NATIONAL ADVISORY COMMITTEE I: gi FOR...

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AERODYNAMIC STUDY OF A WINGFUSELAGE COMBINATION EMPLOYING

A WING SWEPT BACK 63: - EF!?ECTS AT SUBSONIC SPEEDS Cl? A

CCNS7llNT-CXlRD EILEVa CNA WING CAMERED ANI TWISTED

FOR A UNIFORM LOAD AT A LIFT COEFFICIENT Q? 0.25 R

By J. Lloyd Jones and Fred A. Demele

Ames Aeronautical Laboratory Moffett Field, Calif.

NATIONAL ADVISORY COMMITTEE I: gi FOR AERONAUTICS ‘. ,

NACh RM A9127

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AERODYNAMIC STUDYOFAWING+USELMZ COMBINATIONEMPIQYII?GAKlXG mm BACK 63Q.- EFFECTSAT SUBSONIC SPEZDS OF A CONSTAXT-

.CEORD EIEVON ON A WING C- Al!JDTWIS!EDFCRA IJNICFORM LOAD AT A LI??T COEZ‘FICIERT OF 0.25

By J. Lloyd Jones and Fred A. Dem4.e

A cambered and twisted wing having a leading edge swept back 63O and equipped with constant-chord elevens was tested in combination witha slender fuselage to determinethelongitudinal and lateral control afforded by the elevens from a Mach number of 0.20 up to a Mach number of 0.93. The tests were performed at &Reynolds number of 2.0 IllilliOn. Data are presented showing lift, drag, pitchingqoment, and rolling+ao~nt characteristics of the model for various eleven deflections, andhinge- nt characteristics of the eleven. Data from the tests have been applied to the calculation of the longitudinal-stability and -control characteristics of a hypothetical airplane geometrically similar to the model.

With the elevens undeflected, the model was longitud3w unstable about the one-quarter point of the WFng mean aerodynamic chord at lift coefficients above about 0.50. The elevone had sufficient pitching- momentandrolUng+nome nt effectiveness for all lift coefficients at which the model was longitudinally stable. At low speeds, the lift coefficient at which static longitudinal tistability occurred was decreased by increasing negative eleven deflection. Increasing the Mach number increased the pitching- nt effectiveness at lift coefficients above 0.20, but reduced the roll.iug+nomnt effectiveness of the elevons.

INTRODUCTION

A coordinated research program has been undertaken by the Ames Aeronautical Laboratory for an aerodynamic investigation of a wing- fuselage combination employfng a wing having the leadirg edge swept back 63O. Aerodynamic characteristics of such a wing with no camber or twist have been presented in references 1, 2, 3, and 4. Reference 1 includes low-speed data on the effectiveness of a constant-chord eleven, and reference 2 reports the Mach nuviber andReynolds number effects on the effectiveness of the ssme eleven.

2 NACA RM A9127

Camber and.twist have been incorporated in the wing in an effort to improve the flow near the wing tips where , as was evident from early investigations,. loss of lift occurred even at very low angles of attack. Aerodynamic characteristics of such a wing, cambered and twisted to support a uniform distribution of Lift over its surface at a lift coefficient of 0.25 and a Mach number of.l.5, have been presented in references 5 and 6.

This*report presents the results of tests in the Ames U-foot pressure wind tunnel of the effectiveness and hinge moments of constant- chord elevons at Mach numbers ranging up to 0.93. The elevens extended over tha outer 50 percent of the span of the canibered and twisted wing, which is described in reference 6, and had the ssme plan form as ths elevons on the model used for the tests reported Fn references 1 and 2.

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speed of sound, feet per second

wing epanmeasured perpendicular to plane of symmetry, feet

local chord laeasured parallel to plane of symmetry, feet

wing mean aerodynamic-chord

drag coefficient

hinge-moment coefficient hIng e moment 2q x area moment of eleven

about eleven hinge axis

lift coefficient

rolling+noment coefficient (

rolling moment sm - >

dsmping-moment.ccefficient ti roll; the rate of change of rolling -moment coefficient C2 with m-tip helix angle pb/2V, per radian

pitching-moment cdefficient about the one-quarter point of the

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normal acceleration factor

angulsr velocity in roll, radians per second

dynamic pressure , pounds per square foot

Reynolds number y ( 1

wing area, square feet .

free--stream velocity, feet per second

ainking speed, feet per second

gliding speed, miles per hour

lateral ordinate, feet

angle of attack of root chord line, degrees

angle of twist with reference to root chord (positive for washin), degrees

angle of attack of root chord line, uncorrected for tunnel-wall interference, degrees

eleven deflection measured In planes perpendiculsr to the eleven hFnge axes (positive downward), degrees

eleven deflection uncorrected for angular distortion due to load, degrees

left eleven deflection uncorrected for angular distortion due to load,degreee

right eleven deflection uncorrected for angular distortion due to load,degreee

arithmetic sum of positive and negative elevonpeflections, degrees

4 a- & NACA RM A9127

arithmetic sum of positive and negative eleven deflections uncorrected for angular distortlon due to load, degrees

P coefficient of viscosity of air, slugs per foot-second .

