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1. Report No. 2. Government Accession No.
NASA CR-2144
4. Title and Subtitle
AII_CHAFT tIANI)LING QUALITIES DATA
7. Author(s)
Robert K. tleff!cy and Wayne F. Jewell
9. Performing Organization Name and Address
Systems Technology, Inc.
ttawthmme, Califon_ia 90250
12. Sponsoring Agency Name and Address
National Aeronautics m_(t Space A(ln_inistration
Washington, I).C. 20546
3. Recipient's Catalog No.
5._Report Date
l)ecemoer 1972
6. Performing Organization Code
8. Performing Organization Report No.
Technical Report 1004-1
10. Work Unit No.
11. Contract or Grant No.
NAS 4-1729
13. Type of Report and Period Covered
Contractor Report
14. Sponsoring Agency Code
15. Supplementary Notes
16. Abstract
Available information on weight and inertia, aerodynamic derivatives,
control characteristics, and stability augmentation systems is documented
for 10 representative contemporary airplanes. Data sources are given
for each airplane. Flight envelopes are presented and dimensional deriv-
ativcs, transfer functions for control inputs, and several selected handling
qualities parameters have been computed and are tabulated for 10 different
flight conditions including the power approach configuration. The airplanes
documented are the NT-a3A, F-104A, F-4C, X-15, HL-10, Jetstar,
CV-880M, B-747, C-5A, andXB-70A.
17. Key Words (Suggested by Author(s))
llandling qualities, transfer functions, stability
derivatives, airplanes, control systems
18. Distribution Statement
Unclassified- Unlimited
19. Security Classif. (of this report} 20. Security Classif. (of this page)
Unclassified Unclassified
21. No. of Pages
343
22. Price*
$6.00
' For sale by the National Technical Information Service, Springfield, Virginia 22151
I. INTRODUCTION .
ii. NT-33A • • •
III. F- IO_A . . .
IV. F-4C • • •
v. x-15 • • •
VI. HL-IO • • •
VII. LOCKHEED JETSTAR
viii. CONVAIR 880M .
D{. BOEING 747 • •
x. C-_A ....
XI. XB-70A • • •
TABIZ OF COI_TERTS
• I
• • • • • • •
32
• . . 61
• . . Io8
• • • 137
• . . 166
• • • 193
• . . 210
• . . 2_3
• • • 273
APPENDIX A. AXIS SYSTEMS, SYMBOLS, COMPUTER MNEMONICS,
AND DERIVATIVE DEFINITIONS ....
APPENDIX B. TRANSFORMATION OF STABILITY AXIS
DERIVATIVES TO BODY AXIS ......
APPENDIX C. EQUATIONS OF MOTION AND TRANSFER FUNCTIONS . . .
• . A-I
.... B-I
• C-I
iii
SECTION I
INTRODUCTION
The purpose of this document is to provide handling qualities investi-
gators with readily usable data on several representative contemporary air-
craft. Included are those data required to obtain transfer functions relating
the aircraft's response to control inputs. An analytical description of the
aircraft's stability augmentor is also given.
For those aircraft for which complete information was available, the
following summarizes the contents and presentation:
7. Flight conditions for which computations are made
including:
a. Configtu_ations (e.g., fuel load_ flaps,
gear, etc.)
b. Mach/altitude combinations
2. General arrangement
3- Control system description
4. Stability augmentation description
5- Tabulations and/or plots of non-dimensional stability
derivatives for trimmed flight
6- Dimensional, mass, and flight condition parameters
7- Dimensional stability derivatives
8. Transfer functions for control inputs
9" Selected_andling qualities parameters
10. Data sources
A page number cross index is presented in Table I-1.
The intention has been to make this report completely self-consistent
insofar as symbols, nomenclature, definitions, etc. The system used is
described in three appendices. Appendix A covers axis systems, symbols
and notation, and definitions of nondimensional and dimensional stability
derivatives. Appendix B gives the axis system transformations for the
derivatives. Appendix C includes the aircraft equations of motion and
transfer functions used herein.
