AEDC.TR.80.60

P T

Captive Aircraft Testing at High Angles of Attack

R. W. Butler and J. P. Christopher ARO, Inc.

August 1981

Final Report for Period October 1976 ' September 1978

Approved for public release; distribution unlimited.

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AEDC-TR-80-60

4 T ITLE (and Subrllte)

CAPTIVE AIRCRAFT TESTING AT HIGH ANGLES OF ATTACK

? AU THOR(a)

R. W. B u t l e r and J . P. C h r i s t o p h e r , ARO, I n c . , a S v e r d r u p C o r p o r a t i o n Company

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wind tunnel tests captive tests wind tunnel models computer applications angle of attack maneuverability captive flight

20 ABSTRACT (Cenl~ue ~ reverse side I t ne©ensaty and Identity ~ bJock n m b e ~

A captive testing technique capable of simulating aircraft maneuvers at high angles of attack has been developed and tested in the Aerodynamic Tunnel 16T of the Arnold Engineering Development Center (AEDC). The captive technique uses a closed loop system between wind tunnel, model, and computer. Use of the wind tunnel as an analog forcing function eliminates the requirement for a con- ventional static data matrix in motion simulation. Various high

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a n g l e - o f - a t t a c k m a n e u v e r s ( r u d d e r r o l l , a i l e r o n r o l l , w i n d - u p t u r n , e t c . ) w e r e g e n e r a t e d w i t h t h e c a p t i v e s y s t e m and show good c o r r e l a - t i o n w i t h f l i g h t d a t a .

AFSC A~ Id AFB T •nn

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A E D C-T R -80 -60

PREFACE

The work reported herein was conducted by the Arnold Engineering Development Center (AEDC), Air Force Systems Command (AFSC). The results of the research were obtained by ARO, Inc., AEDC Group (a Sverdrup Corporation Company), operating contractor for the AEDC, AFSC, Arnold Air Force Station, Tennessee, under ARO Project Number P32A-K9A. The Air Force project manager was Mr. A. F. Money, AEDC/DOT. The manuscript was submitted for publication on October 13, 1980.

Messrs. R. W. Butler and J. P. Christopher are currently employed by Calspan Field Services, Inc., AEDC Division.

C O N T E N T S

A E D C-T R -80-60

Page

1.0 I N T R O D U C T I O N " ° ° " ° ° ' " ° " " " ' " ° ° ° " ° ° ° ° ° ° ° ° ° o ° ° ° ° ° - ° , ° ° ° , . ° ° ° . . . . o , ° , ° , ° , 5

2.0 A P P A R A T U S

2.1 Test Facility .......................................................... 6

2.2 Test Conditions ....................................................... 6

2.3 Model Description ..................................................... 6

3.0 P R O C E D U R E

3.1 Captive Motion Simulation ............................................. 7

3.2 Precision o f Data ...................................................... 8

4.0 RESULTS AND DISCUSSION

4. ! General ° " ° " ° " ° ° ° ° " " ° ° ° ° * ° " " ° ' • • • ' • • • • • ° , o • ° ° ° • ° ° ° . , ° ° ° o • ° • ° ° ° , ° ° o ° o ° . 9

4.2 Lateral/Directional Commanded Motion ................................. l 0 4.3 Wings-Level Stalls ..................................................... I 1

4.4 Wind-Up Turn to Stall ................................................. 11

5.0 C O N C L U D I N G REMARKS ............................................... 12

R E F E R E N C E S • .-........................................................ 13

I L L U S T R A T I O N S

Figure

I. Aircraft Model Used in Captive Demonstrat ion ................................ 15

2. Aircraft Model Installed in Tunnel 16T Test Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . ! 6

3. Aileron Drive Mechanism Developed for Captive Model . . . . . . . . . . . . . . . . . . . . . . . . ! 7

4. Captive Closed Loop Data Acquisition System ................................ 18

5. Tunnel 16T Angle-of-Attack and Angle-of-Sideslip Capabilities . . . . . . . . . . . . . . . . . . 19

6. Estimated Uncertainties in Wind Tunnel Parameters . . . . . . . . . . . . . . . : . . . . . . . . . . . 20

7. Full Lateral Stick Roll Maneuver ............................................ 21

8. Full Rudder Pedal Roll Maneuver ........................................... 23

9. O n e - " g " Stall Maneuver ................................................... 25

10. Abrupt Stall Maneuver .................................................... 27

I I. Wind-Up Turn Maneuver .................................................. 29

T A B L E

1. Characteristics o f Aircraft Simulated in Captive Testing . . . . . . . . . . . . . . . . . . . . . . . . 31

3

A E D C-TR -80-60

Page

APPENDIXES

A. Euler Equa t i ons o f Mot ion ................................................. 35

B. Simulated Flight Con t ro l System ........................................ . . . . . 39

N O M E N C L A T U R E ....................................................... 43

4

A E DC-TR-80 -60

1.0 INTRODUCTION

During the past five years a number of experiments have been conducted at the Arnold Engineering Development Center (AEDC) for determining the feasibility of simulating aircraft motion in the wind tunnel. This motion simulation technique is referred to as Captive Aircraft Testing. With this type of testing aircraft maneuvers may be simulated in the wind tunnel without the need of the conventional complex and costly static data matrices. The wind tunnel acts as the static analog forcing function, providing an infinite static data matrix which is independent of model configuration. Maneuver limitations are purely a function of wind tunnel model-positioning hardware (support travel) and model- associated hardware (remote control surfaces). With the captive testing tool experimentalists may examine the flight characteristics of aircraft in a more accurate and efficient manner than heretofore possible through conventional motion simulation techniques.

