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RESEARCH MEMORANDUMfor the

Bureau of Aeronautics, Department of the Navy

HIGH-SPEED W_UTD—MgNEL =STIGATION OF TIE LONGITUDINAL

STABILITY AND CONTROL CHARACTERISTICS OF A 0.10-SCALE

MODEL OF THE CrRU ,RPN )T9F-2 AIRPLANE

TED NO. NACA DE301

By

Edward C. $Polhamus and Thomas J. King, Jr.

Langley Aeronautical LaboratoryLangley Field, Va.

This amen[ ^ontains classlfled informationa[fe the :^a[fonal D_ ae of the L'nited

St s !tidn t e me o the Eapion ct,

person is pr ited by taw. Tso classifl . ay be imparted

only to p e r s o n s in 04 .61bury and navalPWPJmwrvices of the UNt Cates, appropriate

civilian officers and employees of the Federal

Government who have a legitimate Interest A/ n A ^^^^therein, and to UNted States cltlzens

NAT eNAL ISO RY CYFOR AERONAUTI To t*

WASHINGTON * f* of to ,Ii(111aIJ i worw 11"111fte

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!c NATIONAL ADVISORY COMMlJYTEE FOR AERONAUTICS

RESEARCH MEMORANDUM

for the

Bureau of Aeronautics, Department of the Navy

HIGHSPEED WIND--TUNNEL INVESTIGATION OF THE LONGI!I!UDDAL

STABnITY AND CONTROL CHARA-ISTICS OF A O.lOSCALE

MODEL OF THE GRUMMAN XF9F-2 AlRPLAIJE

TED NO, NACA DE301

BY ~dwmd C. Polhamus and Thomas J. King, Jr.

SUMMARY

An investigation was made in the Langley high-speed T-by lO-foot tunnel to determine the high-peed longitudinal stability end con&o1 characteristics of a O.lO-scale model of the Gn n9F-2 airplane in the Mach number range from 0.40 to 0.85.

The results indicated that the lift and drag force breaks occurred at a Mach number of about 0.76. The aerodynamic-center position moved reer- wsrd,after the force break and control position stability was present for all Mach numbers up'to a Mach number of 0.85 except for a slight instability around a Mach num%er of 0.80.

l l !JTRODUCTION

At the request of the Bureau of Aeronautics en investigation of the hi&-peed longitudinal stability snd control characteristics of a O.lO-scele model of the Grumman XFgF-2 airplane was conducted in the Langley high-speed 7-by lO-foot tunnel.

This paper presents the results of the longitudinal stabili-ty and control investigation at Mach nwnbers ranging from 0.40 to 0.85. Pitch tests were conducted for various elevator end stabilizer settings and included tests to determine the effect of adding a horn balance to the elevator. The effect of wing-tip tanks on the longitudinal stability characteristics of the model was also investigated.

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coE3FIcIENTsANDsyMBoLS

The stability system of sxes used for the presentation of the data, together with an indication of the positive forces, moments and angles, is presented in figure 1. Pertinent symbols sre defined as follows:

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v

S

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b

M

a

R

CL

a

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it

lift coefficient (Lift/qS)

drag coefficient (Drag/qS)

pitching-moment coefficient measured about the 2Fpercent mean- geometric-chord position (Pitching-moment/qSc')

free-stream dynamic pressure, pounds per square foot

mass density of air, slugs per cubic foot

free-stream velocity, feet per second

wing sxea, square feet

wing mean geometric chord (M.G.C.), feet

wing span, feet

Mach number (V/a)

velocity of sound., feet per second

Reynolds number y ( >

absolute viscosity, pound-seconds per square foot

angle of attack of model, measured from the X-axis to the fuselage reference line, degrees

elevator deflection with reference to stabilizer chord line, degrees

stabilizer setting with reference to fuselsge reference line, degrees

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APPARA!TUSAJYDMEt!HODS

Tunnel and Model

The tests were conducted in the Langley high-speed 7-by lo-foot tunnel, which is a closed re,ctangular tunnel of the return-flow type w a contraction ratio of 15.7 to 1.

ith

The O.lO-scele model was constructed at the David Taylor Model Basin, Carderock, Maryland. It was constructed entirely of steel. Details of the model as tested ere presented in figures 2 and 3. The model was tested through a Mach number rarige of 0.40 to 0.85 at various angles of attack on the sting support shown in figure 4, Elevator and stabilizer effectiveness and the effect of the wing-tip tanks (see table I) were investigated throughout the Mach number end angle-of-attack ranges.

