_,--' i_ ,.

-:: NASA Contractor Report 159098 ,..:V.i

-!•- 19820007204_' 3 1176 00156 6034 ,. -"

• •'_: Accelerated Development and Flight" Evaluation of Active Controls Conceptsl !

: • for Subsonic Transport Aircraft -i .-I

:. Volume II - AFT C.G. Simulation andI -!

._: Analysis!' "

I "iI

-- Lockheed - California Company:'- Burbank, California4

L

Contract NAS 1-14690

::[_._] September 1979

: " _se of its significant early commercial potential this informatio_een- - deve_r a US. Government program is beingdissemin_hin the United:; States]n _1 publication. This informati_licated and used

" _ed. Releaseof this, _h I be made subject to those,! r__val and appro-:_ __his

.:; _o_part.

. r' ..7 1£1/9

LANGLEY RESEARSH CENTER

, ! LIBRARY,NASA- : H_AMF_TO_N,YJRGIB.IA

i'i: NationalAeronautics and

i: : _ SpaceAdministration

! "2i" LangleyResearch.Center ,.Hampton,Virginia23665

AC 804 827-3966

ii

https://ntrs.nasa.gov/search.jsp?R=19820007204 2020-07-28T22:21:34+00:00Z

°.

3,

1. REPORT NO,i i

2. GOVERNMENT ACCESSION NO. 3. RECIPIENT'S CATALOG NO.

., NASA CR.I_0_84. TITLE AND SUBTITLE 5. REPORT DATE

ACCELERATED DEVELOPMENT AND FLIGHT EVALUATION OF September 1979ACTIVE CONTROLS CONCEPTS FOR SUBSONIC TRANSPORT 6.PERFORMINGORG CODEAIRCRAFT VOLUME 2 - AFT C.G. SIMULATION & ANALYSIS

7. AUTHOR(S) 8. PERFORMING ORG REPORT NO.

D. M. Urie, et al LR 2900_-210. WORK UNIT NO.

9. PERFORMING ORGANIZATION NAME AND ADDRESS

LOCKHEED-CALIFORNIA COMPANYP.O. BOX 551 11. CONTRACT OR GRANT NO:

BURBANK, CALIFORNIA 91520 NASI-1469013. TYPE OF REPORT AND PERIOD

112. SPONSORING AGENCY NAME AND ADDRESS COVERED

National Aeronautics & Space Administration Contractor ReportLangley Research Center 14. SPONSORING AGENCY CODE

Hamoton. Virginia 2_66515. SUPPLEMENTARY NOTES

Final Report - Volume 2 (see CR 159097-Volume i)!-

16. ABSTRACT

A study was conducted to investigate Relaxed Static Stability (RSS) and sta-bility augmentation with active controls applied to subsonic transport air-craft. Analytical and simulator evaluations were done using a contemporarywide body transport as a baseline. Criteria for augmentation system perform-ance and unaugmented flying qualities were evaluated. Augmentation controllaws were defined based on selected frequency response and time historycriteria. Flying qualities evaluations were conducted by pilots using amoving base simulator with a transport cab. Static margin and air turbulenceintensity were varied in tests with and without augmentation. Suitability ofa simple pitch control law was verified at neutral static margin in cruiseand landing flight tasks. Neutral stability was found to be marginally ac-ceptable in heavy turbulance in both cruise and landing conditions.

17. KEY WORDS(SUGGESTED BY AUTHOR(S)) 18. DISTRIBUTIONSTATEMENT

Active Controls

Relaxed Static Stability FEDD DocumentStability AugmentationFlying QualitiesFli_ht Simulat_Qn19. SECURITY CLASSIF. 20. SECURITY CLASSIF.(OFTHISPAGE) 21. NO. OFPAGES[22. PRICE"

(OF THIS REPORT) Unclassified 87Unclassified

d/Tf---7/ww

ACCELERATED DEVELOPMENT AND FLIGHT EVALUATIONOF ACTIVE CONTROIS CONCEPTS FOR SUBSONIC TRANSPORT AIRCRAFT

VOLUME 2 - AFT C .G. SIMULATION AND ANALYSIS

LOCKHEED-CALIFORNIA COMPANYCOORDINATED BY: D.M. URIE

SUMMARY

This task entailed simulation of Relaxed Static Stability (RSS) configura-

tions of a contemporary wide body subsonic transport, the Lockheed L-lOll.

Task objectives were to select criteria for augmentation system performance and

unaugmented flying qualities and to define control laws for an augmentation

system suitable for a derivative L-lOll configuration with a smaller horizontal

tail. These objectives have been attained.

Design of augmentation control laws was accomplished using frequency

response criteria. These criteria used the current transport modal character-

istics as a standard for acceptable performance. Control laws so defined were

then evaluated against normalized time history response envelopes of the cur-

rent airplane.

Pilot-in-the-loop flight simulations were conducted on a moving base simu-

lator with an L-lOll cab. Static margin and air turbulence intensity were

varied with and without augmentation. These tests showed that lagged pitch

rate damper provided flying qualities equivalent to the baseline airplane at

aerodynamic stability levels down to neutral in heavy turbulence. Modified

control laws resulting in quicker and slower time response to control input

were evaluated. No clear preference was found, thus indicating that there is

wide latitude in satisfactory dynamic control response.

Unaugmented simulations demonstrated that flying qualities of configura-

tions with as little as 3 percent static margin are acceptable in cruise and

approach flight conditions even with heavy turbulence.

The aft C.G. simulation results provide sufficient basis for proceeding

to flight evaluation of the defined augmentation control laws with RSS and a

small horizontal tail.

iii

TABLE OF CONTENTS

Section Page

SUMMARY iii

LIST OF FIGURES vii

LIST OF TABLES ix

SYMBOLSANDABBREVIATIONS .xi

1.0 INTRODUCTION i-i

2.0 DISCUSSION 2-1

2.1 AUGMENTATIONSYSTEMDEVELOPMENT 2-1

2.2 UNAUGMENTEDCHARACTERISTICS 2-3

2.3 AUGMENTATIONSYSTEMDEFINITION 2-10

2.4 AFT C.G. FLIGHT SIMULATIONOBJECTIVES 2-272.5 SIMULATIONMATH MODEL 2-31

2.6 SIMULATORDESCRIPTION 2-34

2.7 DATA ACQUISITION 2-46

2.8 DESCRIPTIONOF TEST CONFIGURATIONS 2-46

2.9 TEST CONDITIONS 2-46

2.10 PILOT EVALUATIONOF UNAUGMENTEDFLYING QUALITIES 2-51

2.11 PILOT EVALUATIONOF FLYING QUALITIES WITHAUGMENTATION 2-59

2.12 AUGMENTATIONRELIABILITY 2-71

2.13 STALL DYNAMICS 2-73

3.0 CONCLUSIONS 3-1

4.0 RECOMMENDATIONS 4-1

V

LISTOF FIGURES

Title Page

l-1 L-1011-1 General Arrangement 1-2

i-2 L-1011-RE General Arrangement 1-3

2-1 Linear Aerodynamic Model Accuracy: Approach Pitch Rate 2-h

2-2 Linear Aerodynamic Model Accuracy: Cruise Pitch Rate 2-5

2-3 Small Tail C* Time History for the Unaugmented Approach 2-6

2-h Small Tail Pitch Rate Unaugmented Approach 2-7

2-5 Small Tail C* Time History Unaugmented Cruise 2-8

2-6 Small Tail Pitch Rate Unaugmented Cruise 2-9

2-7 Characteristic Roots Unaugmented Standard Tail Approach 2-11

2-8 Characteristic Roots Standard•Tail Cruise withMach Trim 2-12

2-9 Characteristic Roots Unaugmented Small Tail Approach 2-13

2-10 Characteristic Roots Unaugmented Small Tail CruisewithMach Trim 2-1h

2-11 L-1011-RE Augmentation System Block Diagram 2-16

2-12 Effects of Gain and Lag Time Constant Approach C.G.at 0.25 _ 2-17

2-13 Effects of Gain and Lag Time Constant Approach C.G.at 0.35 _ 2-18

2-14 Effects of Gain and Lag Time Constant Approach C.G.at 0.h0 _ 2-19

2-15 Effects of Gain and Lag Time Constant Cruise C.G.at 0.25 _ 2-20

2-16 Effects of Gain and Lag Time Constant Cruise C.G.at 0.35 _ 2-21

2-17 Effects of Gain and Lag Time Constant Cruise C.G.at 0.h0 _ 2-22

2-18 Effect of C.G. Approach Lag Time Constant = 0.8 Seconds 2-23

• 2-19 Effect of C.G. Cruise Lag Time Constant = 0.h Seconds 2-24

2-20 Effect of C.G. Approach System 1 2-25

vii

LIST OF FIGURES (CONTINUED)

Title Page

2-21 Effect of C.G. Cruise System 1 2-26

2-22 Landing Approach C* 2-28

2-23 Cruise C* 2-29

2-24 Control Surface Designation and Axes System 2-32

2-25 Engine Dynamics and Derivation of EPR and Thrust 2-33

2-26 Turbulence Model 2-35

2-27 Longitudinal Control Systems 2-36

2-28 Lateral Control Systems 2-37

2-29 Column to Stabilizer Gearing (J Curve) 2-38

2-30 L-1011-500 Pitch Feel Spring Rate 2-hl

2-31 Pitch Feel Force System 2-h2

2-32 Cockpit Displays 2-h3

2-33 Lateral Flight Director 2-hht

2-34 Longitudinal Approach Flight Director 2-h5

2-35 Handling Qualities Rating Scale 2-h7 .

2-36 Reduced Altitude Best Cruise 2-52

2-37 Cruise Flying Qualities L-lOll-RE No Stability Augmentation 2-53

2-38 Effect of Altitude on Cruise Flying Qualities 2-55

2-39 Approach Flying Qualities L-1011-RE No Stability Augmentation 2-56

2-40 Approach Flying Qualities (Unaugmented) 2-57

2-41 Cruise Flying Qualities (Unaugmented) 2-58

2-42 Pilot No. 1 Evaluation of Cruise Augmentation 2-60

2-43 Pilot No. 2 Evaluation of Cruise Augmentation 2-61

2-44 Pilot No. 3 Evaluation of Cruise Augmentation 2-62

2-45 Time Histories of Flight in Turbulence; Effect of

Augmentation 2-63

2-46 Aircraft Response to HeavyTurbulence, Cruise Configuration 2-65

2-47 Control Activity in Heavy Turbulence, Cruise Configuration 2-66

2-48 Pilot No. 1 Evaluation of Approach Augmentation 2-67

2-h9 Pilot No. 2 Evaluation of Approach Augmentation 2-68

2-50 Pilot No. 3 Evaluation of Approach Augmentation 2-69

2-51 Augmentation System Authority Requirements Data 2-70

2-52 Probability Approach - Handling Qualities Equivalence 2-72

viii

LIST OF TABLES

Table Page

i-i Airplane Geometry i-4

2-1 L-1011-RE Augmentation System Characteristics 2-30

2-2 Recorded Parameters - Analog Strip Chart Recordings 2-48

2-3 Recorded Parameters - Digital Line Printer Recording 2-49

2-4 Statistical Data Parameters 2-50

_x

LISTOF SYMBOLSAND ABBREVIATIONS

SYMBOL DEFINITION

C* Cockpit Normal Acceleration plus UC TimesPitch Rate Divided by Gravitational Acceleration

Mean Aerodynamic Chord (MAC)

c.g. Center of Gravity

ELOC, EG.S. Localizer and Glideslope Errors

EPR Engine Pressure Ratio

FCOL Column Force

H Altitude

IFR Instrument Flight Rules

K Stabilizer Feed Forward Gain

KQ Pitch Damper Gain

Lu,Lw Specified Altitudes Used to Define TurbulenceProperties

M Mach Number

NZ Normal Load Factor at the c.g.