Evolution of Captive Aircraft Testing at AEDC is documented in Refs. l, 2, and 3. The initial system development and checkout (Ref. 2) were conducted as a pilot test in the AEDC Aerodynamic Wind Tunnel (4T) using an A-7 aircraft model with fixed control surfaces. The success of the 4T experiment led to a series of additional tests (Ref. 3) in the AEDC Propulsion Wind Tunnel (16T). These later 16T experiments utilized an F-15 aircraft model with remote control horizontal stabilizers which provided the capability of simulating aircraft maneuvers in the longitudinal plane. Correlation of the F-15 longitudinal captive generated motion with flight test data was excellent.

Due to the success of the above experimental captive programs a new 16T captive system oriented toward user testing has recently been developed. The new system is designed for an aircraft model with a full complement of aerodynamic controls (rudder, elevator, and" aileron). In addition, computer software used in tunnel/computer interfaces has been simplified to make feasible the use of the captive system on a routine basis.

\ .

Documentation of the capabilities and validity of the captive system is accomplished with direct comparisons of F-15 captive versus flight motion. Complex longitudinal and lateral/directional high angle-of-attack (AOA) maneuvers are utilized in the documentation. These wind tunnel/flight data comparisons, along with other work accomplished during system development, are reported herein.

A E D C-T R -80-60

2.0 APPARATUS

2.1 TEST FACILITY

The AEDC Propulsion Wind Tunnel (16T) is a variable density, continuous flow tunnel

capable of being operated at Math numbers from 0.2 to 1.6 and stagnation pressures from

120 to 4,000 psfa. The maximum attainable Mach number can vary slightly depending upon

the tunnel pressure ratio requirements with a particular test installation. The maximum

stagnation pressure attainable is a function of Math number and available electrical power.

The tunnel stagnation temperature can be varied from about 80 to 160°F, depending upon

the available cooling water temperature. The test section is 16 ft square by 40 ft long and is

enclosed by 60-deg inclined-hole perforated walls of six-percent porosity. Additional

information about the tunnel, its capabilities, and operating characteristics is presented in Ref. 4.

2.2 TEST CONDITIONS

Captive tests were conducted at M,. = 0.3 to 0.9 at corresponding total pressure levels

from 2,000 to 1,000 psf, respectively. Within this Mach number and total pressure range the

model Reynolds number, based on the wing mean aerodynamic chord, ranged from 1.6 to

!.9 x 106. Flow angularity corrections were applied to both angle of attack and sideslip as a

function of Mach number. Model blockage corrections were not considered to be important in this test due to the low ratio of model to tunnel cross-sectional area (approximately I percent at 90 deg AOA).

2.3 MODEL DESCRIPTION

The aircrafl model used in this captive testing was a ri'gid (all steel) I/20-scale F-15. Basic

dimensions of the model are shown in Fig. I. The model was equipped with production snag

stabilators and raked wingtips. Engine inlets were maintained in the drooped configuration,

nose down, representative of the F-15 high-AOA flight configuration. The inlets were

plugged just inside the inlet lip, by a flat plate normal to the duct. The model installed in the Tunnel 16T test section is shown in Fig. 2.

Primary aerodynamic control surfaces (horizontal stabilizers, rudders, and ailerons) of

the F-15 were incorporated in the wind tunnel model. The surfaces were remotely controlled and movable over the full range of the production aircraft (horizontal stabilizer + 15 to -25

deg, ailerons __. 20 deg, and rudders _+ 30 deg). The horizontal stabilizers could be operated

A E DC-T R -80-60

both differentially and simultaneously. Control surface positions were obtained directly

from the surface through a strain-gage/flexure arrangement as shown for the aileron in Fig. 3. The direct readout minimized the hysteresis and general looseness often occuring in mechanical control surface readouts. Also shown in Fig. 3 is the unique aileron drive system

developed at AEDC for obtaining remote aileron control on small models with thin wings." The system is designed around a rhombus with a screw jack actuation. The flex cable

actuating the screw jack is driven by a servo housed in the wing root region. The overall performance of the system was good during the course of testing.

it should be noted that the wind tunnel model used in testing was tailored to the F-15

production aircraft. F-15 flight test data presented in Section 4.0 of this report are for an

F-15 aircraft modified for high-AOA testing. The differences between the wind tunnel model and the modified flight test aircraft are as follows: (1) a spin recovery chute was installed on the flight test aircraft, slightly altering the fuselage afterbody; (2) the mass

balance booms on the vertical tail of the flight test aircraft were removed; and (3) the flight test aircraft utilized a test noseboom with flow indicators (vanes) for monitoring angle of attack and sideslip.