The variation of test Reynolds number with Mach number for average test conditions is presented in figure 5. The degree of turbulenc;e oT the tunnel is not known but is believed to be smell because of the high contraction ratio. Experience has indicated that for a model of this size constriction effects should not invalidate the test results at corrected Mach numbers ‘below about 0.91.

Support System

A sting support system was used to support the model in the tunnel and a photograph of the test setup is presented as figure 4. The sting extended from the rear of the fuselage to a vertical strut located behrind the test section. This strut was mounted. on the tunnel balance system ad. was shielded from the air stream by a streamline fairing. The tare f'orces end. moments produced by the sting were determined 13s mounting the moilel on two wing stings, which were also attached to the vertical strut, and teding the model with and. without the center sting. With the center sting in

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place the duct flow was by-passed through a hole in the underside' of the aft portion of the fuselage while without the center sting the flow was exhausted out of the rear of the fuselage. Therefore, the corrected. data represent the, conclition with flow out of the rear of the fuselage. Angles of attack were changed by the use of interchangeable couplings in the stings aft of the model. The deflections of the support system under load were determined from static loading tests.

Corrections

The test results have been corrected for the tare forces and moments produced by the support system. The jet-boundary corrections were computeif from the following equations, which were determined by the method of reference 1:

CL = s + 0.26~~

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where the subscript m indicates measured value.

The drag has been corrected for the buoyancy produced by the longi- tuainal static-pressure gradient in the tunnel, end. the dynamic pressure and Mach number have been corrected for blocking by the model end its wake I)y the method of reference 2.

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RESULITS AND DI;SCXJSSION

Basic Data

The basic data for the various configuratiwns tested are presented in the following table:

--~ Type of test

Elevator effectiveness



Stabilizer

Tip tanks Elevator

On Plain

Off Plain

On Horn balanced

On Plain

Figure

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7

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9

General Aerodynamic Characteristics

A summery of some of the more important aerodynamic parameters as evaluated from.the basic data is presented.in figures 10 and 11.

Lift and drag.-The lift-curve slope (fig. 10) increases with Mach number up to a Mach nwnber of about 0.76, above which there is a rapid reduction. The end-plate effect of the wing-tip tanks increased the lift- curve slope by about 15 percent at a Mach number of 0.40 and by about 20 percent at the force breek Mach number of 0.76. Above the Mach number for lift force break the effect of the tip tanks rapidly diminished. Low- speed wind-tunnel tests of a similar configuration (reference 3) indicate a tank effect of the same order of magnitude as that obtained at the lowest Mach number in the present investigation.

The drag at zero lift (fig. 10) decreases slightly with increasing Mach number up to the drag force break Mach number of about 0.76. This effect has been observed on other models and has been attributed to the increase in Reynolds number with increasing Mach number. Above a Mach nrnnber of 0.76 there is a rather rapid drag rise due to the.onset of compressibility effects. The wing-tip tanks added a constant drag- coefficient increment of about 0.002 up to a Mach number of about 0.76. The rate of increase of drag coefficient with Mach number above the drag break Mach number was somewhat greater with the wing-tip tanks on than with them off.

6 NACA F&I No. SL~KI~6

The msxlmum value of the lif%-drag ratio (fig. 10) is fairly constant up to a Mach nmber of about 0.75, above which it decreases. Except at the low Mach numbers the maximum lift-drag ratio is greater for the tip- tanks-off condition. The effect of the wing-tip tanks on the variation of the lift-dragratio with lift coefficient for sever&L Mach numbers is presented in figure 11. Only at the lower Mach numbers in the high lift range where the induced drag is important is there an increase in the lift- drag ratio due to the end-plate effect of the wing-tip tanks.

Control effective-.- The elevator an& stabilizer effectiveness increased slightly with Mach number up to 0.76 and 0.80, respectively, above which they both decreased rapidly. The horn balance (fig. 3) had a negligible effect on the elevator effectiveness (figs. 6 end 8) as did the wing-tip tanks (fig. 7). The ratio of elevator to stabilizer effectiveness at a Mach ntmiber of 0.40 is about 0.76. Although this ratio seems high when compared with sn estimated value of flap lift effectiveness of about 0.62 (deference 4), it is in good agreement with low-sped wind- tunnel tests of a larger scale model of the same airplane (references 5 =a 6).

