static and dynamic stability derivatives model ofa wind-tunnel investigation has been made to...

STATIC A N D DYNAMIC STABILITY DERIVATIVES OF A MODEL OF A JET TRANSPORT EQUIPPED WITH EXTERNAL-FLOW JET-AUGMENTED FLAPS

by Delmu C.Freeman, Jr., Sive B. Grafton, und Richurd D’Amato

Ldngley Reseurcrb Center Lungley Stution, Hampton, Vu,

N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D . C. S E P T E M B E R 1 9 6 9

https://ntrs.nasa.gov/search.jsp?R=19690026880 2020-04-21T12:05:03+00:00Z

TECH LIBRARY KAFB, NM

0332205

1. Report No. 2. Government Access ion No.

NASA TN D-5408 I 4. T i t l e and Subti t le

STATIC AND DYNAMIC STABILITY DERIVATIVES OF A MODEL OF A JET TRANSPORT EQUIPPED WITH EXTERNAL-FLOW JET-AUGMENTED FLAPS

7. Authods)

Delma C. Freeman, Jr., Sue B. Grafton, and Richard D'Amato

9 . Performing Organizat ion N a m e and Address

NASA Langley Research Center

Hampton, Va. 23365

2. Sponsoring Agency N a m e and Address

National Aeronautics and Space Administrat ion

Washington, D.C. 20546

5 . Supplementary Notes

6. Abstract

~ ~3. Recip ient 's Cata log No.

~~

5. Report D a t a September 1969

6. Performing Orgonizot ion Cod

8. Performing Organizat ion Rep

L-6521 ~

I O . Work U n i t N o .

121-01- 11-03-23 11. Controct or Grant N o .

13. T y p e of Report and P e r i o d C

Technical Note

14. Sponsoring Agency Code

A wind- tunne l investigation has been made to determine the static and dynamic stability derivatives of a

model of a large jet t ranspor t equipped w i th external-flow jet-augmented flaps. The tests were conducted in thi

Langley ful l-scale tunne l , and a model powered by scale-model, compressed-air-driven turbofan engines was u

The resul ts of t h e investigation showed that blowing on the flap system increased the l i f t -curve slope,

delayed the stall, and increased t h e maximum l i f t coefficient. The data also showed that a l l model conf igurat io

generally had static longi tud ina l stability over the test angle-of-attack range except those at t he h ighe r flap-

deflection angles where the effects of power were destabilizing. The resul ts also showed that t he model had pc

t ive damping in pitch, rol l , and yaw th roughout the test angle-of-attack range up to and s l ight ly beyond t h e

stal l for a l l test conditions. The application of power in the flap system resulted in appreciable increases in r

damping but produced essential ly n o effects i n pitch and yaw damping. There were essential ly n o effects due

to osci l lat ion frequency on t h e damping derivatives.

17. K e y Wards Suggested by Author(s) 18. Distr ibut ion Statement

let-augmented flaps Unclassified - UnlimitedDynamic sta bi I i ty

19. Security C lass i f . (of +his report) M. Security C l a s s i f . (of th is page)

Unclassified Unclassified

STATIC AND DYNAMIC STABILITY DERIVATIVES

O F A MODEL O F A JET TRANSPORT EQUIPPED WITH

EXTERNAL-FLOW JET-AUGMENTE D FLAPS

By Delma C. Freeman, Jr., Sue B. Grafton, and Richard D'Amato

Langley Research Center

SUMNIARY

A wind-tunnel investigation has been made to determine the static and dynamic stability derivatives of a model of a large jet transport equipped with external-flow jet-augmented flaps. The tests were conducted in the Langley full-scale tunnel, and a model powered by scale-model, compressed-air-driven turbofan engines w a s used.

The resul ts of the investigation showed that blowing on the flap system increased the lift-curve slope, delayed the stall, and increased the maximum lift coefficient. The data also showed that all model configurations generally had static longitudinal stability over the test angle-of-attack range except those at the higher flap-deflection angles where the effects of power were destabilizing. The resul ts a lso showed that the model had positive damping in pitch, roll, and yaw throughout the test angle-of-attack range up to and slightly beyond the stall for all tes t conditions. The application of power in the flap system resulted in appreciable increases in rol l damping but produced essentially no effects in pitch and yaw damping. There were essentially no effects due to oscillation frequency on the damping derivatives.

