contractor report v. · 2-1 linear aerodynamic model accuracy: approach pitch rate 2-h 2-2 linear...
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_,--' 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 •
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
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/\)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.
2-73
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.
3-1
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.
h-1
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