
Heading,

deg

-2

-- Flight

r o - ..... Simulator

I ...- " °_ ! _ .._ Command

jt_.1_ I, - -. _,,_ heading i

I I I I 1 I

Bank

angle,

deg

10

0

-5

n _ n _

s

I I

1.5 m

Roll rate,

deg/aec

1.0

.5

0

_o5

- 1.0

.6 m

Yaw rate,

deg/aec

.4

.2

0

-.2

- .4 0

I

10 20 30 40 50 60

Time, sec 960313

Figure 5. Comparison of flight and linear simulation heading step response using the default gain set.

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American Institute of Aeronautics and Astronautics
rated the response marginal at best. Lateral-directional

control gains were modified to improve the

performance. The heading error feed forward, lateral

feed forward, turn coordination feedback, and bank

angle feedback ( Klc, gphid, Klat, and Kpr ) gains were

selected for modification (fig. 3). The Kphid and Kpr

gains were used as damping adjustment parameters,

while the Klc and Kla t gains affected the initial

response. Six gain sets were flight tested, but only two

are presented here: the default set and T6 gain set. Gain

set T6 gave the largest roll rates per degree of

commanded input and was the gain set used for the PCA

landing.

Using linear root locus analysis and relying on

nonlinear simulation runs, the gains were modified. The

control system with these gains was then tested in flight

by inputting the variables using the multifunction

control and display unit. Flight conditions were H =

12,000 ft, V = 175 kn, flaps = 28 °, slats = extend, c.g. =

23 percent m.a.c., W = 360,000 lb, gear down, and

default gains.

Figure 6 shows the flight comparison between the

default gains and T6 gains at a flight condition of

175 kn, 12,000 ft, 28 ° flaps, and gear down. A heading

change of 5° was commanded, and the time to reach

90 percent of the final value (rise time) was calculated.

The rise time for the default gain set was 19.5 sec, and

the T6 gain set was 12 sec. The heading performance

was improved by 38 percent using the T6 gain set. The

maximum body axis roll rate in figure 6 increased

77 percent using T6 gains but at the cost of reduced

dutch roll damping (note the roll and yaw rate traces).

Figure 6 also shows the EPR of the left and right

engines for both gain sets. The T6 gain set commanded

more engine activity than the default gains. The pilot's

comments on the T6 gain set response were much more

favorable than those regarding the default gain response.

The comments made were "I could feel the response

'kick-in' faster with T6 compared to the default set" and

"I did not have to wait as long for the airplane to 'catchup' to my input command." The T6 gain set was used

for the PCA landing.

One control law change that should improve the

heading response without decreasing the dutch roll

damping would be to apply a lead-lag compensator in

the pilot's forward path (fig. 7) while using the default

gain set. This method would have the same closed-loop

poles as the default gain set, with increased lead to the

input signal. Unfortunately, there was no way to add a

lead-lag filter to the input signal without a new control

law release. The next phase of flight test will include an

equivalent lead-lag filter. One possible drawback would

be engine saturation, and further testing is needed in this

area.

After the lateral-directional axis response was

improved, and without any longitudinal axis changes, a

successful landing was performed using the PCA

control system. In a trimmed 28 ° flaps condition, a

landing was performed without any control surface

movement, simulating a total hydraulics pressure loss.

The results are shown in figure 8 with the landing

occurring at time = 83 sec. During the approach, a large

thermal upset caused the airplane to bank over to an

angle of approximately 8° just after a lateral command

change (time = 60 sec). The lateral-directional axis

controller commanded a restoring signal to remove the

error without pilot inputs. The 7° upset at time = 12 sec

was caused by the pilot's commanded input as shown in

the heading trace (time = 5 sec). Angle of attack, _,

decreased just before landing because of ground effects.

The gust rejection characteristics are less than desirable

(note the bank angle trace). The design criterion was to

have the bank angle error less than +_5° during the

approach and landing task. Incorporation of a lead-lag

filter with the reduced default gains should help with the

gust sensitivity problem by providing increased lead

response without reducing the dutch roll damping.