X
I
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o
i
i
<_ o_ _ _
_ _ _o_
o_ _ _o _ _ _''_'"
_ aa
. _. . _
The aircraft considered in this report span a wide range of sizes, speeds,
and uses. In each case_ transfer functions and handling qualities parameters
were computed for flight conditions which were selected to cover the flight
regimes of interest. A nominal configuration (generally cruise) was picked
for all up and away flight conditions. For this nominal configuration, plots
of trimmed non-dimensional aerodynamic force and moment coefficients are
presented. Also, in most cases, a power approach case is presented along
with a tabulation of aerodynamic coefficients. The coefficients are based
on rigid wind tunnel data 3 estimated flexible data_ or flight test results,
depending upon availability. This is indicated by the words "rigid, "
"flexible," and "flight" on each aero data plot. Also, the axis system is
indicated by "stability" for a body-fixed stability axis system or 'body"
for a body-fixed system aligned with the F.R.L. (Further clarification of
axis systems used is given in Appendix A.) Descriptions of control systems
and stability augmentation systems are given along with transfer functions.
Where a longitudinal control system has a significant effect on the equations
of motion (as with a bobweight) the stick-free transfer functions and handling
qualities are given.
Transfer functions are always given for body axis motion quantities.
Handling qualities parameters are also given in the body axis. All accelera-
tion transfer functions (az and 4) are for the pilot's position. Thrust
transfer functions do not include any engine response characteristics.
A substantial portion of this report is in the form of computer printout.
The mnemonics used in this printout are defined in Appendix A.
The handling qualities parameters given in this report represent only a
small fraction of those developed over the years. The majority presented here
are used in past and present versions of MIL-F-8785. Although only SAS-off
values are shown, the definitions given in Appendix A are general and could
be used in conjunction with the HAS-on transfer functions to yield SAS-on
handling qualities parameters.
While complete coverage of each aircraft including only the "latest" and
'_oest" data would be desirable, the major criterion used was that the data be
accessible to the author. This is why only isolated flight conditions are
given for some aircraft, and also why, as those people more intimately familiar
3
with each particular aircraft will recognize, the data presented mayrepre-
sent an early estimate in the design process and perhaps the "nominal config-
uration" is one which never left the drawing board. The data have been reviewed
and, although not all those presented indicate unquestionable trends, those
data knownto be based on only early "guesstimates" or showing unreasonable
trends have been deleted. In somecases data were estimated by the author.
As to how well the data can be expected to match the flying aircraft, it is
assumedthat those for whomthis document is intended knowwell the difficulties
of obtaining derivatives from flight test data. Every attempt has been made
to insure reliable translation, interpretation, and transcription of the data
from their source documents.
The manufacturers of the aircraft described herein can not be held account-
able for the information presented, nor would they be bound to concur in any
conclusions with respect to their aircraft which might be derived from its use.
4
 -33A ACXSXOmm
"The NT-33A variable stability airplane (Serial No. 51-4120)
is an extensively modified T-33 jet trainer. The elevator,
aileron and rudder controls in the front cockpit are disconnected
fromtheir respective control surfaces and have been connected
to separate servomechanisms that make up an 'artificial feel'
system. In addition, the elevator, aileron and rudder control
surfaces have been connected to individual servos which can be
driven by a number of different inputs. These servos receive
their electrical inputs from the artificial feel system (pilot's
commands, position or force), attitude and rate gyros, accelero-
meters, dynamic pressure, _ vane and _ probe. This arrangement,
through a response-feedback system, allows the normal T-33
derivatives to be augmented to the extent that the handling
qualities of many existing airplanes, future airplanes or hypo-
thetical research configurations, can be simulated. The original
T-33 nose section has been replaced with the larger nose of an
F-94 to provide the volume required for the electronic components
of the response-feedback system and the recording equipment."*
Transfer functions are given for only the primary surfaces and engine
thrust although the NT-33A also has other control surfaces and a range of con-
trol crossfeed and feedback combinations.
Aerodynamic data, for the most part, was taken from AFFDL-TR-70-71. However,
longitudinal data for the high lift configuration was obtained from LAL 127
andMach number derivatives from NACA-RM-7116.