Aerodsnemic center.-The r&e of change of pitching-moment coefficient with lift coefficient at a constant Mach ntrmber C qL (fig. 10)

-is a measure of the aerodynamic-center location relative to the assumed center-ofwavity position in percen$ of the mean geometric chord. The aerodynamic center is seen to move forward from about 10 percent behind the center of gravity at a Mach number of 0.40 to about 8 percent at a Mach number of 0.76 and then moves rapidly rearward. The wing-tip tanks shifted the aerodynamic center forward. by about 1 percent throughout the Mach number range.

Estimated Flight Characteristics

-.-The variation of the lift coefficient required for level flight with Mach nmber for various wing loadings ena Kltituaes is presented in figure 12. Also presented are the lift-drag ratios corre- sponding to the level-flight lift coefficients. The variation of these lift-drag ratios illustrates the advantage of flying at the higher altitude.

Static i0ngituainal stability.- The variation of the elevator position with Mach number is presented in figure 12 for various wing loadings and altitudes. Control position stability is present throughout the M&h number range except for a smell region around a Mach number of 0.80. How- ever, the trim changes associated. with this instability ere very slight.

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Maneuvering stability.- The response of an airplane to gusts and rapid control deflections is governed by the location of the aerodynam ic center and the damping of the horizontal tail. Inasmuch as the damping of the tail will be stabilizing and the aerodynam ic-center position (fig. 10) is never less than 8 percent of the mean geometric chord behind the assumed centem f-gravity position, maneuvering stability will be present throughout the Mach number range tested..

CONCLUSIONS

Based on high-speed. wind-tunnel tests of a O.lO-scale model of the Grumman XFgF-2 airplane in the Mach nuriber range from 0.40 to 0.85, the following concltisions have been reached:

1. Lift and drag force breaks occurred at a Mach number of about 0.76.

2. The position of the aerodynam ic center moved forward only about 2 percent of the mesn geometric chord up to the force break and moved resrwsrd rapidly after the force break.

3. Control position stability was present throughout the Mach number range except for a small region around a Mach number of 0.80. However, the trim ohanges associated with this instability are very slight.

4. W ing-tip tanks had m inor effects on the stability characteristics but decreased the lift-drag ratio at the higher Mach numbers.

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8. . . A NACA RM No. s-6 5. The horn bslance had a negligible effect on the elevator

effectiveness.

Langley Aeronauticel Laboratory National Advisory Committee for Aeronautics

Larigley Field, Va.

Edward C. Polhamus Aeronautical Resesrch Scientist

Thomas J. King, Jr.

@&Y&e&/u Aeronautical Research Scientist

Approved.: Thomas A. Harris

Chief of Stability Research Division

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REFERENCES

1. Millie, Clarence L., Polhamus, Edwad C., and Gray, Joseph L., Jr.: Charts for Determining Jet43oundary Corrections for Complete Models in 7-by lo-Foot Closed Rectangular Wind Tunnels. NACA AKR No. L5G31, 1945.

2. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, with Gonsideration of the Effect of Compressibility. NACA RM No. ~7~28,~ 1947.

3. Tucker, Warren A., and Goodson, Kenneth W.: Tests of a l/wcale Model of the Republic XP-84 Airplane (Arxtw Project ~~378) in the Langley 300 MPH 7-by lo-Foot Tunnel. NACA MR NO. L6F25, Arqy Air Forces, 1946.

4. Sears, Richard I.: Wind-Tunnel Data on the Aerodynamic Characteristics of Airplane Control Surfaces. NACA Acx NO. 3~08, 1943.

5. Shm, William H., andwatriss, Frederic W.: Wind. Tunnel Tests on Grm Aircraft Engineering Corporation Design 79, Fourth Series. M.I.T. Wind Tunnel Rep. No. 762, Feb. 1947.

6. Tamburello, V., and Beek, Charles R.: Wind-Tunnel Tests of a 1/54%ale Powered Model of the XF9F-2 Airplane. Part I - Longi- tudinal Characteristics, Stabilizer Tdsts -TED No. rmB 301. Rep. C-131Aero 758, David. Taylor Model Basin, July 1948.

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Figure 3.- Drawing of the stabilizer and horn-balanced elevator of the O.lO-ecale model of the XFgF-2 airplane.

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Figure 4.— The 0.10—scale model of the XF9F-2 airplane mounted on the sting support.