INTRODUCTION

A great deal of interest has been shown in the external-flow jet-augmented-flap concept as a means of achieving high lift coefficients. In this concept, the jet efflux from pod-mounted engines is deflected upward to blow over the flaps and through s lots of the flaps and, thus, induces very high lift on the wing. Early experimental investigations of external-flow jet-augmented flaps on general research models (for example, see refs. 1to 5) have demonstrated that desirably high lift coefficients can be generated with this system. Recently, a program has been started at the Langley Research Center to investigate the application of the external-flow jet-augmented flap to a large jet t ransport with high bypass-ratio turbofan engines. The resu l t s of conventional static wind-tunnel tests of this configuration are reported in reference 6 and show that the external-flow jet-augmented flap offered a promising means of achieving improved take-off and landing

L l l l lI I I I 1111111111111 11111111.11111.1.1111.11.11.1 .II--I .111--1.1 I

performance for large jet transports. Because of these promising results, a program has been initiated to evaluate the dynamic stability, flight characterist ics, and general piloting techniques of such a configuration. This work is to be conducted with a fixed-base simulator requiring aerodynamic inputs in the form of static and dynamic stability derivatives of the particular configuration under study. As par t of the overall program, the present investigation was undertaken to measure the static and dynamic stability derivatives of the jet-transport model with an external-flow jet-augmented’flap.

The model used in the investigation w a s powered by four high-bypass-ratio turbofan engines and could be equipped with double-slotted trailing-edge flaps for use in an external-flow jet-augmented-flap system. The flap configurations tested represented possible landing-approach and take -off configurations. The dynamic stability derivatives were determined in pitching, rolling, and yawing forced-oscillation tests at two different frequencies over an angle-of-attack range. In order to help interpret the dynamic data, the static longitudinal and lateral stability characterist ics of the model were also determined and are presented.

SYMBOLS

The longitudinal data a r e referred to the stability-axis system and the lateral data are referred to the body-axis system. (See fig. 1.) The origin of the axes w a s located to correspond to the center-of -gravity position (0.25 mean aerodynamic chord) shown in figure 2.

In order to facilitate international usage of data presented, dimensional quantities a r e presented both in the U.S. Customary Units and in the International System of Units (SI). Equivalent dimensions were determined by using the conversion factors given in reference 7.

b wing span, feet (meters)

aCL cLa lift-curve slope, aa,

C local wing chord, inches (centimeters)

-C mean aerodynamic chord, feet (meters)

FD drag force, pounds (newtons)

FL lift force, pounds (newtons)

FX force along X-axis, pounds (newtons)

2

FY

F Z

f

it

k

MX

MY

MZ

P

q

qca

r

S

T

lateral force, pounds (newtons)

force along Z-axis, pounds (newtons)

frequency of oscillation, cycle/second (1 cycle/second = 1hertz)

horizontal-tail incidence angle, degrees

reduced-frequency parameter, wb/2V o r d / 2 V

rolling moment, foot-pounds (meter -newtons)

pitching moment, foot-pounds (meter-newtons)

yawing moment, foot -pounds (meter-newtons)

rolling velocity, radians/second

pitching velocity, radians/second

free-s t ream dynamic pressure, pV2/2, pounds/foot2 (newtons/meterZ)

yawing velocity, radians/second

wing area, feet' ( m e t e d )

thrust, pounds (newtons)

f ree-s t ream velocity, feet/second (meters/second)

x,y,z body reference axes

Xs,Ys, Zs stability reference axes

a! angle of attack, degrees o r radians

a! rate of change of angle of attack, radians/second

P angle of sideslip, degrees o r radians

3

V

--

ra te of change of angle of sideslip, radians/second

elevator deflection, positive when trail ing edge down, degrees

6f 1 forward trailing-edge flap-segment deflection, degrees

6f 2 aft trailing-edge flap-segment deflection, degrees

6sl leading-edge slat deflection, degrees

P air density, slugs/foot3 (kilograms/meterS)

angle of roll, degrees o r radians

w angular velocity, 2 d , radians/second

cx=-FX cz =-F Z qms qms

m

cy=-FY qcos

r b‘lr - rb cy, =-aCY

a-2 v

a-2 v

acz aCY czp = ap CYp = ap 4

I

TESTS, EQUIPMENT, AND TECHNIQUE

Wind Tunnel

The tests were made in the 30- by 6O-foot (9.1- by 18.3-meter) open-throat test section of the Langley full-scale tunnel with the model mounted about 10 feet (3.05 meters) above the ground board. The model was so small in proportion to the tunnel tes t section that no wind-tunnel wall corrections were needed o r applied. Normal corrections for flow angularity were applied.