3.0 PROCEDURE

3.1 CAPTIVE MOTION SIMULATION

Captive testing is accomplished using a closed loop system of model, wind tunnel, and

digital computers as shown in Fig. 4. The wind tunnel model is installed on a six-component " internal strain-gage balance which provides static forces and moments. The tunnel acts as an

analog forcing function to the model and also provides the AOA and angle of sideslip envelope shown in Fig. 5 via the model-positioning system. Online digital computers control tunnel/model hardware and perform calculations required for conducting motion simulation.

Normally the model is positioned in the tunnel at some trimmed AOA and sideslip angle for a specific flight condition. Upon initiation of the maneuver, appropriate model control surfaces are deflected and static forces and moments are measured. The static aerodynamic

measurements are input to an online digital computer where correctioffs for aircraft flexibility, tunnel flow angularities, etc. are applied. These static aerodynamics are now combined with additional external forces (dynamic and thrust characteristics) and are used

in solving the Euler equations of motion presented in Appendix A. The computed solutions

to the Euler equations are used in controlling the orientation of the model through a point

!

A E D C - T R ~ O ~ O

prediction technique. The technique involves using the last two successive measured values

of each static aerodynamic coefficient to predict the magnitude of the coefficient over the next prediction interval. The prediction interval used in the subject test was 0.05 sec. The predicted coefficients are used to calculate the new model AOA and sideslip by integrating

the equations of motion every 0.005 sec over the prediction interval. The model-positioning

system is then commanded to move the model to the new angles, where the static aerodynamic forces and moments are remeasured. The above process is repeated until a complete maneuver is generated.

The aircraft Mach number is calculated at each prediction interval, and the wind tunnel

Mach number is adjusted to within +_ 0.003 of the calculated value. Thus, the aerodynamic coefficients are measured at the correct Mach number throughout the maneuver. Also, the

aircraft thrust is calculated and modified with each prediction interval by mathematically modeling the simulated aircraft engine/inlet installed thrust as a function of Mach number, altitude, angle of attack, and angle of sideslip.

Specific maneuvers are programmed through the aircraft flight control system. Normally longitudinal and lateral stick position and rudder pedal position as a function of time are the

inputs used in simulating maneuvers. As an option the control surface deflections may be

programmed directly, thereby bypassing the aircraft flight control system. For this test the

F-15 flight control system (including the control Augmentation System, AS)was modeled and utilized as presented in Appendix B.

3.2 PRECISION OF DATA

Maneuvers generated using the captive technique are subject to error from several sources including tunnel conditions, the angle-positioning system, balance measurements, and extrapolation tolerances allowed in the predicted coefficients. The uncertainties (combinations of systematic and random errors) in the basic tunnel parameters are shown in Fig. 6. These data were estimated from repeat calibration of instrumentation and from the

repeatability and uniformity of the test section flow during tunnel calibration. Uncertainties in the instrumentation systems were estimated from repeat calibration of the systems against

secondary standards whose uncertainties are traceable to the National Bureau of Standards

calibration equipment. The instrument uncertainties are combined using the Taylor series method of error propagation described in Ref. 5 to determine the uncertainties of the reduced parameters.

The maximum uncertainty in the model-positioning system was _+ O. 1 deg for settings in AOA and angle of sideslip. Based on the above uncertainties and those resulting from

8

A E DC-T R -80-60

balance inaccuracies and coeMcient prediction tolerances, it is estimated that over the

maneuver time span presented the maximum error occurring in the angular position data, 0,

~,, and ~ will be less than _+1 deg. These estimates were calculated from the following equations, which assume no inertia or aerodynamic coupling.

~.~I J~ A~17. AM x = = ~ ~ - - I ly ' l z .-:

4.0 RESULTS AND DISCUSSION

4.1 GENERAL

The usefulness and validity of a new wind tunnel testing technique such as Captive

Aircraft Testing can best be ascertained through comparisons with flight test data. The flight

test data used in this report were obtained from the F-15A approach-to-stall/stall/post-stall

program (Ref. 6) conducted at the Air Force Flight Test Center. This flight program was selected because it encompasses most of the high-AOA maneuver capability of modern fighter aircraft. The wind tunnel simulation of these high-AOA maneuvers will provide a

rigid check on the captive testing systems capabilities. The validity of the captive system for less demanding, low-AOA maneuvers has been previously demonstrated (Ref. 3).

As in conventional aircraft motion si/nulation, the wind tunnel-measured external rigid forces and moments in captive testing were corrected for airframe flexibility. These corrections were modeled as functions of Mach number, altitude, angle of attack, and load

factor. The flexibility math model was formulated with assistance from both the F-15

airframe manufacturer and the Air Force F-15 System Projects Office (SPO). Due to its complexity, the flexibility model is not included herein. In addition to flexibility corrections,

the aircraft measured pitching moment (Cm) was adjusted for inlet rotation (previous wind tunnel tests have shown Cm to be a function of inlet angle). Inlet angle on the flight vehicle is

scheduled as a function of angle of attack; therefore, corrections to Cm were included when the model fixed inlet angle of 1 ! deg did not correspond to the s~heduled angle.