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Figure 8.- Effect of elevator deflection on the aerodynamic characteristics in pitch 0: the O.lO-scale model of the XFgF-2 airplane. Horn-balanced elevator; tanks on; it = 0 .

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Y' -4 :2 o .2 4 .6 .8 10 L iff coefficienf , CL

S,,(degl - 545 + -e- -4% b_ Tci/off

// /,..I/ ,I 88 o- A. . . 0

00 ..” i 0 00 0

6 I I ldifl I I )

I/Y11 I I

4 0

% -4 :2 0 .2 4 .6 .8 1.0

Lift coefficient, CL

% R 3=

(id M = 0.775. (h) M = 0.800. Figure 8.- Continued.

t? -b ti‘ -s P a ko 4 3 P

.I?

.i

0

:i 20

T2 I I I I nIXI- I

12

6

a

0

-++ 72 0 .2 4 .6 .8 LO Lift coefficienf, CL

(i) M = 0.825.

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0 *a-

4.. i : t 0. l a l a


= ~.zr-yvs.,~ . . . . :&A . . . 0

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cn 9 :/ c

p 9

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0 0 m 12 $ 12

% $ $8 $ 8

5 2 b o 6 4 Q4

‘“, %O BO * ? 3 -4 t91BwmEwr* -? -4 - -7 ’ -4 2 0 .2 6 .6 .8 LO

Liff coefficient, CL

-4 72 0 .2 A .6 .8 LO Lift cocfficient,C~

(a) M = 0.400.

Figure 9.- Effect of stabilizer setting on the model of the XFgF-2 airplane.

(b) M ='0.500.

aerodynamic characteristics in pitch of the O.lO-scale Plain elevator; tanks on; 6, = 0'.

. I

Y .2 .$ $ ./ $ % 0 P 3 :/

F! . . 0 . .

l too l .0 l : :

. . t: A. . . me : . : l : 00 I

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-4 2 0 2 4 .6 ~3 10 Lift coefficient, CL

$8

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.08

-sf :2 0 .2 4 .6 .a A0 LifS coefficient, CL

(c) M = 0.600.


(a) M = 0.650.

8 .,,. . . . . f .

b 00 bb

I

-1 I ,.” I I I I I I

.20

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.08

.04

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Y - \ 74 2 0 .2 4 .6 .8 LO 74 72 0 .2 .4 .6 .8 LO Lift coefficienf, CL Lift coefficient, GL

(e) M = 0.700. (f) M = 0.750.

/2

8

4

0

-4


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./! $ $

.08 $

.04 $ 9

0

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I T -4 :4 :A! 0 .2 .4 .6 .8 10 :4 :z 0 .2 4 .6 .8 LO

Lift cocfficien+, & Liff coefficienf, CL

(g) M = O-775* (h) M = 0.800.

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./2

.08

.04 0

it, /Hey1 -0 + -2

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b

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(i) M = 0.825.

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: :: l

: *a.

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00

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(j)M = 0.850.


./6 G

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0

86

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Cm Je

und TO26 Cm.

'~036

\ 7044

0 cm

ci T/Q

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- .

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.:. . . 07-T

. . l

.J+ 5 .6 .7 .8 .9 .4 .5 -6 .7 ..a .s M7c4 m777~e~ M ' A&CA nw77~e~ A4

Figure lo,- Effect of Mach number on various aerodynamic characteristice, plain elevator.

.: .

.: .

*.

NACA RM No. 2~8~1-6

. : .

. . . . :

l : /6

/z

8

4

a 0 .z 4 .6 .B 0 .Z .-a -6 .8

L/f7 coefic/enf, G

22 I A!4-/)7Li- I I I

I I

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,I t-t I/-TV I I I I u\ /6

$ 1 /z 8 -w .< 44

/6

0 .z 4 .6 .8 0 .Z ‘4 - -6 A L/ft co~fficbe77f, CL

Figure 11.- Effect qf wing-tip tanks on the lift-drag ratio of the O.lO-ecale model of the XFgF-2 airplane, it = O", plain elevator.

I .:

‘: ..i NACA RM No. sL8K16

. . .

. . : . .

l . . .

. .

4 .

.8

.6

4

2

0

--- ---- ---

40 Seu /eve/

4 .5 .6 :7 .6 .9 LO Mach number, N T

Figure 12.- and CL

Variation with Mach number of elevator position L/D for level flight of the XFgF-2 airplane. Plain

elevator; tanks on; it = 0'.

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