PLA Engine Power Lever Angle

PLADOT Rate of Change of Engine Power Lever Angle

Q Aircraft Pitch Rate

RT Yaw Rate

S Laplace Operator

SAS Stability Augmentation System

SH Horizontal Tail Area

xi

SYMBOL DEFINITION

3

T Engine Thrust

TURBU Disturbance Velocities from Turbulence in the

TURBV Forward, Side and Vertical DirectionsTURBW Respectively

U, V, W Forward, Side, and Vertical Components of the "Aircraft velocity vector

U Constant Velocity in C*c

V Equivalent Airspeede

W Aircraft Gross Weight

6AT Aileron Deflection

6C0g Column Deflection

6F Flap Deflection

6H Stabilizer Deflection f,

6R Rudder Deflection

6Sp Spogler Deflection

6w Control Wheel Angle

Damping Ratio

8 Aircraft PitchAttitude

TLAG Lag Time Constant

T Washout Time Constantwo

Heading

_d Damped Frequency

_n Natural Frequency

xii

SECTION i

INTRODUCTION

Contract NASI-14690 is a program whereby Lockheed is investigating the use

of active controls in the L-1011 for increased energy efficiency with applications

in the commercial air fleet as early as 1980. The program investigates the use

of maneuver load control, elastic mode suppression and gust alleviation with

increased wing aspect ratio; and of augmented stability with more aft c.g. and

smaller horizontal tail. The augmented stability permits relaxation of conven-

tional static stability margins leading to the use of a substantially smaller

tail, with significant weight and drag savings. The expected energy efficiency

improvement attributable to the small tail is B% to B-1/2%.

Three tasks are defined for the contracted program:

• Task 1 - Flight Testing of load alleviation systems on an L-1011aircraft.

• Task 2 - Design and pilot-in-the-loop simulator testing of a longitudinalstability augmentation system. Includes development of criteria forsystems-off characteristics.

• Task 3 - Flight testing and evaluation of a modified L-1011 with extendedwing tips and active controls.

Results of the aft C.G. simulation study, Task 2, are reported in this volume of the

final report. Results of Tasks 1 and 3 are reported in Volume 1.

In this task three versions of the L-1011 aircraft are utilized. The current

L-1011-1 is used as a basis for flying qualities evaluations. This configuration

is depicted in Figure 1-1. The baseline aircraft for determining the direct effects

of the tail size reduction is the shorter bodied L-1011-500 with extended wing tips.

The reduced energy L-1011-RE shown in Figure 1-2 is the increased span airplane

with a smaller horizontal tail. Table 1-1 lists the dimensional data of the

L-1011-1 and the L-1011-RE. The aerodynamic data representing these configurations

.... /-iii

I.-.'I

71 FT.-7 IN.F_ _ I__ (21.82m)

FUSELAGE ___DIAMETER19FT.-7 IN.\

-- _ _ WING AREA(321.07m 2)

. 155 FT. 4 IN.(47.35m) 55 FT.-4 IN.

FLOOR.E.G.T-_ (,_ibm)15 FT.-2 IN.(4.62m)

29 FT.-9 IN. FT.-0 J(9.07m) (21.34m) 177 FT. 8 IN. --

(54.2m)

Figure l-l. LlOll-i General Arrangement

[_ 56 FT-6.8 INa _ (17.24 M)

19 FT-7 IN.DIAMETERM)

I,_ 164 FT-4 IN.

(50.09M) _ l55 I_T-4 IN(16.87 M)

L' I_1

[2;],

I 160 FT--10 IN-" (49.12 M) =STA STA223 2153

I-'!Lo Figure 1-2. L-lOll-RE General Arrangement

TABLE i-i. AIRPLANE GEOMETRY

L-1011-1 L-1011-RE

WING

*Reference Area, SW 321.07 m2 (3456 ft. 2) 328.97 m2 (3541 ft. 2)

*Reference Wing Chord, c 7.455 m (24.46 ft.) Not definedw

*Reference Wing Span, b 47.24 m (155.0 ft.) 50.09 m (164.33 ft.)w

Aspect Ratic, ARw 6.95 7.63

Taper Ratio, k 0.3 0.26w

Geometric Dihedral, F 7°31 ' inb'd. 7o31 ' inb'd.w 5°30 ' outb'd. 5°30 ' outb'd.

Wing Sweep at 0.25c, Aw 35.0 deg. 35.0 deg.

HORIZONTAL TAIL

Reference Area, Sh 119.1 m2 (1282 ft.2) 74.32 m2 (800 ft.2)

Reference Chord, ch 5192 m (19_42 ft.) 4.65 m (15.27 ft.)

Reference Span, bh 21.82 m (71.58 ft.) 17.24 m (56.57 ft.)

Aspect Ratio, ARh 4 4

Horizontal Tail Volumn, Vh 0.919 0.573

ELEVATOR

Reference Area (included in Sh), Se 11.85 m2 (127.5 ft. 2) 8.05 m2 (86.6 ft. 2)

Span per side, be 9.33 m (30.6 ft.) 7.04 m (23.1 ft.)

VERTICAL TAIL

Reference Area (Above WL 325), S 51.1 m2 (550 ft. 2) 51.i m2 (550 ft. 2)V

Re_erence Chord, _ 6.19 m (20.3 ft.) 6.19 m (20.3 ft.)v

Reference Span (Above WL 325), b 9.05 m (29.7 ft.) 9.05 m (29.7 ft.)V

Aspect Ratio, AR 1.6 1.6v

Vertical Tail Volume, V 0.066 0.060v

*Although actual wing dimensions are different for the L-1011-RE, L-1011-1reference dimensions were retained for aerodynamic computations.

1-4

are the flight validated L-1011-1 data plus supplementary data from Lockheed

funded wind tunnel tests of the extended wing span and the small horizontal tail.

Lockheed preliminary design studies of several advanced subsonic cruise

aircraft have shown that for a usefully wide c.g. range with a horizontal tail

sized for nose-up and nose-down control power, the aft c.g. static margin Will be

approximately zero. This led to the concentration in this task on c.g. locations

ranging from 25% MAC to slightly aft of the neutral point.

All work done in this task was performed using conventional engineering units.

Results are presented in the international system of units except where instrumenta-

tion output is reproduced directly. Output from these sources and their working

units are identified in Tables 2-2, 2-3 and 2-4. Elements of the L-1011 primary

longitudinal control system whose rigging relationships are defined in engineer-

ing units are also shown in their original form.

1-5

SECTION 2

DISCUSSION

2.1 AUGMENTATION SYSTEM DEVELOPMENT

2.1.1 Criteria

The approach to developing an augmentation system for the small-tail L-1011

active controls airplane was to use the current L-1011 in the manual control mode

as the standard of acceptable performance. The small-tail configuration with

augmented stability (L-1011-RE) was designed such that handling qualities are at

least as good as those of the current L-1011.

The L-1011 is designed to meet Part 25 of the Federal Air Regulations. This

specification tends to be of a qualitative nature, however; so a number of other

references are used to formulate a more quantitative set of design criteria.

Among the more widely recognized criteria are:

• Military Specification, Flying Qualities of Piloted Airplanes,MIL-F-8785B(ASG).

• SAE Design Objectives for Flying Qualities of Civil Transport Aircraft,Aerospace Recommended Practice, ARP 8h2B.

• Naval Air Development Center, Proposal for a Revised Military Specification,Flying Qualities of Piloted Aircraft, NADC-ED-6282.

• Wright Air Development Center, Flight Evaluation of Longitudinal HandlingQualities in a Variable Stability Jet Fighter, WADC TR 55-299 and TR 57-719.

These references are used as a source for defining dynamic response requirements,

and in particular the modal characteristics (e.g., short-period and phugoid

frequency and damping).

The technique of identifying dynamic characteristics in terms of well-separated

second-order modes of motion has come under criticism with the development of

highly augmented control systems. This has brought about the development of time

history criteria. Some of the well known time history criteria include:

2-1

• C* Time Response Criterion for Fighter Aircraft, Boeing ReportD6-178hl T/N.

• Longitudinal Handling Qualities Criteria for Large Advanced SupersonicAircraft, NASA CR-137635.

The philosophy adopted in this study is to develop an augmentation system

which generally satisfies all of the above listed criteria. Therefore, the

following handling qualities were considered in the development: l) pitch rate

and C* time histories for comparison with the current L-1011 short-period response

to a step input; 2) frequency response criteria to ensure that oscillatory charac-

teristics are within accepted guidelines and compare favorably with other transports;

3) the time-to-double amplitude criterion which governs long term pitch instability

and is of primary concern in the case of an augmentation system failure; a_d 4)

4.45 Newtons (one pound) column force per six knots speed change away from trim,

which is required by Federal Aviation Regulations.

Development of the augmentation system concentrates on the flight conditions

of primary concern in the simulation study. These conditions are cruise and the

landing approach. Although augmentation system design and analysis is restricted

to these flight conditions, handling qualities can be evaluated at any flight

condition, since a complete flight regime aerodynamic model has been programmed.

This allows investigation of handling qualities at other flight conditions that

may be critical.

The L-1011 landing gear balance requirements dictate an aft c.g. limit of

0.35_ for takeoff and landing, although in-flight c.g. locations aft of this limit

are possible for research purposes. For purposes of this study, the aft c.g. limit

is defined by the neutral point location. Previous studies have shown that

in cases of stability augmentation system failure in the landing approach, IFR

handling qualities become unacceptable for negative static margins greater than

2 or 3%. Considering the destabilizing effect of the small tail and stabilizing

effect of the extended wing tips, the netstability loss for the L-1011-RE compared

tc the current L-1011-1 is 5% at low-speed conditions and in cruise about 3% at

M = 0.80 decreasing to no loss in stability at M = 0.90 and above. Corresponding

neutral point locations are about 0.42_ for the landing configuration and in cruise

from 0.38 to 0.41_. Since the purpose of this study is to investigate the effects

of relaxed static stability, c.g. locations forward of 0.25_ were not planned for the

2-2

flight simulation, and are therefore not considered at this stage of the augmenta-

tion system development. Therefore, the c.g. range of greatest interest in this

study is 0.25 to 0.40_.

2.1.2 Mach Trim

Technically speaking the Mach trim system is part of the L-1011 augmentation

system, although it is not considered to be in this analysis, inasmuch as the

current L-1011 is equipped with Mach trim. Its purpose is to give a satisfactory

stable stick force gradient with velocity at high speed to comply with the FAR

Part 25 requirements, i.e., 4.45 Newtons (1 lb)/6 knots away from trim speed.