Dynamic characteristics of the F-15 aircraft were included in the simulation as a math model. The dynamic data utilized were formulated from wind tunnel, flight, and analytical

predictions. Their validity may be justified only from the fact that in previous motion simulation studies they produced reasonable results. The dynamic data matrix was

formulated with inputs received from both the F-15 airframe manufacturer and the SPO.

As previously noted in Section 3.1, the aircraft flight control system (including control augmentation system) as modeled in this experiment is presented in Appendix B.

' 9

A E D C-T R -80-60

Simplifications in the control system incorporated for captive testing were: (1) flight was

limited to subsonic conditions, (2) all time constants less than 0.05 sec were omitted, and (3)

mechanical inputs were limited to slick and rudder positions (i.e., stick dynamics were not

included).

The aircraft mass, inertia, and cg position simulated in the captive generated maneuvers

are identical to those for corresponding flight maneuvers in Ref. I. Table 1 list,, these

parameters for the simulated manuevers included herein.

4.Z LATERAL/DIRECTIONAL COMMANDED MOTION

i

Captive maneuvers generated in the lateral and directional planes of ,notion are

presented for comparison with flight motion in this section. Because of their simplicity the

F-15 full lateral stick roll (Fig. 7) and full rudder pedal roll (Fig. 8) were selected for initial

comparison. These maneuvers are considered simple from a motion ,,imulation viev,,point

because they occur at aircraft angles of attack below stall ( 10 to 20 deg) and are for the most

part commanded motion resulting from control surface deflections (aileron and rudder).

The full lateral stick roll ,notion presented in Fig.7 is in good agreement with flight data.

The lateral/directional parameter roll rate (p), yaw rate (r), and roll angle (PHI) are near-

overlays of the flight data. Small discrepancies occurring in pitch rate (q) may be attributed

to small horizontal stabilizer movements which were not modeled in the captive system via

the longitudinal stick input. The full rudder pedal roll of Fig. 8 also shows good agreement

between captive and flight motion. Again, the small differences occurring in the Iongitudinal

and directional m6tion rates ,nay be attributed to discrepancie,, in modeling of the

longitudinal stick and rudder pedal.

The rate of data acquisition experienced in generating the captive motion of Figs. 7 and 8

was 3.0 to 3.5 min of wind tunnel time per i sec of flight time. The limiting factor was

positioning of the model control surfaces. The 3.5 rain/see rate should be considered typical for maneuvers possessing large control deflection rates.

A hidden attribute in captive testing which often plays a significant part in accurate

motion simulation is the inherent modeling of control surface interactions (rudder,

horizontal stabilizer, etc.). For control surfaces in close proximity which experience large

deflections, as in Fig. 8, interactions often become significant and thereby affect vehicle aerodynamics. Conventional motion simulation cannot sufficiently account for such

interactions because of the requirement for unrealistically large data matrices and

corresponding wind tunnel programs.

10

A EDC-TR-80-60

4.3 WINGS-LEVEL STALLS

One of the first high angle-of-attack maneuvers flown in most flight test programs is the

l - " g " stall. Although this maneuver is normally considered simple in the flight program, it sometimes causes problems when one attempts to reproduce it via motion simulation. The

problem area is associated with uncommanded lateral/directional motion occurring at high

angles of attack. This motion may be seen in Fig. 9, where lateral/directional parameters Beta, P, and R become oscillatory while both lateral stick and rudder pedal are fixed. The oscillatory motion is believed to result from separated and asymmetric flow over the aircraft

resulting in nonlinear and hysteresis effects in the aircraft aerodynamics. Unless these aerodynamic phenomena are both defined and modeled properly in the data matrices of motion programs, the chances of simulating these uncommanded motions are poor. As seen

in Fig. 9, the captive system does a good job of simulating these motions as a result of inherently accounting for nonlinear and hysteresis effects. The fact that the motion is

I

approximately 90 deg out of phase is insignificant since the onset of the excitation (flow separation) is not always predictable.

An abrupt or more accelerated wings-level stall is shown in Fig. 10. As observed in the l - " g " stall the uncommanded lateral/ directional motion (wing rock) commences in the 20-to 25-deg. angle-of-attack range. An improvement in the onset of the uncommanded motion is noted in Fig. 10, where the phase angle between the flight and simulated oscillations is near zero.

Data acquisition rates for the I-"g" stall and abrupt stall captive maneuvers of Figs. 9

and l0 were approximately 1.8 rain and 3.4 min respectively, of wind tunnel time per second of flight time. The improved acquisition of the l - "g" stall may be attributed to the reduced control deflection rates (see longitudinal stick). The data acquisition rate of Fig. l0

corresponds to that presented in Figs. 7 and 8, which also show high conti'ol deflection rates.

4.4 WIND-UP TURN TO ST, ALL

An accelerated stall resulting from a wind-up turn maneuver is shown in Fig. 1 I. The maneuver is initiated with the aircraft in a 3.8-"g" (Nz) banked turn at 0.75 Math number.

The near-4-"g" turn is maintained by increasing angle of attack (via longitudinal stick position) as Mach number decreases until maximum control surface and/or lift coefficient is reached. As experienced with previous high-AOA maneuvers an uncommanded wing-rock

motion commences as aircraft angle of attack passes approximately 20 deg. The captive

system is successful in simulating these motions although, for reasons adressed in Section 4.3, the phase angle is different.