6
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NT-33A
PITCH AXIS
Variable
Feel
Input
FST(Ib) _..026
I
s z +.89s + 22.5
Variable
Stability
Input
_]ST(in) ] -I0 _._---- _.3 Be(rod)
ROLL AXIS
Variable Variable
Feel Stability
Input Input
FLAT (ib) I OST _,n I0 _a(rad)sT
YAW AXIS
Variable
Feel
Input
FpED(Ib)_
I
78
Variable
Stability
Input
I_PED(in) I 2.34 _.= _._ 8r(rad)
Feel system parameter values shown correspond
to the "Front Seat Engage" mode (normal NT-33)
Figure 11-3. NT-33A Control System
9
TABLE 11-I
Power_oach Non-Dimensional Stability Derivatives
h = sea level
VTo = 228 ft/sec = 139 kt
oo = 2.2 °
Longitudinal
cL = .813
cD = .139
CLm = 5.22/rad
CD_ = .94/rad
% = --.401/rad
Cmq : -mo/raa
:
CL5e = .34/rad
Cm6e = -.89/rad
Lateral-Directional
(Stability Axis )
cy0 = -.72/r 
Cn_ = .049/rad
C_0 = --.127/rad
C_p = -.O7/rad
C_p = -.045/rad
C_r = .20/rad
Cnr = --.16/rad
Cn5 a = --.O09/rad
C_5 a = .14/rad
CYSr = .17/rad
Cnsr = -.O73/rad
C_5 r = -.OO2/rad
10
Clo
(deg)
14
12
10
8
6
w
B
n
D
D
4-
Z-
0 o
' SL NT-33A
.... 20,000 ft 13700 Ib
------ 40,O00ft .263
Rigid
\
11
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F-r'-
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0
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(red -I )
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0
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!
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13700 Ib
Rigid
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..... 20,O00ft
-------- 40,000 ft
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I (_ I I I
/
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0
0
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Cmq
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-8
-12
-16
Mach
• _.2 .4 .6 .8
Cm&
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14
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1.0
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| _ .4 Mach .6
NT-33A
13700 Ib
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Rigid
SL
..... 20,O00ft
------- 40,O00ft
CD M
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Mach
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0
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Moch
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15
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Rigid
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Mach
0 o
Moch
.2 .4 .6 .8
-.4
Cm_ e
(rod -j )
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1 _;
0
0
Cy_
(rad-i) -.4 --
Mach
.2 .4 .6 .8
1 I I I
SL
..... 20,O00ft
------- 40,O00ft
NT-33A
13700 Ib
Stability Axes
Rigid
Cn/_
(rad -i )
.2
0
-.2
I I I I
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f.®." /
17
Mach
0 .2 .4 .6 .8
0 I t l
c.tp
(rod "I)
-.2
-.4
m6
Boa
- - SL
..... 20,O00ft
--- "-"- 40,000 ft
NT-S3A
137001b
Stability Axis
Rigid
.O4
0
Cnp
(rod "l)
-.04
-,08
_4r _,f
C._r_
Cn r
(rad "l)
.3-
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0
'.1
SL NT-33A
• 20,000 ft 13700 Ib
------- 40,O00ft Stability Axis
Rigid
\',, \.
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C._r
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Mach
Cn r
t9
C}s a
(rad "l )
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.12
' SL
..... 20,000 ft
40,O00ft
NT-33A
13700 Ib
Stability Axis
Rigid
0 I I I I
0 .2 .4 .6 .8
Mach
.01
0
Cn8 o
(rod -j )
-.01
-.02
Mach
.2 .4. .6 .8
¢
So is sum of both right and left
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2O
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Mach
0
0
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Cn8 r
(rad"I)
-.08
Mach
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C,_sr
(rad "l )
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0
0
, SL
--- ?_O,O00ftm *==
---- _- 40,O00ft
NT-33A
13700 Ib
Stability Axis
Rigid
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3O
NT-33A DATA SOURCES
Hall, G. Warren, and Ronald W. Huber, System Description and
Performance Data for the USAF/CAL Variable Stability T-33
Airplane, Air Force Flight' D_nsmics Laboratory Rept. No.