For purposes of this analysis, the effects of the Mach trim system have been

incorporated into the basic airframe speed derivatives.

2.1.3 Linear System Models

Control system analysis was performed using a linearized aerodynamic model.

Figures 2-1 and 2-2 show that the linear system model gives an accurate represen-

tation of the airplane response at both low and high speed conditions. Pitch rate

time histories obtained with the linear system models are nearly the same as those

from digital computer program solutions with complete nonlinear aerodynamic model.

2.2 UNAUGMENTEDCHARACTERISTICS

2.2.1 Time Histories

The C* and pitch rate time histories of the small-tail L-1011 were determined

for comparison with criteria which were delineated as study guidelines. These

guidelines are defined by time history envelopes which represent response

characteristics of the current L-1011 for a wide range of flight conditions.

The comparison is shown in Eigures 2-3 through 2-6. These figures show that

the time history characteristics of the small-tail L-1011 are generally within

guidelines except for c.g. locations aft of 0.35 _, where the response falls

below the lower boundary.

2.2.2 Characteristic Roots

Characteristic roots of the standard-tail L-1011 were evaluated to establish

minimum frequency and damping requirements of the L-1011-RE, and to determine

stability characteristics of the baseline airplane for the c.g. range of interest

2-3

1.2LINEARRESPONSE

NO'NLINEAR_

/ RESPONSE _

/0.8 /o /

'" /_3'" II /

uJ 0.6 /'_ /n,-•,- /

II--

_" 0.4 // LANDING MASS

/ Ve = 135 KEAS/ C.G. = 0.25

0.2 / _ SMALL TAIL/ aF = 33°

/ GEAR DOWNI/0

-6INITIAL TRIM

U.I

I

"' _ -7 NONLINEAR RAMP-N I

- \.,,Iw

_ LINEARSTEPi.-z

-80 1 2 3 4 5 6

TIME - SEC

Figure 2-1. Linear Aerodynamic Model Accuracy: Approach Pitch Rate

1.2I

• MID CRUISE MASSM = 0.85

H = 10,058 M

C.G. = 0.253

1.0 SMALL AlL m

x

0.8ow.I

"' NONLINEARQ.

I 0.6 RESPONSE

_- LINEA-R";" 0.4 RESPONSE

0.2

0

-1

L_UJ

n-QuJ IN :Z '_ INITIAL TRIM"J_ -2

I-_<_ NONLINEAR RAMP

'_ LINEAR STEP-3 I

0 1 2 3 4 5 6

TIME - SEC

Figure 2-2. Linear Aerodynamic Model Accuracy: Cruise Pitch Rate

2-5

1.6 LANDING MASS

Ve = 135 KEASSMALL TAIL

CRITERION _F = 330UPPER BOUNDARY GEAR DOWN

1.4 ,._G_ _ _=, _/i

// \k

1.2 // _ \• • --._ , m k

CRITERION

1o0 LOWER BOUNDARY

"- / '

0.6 _

o.,y/_/0.2

?0

0 1 2 3 4 5

TIME - SEC

Figure 2-3. Small Tail C* Time History for the Unaugmented Approach

2-6

2.5LANDING MASS

f'-'. _ Ve = 135 KEAS/ _ SMALL TAIL

I \ _F = 330

I \ GEARDOWNI \

2.0 III \ CRITERIONUPPER BOUNDARY

I \I \

O, I \

"' 1.5I-.< I \= I \o

C.G.=

- I<_ CRITERION

1.0 I LOWER BOUNDARY --

o I

-,z z / 0.40_

0.5

,y0t0 1 2 3 4 5

TIME - SEC

Figure 2-4. Small Tail Pitch Rate UnaugmentedApproach

2-7

1.6 MI[) CRUISE MASSM = 0.85H = 10058 MSMALL TAIL

f_

1.4 // _\I _, CRITERION

\ UPPER BOUNDARY

I Xk!

I1.2 I

I ,,I \\I

1.o F "_.= _° ._3 / ITERION JN 0.8 /

, /,z /

// 0.40 _ /0.6

I '/

o2 // /

0

0 1 2 3 4 5

TIME - SEC

Figure 2-5. Small Tail C* Time History Unaugmented Cruise

2-8

3.0 / .I

\ MID CRUISE MASS

" i \ M = 0.85I \ H = loo58MSMALL TAIL

25 I \\/ \ CRITERIONI \ UPPERBOUNDARY/ \I \I \\

2.0 I C.G. 0.25 C \

I _ \\

CRITERION k N

_,__._E: BOONDARY

1.0 __..0.5

00 1 2 3 4 5

TIME - SEC

Figure 2-6. Small Tail Pitch Rate Unaugmented Cruise

e

2-9

in the flight simulation (0.25 to 0.50_). Figures 2-7 and 2-8 show the effect of

c.g. location on characteristic roots for the low and high speed conditions.

The short-period frequency and damping values at mid c.g. were used as the minimum

acceptable in the augmentation system design. Mid c.g. data for the landing approach

(Figure 2-T) show a damping ratio of 0.57 and frequency of 0.86 rad/sec (0.14 Hz);

corresponding values in cruise (Figure 2-8) show a damping ratio of 0.45 and fre-

quency of 1.65 rad/sec (0.26 Hz). In the landing approach, unstable roots do not

appear until the c.g. reaches 0.50_, and in cruise a low frequency instability

appears for c.g. locations aft of 0._0_.

Figures 2-9 and 2-10 show the effects of c.g. location on characteristic roots

for the unaugmented small-tail airplane. At the mid c.g. point damping values are

essentially the same as for the standard-tail airplane; however, frequencies are

somewhat reduced: 0.75 rad/sec (0.12 Hz) in the landing approach and 1.5 rad/sec

(0.24 Hz) in cruise. In the landing approach, the configuration becomes increas-

ingly unstable for c.g. locations aft of 0.40_, and in cruise there is a low-

frequency instability for c.g. locations aft of 0.37_. These results suggest that

0.40_ may be considered as an aft c.g. limit for the small-tail airplane.

2.3 AUGMENTATION SYSTEM DEFINITION

2.8.1 Design Approach

The design philosophy for the L-1011-RE augmentation system was based on

consideration of the following characteristics of the unaugmented small-tail

airplane:

1. With few exceptions, results from the piloted flight simulation showgenerally acceptable handling qualities (pilot rating _ 6.5) for c.g.locations as far aft as 0._0_ in the landing approach and 0.35_ in cruise.In cruise with the c.g. at 0.38c, one pilot rated the airplane a 7 in alllevels of turbulence and another pilot rated theairplane a 6.5.(See Section 2.10.)

2. Normalized time history characteristics are within the selected criteriaboundaries except for c.g. locations aft of 0.35_.

3. The angular frequency characteristics are unacceptably low at mid c.g.,compared to the standard-tail L-1011, and continue to degrade as thec.g. moves aft.

2-10

!.2 LANDING MASS

Ve = 135 KEAS0.4 BIG TAIL

_F = 330

GEAR DOWN

1.0 / 0.5/

0.8 /_'o._,_

.,., 0.7\ ._ 0.6I

0.8_

0.4 _

0.2 _

0 0.50 i 0.50 0 5-1.0 -0.8 -0.6 -0.4 -0.2 0 0.2

_'_n - SEC-1

Figure 2-7. Characteristic Roots Unaugmented Standard Tail Approach

2-11

3.2 IMID CRUISE MASSM = 0.85

H = 10,058M 0.2BIG TAIL

0.460.44 _ 0.48

2.8 0.4_ _0.500.1 j 0.25C = C.G.

2.4

o I0 0.1

2.00.4

.\

_ 1.6 _ "

3

1.2 l L;1M01_ AFT

d o

0.0.8

0.50 0.46 0.44 0.44 0.46 0.500 .............

--2.0 --1.6 --1.2 -0.8 -0.4 0 0.4

_'_on -- SEC-1

Figure 2-8. Characteristic Roots Standard Tail Cruise With Mach Trim

2-12

1.0 LANDING MASS

Ve= 135 KEAS0.4 \ SMALL TAIL

\ I(_F= 33°

0 5 _ GEAR DOWN

0.6

0.7 \ G. = 0.25C"\0.6 ',

"' 0.8 .30I"o3

0.4 .

0.2 .

0.46 :44 0.460

-1.0 --0.8 -0.6 -0.4 -0.2 0 0.2

_b:n -- SEC-1

Figure 2-9. Characteristic Roots Unaugmented Small Tail Approach

2-13

3.2: MID CRUISE MP,SSM = 0.85H = 10,058 MSMALL TAIL 0.2

2.8 0.41 0 42,.._._.43

0.37

I 0.1C.G. = 0.25C

2.4

o I0 0,1

2.0

0.4ILl •

1.6

"O

3 0.6

C.G. = 0.25_"0.7

1.2

0.8

0.35

0.37

0.4 0.38

SEE INSET

.0.39

0.420.43 0.42 0.41 0.40 0.395 _0.3950.4010.41

0-2.0 --1.6 --1.2 -0.8 -0.4 0 0.4

_'_n - SEC-1

Figure 2-10. Characteristic Roots Unaugmented Small Tail Cruise With Mach Trim

2-14

Based on these findings, it was concluded that good handling qualities could

be achieved with a simple augmentation system, which would be highly reliable.

• This system was conceived as a lagged pitch rate damper to provide the necessary

short-period frequency and damping characteristics to suppress turbulence effects.

In addition to this, a washed-out stick feed forward loop was designed to provide

the flexibility of adjusting the C* and pitch rate time history characteristics

without affecting system stability. A block diagram of the L-1011-RE augmentation

system is shown in Figure 2-11.

2.3.2 Modal Characteristics Procedure

The pitch damper was analyzed by plotting the locus of short-period charac-

teristic roots as the damper gain and lag time constant were varied. Root loci

were calculated for c.g. locations of 0.25, 0.35, and O.hOc. Results of these

calculations are plotted in Figures 2-12 through 2-17. Data for the landing

approach in Figures 2-12 through 2-14 show that a lag time constant of 0.8 seconds

and a gain of at least 0.6 will give good frequency and damping characteristics

for the c.g. range 0.25 to 0.h0_. Corresponding data for cruise in Figures 2-15 to

2-17 show that a lag time constant of 0.h seconds and a gain of at least 0.3 will

be required.

Figures 2-18 and 2-19 summarize the effects of c.g. on the small-tail L'lOll

with pitch damper lag time constant fixed at 0.8 seconds for the landing approach

and 0.4 seconds for cruise, respectively. These data show that increased gain tends

to decrease the effects of c.g. but at the same time degrades the damping. Using

characteristics of the unaugmented standard-tail airplane as a reference, it was

decided that a good compromise would be a gain of 0.8 for the landing approach and

0.h for cruise.