Apparatus and Model

The investigation was conducted on the four -engine, high-wing, jet-transport r m ~ M illustrated by the three-view drawing of figure 2(a). The dimensional characterist ics of the model are given in table I. The wing had an average leading-edge sweep angle of 28' and incorporated leading-edge slats and double-slotted trailing-edge flaps. A detailed sketch of the flap assembly and engine-pylon arrangement is shown in figure 2(b). The wing airfoil section was the same as that used in reference 6. The forward and aft flap-deflection angles tested represent approximately a landing-approach condition (6f1/6f2 = 3Oo/6O0) and take-off conditions (6f1/6f2 = 2Oo/4O0 and 6f1/6f2 = 10°/200). The slat-deflection angles represent these same conditions. Photographs of the model and flap system are shown in figures 3(a) and 3(b), respectively.

To facilitate model configuration changes and to insure accurate flap-deflection angles, the wing of the model was designed with removable trailing edges. To convert the model from the clean configuration to each of the flap-deflected configurations, the clean trailing edges w e r e replaced with trailing-edge flaps constructed with fixed gaps, overlaps, and deflection angles. The leading-edge slats were designed so that they could be fastened to the wing leading edge at fixed positions when desired.

The model engines represented high-bypass-rativurbofans and were installed at -3' incidence so that the jet exhaust impinged directly on the trailing-edge flap system. The engine turbines were driven by compressed air and turned fans which produced the desired thrust.

All dynamic force tests were made with a single s t ru t and sting support system and a strain-gage balance. Sketches of the forced-oscillation test equipment are presented in figure 4, and the equipment is described in reference 8. The static force tests were made on a conventional sting which entered the rear of the fuselage.

Tests and Procedures

Calibrations were made to determine the engine-installed thrust as a function of engine speed in revolutions pe r minute with the model at an angle of attack of 0' and with

5

Illll I l l I I I I I I

the trailing-edge flaps and leading-edge slats undeflected. The aerodynamic tests were made by setting the engine speed to give the desired thrust at an angle of attack of 0' and then maintaining this engine speed as the model was tested through a range of angles of attack.

The thrust calibrations were made at the free-s t ream dynamic pressure used in the present tests, 7.1 .lb/ft2 (340 N/m2). The value for the thrust used in computing the thrust coefficients for the forward flight tests is the difference between the longitudinal force with power on and the longitudinal force with power off and flow-through nacelles, both for the same free-s t ream dynamic pressure. The longitudinal force with power off and flow-through nacelles was not actually measured on the present model but was computed by subtracting the increment of drag coefficient due to the windmilling of rotating pa r t s f rom the drag coefficient of the model in the windmilling condition. The drag of the windmilling pa r t s was assumed to be the same as that measured for s imilar but 60-percent la rger scale-model engines in an unpublished investigation. There is the possibility of a smal l e r r o r in this procedure if the drag coefficient of the windmilling par t s is not exactly the same on the larger and smal le r engines. Even if the assumed drag coefficient of the windmilling par t s of the present engines were 100-percent different, however, the e r r o r in thrust coefficient would only be about 5 percent for the lowest thrust coefficient of the tests and would be even smaller for the higher thrust coefficients. The thrust calibrations were made through a range of engine speeds up to 60 000 revolutions pe r minute, at which speed the fans developed, in the static condition, their rated thrus ts of approximately 30 pounds (133.44 newtons) each.

Dynamic-force tests were made to determine the longitudinal and lateral oscillatory stability derivatives of all model configurations with power off and for thrust coefficients CT of 0.38, 0.78, and 1.70. These force tes t s were made over an angle-of-attack range f rom -4' to 24' for the three flap-deflection combinations previously discussed. The dynamic stability derivatives werg measured for an amplitude of 5.5' and for frequencies of 0.5 and 1.0 cycle pe r second corresponding to values of the reduced-frequency parameter k of 0.023 and 0.045, respectively, for the pitching t e s t s and 0.160 and 0.321, respectively, for both the rolling and yawing tests. In order to help interpret the dynamic data, static-force tests were also conducted to obtain the static longitudinal and lateral stability character is t ics of the model. All tests were conducted with the rudder undeflected and with the tail incidence and elevator-deflection angle se t to give longitudinal t r i m in the operational angle-of-attack range of the airplane.