AFFDL TR-70-71, Aug. 1970
Tests of a I/5 Scale Wind Tunnel Model of the TP-80C Trainer,
Lockheed Aerodynamics Laboratory Rept. No. LAL 127, Jan. 23, 1948
Cleary, Joseph W., and Lyle J. Gray, High Speed Wind-Tunnel Tests
of a Model Pursuit Airplane and Correlation with Flight-Test
Results, NACA-RM-7116, .Jan. 21, 1948
Statler, Irving C., et al, The Development and Evaluation of the
CAL/Air Force Dyuamic Wind Tunnel Testing System_ Part l--
Description and Dynamic Tests Of an F-80 Model, A_'_'DL-TR-66-153,
Feb. 1967
Flight Manual_ USAF Series T-33A Aircraft, T. O. IT-33A-I.
31
SECTION III
F-IO4A
32
F- 104A BACKGROD'AID
The F-IO4A is a single place_ lightweight_ supersonic air superiority
fighter powered by a single turbojet engine with afterburner. The wing has
a full span leading edge flap. Trailing edge flaps have a blowing-type
boundary layer control system. Control is provided by conventional ailerons
and rudder and an all-movable stabilizer. Pitch_ roll_ and yaw dampers are
incorporated_ however their effect is not shown here. Pitch and roll con-
trols are fully irreversible while the yaw control is a cable-actuated rudder
without boost. A bobweight is used in the longitudinal feel system. Its
position is assumed to be at the pilot's location.
The primary source of data was LR 10794. Drag information was obtained
from LR-12873.
The nominal configuration used here is the combat loading for the F-]O4A
based on actual weight and balance data. The PA configuration is a typical
loading at flight manual approach speeds.
33
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35
PITCH AXIS
F-104A
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32.2 F B
Oz
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az assumed to be at pilot location
G
57.3
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F LAT --I I
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2.0
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Figure 111-3. F-IO4A Control System
36
Power Approach Non-Dimensional Stability Derivatives
h = sea level
VTo = 287 ft/sec = 170 kt
_o = 2"3°
_s = --7.1°
Longitudinal
CL = .735
% = .263
CL= = 3.44/tad
CDa = .45/rad
Cm_ = -.6_/ra_
Cma = --I.6/rad
Cmq = ->.8/ra_
Cg_s = .68/tad
Cm_s = --1.4g/rad
37
Lateral-Directional
(Stability Axis )
Cyp = -1.17/rad
cn6 = .5o/r_
C2p = --.175/tad
C_p = --.285/tad
Cnp = --.14/rad
C_r = .26_/rad
Cn r = --.7_/rad
Cnsa = .O0_2/rad
C£5a = .039/rad
Cy_r = .2OS/ra_
C_Sr : .045/rad
C_ r = --.16/rad
CY_d = • 0325/rad
CnSd = --.025/tad
Cg5 d = -.O044/rad
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59
F- 104A DATA SOURCES
Stabi? [ty and Control and Handling Qualities_ F-IO4A, Lockheed Rept.
No. LR 10794, 12 Dec. 1955
Andrews, William H., and Herman A. Rediess, Flight-Determined Sta-
bility and Control Derivatives of a Supersonic Airplane with a
Low Aspect-Ratio Unswept Wing and a Tee-Tail, NASA Memo 2-2-59H,
A!or. '1959
Performance t F-IO4D, Lockheed Rept. No. LR-12873, I May 1958
Flight Manualt F-IO4A and F-IO4B USAF Series Aircraft, T. O. IF-IO4A-I,
15 Dec. 1961
Technica ! Manual t Flight Controls t USAF Series F-IO4A and F-I04C
Aircraft, T. O. 1F-104A-2-8, 15 Mar. 1960
6O
SECTION IV
F-4C
61
F-_C BAC_DROUND
The F-4C is an Air Force tactical fighter whose primary mission is
all-weather air-to-air missile combat. Lateral control is achieved by
ailerons in combination with spoilers on a swept wing. A swept stabilator
provides longitudinal stability and control. Directional stability and
control is accomplished through a conventional fin-rudder combination.