II

A E D C -T R -80-60

The commanded motion ALPHA, THETA, and N7 of Fig. I I are in good agreement

with flight. A high-"g" maneuver such as the wind-up turn normally results in a rapid bleed-

off in Math number for constant throttle settings. An example is the maneuver in Fig. I l

where Mach number decreases from 0.75 to 0.35 in 13 sec flight time. In captive testing the

tunnel Math number is maintained within _0.005 of the calculated Math number/~t all

times. The continuous wind tunnel Mach number adjustment was accomplished for the

wind-up turn maneuver without delay in data acquisition. The data acquisition rate was

approximately 3 rain. tunnel time per second of flight time.

5.0 CONCLUDING REMARKS

As a result of the captive test experiment, the following conclusions and observations are

noted:

I. Accurate simulation of complex high-angle-of-attack aircraft maneuvers may be

obtained with captive testing as verified by wind tunnel/fl ight correlations.

. The fabrication of remote control surfaces on small wind tunnel models (1/20

scale) is feasible, and the reliability and accuracy of these control surfaces are

acceptable for captive testing.

. Uncommanded motion (wing rock) near aircraft stall was simulated accurately

with the captive system. The inherent inclusion of all aerodynamic nonlinear and

hysteresis phenomena in captive testing is believed to be an important factor in

perturbating this motion.

4. Wind tunnel time required for generating captive maneuvers ranged from 1.8

rain to 3.5 min tunnel time per second of flight time.

12

A E DC-TR-80-60

REFERENCES.

1. Milillo, J. R. "Post-Stall Testing of Aircraft with a Wind Tunnel Captive System." AE DC-TR-72-126 (AD75146 !), November 1972.

2. Butler, R. W. "Evaluation .of a Wind Tunnel Technique to Determine Aircraft Depart ure Characteristics" AEDC-TR-73- ! 83 (AD776317), March 1974.

3. Butler, R. W. "A Wind Tunnel Captive Aircraft Testing Technique." AEDC-TR-76-22 (AD-A023690), April 1976.

4. Test Facilities Handbook (Eleventh Edition). "Propulsion Wind Tunnel Facility, Vol. 5." Arnold Engineering Development Center, June 1979.

5. Abernethy, R. B. and Thompson, J. W., Jr. "Handbook m Uncertainty in Gas Turbine Measurements." AEDC-TR-73-5 (AD755356), February 1973.

6. Wilson, D. B. and Winters, C. P. "F-15A Approach-To-Stail/Stali/Post-Stall Evaluation." AFFTC-TR-75-32, January 1976.

13

INBOARD PYLON

STAT I ON

! 5.'/

16.940

i r

DIMENSIONS IN INCHES

25.688 i

Figure 1.

35.235 38.250

Aircraft model used in captive demonstration.

J '.054 8.171 _±

~> m

c) c~

~0 ¢)

o

. . . . . . , : : . :

" : : " : . . . . . . . . . . . . . . . . . . . . . .

. . . . . . o . . , . ,o ~.,:: , _

l *

. . , . . , . : .

, e e ~ t

w

Q =lp

==° * "o -~. o >

. . . : . . . . . . . . • . . . . : ~ tQg~ t

0 W

' i ; ~ ̧' J i

A E D 8 4 9 0 - '

m E~ C~

;0 ~o O

Q

Figure 2. Aircraft model installed in Tunnel 16T test section.

• J" , S

\

Strain Gage

Flexure ~ .....

C)

Screw J a c k

@

Figure 3. Aileron drive mechanism developed for captive model.

m cJ

o & o

O0

II C o n t r o l F u n c t i o n s • A i l e r o n s • R u d d e r s • E l e v a t o r s

• A n g l e o f A t t a c k • Angle of Sideslip

~r

i

G r a p h i c s

M e a s u r e m e n t s • F o r c e s & Moments • Tunnel Conditions

E q u a t i o n s o f

M o t i o n

Math Models • Dynamic Model • E n g i n e Model • C o n t r o l S y s t e m • F l e x i b i l i t y

~> m 0 C~

~0 do 0 & o

Figure 4. Captive closed loop data acquisition system.

A E DC-T R-80-60

ROLL ANGLE, dog

"o \

l 0 to 0 o I ~

\ ~ 9 o 1 - - - - . . . . . ~ / o

/

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-80~ ~ 8 0

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Figure 5. Tunnel 16T angle-of-attack and angle-of-sideslip capabilities.

I00

%

19

A E D C -T R -80-60

Mach Number Uncertainty

0.014

0.012

O. 010

0.008

0.006

i 0.004

0.002

0

I I

Static Pressure

Uncertatnty, pst

-PT = 3,000

• - [ - -

Oynamlc Pressure

Uncerlalnty, psf 2

I

o I

-PT - 3,000

0.06

0.0~

Reynolds 0.04 Number

U ncerta, lnty x lO-Otft 0.03

0.02

f ~ PT - 3,000

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i

f

o.ol I 0 0.2 0.4 0.6 0.8 1.0

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I

1..2 1..4 1.6 MACH NUMBER

Figure 6. Estimated uncertainties in wind tunnel parameters.