The force tests were conducted at a dynamic pressure of 7.1 lb/ft2 (340 N/m2) which corresponds to a Reynolds number of 0.543 X lo6 based on the mean aerodynamic chord of the model.

6

RESULTS AND DISCUSSION

Static Stability Derivatives

Longitudina1.- The static longitudinal stability derivatives of the model are presented in figure 5. In general the data show that blowing on the flap system increased the lift-curve slope, delayed the stall, and increased the maximum lift coefficient. A maximum lift coefficient of 4.4 was achieved by using the landing-approach flap system (6f1/6f2 = 3Oo/6O0) with maximum thrust (CT = 1.70). The data show that all the model configurations generally had static longitudinal stability over the tes t angle-of -attack range for the power-off condition. The effects of power were smal l at the lower flap-deflection angles, but at the higher deflection angles these effects were destabilizing and actually caused the model to become unstable at negative angles of attack. The horizontal-tail incidence and elevator deflection used in the investigation were approximately correct for providing longitudinal t r im.

Lateral.- The static lateral stability derivatives of the model a r e presented in fig-~~

ures 6 and 7. The data of figure 6 show that the variation of Cz, Cn, and C y is generally linear over the range of sideslip angles measured (p = k5') at all but the highest angle of attack tested (CY = 25.20), which was generally near, or above, the stall. The data presented in figure 7 were determined f rom the incremental differences in C2, Cn, and C y at angles of sideslip ranging from -5O to 5O. These data show that the clean configuration had directional stability (+CnP> and positive dihedral effect (-Cz P> fo r angles of attack up to about 15O. Deflecting the leading-edge s la ts and trailing-edge flaps increased the level of directional stability for the power-off condition, but application of power generally caused a reduction in directional stability. Generally, except a t 6f1/6f2 = O o / O o and Sf1/6f2 = 1Oo/2O0, the model had positive dihedral effect (-Cip) throughout the test angle-of-attack range. Fo r the higher flap-deflection angles (6f1/6f2 = 2Oo/4O0 and 6f1/6f2 = 3Oo/6O0), the effect of power w a s to markedly increase the positive dihedral effect.

Dynamic Stability Derivatives

Pitching.- The variation of oscillatory pitching derivatives with angle of attack is presented in figures 8 and 9. The data show that the model had positive damping in pitch (negative values of Cmq + Cmb) throughout the test angle-of-attack range for all test conditions. The data also show that there were essentially no effects of power on the pitch damping for angles of attack below the stall. At the stall changes in the pitch damping occurred, which might be associated with separation effects on the wing o r possibly with some alteration of the downwash at the horizontal tail. At the higher angles of attack, the power generally increased the pitch damping for all flap-deflection angles. This increased damping probably can be attributed to the delay in the stall associated with the

7

application of power. The data of figure 9 show that oscillation frequency had no appreciable effect on the pitch damping in the normal operating angle-of-attack range of this configuration.

Rolling.- The variation of the lateral oscillatory rolling derivatives with angle of attack are presented in figures 10 and 11. These data show that the model had positive damping in roll (negative values of Clp c C l - sin a!) throughout the test angle-of-attackP range for all tes t conditions. The effect of power was to increase the roll 'damping. This result was expected since the damping in roll is dependent upon the lift-curve slope, and the data of figure 5 show that substantial increases in lift-curve slope occur with increased power. The relationship between the increase in lift-curve slope and the increase in roll damping due to the power is indicated in figure 12. These data show considerable scat ter but indicate that the percent increase in rol l damping due to power was only about one-half as great as the percent increase in lift-curve slope. The probable reasons for this resul t are that the comparison is made fo r the complete airplane ignoring any effects of the vertical tail and that most of the power-induced lift, and consequently the power-induced increase in lift-curve slope, is produced on the inboard sections of the wing where the moment a r m fo r producing rolling moments is small.

Yawing.- The variation of the lateral oscillatory yawing derivatives with angle of attack is presented in figures 13 and 14. These data show that the model had positive yaw damping (negative values of Cnr - CnB cos a) throughout the test angle-of-attack range for all test conditions. For angles of attack below the stall, there appeared to be only small effects of power on the yaw damping. The data a lso show that the roll-dueto-yawing parameter (Clr - Clb cos a)was positive and increased with increasing angle of attack up to the stall. The data of figure 14 show a slight effect of frequency for the power -off condition.