Figures 2-20 and 2-21 show the effect of c.g. location on characteristic roots

of the small-tail L-1011-RE with the basic augmentation system (System 1). These

data show that the configuration has good short-period frequency and damping

characteristics for the complete c.g. range, compared to the unaugmented standard-

tail L-lOll with mid-c.g. (0.25c). Also, all other roots are stable for c.g.

locations aft to 0.40_. It is noteworthy that the augmentation system significantly

increases the frequency over that of the unaugmented small-tail airplane, and also,

because of the pitch damper lag, suppresses the low frequency instability of the

unaugmented airplane in cruise.

2-15

roI

I--.,

PITCH DAMPER

KQ

tWOS "fLAGS + 1

.'twoS + 1

\ _H_,_ / _ AIRFRAME ,_ Q

SML. TAIL

_COL _ FCOL

STD. TAIL TRIM

-14

6H -_

+1

Figure 2-11. L-lOll-RE Augmentation System Block Diagram

1.5

" LANDING MASS0.5 0.4 Ve = 135 KEAS

0.6 X _ . SMALL TAIL/ \\ X \ _F =330

/0.7 \ _ \ GEAR DOWN/ -" K^ = 1.0 _

ux ,,)\o_ \ \ !XL\ \1.0 _ \ X 0.6\ X0.8_ X

_ =__ \"o,. \ _oo_oo_o_o_--0.6__ _ _ _. _ STANDARD TAIL

-o _ _ 0.4 _.,....,_ 0.2 \

0.5 1,1 o._/-,,,,,\I_\ - _0.8 / "flag= 0 _ _ . \

y 1.o \1.o• "

0--1.5 -1.0 -0.5 0

_'_n -- SEC--1

Figure 2-12. Effects of Gain and Lag Time Constant Approach C.G. at 0.25_

2-17

1.5 0.4 LANDING MASS

\ Ve= 135 KEAS. _'"1 0.5 \ SMALL TAIL

.Fo_ \ \ _F=_o_ _ _ _ GEAR DOWN

0.7 _ .

1.0 0._

/ _ \0.6\ _'_\ \ (_ UNAUGMENTED-a / _ \ 0.6 \ STANDARD TAIL

3 0_8 "_ _ \\ C.G.= 0.25C"

\ oo_ _ "_o_"o2_,,\\\_ o_ o4:.o_oo_oo\ \

- _/_oo _ \\ \\

_1.0/'8 "rlag=O _,_ _0

-1.5 --1.0 -0.5 0

_'_n - SEC-1

Figure 2-13. Effects of Gain and Lag Time Constant Approach C.G. at 0.35_

2-18

1.5LANDING MASS

Ve = 135 KEAS

/oo\1o o_\ ___

0.8_ KQ---I_01_ 1_0 ' __

m 0.8 MENTED "

oo\ o=o,co._ _ \o.o o._._ \, \_ \

• \_"ag=0"6 __ __

• "_0

-1.5 -1.0 -0.5 0

_'_on -- SEC-1

Figure 2-14. Effects of Gain and Lag Time Constant Approach C.G. at 0.40_

2-19

4.0MID CRUISE MASSM = 0.85H = 10,058 MSMALL TAILWITH MACH TRIM

0.4

/_0.63.0 _

O7

\

._ .0 ,3

2"_,o;S_ANOAROTA,UNAUGMENTED L

1_b 0.0

1.0

0-3.0 -2.0 --1.0 0

• _'_n -- SEC-1

Figure 2-15. Effects of Gain and Lag Time Constant Cruise With C.G. at 0.25c

2-20

r MID

3.0 . j M = o.C8R5UISEMASS0.4 H = 10,058 M

/ _ ^ _ \ SMALL TAIL

0.6 u.o \ WITH MACH TRIM0.7o. \°° i. \,

/ \ _ / \ \ \ UNAUGMENTED

, \\_\c•_= I \\- \\o,_\\\

",0.4I "1.o

0.2 _

0.4 "

0.1

0

-3.0 -2.0 -1.0 0

J'_n -- SEC-1

Figure 2-16. Effects of Gain and Lag Time Constant Cruise C.G. at 0.35_

2-21

3.0MID CRUISE MASSM = 0.85

J 0.4 H = 10,058 MSMALL TAIL

P 0.5 WITH MACH TRIMb

0.7

2.0 0.6 =KQ_

0.8

0.5i 0.5

w

I 0.4 _) UNAUGMENTED-o STANDARD TAI L3 "flag= 0.2 C.G. = 0.25C"

0.3-- 0.3

1.0 0.4 0.30.4

Figure 2-17. Effects of Gain and Lag Time ConstantCruise C.G. at 0._0c

2-22

1.5

0.4 LANDING MASS

Ve = 135 KEAS0.5,

SMALL TAIL

_ KQ_,=I_ _F = 33°0.6, GEAR DOWN

o, Ioll .o_1.0 "

0.8 ,_1.0 _0_8_ °-u.I_ .8

I..o _ "_0"6 k _)UNAUGMENTED. C.G. = 0.25C

0.5 C.G. = 0.40C"

0-1.5 -1.0 -0.5 0

_kon -- SEC-1

Figure 2-18. Effect of C.G. Approach Lag Time Constant = 0.8 Seconds

2-23

3.0

0.4

_./ , 0.5 _ MID CRUISE MASS

\ \ M=O.S_0 6 \ | \ H = 10,058 M"\ _KQ = 0.6 \ SMALL TAIL

\ \,\ \ w,_._c.._,M0.7 7lag= 0.4 SEC.

o., \\T " \0.5 / ,_.4 I \ \ "

I x • " \ I STANDARD TAIL

_ _o_\_ c.o.=o.,_

.:..IL\. \ \ ',1.0 C _

0-3.0 -2.0 -1.0 0

_'_n - SEC-1

Figure 2-19. Effect of C.G. Cruise Lag Time Constant = 0.4 Seconds

2-2b,

1.6LANDING MASS

V e= 135 KEASSMALL TAIL

_F = 33oGEAR DOWN

1.4 0.4

0.6

\

25_"

1.0 0.7 .. /'0___ _ "N

2 TAI L0.6

0.4

0.2 \ 10"2"=! 0.40

0.25 0.30 0.35 0.38 0.40 0.42 0.44 0.45 0.45 0.44 0.42 42 0.460 ....... I ...... • ..... , u ....

--1.0 -0.8 --0.6 -0.4 -0.2 0 0.2

_'_on -- SEC-1

Figure 2-20. Effect of C,G. Approach System 1

2-25

2.8

MID CRUISE MASSM = 0.85H = 10,058 MSMALL TAI L

0.4 WITH MACH TRIM

2.4 _ I

0.62.0

0.7

1.6

0.8UJ

I

3 1.2

0.8

\

_ 25 .,-0.43

0.4

0.25 0.30 0.35 0.37 0.40 0.430 .. i .................

-2.0 -1.6 --1.2 -0.8 -0.4 0 0.4

_kon -- SEC-1

Figure 2-21. Effect of C.G. Cruise System 1

2-26

2.3.3 Time History Analysis

C* step response time history characteristics of the augmented system with

the selected pitch damper gains and lag time constants are shown in Figures 2-22 and _

2-23. Data for the landing approach in Figure 2-22 show the C* time histories are

within the prescribed boundaries except for c.g. locations forward of 0.25_. Data

for cruise in Figure 2-23 show the Cw time histories are nearly centered between

the upper and lower boundaries for intermediate c.g. positions. In order to evaluate

the effects of C* in the flight simulation, an attempt has been made to match the

upper and lower C* boundaries with the washed-out stick feed forward loop. In each

case, that is for the landing approach and cruise, it was found that an upper

boundary match was facilitated by reducing the pitch damper gain in addition to the

stick feed-forward manipulation; gains and time constants for this system are

identified as System 2. A lower boundary match was achieved by reducing slightly

the stabilizer-to-column gain; this is identified as System 3.

Augmentation system characteristics are summarized in Table 2-1.

2.4 AFT C.G. FLIGHT SIMULATION OBJECTIVES

The L-1011-RE configuration has been developed to optimize performance and

fuel economy. Flying qualities analysis and testing has been performed primarily

to support the augmentation system design. Because of the lack of proven flying

qualities design criteria for augmented aircraft and because of a need to establish

minimum flying qualities requirements for the unaugmented aircraft, a flight

simulation program was necessary to supplement the control system analysis.

Following is a list of the major objectives of the flight simulation program.

I. Evaluate minimum acceptable stability limits for operation with augmenta-tion off.

A. In high altitude cruise

B. During landing approach

C. In o_her flight conditions that may be critical

II. Evaluate augmentation system characteristics and optimize where possible.

A. In high altitude cruise

B. During landing approach

C. In other critical conditions

2-27

1.6FLIGHT CONDITIONS:

Ve = 135 KTSH = S.L.LANDING MASS

(5F = 33°1.4 GEAR DOWN

,_ SMALL TAIL_eQt

\ ,, ICG = 0.25C / ..• \ "..UPPER BOUNDARY

/. \"• \1.2

/:" -.

/.." \ "," 0.35C

1.0 0.40Ce •

i .@0.8

III ) •N I,"--I

a:: / ."O "Z

0.6/ ,"

.-"LOWER BOUNDARY

e•

0.4

SYSTEM K O TLAG K8 'Two1 _ 0.8 0.82----m 0.4 0.8 0.2 5.03 --_-- 0.8 0.8 --0.4 2.0

_t •"

0.2

••°

)*

00 1 2 3 4 5

TIME --SEC

Figure 2-22• Landing Approach C*

2-28

1.6 FLIGH]: CONDITIONS:M = 0.85H = 10058 MMID CRUISE MASSSMALL TAI L

.....• . UPPER BOUNDARY WITH MACH TRIM

14 _ Ir_\ ..:i \ Q

._ '_ •

J \ "...i \ ".

1.2 _ %. ",

/ _ \ ..

\ •

"_' • °a

_ _ °•°

1.o .T=_,,=_..=_----._

I

o /oo:o.%

°- A:¢m 0.8

.JN_;_:zO_ / ""0.6

f/ i" 1 _ 0.4 0.4." 2----u 0.2 0.4 0.5 5.0

3 --_-- 0.4 0.4 -0.2 1.0

0.4 J

0.2

d

I

!0

0 1 2 3 4 5

TiME - SEC

Figure 2-23. Cruise C*

2-29

PoI

L,3o TABLE 2-1. L-lOll-RE AUGMENTATION SYSTEM CHARACTERISTICS

LANDING APPROACHSYSTEM CRUISE SYSTEM

1 2 3 1 2 3PITCH UPPER LOWER PITCH UPPER LOWER

DAMPER C* C* DAMPER C* C*

KQ- (SEC) 0.8 0.4 0.8 0.4 0.2 0.4

_'lag- (SEC) 0.8 0.8 0.8 0.4 0.4 0.4

K_ -- 0.2 --0.4 -- 0.5 -0.2

TWO -- (SEC) -- 5.0 0.2 --- 5.0 1.0

III. Evaluate controlability of the transient response due to augmentationsystem failure.

IV. Evaluate the influence of atmospheric disturbances on flying qualities,both augmentation on and off.

V. Determine augmentation system authority requirements.

2.5 SIMULATION _TH MODEL

The simulation math model was programmed on a hybrid system using both analog

and digital computing equipment. The digital computer was used for storage and

retrieval of nonlinear aerodynamic and engine data and for integration of the

equations of motion. The analog computer was used to simulate control system

dynamics and cockpit control forces and to display the simulation outputs on strip

chart recorders. Additional parameters were recorded digitally, such as touchdown

conditions and RMS errors on final approach.