20

A E D C - T R -80-60

<{ "r Q. J <{

5 0

4 0

3 0

2 0

I 0

0

2 0

o W I N D T U N N E L

- - - F L I G H T

i ! ' ' r ,

i I

I I I

I

~" 0 QD

I , I I - 2 o I l

.,~ol I I I I 2 0 -

- I 0 ' , I I I I I

-,o[ 2 o o , . _ _ . _ ! i I I _ / I I .°'o~-. I

.~ool , I I J J J I ~ , - - I I

- , o o [ I I , 0 I 2 3 4 5 6 7 8 9 I 0

-TIME, s e c

Figure 7. Full lateral stick roll maneuver.

21

A E DC-T R -80-60

z o

0 WIND T U N N E L

- - - - - - F L I G H T

1.0

0 . 8

r,~m

< 3 5

N Z

" I I 2 J- \ /

ol " J 4

= o 2

---P o ~ ~ -----" Ob~ Z 0 J - 2

r

.....I

J

2 "

; ; -2 _1 - 4

I - - - - J

2

~ o

2 0 5 I 0 15

T I M E , se¢

Figure 7. Concluded.

~3

A EDC-TR-80-60

• ~ ,

I1.

- I

¢¢

I - - bJ m

O.

0

n-

50

4 0

30

2 0

I O ° - < > - "

0

20

I 0

- I 0

- 2 0 I

WIND T U N N E L F L I G H T

;

i

I I

, I I 0 0

- I 0 0

2O

O ( ~ ~ . . ~ . s _ ~ . ~ " ~ 1 ~ ' ( ~ " `~

30

2O

IO " ~

.,2ir l

I I I

I

" \

, I I . . . . , i' 2°°1 I I ! ~ - ' l

• T- OI P ! ~ . . - - ~ ' ° ' ~ I I 1 " "

,o2 i - j

= I I ] ~ - i o o I I 0 I 2 3 4 5 G 7 a

TIME, sec

i I

I

J I

!

I

9 I 0

Figure 8. Full rudder pedal roll maneuver.

23

A E DC -T R -80-60

"r" O <[ :Z

0 .6

0 . 4

"O WIND T U N N E L

F L I G H T

3 f bJ

<[

2q N

z I

0

. J < 4 z

~ m 0 I z o "J - 2

2

< O(

_1 - 4

4

o. O( n*

- 2 0 I 2 3 4 5 6 7

T I M E , s e c

Figure 8. Concluded.

24

A E D C-T R -80-60

WIND TUNNEL FLIGHT

"!-

--I

4 0

2 O( ~ ~ F - ~ ) 0 ~

J ~ a"~" ~ ' ~ ~

0

I 0

I-- O, J

I

- I 0 I

~---~ ~ - -~ .~ -~ . . . _ . =--o<.

I i - - , i \

I I \ \ J

- 5 0 ~

0

io I I . - .

. ao l I I I I [ I / ! | I

I 0

o= Oq

- I 0

2 0 0

o.

- 2 0 0

¢z I-- Lg q- I--

I 0 0 I

I J I , - I 0 0 0 2 4 6 8 I 0 12 14 16 18 20

T I M E , sec

Figure 9. One-"g" stall maneuver.

2S

z

Z

N Z

O WIND T U N N E L

- - - - - F L I G H T

" ' 3 9

~ o 3 8 ..11( < 3 7 I I I I

I -k - < , ~ .

2

6

4 m ~ v ~__u 2

o - I - 2

I l l l m=,< ~..( r.,~ -~-(

_m._w.w_.ZZ.EE

A E DC-TR -80-60

~ . , I I I 0 2 4 6 8 10 12 14 16 18 2 0

T I M E , sec

Figure 9. Concluded.

26

A E D C -T R -80-60

o WIND TUNNEL - - - - - - FLIGHT

-i- o . _1

5O

30 I ~ .-o--o-~"m--" 2O ~

10 I

0 I

IO

i-- tO Oq m

- I 0

'1 I

I I

..b.. :,.R

. ° ~ L _ L _ ~ __1 ~ ! I /1 I F ' ~

0

2 0

.P '~ I ' ~ \ 1 I

ol "~- - I _ _ _ _ L _ _ _ _ _

I O I I I I t

o ! l , : l ~_ I I_ I ~

2 0 0 I I

~ o! ~ I I _' I I J ~ I _ 1 .~ool I I

~ ' ° ° ! ! ~ , , I , , I , . , o = l I I I I--

- I 0 0 ,

0 I 2 3 4 5 6 7 8 9 I0 TIME, sec

F igure 10. A b r u p t stal l maneuver .

27

A E DC -T R -80-60

O WIND TUNNEL

FL IGHT

• I- 0 .5

X 0.4

0.3 i o35 0 ~ ~ a I - o 34 I - - - - _ I X < 33

N Z

0r-T" I I I 1 6

z 4 5 . , ,~ ~ 2

o, - 2

,.=,~ - i

00,.0,

0 2 4 6 8 I0 T I M E , sec

Figure 10. Concluded.

28

A E D C - T R - 8 0 - 6 0

< ' T

0 . . - I

50

40

50

20 I (

I0

o WIND TUNNEL FL IGHT

O

2 0

m - I0

-20 L

-I001

0

-T- o.

l i J P I l , ! I !