Landing speed is reduced by full span leading edge flaps and inboard plain
trailing edge flaps in conjunction with blowing-type boundary layer control
(BLC). Boundary layer control is automatically induced when full flap
deflection occurs.
Features distinguishing the USAF F-4C from its Navy counterpart, the
F-4B, are:
• Lack of drooped ailerons with flaps down resulting in
higher landing speeds.
• Dual flight controls resulting in slightly increased
control system inertia.
• Wing bumps to house larger main gear wheels resulting
in a slight drag increase.
Data included here was obtained primarily from MAC Report No. 9842.
Special emphasis is placed on the longitudinal control system because of
its relative complexity when compared to other aircraft. Figure IV-4 has
been addqd to help illustrate this system. Also, care has been taken to
0
retain s_m% of the control system nomenclaure used by the manufacturer,
e.g., qB and PBF (see Fig. IV-5).
The Stability Augmentation block diagrams are shown in Fig. IV-7. The
roll SAS described is not included in lateral directional SAS on transfer
functions since it is faded out with the lateral control stick out of neutral
position.
62
@
(9
i-.-]
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68
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64
PITCH AXIS
FST(Ib__ _-_.0369s2 +.208s
F-4C
Feel System
_SsAs (rad)
I _ Gearing / Actuator
18sT(in)_ _
.0569qePsF
,:_1 I _'-_I- _ -I°_'+'1
_-- See Fig and for feel system details
Bobweight
az ts---_-_) of JtB= 39.3 ft
+ .OI57qBPBF -*
8s(rad)
ROLL AXIS
8OsAs (rad)
Feel Spring Gearing /
_;T,,_ -I 2"961 =1_ i-_
AR! Gain
C38r {-.46 SAS OFF
= _-.69 SAS ON
CLEAN O ,/
Sa(rad)
Spoiler
-__ 8sp( ra d )
PA ARI
__r I___--
_-:1_ <_rAm(rod)
YAW AXIS 8rARl(rad) 3rsAs (rad)
\ / Rudder
Feel Spring Gearing \ / Flexure
.._j_'_ SPED(i n) .___.J__
_"_°_'_- I _ I _I_ I- _ - l""txl -
K mR G air
V<235KIAS 36.61b/in -11.5deg/in
V>220KIAS 8.51b/in -6.5deg/in
8r(rad)
_-- See Fig
Figure IV-3. F-hC Control System
65
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66
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©
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q
.6
-4C
/ _---- 35,00Oft 389:)4 Ib
.... 55,00Oft .289
I I I I 1
O0 4 8 12 1.6 2.0
Mach
PBF
(ft z)
J.I m
1.0-
8
I
0
I I L 1 I
.4 .8 1.2 1.6 2.0
Mach
h
(ft)
60k
40k
20k
Viscous Damper Off Stop
3.03 Ib/in/sec
"l//f/[_ I
Viscous Domper On Stop
b=_
I I I I I
SLo" .4 .8 1.2 1.6 2.0
Mach
Figure IV-5. F-4C Feel System }arameters
6?
I
I
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!