2.5.1 Aerodynamic Model

Aerodynamic data were input in the stability system of axes shown in Fig-

ure 2-2h, with an indication of signs of the parameters. The various components

of aerodynamic forces and moments were programmed over the complete range of air-

speed and altitude within the operational limits of the airplane. A complete non-

linear flexible model of longitudinal aerodynamics was programmed, since the

evaluation was concentrated on longitudinal flying qualities. A simplified lateral-

directional model was employed, to be representative of the airplane response at the

test conditions, but not containing a complete description of all systems.

Engine forces and moments were computed in the body system of axis, with each

engine of the three controlled by a separate throttle. An engine dynamic model was

programmed as shown in Figure 2-25, with time variations of Engine Pressure Ratio,

EPR, adjusted to match flight-test-derived engine accelerations and decelerations.• &

The varlatlon of thrust with altitude and Mach number was also derived from flight

test data. Aerodynamic forces and moments were transformed to body axis and combined

with the engine effects and center-of-gravity corrections to compute the total body

accelerations and moments using standard 6 degree-of-freedom equations of motion.

2-31

x'*b

/ -AZIMUTH REF

N AND rTOP VIEW

INB'DAILERON

SPOILERS 71 THRU 6

STABILIZER

6 ELEVATOR8OUTB'D AILERON

L AND p+8a

HORIZON

AFT VIEW l'

l_°' 1REF LINE. L, Nz

RELATIVF_'_.,__.__ .Z YWIND

HORIZON 11RUDDER

SIDE VIEW

%

w ._Z H

Figure 2-24. Control Surface Designation and Axes System

2-32

kzPLAP

H M M H

POSITION "_I S I EQUATION(DEG) " -16.0

4.5 DEG

I

Figure 2-25. Engine Dynamics and Derivation of EPR and Thrust

Air turbulence was simulated by inserting random velocity inputs in the

aerodynamic equations. Magnitudes and filtering of the input velocities were

controlled according to the Dryden form of the random turbulence equations,

presented in Figure 2-26. In the basic Dryden model the characteristic lengths

are reduced as a function of height near the ground. In this study the lengths

have been held at 305 m (I000 ft.). As a result the peak velocity gusts simulate

vertical and horizontal wind-shear bursts on landing approach. Flying qualities

were evaluated both in approach and cruise flight conditions in levels of turbu-

lence from still air to heavy turbulence. Heavy turbulence is defined, for this

study as 3.7 m/sec (12 fps) RMS in cruise and 2.7 m/sec (9 fps) in landing approach.

These levels were obtained from various NASA simulation reports and from observa-

tion of the load factor and glide slope excurions caused by turbulence.

2.5.2 Control System Model

The control systems programmed in the simulator included a complete dynamic

model of the longitudinal control system, and a simplified model of the lateral and

directional systems. Also included were trim controls, and flap and gear control

systems. Figures 2-27 and 2-28 are biock diagrams of the analog models of the

three primary flight systems. The longitudinal system included the "J-curve" of

stabilizer-to-column gearing of the L-1011"l airplane for tests with standardtail

size and modified "J-curve" with increased deflection limits (Figure 2-29) for

tests with reduced tail size. All other components in the system were held constant,

except for the addition of stability augmentation for some tests. The standard

L-1011-1 trim system was used for all tests. The autopilot system was not simu-

lated for these tests, since manual pilot control was assumed for all testing. It

should be noted, however, that the autopilot provides a separate backup mode in the

event of augmentation failure, and the augmentation-off conditions simulated in

this program could occur only after complete failure of both autopilots as well as

the augmentation system

2.6 SIMULATOR DESCRIPTION

2.6.1 Computing Equipment

The flight simulation equipment consisted of an Electronic Associates, Inc.

8400 digital computer, an Electronic Associates, Inc. 7800 analog computer, and a

2-34

LONGITUDINAL:

TURBU;W'N"!1/Y ;U 1 + LU SU

LATERAL:

,s/TURBV = W.N. LW _2+_ SU

VERTICAL:

q_Lw ')

L4L/"_ 1+ U S

TURBW = W'N" _ _--U ( " LW )21+ -_-- S

Lu = LW = 305 M

W.N. = WHITE NOISE

Figure 2-26. Turbulence Model

2-35

+, +

FCoL'LB'.Ho'IO OLBOE--lS('--z--/DEG DEG CONTROL WHEEL STEERING

o_-_ 0.8(DEG--'-_-,4._L_s--K) 4s+1LARGE TAIL

J(_H _///_ _L _'O u --C +./_- "" 1(INCHES) +d -14 ' vV_x_ / v 0.072S+1 _ 6H"I''

" SM..___L TAIL | r

,i.F_.-i_vlURN 0.5 IN (5H [ K(_+2. I rwoS+I

Fc(LB) 0.0125 1IN/LB I_)-* K(_ :"• fLAG S+1

MANUAL CONTROL I CRUISE LANDINGI PITCH PITCH LOWER

DAMPER DAMPER C *

KQ 0.4 0.2 0.4 0.8 0.4 0.8

I TLAG 0.4 0.4 0.4 0.8 0.8 0.8K(5 0.5 -0.2 0.2 -0.4

•I 'wo 02IAUGMENTATI. ONSYSTEM

Figure 2-27. Longitudinal Control Systems

{IN.)_RP_ 9.2°/IN.

DIRECTIONAL COORDINATION

RT '(DEG/S) 4S 1 + 1 1 a R(DEG)4S + 1 2S + 1 0.016S + 1 0,1S + 1

q LIMIT

SAT(DEG) _Ka + 30 DEG IF KEAS <164 KTS OR _F>3 DEG3S+1 + 8 DEG IF KEAS>164 KTS AND SF< 3 DEG

KR Ka

ROLL CONTROL (_F<3 1.35 0.25

1.5 _spIDEG) _F>30 6 68w(DEG ) 0.5 '_ 0.05S + 10.05S + 1

I

Figure 2-28. Lateral Control Systems

-18 -- -14

-16-12 --

__-14-_.p.

.-_ _.-11-

<_ -12 -_ _ -

'" -10 -_i

,,, a.J

"< -8-_ WlTHCAB'E/J'/ / / -= ,< -,+- STRETCH/i / TR,M / /'" // /"_ //N _ WITHOUT CABLE"J --6 - N

_ STRETCH. / / / / /

I-- -4 ---J --4

<_ CABLE STRETCHo

+oo -0 _ _ AFTFWD STOP

2 -- STOP

_1 I I I I I I il0 2 4 6 8 10 12 14

COLUMN POSITION INCHES

Figure 2-29. Column to Stabilizer Gearing (.JCurve)

2-38

trunking rack to interconnect the computers and other peripheral equipment such as,

a sound generator, visual system, motion system and the cockpit instrument and

controls.

2.6.2 Motion System

The motion system used in this simulation is a four degree-of-freedom system,

providing pitch, roll, vertical and lateral motions. The motion system provides

completely independent motion in each degree of freedom, such that full excursion

is available in any axis, independent of the excursions in the other axes. Deflec-

tion, rate and acceleration limits in each axis are presented in the following

table.

ACCEL RATE DEFL

PITCH +25°/sec2 +15°/sec +15°

ROLL +_70°/sec2 +._15°/sec +_15°

VERTICAL +0.8 g, -1.0 g +--0.3048m/sec (1 fps) +__0.3048m (i.0 ft.)

LATERAL +-0.2g +-0.3048 m/sec (1 fps) +-0.3048m (1.0 ft.)

Because of the importance of air turbulence in this evaluation, motion system

gains were optimized to present the most realistic turbulence simulation possible

within the limits of the actuators.

2.6.3 Control Column Force Simulation

The forces experienced by the pilot in the simulator are supplied by a

hydraulic column-loader which is a conventional closed-loop servo system with posi-

tion feedback and a high forward loop gain, to give high column response. The model

consists of a second order system having position and rate feedback as follows:

F___p= s__2+ Kc) + Ks.+KdX K

KS = spring rate

l

Kd = detent spring rate

KV = viscous friction

2-39

K = coulombfrictionc

l/K = system mass

X = stick position

Fp = pilot force

K was varied as a functionof trim stabilizerpositionand Mach number,as showns

in Figure 2-30. All other parameterswere held constantthroughoutthe analysis.

Figure 2-31 presentsa block diagramof the model used to generate columnforces.

2.6.4 CockpitInstrumentation

The flight instrumentdisplaysare situatedin the pilot's stationinstrument

panel as shown in Figure 2-32. Where possible actual aircraft instruments were

used, and in other cases galvanometerand synchro-drivendisplayswere used with

instrumentfaces representativeof flight hardware. Two displaysnot present in

productionaircraft,a feel force gauge and a vertical accelerometer,were included

for evaluationpurposes. For approachtesting,a cross-pointerflight director

displaywas availableon the Attitude indicator. Equationsused to drive the

cross pointersare representativeof the productionL-1011 system and are shown

in Figures2-33 and 2-34.

2.6.5 Visual DisplaySystem

The visual system is a single-windowtelevisionsystemwith a 25-inchTV

monitormounted on the pilot'sglare shield. The source of the displayedimage is

a three dimensional1500:1 scale mode! of the Palmdale,Californiaairportand

surroundingterrainmounted on a continuousmoving belt. The monitor image is

generatedby a closed circuittelevisionchannel,the camera of which is mounted on

a servo-controlledcarriagethat moves across the width of the model belt and at

right angles to its surface,thus providinglateraland vertical displacementof

the image. These movements,along with model belt motion,present the true position

of the aircraft,relativeto the airportrunway. A servo-controlledprism-mirror

system,attached to the camera,providespitch, bank, and heading displacements.

Figure 2-32 shows the view presentedto the pilot at a position 30.5 m (i00 ft.)

above the runway and I0 seconds from touchdown.

2-b,O

60

_ M = 0.86

_z 50

m

.Jo

_ 40

o

___ __ 0.70I 30

"I-

0.60

_z 20

=. 0.50

'" 0.40III •

" 10 _ _- _0.30

0-1 0 1 2 3 4 5 6 7 8 9 10

_H TRIM (DEG)

!