Concluding Remarks

An emergency backup control system using engine

thrust only was designed and flight tested on a large,

civil transport MD-11 airplane. This report describes

the longitudinal and lateral-directional propulsion

controlled aircraft (PCA) control system and compares

simulation and flight-test results. Flight data

comparisons with linear models were shown with the

emphasis on the lateral--directional sluggish response.

The control system was designed from the onset with

the flexibility to change several control gains in flight.

The longitudinal control system performed well

during the high-altitude operations, low approaches, and

landing. The pilots rated the longitudinal characteristics

good with minor pilot compensation needed. The

performance was satisfactory; therefore, changes were

not required during the first phase of flight test.

The flaps were set at 28 ° (takeoff flap position) to

obtain low landing speeds. Other flight conditions were

flow, such as 0.0 ° flaps and a range of center-of-gravity

(e.g.) positions (23- to 31-percent m.a.c.), altitudes, and

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Heading,

dog

86

................ _..-':___. _:_--_:7:.. "___-_-.-

90 percent

82 l ji_... " " command I "

80

78

Bank angle,

deg 0

Roll rate,

deg/sac

Gains

T6

- - - Default

Yaw rate,

deg/sec

EPR

I I I I I I

-5

m

-I

L

!

-'s/ I ! I I I

- 1.0 J

1.5 F _t engine

I \ f _ tr _ Left engine

1.2 , t t I

0 10 20 30 40 50 60

Time, sec 9603_4

Figure 6. The flight heading step response using the default and T6 gains.

10

American institute of Aeronautics and Astronautics
Pilot

Lead-lag Limit

_____ Kleft =- _-I

Left engine

command, Ib -F___ _i __., __

_1 command, Ib

-_<_ -" Bank an[_g]glerate

Bank angle

Figure 7. Lateral-directional control law B.D. phase 2 lead-lag controller.

960315a

11

American Institute of Aeronautics and Astronautics
Flightpath

angle,

deg

Heading,

deg

Bank

angle,

deg

Control

surface

position,

2

0

-2

-4

- 132

- 134

- 136

- 138

10

5

0

-5

10

I _ Ycmd

I I I I I I I I I

_ _cmd

ii % *" "_ _ I %

"'i I I J_ I "'t--" I I i

f -- Aileron

.... Elevator

----- Rudder

_'_r'-'_"

I I I I I I I I I

deg

Altitude,

ft

Angle

of attack,

deg

EPR

Sink rate,

fl/sec

-10

IOO0

500

I

0

6

4

2

-- Left

1.5 [-- .... Tall ,- II I I I I I 1 I I 1 .8

0

-10

- 20 0 10 20 30 40 50 60 70 80 90

Time, sec 960316a

Figure 8. Pilot-commanded landing with 28 ° flaps and an airspeed of 175 kn.

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American Institute of Aeronautics and Astronautics
speeds. Low approaches to 50 ft above ground level

were flown with 0.0 ° flaps, gear down, and airspeed of

195 kn. These cases were not allowed to touch down.

Although the 0.0 ° flap approach speeds would have

been pushing the upper limitations of a normal MD- 11

landing (204-kn tire speed), during an emergency, these

conditions would be acceptable. The PCA flying

qualities with the flaps at 0.0 ° were well-behaved.

The original lateral-directional response was

considered too sluggish near the ground because of

wind gusts. Control law changes used to improved the

response were shown and compared to the initial

system. The roll response increased approximately

77 percent with a new gain set (T6), but the dutch roll

damping decreased. Future work may include a lead-lag

filter with the default gains for improved response. The

default gain set has greater dutch roll damping and will

inherently improve gust rejection. Allowing for gain

changes in flight-improved, lateral-directional response

without the need for time-consuming and expensive

control law updates.

This backup control system could be used in the event

of the airplane suffering a major primary flight control

system failure, such as a total hydraulic pressure loss.

The PCA system has limited control power which may

not be sufficient to handle surface hardovers or large

mistrim configurations. However in the absence of large

mistrim configurations, the PCA system provides a

method for returning the airplane to the airport and

landing without the aid of aerodynamic control surfaces.

The PCA system changes a flight situation where there

is an extremely high work load (using manual throttle

inputs) to a viable piloting task.