i
- I, 00
I_:
//,,,///7i °/
000000
oOOq
LO
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68
F-4C
PITCH SAS
E}(rad/sec) _1 .15s
- I s+l
_SsAs(rad)
ROLL SAS
PG(rad/sec) _I -.265 I _--- 8asAs(rad)
P6 = P (Roll rate gyro assumed aligned with FRL)
Note." Roll GAS faded out with lateral control
out of neutral
YAW SAS
rG (radlsec)
' (ftlsecz)ay
_I,s
_1 S+'5
I--I .0168
_rsAs (rad)
rG = r cos(-I.5 °) +p sin (-I.5 °)
I
ay = ay + 9.9 _"-.391b
Yaw rate gyro inclined 1.5° below FRL and
lateral accelerometer at ES. 198.0and W.L.23.0
Figure IV- 7. F-4C Stability Augmentation
69
TABLE IV- I
F-_C
Power A_roach Non-Dimensional 8tability Derivatives
h = sea level
VTo = 230 ft/sec = 136 kt
% = 11.7°
_s = -9 .1°
Longitudinal
CL = .915
CD = .242
CL_ : 2.8/rad
CD_ = .555/rad
Cm_ - .098/rad
Cm& - .95/rad
Cmq = -2.0/rad
CL5 s = .24/rad
C-mss = --.322/tad
CD5 s = --.14/rad
Lateral-Directional
(Stability Axis )
Cy6 : --.655/rad
Cnl3 : .199/rad
C26 = -.156/rad
C£p = --.272/rad
Cnp = -.013/rad
C_r = .20_/rad
Cnr = -.320/rad
CYSa = --.0359/rad]
Cnsa = --.O041/rad I
= .o 7/r aj
CYSr = .124/rad
Cnsr = --.072/rad
C_5 r = --.O009/rad
Spoiler
Effects
Included
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106
F-4C DATA SOURCES
Bonine, W. J., et al, Model F/RF-4B-C Aerodynamic Derivatives,
MAC Report 9542, I0 Feb. 1964
Crawford, W. N., and G. Nadler, Static and Dynamic Control System
Characteristics for the F-4 Aircraft, MAC Rept. F21_, 16 Dec. 1966
Bridges, B. C., Calculated Longitudinal Stability and Performance
Characteristics of the F-hB/C/D/J' and RF-hB/C Aircraft plus the
AN/ASA-32H Automatic Flight Control System, MAC Rept F934,
19 Apr. 1963
Bridges, B. C., Calculated Lateral-Directional Stability and Perfor-
mance Characteristics of the F-4B/C/D/J and RF-4B/C Aircraft plus
the AN/ASA-32H Automatic Flight Control System, MAC Rept. F935,
3 May 1968
NATOPS Flight Manual_ Navy Model F-4B Aircraft, NAVAIR 01-245 FDB-I,
I Nov. 1966
107
SECTION V
X-15
108
X- 15 BACKGROUND
The X-]_ is a single-place, rocket-powered airplane designed for flight
at hypersonic speeds and extreme altitudes. The airplane is carried aloft
under the right wing of a B-52 and is launched at an altitude of about
45_000 ft and a Mach number of about 0.80. After launch_ the X-J5 performs
a powered flight mission_ followed by a deceleration glide prior to vectoring
for a landing. With this operational technique, the airplane is capable of
attaining a Mach number of 6 and can be flown to and recovered from an altitude
in excess of 300_000 feet.
Flights to high altitudes have been made with all three of the X-J5
airplanes in two configurations: the basic and the ventral off. The basic
configuration is considered here.
Aerodynamic control is provided through conventional aerodynamic surfaces_
with vertical surfaces used for yaw control and the horizontal tail for both
pitch and roll control. All of the aerodynamic control surfaces are actuated
by irreversible hydraulic systems. Control force is provided by bungee for
pilot feel. A conventional center stick is used for pitch and roll control_
and rudder pedals are used for yaw control; however_ a side-located stick is
provided for control of pitch and roll in high-acceleration environments at
the option of the pilot. Most of the X-15 missions have been made with the
side stick_ although the pilots used the center stick on their first flights.
Only the center stick control is shown here.
The augmentation system shown in this report consists of angular rate
feedback loops about all three axes. In addition to the normal p -_$a roll
SAS loop_ there is an r -_5 a feedback known as the YAR loop. The gains for
each SAS loop are manually set by the pilot. The SAS-on transfer functions
given for this airplane assume maximum gain settings for each loop. This may
not have been realistic for actual flights.
The flight conditions considered for this airplane are all for straight
and level trimmed flight. This is definitely unrealistic for this airplane_
however_ the intent here is to show general speed and altitude variation
effects.
109
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Figure V-3. X-15 Control System
112
X-15
PITCH SAS
_(rad/sec) .75 SssAs(rad)
ROLL-YAW-YAR SAS
p (rad/sec)
Roll Gain
Yar Gain
BasAs(rad)
r (rad/sec)
Yaw Gain
.3O BVSAs(rad)
Note:
Gains variable in 10% increments of the
maximum values which are shown above.
(e.g. roll gains selectable are .05,.I0,.15,
.2_0, .25, .30, .35, .40, .45, and .50)
Figure V-4. X-15 Stability Augmentation
1 13
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