Figure 2-30. L-lOll-500 Pitch Feel Spr_ng Rate

roI.i:-"ro

MECHANICAL PATH

toslLIMIT •

RATELIMIT

• . . 8CoLilN)

FC (LB) (SCOL _{_S I _"? I_I SERVOHYDRO t

I POSITION 1,_ iIX DUCER

--;() ";IJ.C..WE: + FUNCTIONi

M_.T_,M E:l°O°L,_E_° ,>_. ;>:,___,,o , .._ _.T_,M

+ tOE,E,, (__ LARGE _ MACH

' _ TA?LLLF_ MACH

Figure 2-31. Pitch Feel Force System

FLIGHT INSTRUMENT DISPLAY

VISUAL RUNWAY DISPLAY

Figure 2-32. Cockpit Displays

2-43

I

°Loo_oEoI

A i '°° h Ir>,_14-oS-'_"RATEIMIT LATERAL-- FLIGHT

p.a _ _"O_kl" LIMIT '_ 0.1S+1 l" DIRI=CTOR_ COMMAND

6.74 DEG/la

(DEG)

4S4S+1

J 3 DE___GGDEG

I r ' j,3.4. 0.562 FT/S=2

RUNWAY 4S+ 1 DEG IHEADING _(DEG)

GLIDE ] +

L__ + 2DEG 60S + 2-- j-= Ny(FT/S )LIMIT 60 S + 1

Figure 2-33. Lateral Flight Director

_ 5S+ 1 (DEG) PITCH/_8 FLIGHT DIRECTOR

+ COMMAND

g 1 DEG COMM = 1 DEG ATT

)_1 30S

A NZ(_G I 30S+1 5+_ SEC

16.1 FPSg +_ 0.25+

11.3 (FT/S)

Gs--,121,-_ o5s+1i i , _a ,NT/

I;J3(0.24) DEG 5 CAPT I cG.S.I _<0"14DEG

v I" _" _'_ SECFADER ,NT I_G.S.I<o.o,DEGi

TRKTRK = (INT)o (t SINCE > 25 SEC)

CAPT• ..

I

Figure 2-34. Longitudinal Approach Flight Director

2.7 DATA ACQUISITION

Output data from this evaluation are in the form of:

i. Pilot ratings and comments on the workload required to obtain satisfactoryaircraft performance for the task being evaluated. Pilot ratings arerecorded in terms of the Cooper-Harper handling qualities scale includedas Figure 2-35.

2. Analog strip chart records of several parameters for each case evaluated.

3. Digital printout of several touchdown parameters for approach tests.

4. Statistical data computed during all approach tests and selected cruisetests.

Tables 2-2 and 2-3 present the parameters recorded on the analog strip charts

and digital recorders, respectively, including symbols and units. Table 2-_ presents

the parameters for which statistical data were computed for the approach tests.

2.8 DESCRIPTION OF TEST CONFIGURATIONS

In the flight simulation program three versions of the L-1011 aircraft were

evaluated. To validate the simulation model, and to serve as a reference for pilot

rating comparisons, the baseline L-1011-1 was evaluated in all flight conditions.

The two primary test configurations were the L-1011-500 with extended wing tips

designated herein as the baseline aircraft, and the L-1011-500 with extended wing-

tips and reduced horizontal tail area, 74.3 m2 (800 ft2) vs ll9.1 m2 (1282 ft2)

for the baseline. This small tail configuration, designated L-101!-RE herein,

was evaluated initially with no stability augmentation, and with three variations

of automatic stabilization. Figures 1-1 and 1-2 present 3-view drawings of the

-1 and the -RE configurations, respectively, and Table 1-1 presents a comparison

of basic dimensional data for the two configurations. The baseline configuration

dimensions are identical to those of the -RE except for the horizontal tail which

is the same as that shown on the -1.

The stability augmentations system Used in the evaluation, and a description

of how it was developed is presented in Section 2-3 of this report.

2.9 TEST CONDITIONS

The flight simulator testing was concentrated on high altitude cruise and

landing approach in varying levels of air turbulence from still air to heavyturbulence.

2-46

ADEQUACY FOR SELECTED TASK OR AIRCRAFT DEMANDS ON THE PILOT PILOTREQUIRED OPERATION* CHARACTERISTICS IN SELECTED TASK OR REQUIRED OPERATION* RATING

EXCELLENT PILOT COMPENSATION NOT A FACTOR FORHIGHLY DESIRABLE DESIRED PERFORMANCE

GOOD PILOT COMPENSATION NOT A FACTOR FOR /'_NEGLIGIBLE DEFICIENCIES DESIRED PERFORMANCE

' FAIR - SOME MILDLY MINIMAL PILOT COMPENSATION REQUIRED FOR (_

UNPLEASANT DEFICIENCIES DESIRED PERFORMANCE _YJ

MINOR BUT ANNOYING DESIRED PERFORMANCE REQUIRES MODERATE ('_DEFICIENCIES PILOT COMPENSATION _J

/

NO DEFICIENCIES I = MODERATELY OBJECTIONABLE ADEQUATE PERFORMANCE REQUIRES CONSIDERABLE

WARRANT F DEFICIENCIES PILOT COMPENSATION

_JIMPROVEMENT

VERY OBJECTIONABLE BUT ADEQUATE PERFORMANCE REQUIRES EXTENSIVETOLERABLE DEFICIENCIES PILOT COMPENSATION

ADEQUATE PERFORMANCE NOT ATTAINABLE WITH•MAJOR DEFICIENCI ES MAXIMUM TOLERABLE PILOT COMPENSATION. (7)

_ CONTROLLABILITY NOT IN QUESTION O

®REQUIRE MAJOR DEFICIENCIES CONSIDERABLE PILOT COMPENSATION IS REQUIRED

IMPROVEMENT FOR CONTROL

MAJOR DEFICIENCIES INTENSE PILOT COMPENSATION IS REQUIRED TORETAIN CONTROL

IMPROVEMENT MAJOR DEFICIENCIES CONTROL WILL BE LOST DURING SOME PORTION OF

MANDATORY H REQUIRED OPERATION _ I

*DEFINITION OF REQUIRED OPERATION INVOLVES DESIGNATION OF FLIGHT PHASE AND/ORSUBPHASESWITH ACCOMPANYING CONDITIONS

PILOT

DECISIONS

COOPER-HARPER REF NASA TND-5153

I

Figure 2-35. Handling Qualities Rating Scale

TABLE 2-2. RECORDED PARAMETERS - ANALOG STRIP CHART RECORDINGS

PARAMETER SYMBOL UNITS

Angle of Attack _ degrees

Pitch Attitude 9 degrees

Pitch Rate qB deg/secAltitude h feet

Vertical Load Factor nzB g's

Pilot Force on Column F lbc

Column Position 8c inches

Stabilizer Position 6H degreesEquivalent Airspeed V knotseMach Number M __

Rate of Climb R/C ft/min

Throttle Position (#2 Engine) 8th degreesEngine Pressure Ratio (#2 Engine) EPR --

Throttle Thrust (All Engines) THR lb

* Glide Slope Error GSE degrees

Localizer Error LOCE degrees

* Distance from Runway Threshold X feetrwy

Bank Angle _ degrees

Slide Slip Angle _ degrees

Roll Rate PB deg/sec

Yaw Rate RB deg/secLateral Load Factor n G's

Y

Heading _ degrees

Wheel Position •6w degrees

Rudder Position bR degrees

NOTE: Additional parameters will be recorded if required to monitoraugmentation system performance.

* These records are active only for approach testing.

2-h8

TABLE2-3. RECORDEDPARAMETERS- DIGITALLINEPRINTERRECORDING

PARAMETER SYMBOL UNITS

Test Number Test -

Run Number Run _

Elapsed Time (From Start of Run) Time seconds

Equivalent Airspeed KEAS knots

* Distance from Runway Threshold XDIST feet

* Distance from Runway Centerline YDIST feet

Rate of Climb HDOT ft/sec

Pitch Altitude THETA degrees

Roll Angle Roll Angle degrees

Vertical Acceleration C.G. ACC ft/sec2

Total Thrust (All Engines) THRJIP

Angle of Attack ALPHA degrees

Center of Gravity Location SCGTOT %MAC/100

Stabilizer Position DELH degrees

Column Position Stick inches

Gross Weight Weight lb

Pitch Rage Pitch Rate deg/sec

Heading Heading degrees

Slideslip Angle Slideslip degrees

NOTE_____:The digital line printer records these data at the end of eachrun and on command from a cockpit switch.

* These records are active only for approach testing.

2-49

TABLE2-4. STATISTICALDATAPARAMETERS

CRUISE

PARAMETER UNITS

C.G. Load Factor G'sPitch Attitude DegreesAirspeed KnotsAngle of Attack DegreesAltitudeRate Ft/sec.Stick Position InchesStick Force Lbs.Pitch Rate Deg/sec.StabilizerPosition DegreesRoll Attitude Degrees

LANDING APPROACH

PARAMETER UNITS

Glidescope error DegreesPitch attitude DegreesAirspeed Knots

Angle of Attack DegreesAltitude Rate Ft/sec.Stick Position InchesStick Force Lbs.Pitch Rate Deg/sec.Localizer Error DegreesRoll Attitude Degrees

2-50

A representative cruise condition, M = 0.83 at 10,058 m (33,000 ft.) altitude,

was used for most of the cruise testing at an aircraft gross mass of 181,437 kg

(weight of 400,000 ibs.). Other Mach-altitude combinations along the best-cruise

speed line of Figure 2-36 were used in a limited evaluation of altitude effects on

augmentation-off flying qualities. The pilots were instructed to attempt to

maintain airspeed and altitude in turbulence and to attempt small heading and

altitude changes for a qualitative evaluation of the workload associated with typical

flight tests in cruise. Additionally, intentional upsets were introduced to evaluate

their ability to recover to the initial flight conditions.

Landing approach testing was initiated at a distance of ten miles from the run-

way threshold in level flight at initial altitude of 457 m (1500 ft.) AGL. An

instrument approach was flown to 91 m (300 ft.) AGL, at which time the visual

presentation of the airport was available for final approach and touchdown. The

initial aircraft configuration was gear up and flaps at I0°. The pilots flew level

to glideslope intercept, at which time the landing gear were extended and flaps

were extended to 22° and finally to 33° as the aircraft descended on the glideslope.

A flight director representative of the L-1011 system was used for the IFR portion

of the approach.

In both cruise and approach tests, aircraft center of gravity was varied from

a mid-cg condition, 25% MAC, to a maximum aft location which was defined by

unacceptable pilot ratings (> 6.5) with the augmentation system off. The cg range

thus defined, was then used for testing with augmentation on.

2.10 PILOT EVALUATION OF UNAUGMENTED FLYING QUALITIES

In order to determine the acceptable range of aircraft center of gravity in

the event of complete augmentation system failure, center of gravity was moved aft

from 25% MAC in small increments until unacceptable pilot ratings were obtained,

both in cruise and in landing approach. This test method was repeated for three

levels of air turbulence from calm air toheavy turbulence. Figure 2-37 presents

pilot ratings of the L-1011-RE (small tail) configuration in cruise for the afore-

mentioned conditions, for two pilots. Both pilots indicated that flying qualities

were relatively good at centers of gravity forward of 30% MAC, but began to deter-

orate aft of this point. At a center-of-gravity of about 38% MAC, both pilots felt

that controllability of the aircraft required an unacceptably high workload for

2-51

40-12 --

35 LOSSINRANGE, % Ve, KTS

0 ( 28o

15 - 25

10 32 313

5 - 39 314

O- O, I I I I I0.40 0.50 0.60 0.70 0.80 0.90

MACH NUMBER

F.igure2-36. Reduced Altitude Best Cruise

2-52

lO I I I I9 PILOT NO. 2 PILOT NO. 1

TURBULENCE I I68 - C) CALM ' _/

[] MODERATEZ

/_ HEAVY "

r-i_

24 28 32 36 40 44 .24 28 .32 36 40 44

CENTER OF GRAVITY, % MAC

IFigure 2-37. Cruise Flying Qualities L-1011-RE No Stability Augmentation

flight of long duration. Aircraft response became sluggish and pitch attitude

or altitude control was extremely difficult. Turbulence did not have a significant

influence on pilot ratings in the cruise condition. In general, ratings were degraded

by one rating unit in heavy turbulence indicating that the sluggishness and inherent

lack of stability of the aircraft was of more importance than disturbances from air

turbulence.