I I i i ~ L / ~ - - I I / ' '. j ~ l ;~./, ~, . , / ' , , f

I I [ -~- ,:--~---"~ I "~--'A ",L~L,',~ I', 'I r il[i -~[F.~l --

: : : ! i I f i i ~ F ~ - , ~ l

200 I I I I

,oof _ _ ' _ ~ ~ ! , ! ' "--.--4 - , o o l i I

0 I 2 3 4 5 6 7 8 9 IO II 12 1:3 14 15 TIME, sec

F i g u r e 1 1 . W i n d - u p t u r n m a n e u v e r .

29

A E DC -TR -80-60

-I- ¢,J <[ : i

0 WIND TUNNEL

FLIGHT

° " I ~ . ~ f . I ! j ! 0.6

0.5 ~ ' ~

0.4

o~ i I

I I i I

! I-Y'~-- |

O

N

a

I -

k- - i

34~ !

32 I ~

N Z

2 , , L

i i i

..J z '~ 6 . ~ ~--~ ~=-~ :~---I--- N " ' 4 I"- F _ . _ .

~., a I o 0 ~ ~ 1 i I I I I ~ x ~ O ~ ~

~= _, I - - I ! I ! i I I

c31=j ~=- .

0 I 2 3 4 5 6 7 8 9 I0 II 12 13 14 15 TIME, sec

Figure 11. C o n c l u d e d .

30

%0

Maneuver

Lateral Stick Roll

Rudder Pedal Roll

One-"g" Stall

Abrupt Stall

Wind-Up Turn

Table 1.

Weight,

ib

33,500

32,900

35,40O

38,900

33,400

Characteristics of Aircraft Simulated in Captive Testing

cg, I x , Iy, [Z'

Per____ee~t__~ Slu~-ft 2 Slug-ft~ S1~-f~_ 2

27.4 22,400 196,100 211,600

28.2 22,200 192,700 208,200

28.5 24,300 192,000 209,600

26.9 28,900 196,300 21.8,400

28.8 22,500 189,800 205,600

I X Z ~ sl_._~-ft2

-1 ,000

0

0

0

0

Power

Military

Military

Military

Military

Military

m O o

O & o

A E DC-TR -80-60

APPENDIX A

EULER EQUATIONS OF MOTION

° 33

A E DC-TR-80-60

APPENDIX A EULER EQUATIONS OF MOTION

The following equations were used in calculating the motion of the aircraft during captive testing.

13 - Gx Rx Ixz l y - It

- I"~ + ~ + ~ (~ + Pq) + Ix qr

ci _ Gy ~RY ~lxz lz - Ix ly + l y + ly (r2 " p 2 ) + iy rp

r - G~. Rz lxz Ix - ly - I'--~" + I"'~- + .... (p - q r ) + lz lz

pq

x + Tx = - - - - + n l g - q w + rv m in

"¢ = Y + n2g - ru + pw m

v v = z Tz - - + ~ + n3g + q v - pv m m

The aerodynamic forces and moments used in the simulation were as follows:

x = x(a, ~)

v=Sb y = yCa, /3) + p ~

4 (Cyrr + Cy~)

z = z(a,/3)

v= Sb 2 Gx = Gx(a , ~) + p ~ C~pp + C~zr + C~/j~)

v** S ~ 2 Gy --- Gy(a , ~) + p - - - - ~ (Cmqq + Cma h)

v . Sb 2 G~ = G~Ca, ~) + P-"-T'- (C~r + C~pp + C.:8) p~

where x(a, ~), y(a, ~), z(a, ~), Gx(a, ~), Gy(a, ~), and Gz(a, ~) are the aerodynamic static forces and moments generated in the wind tunnel.

35

A E DC-TR -80..60

APPENDIX B SIMULATED FLIGHT CONTROL SYSTEM

37 1

LOAD FACTOR: (I

ROLL ACCEL t rod/see

PITCH ACCEL, rad/sec I

MECHANICAL PITCH

LONGITUDINAL STICK POSITION, in i

PITCH TRIM I

INTERCONNECT I C~MPENSATOR LOAD FACTOR ERROR / PITCH TRIM

+AFT -FWD STICK 81ASp in

O.S

PITCH RATIO ADJUST DEVICE

HORIZ STAB COMMAND, deg

I

LATERAL STICK STAB COMMAND, deg

JiB

LEFT,

MEAN STAB DEFL, deg

CAS R I CAS L

RIGHT

STABILATOR ACTUATOR:

U RIGHT STAB OEFLI¢

+TEO -TEU

LEFT STAB DEFL, deg

Figure B-1. Longitudinal control.

~> I l l

c)

::0 & O & C~

o

I AILERON-RUDDER INTERCONNECT

l RUOOER ACTUATOR RUDDER PEDAL P,OelTION ~ ~ IT 2 O RUOOER DEFLECTION, dog

° RT ' " ~ I " ~ ' ~ ~ I , ÷ i : o j + T__(L . LT YAW v " / e O F F - TER

CAS ~ ANGLE OF ATTACK, deg PREFILTER

I 0 [ - ~ ÷ +~PROPORTIONAL+INTEGRAL

LATERAL ACC! ftlsec:

._L.s S + l

WASHOUT ROLL-YAW CROGSFEED

YAW ACC, rod/sH E

YAW RATE, rod/s ic HIGH PASS FILTER

.Ic ) J

GTRUCTURAL FILTER

($/2z) 2 + ~(s/Ez)+t

r n

c)

do Q & C~

Figure B-2. Lateral control (rudder).