References

_David Hughes and Michael A. Dornheim, "United

DC-10 Crashes in Sioux City, Iowa," Aviation Week &

Space Technology, July 24, 1989, p.p. 96-97.

2Burcham, Frank W., Jr., Trindel A. Maine, John J.

Burken, and Drew Pappas, "Development and Flight

Test of an Augmented Thrust-Only Flight Control

System on an MD-11 Transport Airplane," AIAA

96-3742, July 1996.

3Burcham, EW., Jr., John J. Burken, and Trindel A.

Maine, Flight Testing a Propulsion-Controlled Aircraft

Emergency Flight Control System on an F-15 Airplane,

NASA TM-4590, 1994.

4Burcham, Frank W., Jr., and C. Gordon Fullerton,

Controlling Crippled Aircraft--With Throttles, NASA

TM- 104238, 1991.

5Gilyard, Glenn B., Joseph L. Conley, Jeanette Le,

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

a Four-Engine Jet Transport Using Engine Thrust

Modulation for Flightpath Control, NASA TM-4324,

1991. (See also AIAA 91-2223, June 1991.)

6Burcham, Frank W., Jr., Trindel A. Maine, and

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

Airplane Using Throttles for Flight Control, NASA

TM-104255, 1992. (See also AIAA 92-4109.)

7Burken, John J., Trindel A. Maine, Frank W.

Burcham, Jr., and Jeffery A. Kahler, "Longitudinal

Emergency System Using Thrust Modulation

Demonstrated on an MD-11 Airplane," AIAA 96-3062,

1996.

13

American Institute of Aeronautics and Astronautics
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

Public reporting burden lot this collection of information is estimated to average 1 hour pet response, including the time for rovNswing instructions, searching existing data sources, gathering and

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including suggestions for reducing this burden, to Washington Headquarlers Services, Directorate for Information Operations and Re >otis, 1215 Jefferson Dav_s H_ghway, Sure 1204, Arlington.

VA222024302, and to the Officeof Managementand Budget,PaperworkReduCtionProject(0704-0188),Wuhk',gton, DC 20503.

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

July 1996 Technical Memorandum

4.TITLE AND SUBTITLE 5. FUNDING NUMBERS

Flight Test of a Propulsion-Based Emergency Control System on the

MD-11 Airplane With Emphasis on the Lateral Axis

6. AUTHOR(S)

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

Steven Goldthorpe, and Jeffrey A. Kahler

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 505-64-10

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2110

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-4746

11. SUPPLEMENTARY NOTES

Presented as AIAA 96-3919 at the AIAA Guidance, Navigation, and Control Conference,

San Diego, California, July 29-31, 1996.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 08

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

A large, civilian, multiengine transport MD-I 1 airplane control system was recently modified to perform as an

emergency backup controller using engine thrust only. The emergency backup system, referred to as the propulsioncontrolled aircraft (PCA) system, would be used if a major primary flight control system fails. To allow for longitudinal

and lateral-directional control, the PCA system requires at least two engines and is implemented through software

modifications. A flight-test program was conducted to evaluate the PCA system high-altitude flying characteristics and to

demonstrate its capacity to perform safe landings. The cruise flight conditions, several low approaches and one landing

without any aerodynamic flight control surface movement, were demonstrated. This paper presents results that show

satisfactory performance of the PCA system in the longitudinal axis. Test results indicate that the lateral-directional axis

of the system performed well at high altitude but was sluggish and prone to thermal upsets during landing approaches.

Flight-test experiences and test techniques are also discussed with emphasis on the lateral-directional axis because of the

difficulties encountered in flight test.

14. SUBJECTTERMS

Control system failure; Emergency controller using thrust; Hydraulic system

failure; MD-11 airplane; Propulsion-only control

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

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

15. NUMBER OF PAGES

16

16. PRICE CODE

AS3

20. LIMITATION OF ABSTRACT

Unlimited

NSN 7540-01-280-5500 Available from the NASA Center for AeroSpace Information, 800 Elkndge Landing Road, Standard Form 298 (Rev 2-89)

L inthicum Heights, MD 21090; (30 I)621-0390 PrescribedbyANStStd Z39-18

298-102