The effect of cruise altitude on flying qualities with augmentation off was

evaluated by simulated flight at Mach-altitude combinations along the best range

line of Figure 2-36. As altitude was reduced from 10,058 m (33,000 ft.) to 4571 m

(15,000 ft.), pilot ratings improved by two rating units, as shown in Figure 2-38

and the pilots reported a significant reduction in workload to maintain airspeed

and altitude. There is a reduction in range associated with lower altitudes, also

shown in Figure 2-36, which would dictate the altitude reduction available in any

given situation, in the event of total system failure.

In landing approach, a similar test sequence was conducted for three levels

of air turbulence. Figure 2-39 presents pilot ratings obtained from these tests.

It can be seen that pilot ratings in'the landing approach are relatively insensitive

to center-of-gravity location compared to cruise results. Acceptable approaches

were flown as far aft as 4_% MAC (-2% static margin) in moderate turbulence. The

most significant affect was found to be turbulence level, because of the effects of

updrafts or downdrafts and horizontal wind shears on glidepath control. In heavy

turbulence, acceptable glideslope control was marginal at any center-of-gravity,

because of rapid excursions above or below the glideslope, which could not be

controlled to an acceptable level. These results show that the cruise flight condi-

tion dictates the aft cg limits.

In order to determine if tail size effects were restricted to static stability

differences only, the baseline configuration (with large horizontal tail) was

evaluated at comparable stability levels by flying at more aft centers of gravity.

Figures 2-40 and 2-41 present pilot ratings from tests of both configurations

plotted as a function of static margin (rather than center-of-gravity). In land-

ing approach with no turbulence, pilot ratings appear to be a function of static

margin only, but in cruise an additional effect is present. The large-tail air-

craft has improved flying qualities at a comparable stability level relative to

2-54

7 m

6 m

z 5- -- ..... -<< ...-----..................0-....m M = 0.83

_..,.-,-''" M = 0.76I"..JO4- _,,...'0" "

I_ _ M = 0.70

3-- 0 '''IM = 0.64

2--

1 I I I I I I3 4 5 6 7 8 9 10

ALTITUDE, 1000M

I I I I I I10 15 20 25 30

1000 FT

I

Figure 2-38. Effect of Altitude on Cruise Flying Qualities

I'Ul

I 1 I I10 TURBULENCE PILOT NO. 2 PILOT NO. 1

9 - (_ CALM - I I

8 - [] MODERATE_

z_ HEAVY

7 , I

z_ 6 A ._. _-."I "_ jJz< 5 _ fT

o" ..------'c] _ _.._ ]___c].In_c. 4 _ ------"

3@ --O-"" "O C _)'_

2

AFT LIMIT AFT LIMIT1

24 28 32 36 40 44 24 28 32 36 40 44

CENTER OF GRAVITY, % MAC

Figure 2-39. Approach Flying Qualities L-1011-RE No Stability Augmentation

8 -- CALM AIR--PI LOT

7 1 2

/X O SMALL TAIL

6_7 [] BASIC TAIL

L__Z5I.,-

!=

I-_.,.jO4 _ _

3 _ _ -_ [] _ [][]

21 18 15 12 9 6 3 0

STATIC MARGIN, %MAC

ro!

_-_ Figure 2-40. Approach Flying Qualities (Unaugmented)

r_I

co

10 IPILOT

9 1 2

Z_ O SMALL TAIL8 CALM AIR

_7 [] BASIC TAIL

7 _ pf

/#) //_6

,,-5I.-o-1

.....3

18 15 12 9 6 3 0 -3 -6 -9

STATIC MARGIN. %MAC

Figure 2-hl. Cruise Flying Qualities (Unaugmented)

the small-tail aircraft, particularly at lower stability levels. The increased

pitch damping provided by the larger horizontal tail is believed to be the pri-

mary reason for improved controllability in this configuration.

2.11 PILOT EVALUATION OF FLYING QUALITIES WITH AUGMENTATION

2.11.1 Augmentation Performance

The augmentation system developed for this evaluation, described in detail

in Section 2.3, is a lagged pitch damper with stick quickening for pitch response

adjustability. The three systems designated herein as #1, #2, and #3 are:

#1 - lagged pitch damper only

#2 - lagged pitch damper with increased pitch response (C*)

#3 - lagged pitch damper with reduced pitch response(c*)

Each configuration was evaluated in calm air and in heavy turbulence and compared

to the augmentation-off case by three pilots. Figures 2-42 through 2-44 present

pilot ratings for the cruise flight condition. All pilots reported that the

augmentation provided a significant improvement in controllability at aft centers

of gravity in both levels of air turbulence. There is no clear-cut preference for

one system over another, which suggests that the improvement in pitch damping

provided by all systems is more significant than differences in aircraft control

response. The pilots commented that, although they may have preferred the response

of one system over another, they could quickly adapt to any of the systems evaluated.

Figures 2-43 and 2-44 show a direct comparison between the L-1011-1 with the

current operational tail size and the small tail L-1011-RE, for identical flight

conditions flown "back-to-back". In both cases, the L-lOll-RE, with the preferred

augmentation system engaged, was rated equivalent to the L-1011-1 in calm air and

slightly better than the L-1011-1 in heavy turbulence. Figure 2-45 is a time

history of a segment of simulated flight in the cruise flight condition in heavy

turbulence, taken from a strip chart recorder. The effect of stability augmentation

on several flight parameters is shown, demonstrating the reduction in aircraft

disturbance and pilot workload required for control of air speed and altitude.

2-59

[o!G_o

9MID'CRUISE MASS I

NO TURBULENCE M = 0.83 HEAVY TURBULENCE

8 H = 10058 ML-1011-RE

°°z t5

_. AUG OFF

O 4 (OFF

NO. 1

3

25 30 35 40 25 30 35 40

CENTER OF GRAVITY -- % MAC

Figure 2-42. Pilot No. i Evaluation of Cruise Augmentation

• I9 I MID'CRUISE MASS HEAVY TURBULENCE

NO TURBULENCE M = 0.83H = 10058 M

8 L-1011-RE

+ C3L-1011-1

OFF (LARGEZ-- 5_- 1 AUG<_ NO. 1 AU(_n,-.I- NO. 2 NO. 2

;" ."--''""'_ NO. 3 L-1011-1(LARGE TAIL)

3

025 30 35 40 25 30 35 40

CENTER OF GRAVITY, % MAC

Io'_

Figure 2-43. Pilot No. 2 Evaluation of Cruise Augmentation

roI

9 I MID'CRUISE MASS INO TURBULENCE M = 0.83 HEAVY TURBULENCE

8 H = 10058 M ._L-1011-RE )

F

_:. L-1011-1 (LARGE TAIL)

_,___________ ___ .• _o.___ I __¥ No2_-_ AUGNO13 r E

_ _"f_ AUG NO. 1L-1011-1

2 --(LARGE TAIL)

1

i

25 30 35 40 25 30 35 40

CENTER OF GRAVITY - %MAC

Figure 2-_b. Pilot No. 3 Evaluation of Cruise Augmentation

..;.....

AFTCG

3.7M/SECTURBULENCE

,,~

, "

:-,- ...RUISE

M 0.83, 'H-10058 M

• " • • •• I ; I ,

AUGMENTATION SYSTEM NO.1 ENGAGEDI ;" '. ; I "T,~r1~

:-1

-v._,

10 SECONDS ._ -

-t _,...... -+----+----+---+-.--+ .--

PITCHATTITUDEDEG

CONTROLCOLUMNFORCE .'LBS

VERTICALLOADFACTORG'S

ALTITUDE1000FEET

ATTACKDEG

'PITCHRATEDEG/SEC

CONTROL 4COLUMN 2 ,POSITION ~ -1-

IN. -4-

STABILIZER 10ANGLE 5DEG 0

5I

TIME

l\)I0\W Figure 2-45. Simulated Flight in Turbulence; Effect of Augmentation

Figures 2-46 and 2-47 show a comparison of several statistical parameters for the

L-lOll-1 and L-IOll-RE in cruise. These data are in agreement with both the pilot

ratings and time history data in showing a reduction in workload with augmentation

on.

Figures 2-48 through 2-50 present pilot ratings of flying qualities during an

IFR approach and landing in varying conditions of center of gravity and air

turbulence. In calm air, a slight improvement in controllability was noted, but

because the unaugmented small tail aircraft was relatively easy to fly, the rating

improvement was small. In heavy turbulence, a significant improvement in flying

qualities was observed at all centers of gravity. The pilots were able to capture

and track the glide slope with an acceptable level of work load, even in the severe

turbulence conditions. As in the cruise condition, a "back-to-back" comparison

of the L-1011-1 and the L-1011-RE with augmentation engaged showed the two con-

figurations to be equivalent in calm air and the augmented L-1011-RE to be easier

to fly in turbulent air.

2.11.2 Augmentation System Failure

The effect of a sudden failure of the stability augmentation system both in

cruise and in IFR approach conditions was evaluated at aft centers of gravity in

heavy turbulence. In no case was any unacceptable aircraft transient or change in

required pilot technique observed. Insome cases the failure was initiated without

a cue to the pilot, and a few seconds elapsed in each instance before the pilot

became aware that the failure had occurred. It should be noted, however, that these

failures were all of the "soft" type where the augmentation input to the stabilizer

failed to a null condition. Additional analysis is necessary to determine the

possible effects of failures that could cause a "hard-over" input to the stabilizer

from the augmentation system.