LATERAL STICK POSITION, In

+ RT - LT

:E

AILERON -RUDDER INTERCONNECT BOOST

B

YAW RATE, r0d/see

J I L~TERAL L RATIO

NOR IZ 8TAB COMMAND, deg ÷ TED -TEU

AILERON ACTUATOR AILERON

1 EO l ~ I +TIEvioIOEFLECTION'degULEFT

RIOHX_ LATERAL STICK STAB COMMAND t de~l~

HORIZ STAB : COMMAND e deg

AILERON-RUDDER INTERCONNECT

Figure B-3. Lateral control (aileron).

3> I l l o o n

O & O

ROLL RATE, rod/sec

ROLL CAS

LATERAL STICK POSITION I in. 4" RT - LT

LOAD FACTOR. O

PITCH ACCEL, rod/see:

PITCH RATE, rod /see _ . ~ . , - L - -

PITCH GAS

LONGITUDINAL STICK POSITION 4" AFT - FWD

PITCH RATE, rod /sec

,{ 54 deg/eec

PRE-FILTER

s.3s) I 8 ÷ 3 . 3 3 3

PRE-FILTER

WASHOUT STALL INHIBITOR

ANGLE OF ATTACK ! de~,

4,

MEAN STAB

4,,

PITCH TRIM INTERCONNEC1

ANGLE-OF-ATTACK

ROLL RATE GAIN

ROLL CAG

OFF ON • ROLL CA9

tROPORTIONAL INTEGRA

JFF CA8

ON l eOFF

I , E R , . / LEFT__ ERVO

CA8 L ~ 1CASR

3> m

<-)

-11 6) O 6) C~

Figure B-4. CAS control system.

AOA

b

Ce

Cep

Cm

Cmq

Cm&

Cn

Cnp

Cn F

Cv

CY r

cg

Fy

g

Gx,Gy,Gz

Ix,Iy,lz

NOMENCLATURE

A E DC-T R-80-60

Angle of attack

Wing span, ft

Rolling-moment coefficient, Mx/q~Sb

Dynamic damping derivative in roll, OCe/(cgpb/2V=)

Dynamic cross-derivative or rolling moment due to yawing, aCe/(Orb/2Voo)

Lateral acceleration derivative in roll, 0Cd(a/3b/2V~)

Pitching-moment coefficient, My/q=S~

Dynamic damping derivative in pitch, OCm/(q~/2V~)

Longitudinal acceleration derivative in pitch, 0Cm/&~c'/2V®)

Yawing-moment coefficient, Mz/q®Sb

Dynamic cross-derivative of yawing moment due to OCn/(Opb/2V®)

Dynamic damping derivative in yaw, 8Cn/(Orb/2Voo)

Lateral acceleration derivative in yaw, 0Cn/(0/~b/2V®)

Side-force coefficient, Fy/q®S

Dynamic derivative of side force due to yawing, 0Cv/(t~rb/2V~o)

Lateral acceleration derivative in side force, 8Cy/(8/~b/2V®)

Mean geometric chord, ft

Aircraft center of gravity

Side force, Ib

Acceleration due to gravity, ft/sec 2

Components of the aerodynamic moments about the XB, YB, and ZB axes, respectively, ft-lb

Moments of inertia about Xa, YB, and Za axes, respectively, slug-ft 2

rolling,

43

A ED C-TR-80-60

|xz

Mx

My

Mz

m

Nz

nl,n2,n3

PT

p,q,r

q,,~

Rx,Ry,Rz

S

Tx,Tz

UjVjW

V~

XB,YB,ZB

x,y,z

ot

Q

NOTE:

Product of inertia, slug-ft 2

Rolling moment, ft-lb

Pitching moment, ft-lb

Yawing moment, ft-lb

Aircraft mass slugs

Load factor

Direction cosines of body axis relative to earth axis

Tunnel total pressure, lb/ft 2

Aircraft roll, pitch, and yaw rates about the Xn, YB, and ZB axes, respectively, deg/sec

Dynamic pressure, lb/ft 2

Moment contributions from engine thrust about the XB, YB, and ZB axes, respectively, ft-lb

Wing area, ft 2

Components of engine thrust along the XB and ZB axes, respectively, ft/sec

Linear velocity components along the XB, YB, and ZB axes, respectively, ft/sec

Free-stream velocity, (u 2 + v 2 + w2) ~/2, ft/sec

Right-hand body-axis Cartesian coordinates, X positive forward

Components of the aerodynamic forces along the XB, YB, and ZB axes, respectively, Ib

Angle of attack, tan- ~ w/u, deg or rad

Angle of sideslip, sin- 1 v/V=, deg or rad

Simulated air density, slugs/ft 2

Dot over symbol indicates derivative with respect to time.

44