SUMMARY OF RESULTS

The results of an investigation to determine the static and dynamic stability derivatives of a model of a large jet transport equipped with an external-flow jet-augmented flap may be summarized as follows:

1. The static longitudinal data show that blowing on the flap system increased the lift-curve slope, delayed the stall, and increased the maximum lift coefficient. The data a lso show that all the model configurations generally had static longitudinal stability over the test angle-of-attack range and that at the higher deflection angles the effects of power were destabilizing.

2. The effects of power generally were to reduce the directional stability and to increase the positive dihedral effect especially for higher flap-deflection angles.

8

3. The model had positive damping in pitch, roll, and yaw throughout the test angleof-attack range up to and slightly beyond the stall for all test conditions.

4. The application of power in the jet-augmented-flap system resulted in appreciable increases in roll damping but produced essentially no effects in pitch and yaw damping. The percent increase in the roll damping associated with power was found to be about one-half as great as the percent increase in the lift-curve slope.

5. There were essentially no effects due to oscillation frequency on the damping derivatives.

Langley Research Center, National Aeronautics and Space Administration,

Langley Station, Hampton, Va., June 5, 1969, 721-01-11-03-23.

9

REFERENCES

1. Campbell, John P.; and Johnson, Joseph L., Jr.: Wind-Tunnel Investigation of an External-Flow Jet-Augmented Slotted Flap Suitable for Application to Airplanes With Pod-Mounted Jet Engines. NACA TN 3898, 1956.

2. Lowry, John G.; Riebe, John M.; and Campbell, John P.: The Jet-Augmented Flap. Preprint No. 715, S.M.F. Fund Paper, Inst. Aeronaut. Sci., Jan. 1957.

3. Johnson, Joseph L., Jr.: Wind-Tunnel Investigation of a Small-scale Sweptback-Wing-Jet-Transport Model Equipped With an External-Flow Jet-Augmented Double Slotted Flap. NASA MEMO 3-8-59L, 1959.

4. Johnson, Joseph L., Jr .: Wind-Tunnel Investigation of the Static Longitudinal Stability and Tr im Characterist ics of a Sweptback-Wing Jet-Transport Model Equipped With an External-Flow Jet-Augmented Flap. NACA TN 4177, 1958.

5. Fink, Marvin P. : Aerodynamic Characteristics, Temperature, and Noise Measurements of a Large-Scale External-Flow Jet-Augmented- Flap Model With Turbojet Engines Operating. NASA T N D-943, 1961.

6. Parlett, Lysle P.; Fink, Marvin P.; and Freeman, Delma C., Jr. (With appendix B by Marion 0. McKinney and Joseph L. Johnson, Jr.): Wind-Tunnel Investigation of a Large Je t Transport Model Equipped With an External-Flow Jet Flap. NASA TN D-4928, 1968.

7. Mechtly, E. A.: The International System of Units - Physical Constants and Conversion Factors. NASA SP-7012, 1964.

8. Chambers, Joseph R.; and Grafton, Sue B.: Static and Dynamic Longitudinal Stability Derivatives of a Powered 1/9-Scale Model of a Tilt-Wing V/STOL Transport. NASA T N D-3591, 1966.

10

TABLE I.- DIMENSIONS OF MODEL

Wing: Area. ft2 (mz) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.87 ( 0.731) Span (to theoretical tip), in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 93.71 (238.02) Aspect ra t io . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.75 Length of mean aerodynamic chord, in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.22 ( 33.59) Location of quar te r chord of mean aerodynamic chord, referenced to nose of model, in . (cm) . . . . . . . . . . . . . . . 40.54 (102.98) Spanwise station of mean aerodynamic chord, in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19.33 ( 49.10) Root chord, in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19.49 ( 49.50) T i p chord (theoretical tip), in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.54 ( 16.62) Break station chord, in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.08 ( 30.67) Spanwise station of break station, in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20.16 ( 51.20) Sweep of quarter-chord line:

Inboard panel. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24.08 Outboard panel. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25.00

Dihedral of quarter-chord line: Inboard panel. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -3.50 Outboard panel. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -3.50

Incidence of mean aerodynamic chord. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.50 Incidence of root chord. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3.50 Geometric twist:

R o o f d e g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.0 Break station. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -1.5 Tip. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -3.5

Vertical tail: Area. It2 (m2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.22 ( 0.113) Span. in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14.76 ( 37.48) Length of mean aerodynamic chord. in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.95 ( 30.35) Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.240 Sweep angles:

Leading edge. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 36.6 Trailing edge. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29.4

Root chord. in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.22 ( 33.59) T i p chord. in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.58 ( 26.87)

Horizontal tail: Area. f t2 (m2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.73 ( 0.161) Span. in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35.88 ( 91.13) Length of mean aerodynamic chord. in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.28 ( 18.49) Location of quar te r chord of mean aerodynamic chord. referenced to nose of model. in . (cm) . . . . . . . . . . . . . . . 96.09 (244.06) Incidence . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Variable

Engines: Spanwise location of inboard engines. in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17.10 ( 43.43) Spanwise location of outboard engines. in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26.61 ( 67.58) Incidence of a l l engine center l ines relative t o wing-chord line. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -3.00

Moment reference: Longitudinal location. referenced to nose of model. in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40.54 (102.98) Vertical location. referenced to top of fuselage at wing. in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.92 ( 12.49)

Control-surface dimensions: Rudder:

Span. in . (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.53 ( 26.75) Chord. upper end. parallel to water line. in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.67 ( 6.78) Chord. lower end. perpendicular to hinge line. in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.72 ( 6.91) Hinge-line location. percent chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75 Sweep of hinge line. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31.4

Elevator: Span. in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.26 ( 33.69) Chord, outboard, in. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.27 ( 3.23) Chord, inboard, i n. (cm) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.53 ( 6.43) Hinge-line location. percent chord . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73 Sweep of hinge line. deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16.5

11

4 FL

Figure 1.- Axes systems used in presentation of data. Arrows indicate positive direct ion of moments, axis directions, and angles.

12

- I

t . + -\ -.-17.10

t

I_93.71 (238.02) 3 c 94.96 (241.21)

(a) Three-view drawing of complete model.

Figure 2.- Drawing of model used in investigation. Al l l inear dimensions are in inches (centimeters).

I

Leading-edge slat, 0.15 c

i-Aft &men;

0.12 c

Aft-segment Aft-segmentForward-segment Forward-segment 6f1’6f2 gap, Gfb overlap, ($1,

deg’dq percent c percent c

10/20 6.5 0

gap, ‘32, overlap, Of2,

percent c percent c

3.0 0

20/40 ~ 5.0

30/60 3.02.5

(b) Flap assembly and engine-pylon details. (Incidence of t h r u s t l ine to wing-chord l ine may be obtained in table I.)

Figure 2.- Concluded.

1

"0

W C

'" "0 0 OJ E "0

0

'0 E

~ '0 OJ ':; V>

.s::: c. c.

09- E 0'>

:§ .B 0

.s::: c...

15

with

Flywheel

(a) P i tch ing setup.

Figure 4.- Sketches of test setup for oscil latory force tests.

17

(b) Rolling setup.

Figure 4.- Continued.

18

(c) 'Yawing setup.


19

111111111.111111.11.1 I I I .I., . .

1. c

.5

c m

0

-.5

1.5 CL

1.0

.5

c D 0

-.5

I -1.0 . ,

-5 0 5 10 15 20 01. deg

Figure 5.- Effect of power on static longitudinal stability derivatives.

20

25

0 5 10 15 a, deg

(b) hfl/hf2 = 10°/28; it = 8; 6, = -18;6,, = 5 8 .


21

.78

1.

c m

I

_ I

4.(

0 0

c L 1.70 i 3. f i , : I

3. c

2.5

2.0

1.5

1.0

. 5

0

c D ! !

-.5 3

-1.0

-1.5

(c) bf&z = 2Oo/4O0; it = -5'; 68 = -10'; 6,1 = 50'.


22

I.<

cm t 1 s

f

-.5

cT 4.5 0 0

>.A .38

.78 0 1.70 c~ 4.0 I!

3.5 I

3 . 0

~

2.5 ! i

2.0 ii

1.5 t

1.0

.5

cD 0

-.5

-1.0 -5 0 5 15 20 25

(d) 6f& = 3Oo/6O0; it = -9'; 6, = 0'; 6,l = 60'.


23

CY

c"

.5

0

-.5

.02

0

-.02

.05

0

-.05 -5 0 5

B, deg

i i

a, d e -2.8 1.2 5.2 9.2

13.2 17.2 21.2 25.2

/ j ' ,, , #/

I ! , , I

-5

Figure 6.- Static lateral stability derivatives of model.