2.11.3 Augmentation System Authority Requirements

The testing discussed in previous paragraphs was conducted with no constraints

on augmentation system authority or control surface rates other than those of the

primary control system. To determine the limits of authority required from the

system, several simulator runs were conducted in cruise and landing approach in

calm air and in heavy turbulence. Figure 2-51 shows augmentation system input time

2-64

L-1011-1 M = 0.83I H = 10,058 M

MID CRUISE MAS,q25 I L-1011-RE

o_ I L-1011-RE AUG NO. 1

{JI -1

35 I RE

I RE AUG NO. 1

I I I I I I I0.04 0.08 0o12 0.16 0.20 0.24 0.28

LOAD FACTOR DEVIATION (RMS) G'S

I L-1011-1

25 I L-1011-RE

o_ | L-1011-RE AUG NO. 1

u i_ 1

35 I RE

I RE AUG NO. 1

I I I I I I I0.2 0.4 0.6 0.8 1.0 1.2 1.4

PITCH ATTITUDE (RMS) DEG

| L-1011-1

25 I L-1011-RE

o_ I L-1011-RE AUG NO. 1

I -1

35 I RE

I RE AUG NO. 1

I I I I I I I0.004 0.008 0.012 0.016 0.020 0.024 0.028

PITCH RATE DEVIATION (RMS) DEG/SEC

Figure 2-46. Aircraft Response to Heavy Turbulence, Cruise Configuration

2-65

M = 0.83I L-1011-1 H = 10,058M

25 | L-1011-RE MID CRUISE MASS

0_ I L-lOll-RE AUG NO. 1

_J I -1

35 | RE

| RE AUG NO. 1

I I I0 5 10 15

CONTROL COLUMN MOTION (RMS), MM

| L-1011-1

25 | L-1011-RE

o_ | L-1011-RE AUG NO. 1fj"

I-1

35 I RE ,

I RE AUG NO. 1

i i i I I i0 10 20 30 40 • 50 60

CONTROL COLUMN FORCE (RMS), NEWTONS

| L-1011-1

25 | L-1011-RE

I L- 1011-RE AUG NO. 1

I -1

35 I RE

I RE AUG NO. 1

i I i i i I i i0.1 0.2 0.3 0.4 0.5 0.6 0.7

STABILIZER MOTION (RMS), DEG

Figure 2-47. Control Activity in Heavy Turbulence, Cruise Configuration

2-66

° , , rNO TURBULENCE LANDING MASS /

Ve = 135 KTS HEAVY TURBULENCE8 L-1011-RE

7

NO. 3 AUG.6

NO. 2z 5F-=.k.. NO.1o 4,-I

NO. 2

3 NO. 1

25 30 35 40 25 30 35 40CENTER OF GRAVITY, % MAC

!c_ Figure 2-h8 Pilot No. i Evaluation of Approach Augmentation

h)IO'_Co

9 I INO TURBULENCE LANDING MASS

Ve = 135 KTS HEAVY TURBULENCE8 L-1011-RE -

7 f

OFF

6'

L-1011-1

z 5 OF__) LARGE TAIL)NO. 3I:r,.

I-- NO. 3 AUG. _

_" _ NO. 1 NO. 2 AUG.

3 "_._ NO. 2L-1011-1(LARGETAIL)2

25 30 35 40 25 30 35 40,

CENTER OF GRAVITY, % MAC

Figure 2-49. Pilot No. 2 Evaluation of Approach Augmentation

9 I INO TURBULENCE LANDING MASS

Ve = 135 KTS . HEAVY TURBULENCE8 L-1011-RE

7

6 _,NO. 2 AUG.

z 5

• (LARGETA,L)L_,,_. 2 AUG.

: ___ _ OFF _

L-lOll-(LARGE TAIL) ""

2

25 30 35 40 25 30 35 40

CENTER OF GRAVITY - %MAC

!o_ Figure 2-50. Pilot No. B Evaluation of Approach Augmentation

SYSTEM NO.2CG = 0.25 MAC

.. .. ,'

I:·.. ,....

.: ... ,....... :

... " :.: ... ...... - ..,=:.::' :.'=: ..

:::: .:'.,,:: '--'

..... ::.:: ,.......:::

. ..

, i":: :,. :: 'I·:: :::. ..·1· .. ·I .. ·· .. · :.1 ..

,.. I.. ·

. ... -- ·~I ...... ~_.I::

RUN 3 CRUISE NO TURBULENCE

..

..

3 " .", ,'.' .~' .,:..::: :::. :i:: :": :...

2 I :;' ;m·I;.~ '8 .• '! :~; .:: .;

1 c. =::: =:::0 1

;' =-::.' :. ....--'.,•. ·of. - ..... ,.

-1:"':-2~:'"-3"' ,:::: :

liHSAS

DEG

.. '

.1 .,.•

":' ..I .. ·

''I

.

III

.I·:·

'c'-- .. - '::"

.

....n

... I:'f

.

"r,-" '.

,.." . ... ...

".. :: ...I: : I ...

/I :, :I ft' .. r ...

.' , ...I.': I'·

:' i: ..

:w·

I", .:: -: :..: '.::

;; '~. ; ~; ;

I"··'·

I,',·· ....:. ::'::.::1:: ::::.. I.:: :::;

3 :,,:':'12 ... :iI;

," "::.1 ':':o ....

-1 I,,:::;:: ":-1-." ::;:::::,.I"":::':' I ~I. ".

=~hfl~I=t=t'~:bB=f:.~::~:::d+ffi±,'±b'ilij±=tBEt2±,..±,t±=b!1±±±±:±±11··±8:=f±±fi=8=if±fEH=t=tft=Ff=RUN 4 CRUISE HEAVY TURBULENCE

liHSAS

DEG

RUN 9 APPROACH NO TURBULENCE

32

liH 1SAS 0

-2-3

. .':: :~:. !F .;.~:;

~!:: T;T :~:

:;;. :~u

--1! :..

.... I.:· =-.'

I',· ::'- .

[..I:'. :

,,·1-­I~" ....

f'

I

.

L·.

I·

'''-r;: .I,~::' ..

;..;.+-

:.·1:'= :::::",:'":,'"

...

I

...

I

r'Y,··:· :::.:::1·... :.kl· .. ....."':::, .

1"1''': " .. ,-.,~

II 11.&.W I'"

II,.. I··

c::rcc I '.'

:::.[ .

"",+:1::':1':::,"1,' .• '"

'i-I""':!\,

I·,·::':::.. ":::, .. ,F~'"'''' .

TIME

Figure 2-51. Augmentation System Authority Requirements Data

histories from four simulator runs. The two upper curves were obtained in simulated

cruise conditions, while the lower curves came from simulated landing approaches,

starting at ten miles from the airport and continuing to touchdown. The maximum

deflections occur during the approach condition, as expected and significantly

lower amplitudes are required in cruise. Based on these tests, authority limits

of at least + l-l/2° in approach and + 3/4° in cruise are required for adequate

stabilization in heavy turbulence levels.

2.12 AUGMENTATION RELIABILITY

The methodology for determining the acceptability of augmented relaxed static

stability within the philosophy of equivalence must rely on probabilistic analysis.

One approach for relating unaugmentedstatic margin and augmentation performance

with numerical probability is demonstrated in Figure 2-52. This figure presents

the landing approach simulation pilot rating data in terms of probability of

exceedance per flight. Probability is associated with pilot rating by correlating

the simulator ratings as a function of turbulence RMS velocity with a probability

of exceedance gust model of the atmosphere at low altitude. These data were plotted

for the reference big-tail configuration with 12% static margin, for the small tail

configuration with neutral static stability and no augmentation, and for the small

tail with pitch damping operating full time.

The maximum turbulence intensity in which the simulator testpilots say they

would continue a landing approach is 2.7 m/sec (9 fps) RMS. The gust model indi-

cates a probability of approximately l0-4 of encountering turbulence exceeding

this level on one landing approach of about 4 minutes duration. A current unaug-

mented airplane with anaft c.g. minimum static margin in the approach configura-

tion (12% MAC) would receive a pilot rating of 5 in that turbulence based on

Task 2 simulation data.

Using 5 as a baseline value for approach handling qualities pilot rating, it

can be determined from a weighted sum of the unaugmented and 100% augmented simu-

lation data that an augmentation failure rate of 3% would provide a probability of

exceedance equivalent to that of the conventionally stable airplane. That is, the

probability of refusing a landing due to handling qualities difficulty would be no

greater for the small tail airplane with zero static margin and 97% reliable aug-

mentation than for the big tail airplane having 12% static margin. In light of

•2-71

1 I ICONVENTIONAL 12%STATIC MARGIN

EUTRAL UNAUGMENTED-I-

__.JIt.

n-,,, 10-2w

z

Qw

u 10-3XUJ

U.

O>-

"i 10-4 2.7M/SEC (RMS) TURBm

/\)o NEUTRALo. AUGMENTED

100%10-5

'_"_. NEUTRALAUGMENTED97% RELIABILITY

10--62 4 6 8 10

PILOT RATING

Figure 2-52. Probability Approach-Handling Qualities Equivalence

2-72

years of operational experience with the existing L-lOll yaw damper, 3% is an

extremely conservative assumption for failure rate. A more realistic value for

a dual/dual system utilizing state-of-the-art avionics would be lO-4 to lO-5.

This level of augmentor reliability would make the neutrally stable augmented

airplane an order of magnitude less likely to encounter degraded handling

qualities leading to a wave-off decision by the pilot.

The preceding discussion is confined to the landing approach flight

condition. The cruise condition would appear to be more demanding in terms

of augmentation system reliability. This judgment is based on consideration

of Figure 2-37 which shows the dependence of unaugmented handling qualities

on static margin regardless of air turbulence intensity, and of Figure 2-41,

which shows that static stability alone does not organize the simulator pilot

rating data. The greater exposure time at cruise Slight condition must also

be taken into account. The cruise condition simulation data acquired in Task

2 are insufficient for application of this probabilistic technique to deter-

mination of augmentation reliabilityrequirements in this more demanding flight

condition. It is considered that the technique is valid based on the landing

approach results and additional research into cruise is warranted.

2.13 STALL DYNAMICS

Continuous Systems Modeling Program (CSMP) study of aft c.g. stall

recovery dynamics and their influence on augmentation design was originally\

envisioned as an element of this task. CSMP stall study would be of considerable

value applied to augmented stability configurations having significantly large

negative static margin (i.e., -5_ MAC or more). However, it was determined

that for the c.g. range of interest in the L-lOll derivative model of Task 2,

stall recovery is essentially a static control power problem. Therefore, this

element was deferred in favor of other flying qualities analysis.

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SECTION3

CONCLUSIONS

Based on results from Task 2 analytical and flight simulation studies, the

following conclusions can be drawn concerning augmentation system design and

unaugmented handling qualitites criteria.

For conventionally configured subsonic transport aircraft of approximately

neutral static stability:

• The philosophy of providing handling qualities safety equivalent tothat of current aircraft designed to conventional static stabilitymargins is a workable guideline for augmented stability transportaircraft acceptability.

• Expression of current aircraft characteristics in terms of wellknown frequency response and time history parameters is suitablefor augmentation system design criteria.

• Classical control systems analysis methods implemented on productioncomputing techniques are adequate for design of stability augmentation.

• Flying qualities that meet or exceed equivalence criteria can beobtained with simple lagged pitch rate damping.

• Cruise flight is the condition where unaugmented handling qualitiesare most sensitive to relaxed static stability (RSS) and hence isthe flight condition of primary interest in stability augmentationdesign.

• Handling qualities of unaugmented aircraft on landing approach areaffected strongly by turbulence level but are relatively insensitiveto RSS until static margin becomes negative.

• An RSS aircraft with pitch rate damping can have approach handlingqualities better than current aircraft on landing approach.

• Airframe motion in turbulence is less for an RSS configuration

with augmentation than for current transports.

• In turbulence, pitch control surface activity for an RSS aircraftwith augmentation operating is'less than with augmentation off,but more than for current unaugmented aircraft.

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SECTIONh

RECOMMENDATIONS

On the basis of results from the flight simulator testing of an L-lOll

derivative with a smaller horizontal tail and augmented stability, it is

recommended that a flight test program be undertaken. Testing should be planned

to validate unaugmented flying qualities, to test simulator defined control laws,

and to demonstrate the acceptability of a smaller horizontal tail.

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