24

5

.5

cy 0'

-.5

.04

le

3 -2.8.02 3 1.2

3 5.2 A 9.2 tl 13.2 0 17.2 0 21.2 0 25.2

-.02 111

-.04 I /

.05

o i

-.05 -5 5 -5 0 P, deg

(b) 6f1/6f2 = 2Oo/4O0; it = 00;6, = -100; 6,, = 50°.


25

.5

0 i

-.5

.a

.02 I1 0 'I

'i -.02 !

-.04 ! I

.IO I I

.05

0

-.05 -5 0 5

P, deg

(c) 6,,/6,, = 10°/200; it = -5O; 6, = -100; 6,1 = 50'.


26

- I

CT = 0.381

C T = 0.38

-5 0 5 -5 P, deg

(d) 6 f l / b f 2 = 3Oo/6O0; it

Figure 6.-

C T = 0.78

2-r = 0.78

0 P, deg

= -9O; 6, = Oo; 6,l = 60°.

Concluded.

27

I

0 0

C "p

0 5 15 20 25

(a)

Figure 7.- Static lateral-stabi l i ty parameters of model with respect to angle of attack.

28

0

CYp -.01

-.02

.OM

.002

-.002 3 0 3 .

-.OM

,002

-.002

-.004 -5 0 5 10 15 20 25 30

(b) + 1 / 6 ~ = 10°/200; it = 0'; 6, = -10'; 6,1 = 50'.


29

0

CYp -.01

-.02

.OM

.002

-.002

-.004

.002

0

-.002

-.004 I /

-.006

-.008 i l -5 0 5 10 15 20 25 30

(c) bfi , /bf2 = 2Oo/4O0; it = -5O; be = -10'; 6,1 = 50'


30

0

-.02

.004

.M)2

"P 0

-.002

.002

0

-.002

Cb -.004

-.006

-.008 0

(d) +1/6f2

, i i i i i 'T

3 A . 3 8 0 .78

25r 0

= 3Oo/6O0; it= -9'; 6, = 0'; 6,, = 60'.


31

C

0

'm +'m9 6 -50

-100 c T

0 78

20

IO ic + c xq xa

0

c + c zq zli

5 15 20

(a)

Figure 8.- Var ia t ion of osci l latory p i tch ing derivatives with angle of attack k = 0.045.

32

25

0

-50

20 0 0 0 . I

10 0 1. ;

c + c xq *ti

0

-10

0

-20

-40 1

-60 -5 0 5 10 15 20 25

a, deg

( b ) 91/6f2 = 10°/200; it = 0'; 6, = -ioo; 6,, = 50°.


33

c + c xq

c + c Zq za

0

-50

-100

20

10

0

-10

0

-20

-40

-60

-80-5 0 5

T

- 7 8 t70

10 20

34

25

0

-100 0 .38 ,787n

20

10

-10

-20

0

-20

-40 c + c

zq zu

-60

-80

-100 0 5 10 15 20 25 4 deg

(d) brl/dr2 = 3Oo/6O0; it = -9'; 6, = 0'; 6,l = 60'.


35

0

c + c ma

-50 1 30

k

0 0.023 045

20 c t c

xq xa

10

0

0 j : I

-20

-40 hcz t c q za

-60

-80. 5 10 15 a, deg

(a) CT = 0.

Figure 9.- Effect of frequency on dynamic pitching derivatives. 6f1/6f2 = 3Oo/6O0; i t = -9'; 6, = Oo; 6,1 = 60'.

36

0

c + c mq mdr -50

-100

20

10

0

cx + c q

-10

-20

-30

0

-20

-40

-60 c t c

zq za -80

-100

-120

-140

I

kI.

0 0.0

I

I

I I

I

-I ,I

I

I

F . ' I . !

I

-5 0 5 10 20 25

(b) CT = 1.70.


37

CY +Cy.sin a P P

Cn + Cn.sin a P P

C + C sinab

. 2

0

-.2

0 . I 78

0

-.1

-.2

0

-.2

-.4

-.6 -5 0 5 15 20

f igure 10.- Variat ion of lateral osci l latory r o l l i n g derivatives w i t h angle of attack. k = 0.321.

38

25

.2

C + C sinaO yP yP

-.2

0

-.1

-.2

-.3

0

-.2

-.4 C + C sina zP %

-.6

(b) 6f1/6r2 = 1O0/M0;it = 0'; Be = -10'; 6,, = SO0.


39

Cy +Cy sin P P

C n + C n . sin P P

-5 0 5 10 15 20 25

(c) bf1/hf2 = 2Oo/4O0; it = -5'; 6, = -IOo; 6,, = 60'.


40

. 4

2

CY +Cy sin a P P

0

-.2

. 1

0

Cn + Cn-sin a P P

-.1

-.2

-.3

0

-.2

C + C sina *P b

-.4

-.6

I

1 J

0

0 .38

10 15 20 25

(d) 6Qr2= 3Oo/6O0; it = -9O; 6, = Oo; 6,, = 60°.


41

CYP +Cybsina

k

0 0.160 0 .321

0 10 15 20 25 a, des

(a) CT = 0.

F igure 11.- Effect of frequency on dynamic ro l l ing derivatives. bf1/bf2 = 3Oo/6O0; it = -9'; be = 0'; 6,l = 60'.

42

- 4

C + C s ina yP yb - 2

0

.1

0

Cn + Cn.sin a P P

-.1

-.2

-.3

0

-.2

C + C s ina zP 4j - .4

-.6

-.8 -5 0 5

I

111I / /rl

0 0 0 .321

10 15 20 25 a, deg

(b) CT = 1.70.


43

--

1 1 1 1 I I . .,

2.4

2.2 n) C L ~( p o w e r on --

L a e r off 2.0

1.8 CT 6f ll6f 2

deg1deg

1.6 0.78 010 .78 10120

1.70 10120 .78 20I40

1.4 1.70 20I40 .38 30160 .78 30160

1.70 30160 1.2 Open symbolj a = $ '

0 ..Sol id SymbOlj a = 5.

1.0 1.0 1.2 1.4 1.8 2.0 2.2 2.4

CL, ( p o w e r on)

CLa ( p o w e r o f f )

Figure 12.- Increase in roll damping compared wi th increases in l i f t-curve slope due to jet-augmented-flap operation.

44

1.0

C - C cosa y r YB .5

0

0

-.1

Cn - C cosa r "b

-.2

-.3

.5

0 -5

Figure 13.- Variation

78

I

0 5 10 15

(a) 6f1/6f2 = Oo/Oo; it = 0'; 6 , = 0'; bSl = 0'.

of lateral osci l latory yawing derivatives w i th angle of attack. k = 0.321.

45

a-r .5

-. C - C n cosa 1B "r

-.2

-.3

-.4

1.0

C l - C cosa r b 1"5

0-5 0 5

.78 1.70

10 15 20 25

Ib) b f p f 2 = 10°/200; it = 0'; 6, = -10'; 6,, = 500.


46

1.0

0

0

-.1

Cn - Cn cos a r b

-.2

-.3

1.0

C l-C cos a r z r j .5

0 5 10

1I I tTIIIII

cT 0 .78

1.70

15 25

(c) 6f1/6f2 = 2Oo/4O0; it = -5'; be = -10'; 6,l = 50'.


47

I I 1

I Pcy; cy*cos a .5 . I

1I

4I

I II.I

,1 I I

0 II

i o 0 iiA .38

0 .78 I

0 0 1.70 1II

r I

1 I II

1I I I

-.1 I I < iI E L YII 4 1 i

iIIIIIIiIIC - C cosa

7b III

III

sL 'IIiI

-3 5 15 20 25

6(d) b f l / b f 2 = 3Oo/6O0; it = -9'. , e = 0'; bSl = 60°.


48

1. (

CY - CYbcosa r E

. J

(I

0

C - C n C O S Q “r -.1

-.2

1.0

1 0 ,‘

75

k 0.160

CI :321

1 1 1 1 1

0 5 10 15 20 25 a, deg

(a) CT = 0.

Figure 14.- Effect of f requency on dynamic yawing derivatives. 6fl/bf2 = 3Oo/6O0; it = - 9 O ; 6, = @; 6,, = 60’.

49

l l 111 l l l l l l l l l Ill I1 I

II II I

I II Ii I

II

II I

II

III

II I

II III

I

~

T

i

\ .I

iII

IIII I

I IIIIII

\

+

~

0 5 10 15 20 25

(b) CT = 1.70.

Figure 14.- Concluded,

50 NASA-Langley, 1969 -2 L-6521

AERONAUTICSNATIONAL AND SPACE ADMINISTRATION WASHINGTON, D. C. 20546

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static and dynamic stability derivatives model ofa wind-tunnel investigation has been made to...

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