P mass density of air, slugs.p.er cubic foot

MODELABDAppARATlJS

The model used in this investigation was the one used in the tests reported in reference 6.. Photographs of the model are presented in figure 1 and dimens~one are given in figures 2 and 3.

The wing had a leadingedge sweepback of 63O, a taper ratio of 0.25, and an aspect ratfo of 3.5. The streamwise airfoil sections had the NACA 64AOO5 thickness distribution combined with a = 1 man caniber lines. The wing, as developed theoretically by the method given in reference 7, was cambered and twisted to support a uniform dietribut$on of lift over its surface at a 1Fft coefficient of 0.25 and a Mach number of 1.5. %o provide for twisting of the wing under aerodynamb load's, the model wing was constructed with less twist than was indicated by t'heorg, as is described in reference 6.

The elevone were of constant chord and extended mer the outer 50 percent of the span. Each eleven was supported by three hinges and was restrained near the inner -extremity. The ratio of eleven chord to wing chord was 1 to 4 at the wing midsemispan. The elevons had radius noses with no aerodynemic balance. The nose gape were approximately 3/64 inch a and were unsealed. These large gaps were necessary to permit the desired angular deflection.eince the elevone had considerable spanwise curvature. Hinge moments were measured by-means of &-tiE%%s'Lstance strain gage mounted on the restrainbg m&m? ofthe.elevon on the left&m&wing.

The model was sting mounted, and the angle of attack was continu- ously controllable from a remote station during winGtunnel operation. Forces and momante acting upon the model were measured by mane of a wire-resistance strain-gage balance enclosed by the f&&e.

TESTS

Lift, drsg, pitching- nt, rolling- nt, and eleven-hinge- moment data have been obtained throughout sn angle-of--attack range of. -80 to +190. This range was mqre limited at the larger eleven deflections and higher Mach numbers where vibration of either the model or its support or wind-tunnel power limits. were critical. All tests were made at an angle of sideelip of O". The elevone were defiected negatively for longitudinal control and differentially for lateral control as given in the 'following table:

NACA RM A9127

Eleven deflection angles Longitudinal-control Lateral-control

data data

h&l (w3) 6Ru b%3) f& (ded %u b-g

0 0 0 0 -5 -5 10 -10

-10 -10 -15 -15 420

45

The tests were performed at several Mach numbers ranging from 0.20 to 0.93 at a constant Reynolds nuniber of 2.0 million.

CORRECTIONS

The data have been corrected for the effects of tunnel+all interference, constriction due to the tunnel walls, base pressure, and static tares duetothe weight of the model. No correctian has been applied to account for the change of eleven deflection under load upon the force and moment coefficients except when presented as functions of eleven angle. The angle of attack of the.model was nusasured visually by means of a cathetometer; hence, no corrections were necessary to account for deflection of the support equipment. Precision of the force and moment measurements obtained from the strain-gage balance has been discussed in reference 6.

Tunnel47all Interference

Corrections to the data to account for induced tunnel-wall interference have been determined by the method of Glauert (reference 8). Since the ratio of model span to tunnel diameter was small, the total corrections were small, and no account was taken of sweepback or of the differential flap deflections. The following corrections were added:

kL= 0.26 cL

A% = 0.0046 CL2

No correction was applied to the pitching mount.

6 NACA.BM A9127

Constriction

The constriction effects of the tunnel walls have been evaluated by the method of reference 9. No modification of this method has been made to account for the effects of sweepback. The magnitude of the corrections applied to the Mach number and to the dynamic pressure is fllus- trated by the following table:

Corrected Uncorrected q, corrected Mach number Mach number qJ uncorrected

0.930 0.919 1.012

:EE .884 l 798 1.007 1.003 .600 0599 1.002 .200 .2OO 1.001

Base Pressure

The pressure on the base of the model fuselage was measured snd, in an effort to correct for support interference, the drag data were corrected to correspond to a base pressure equal.to the static pressure of the free stream. The baseqressure correction to the drag was lees than 5 percent for Mach-nurribers up to 0.75, and Increased to approximately 20 percent at a Mach number of 0.93. The base-reseure correction reduced the drag. L

Tare e

There were no tsree due to direct air forces on the model--eupport equipment, since thebalance was tithin the model. CorrectLpns were ~_ made for the change in static tares due to angle of attack.

BESULTS AND DISCUSSION

Longitudinal Characteristics

Eleven effectiveness and hinge momsnte.- Angle of attack, drag coefficient, and pitching- nt coefficient as functions of lift coefficient, and hfnge-moment coefficient as a function of angle of attack szre presented in figures 4 to 8, inclusive, for various elevon deflections for Mach nunibers ranging from 0.20 to 0.93. The angle of attack for zero lift became more positive as the eleven was deflected upward and the minimum drag coefficient was increased considerably by negative eleven deflections greater than -50.

NACA RM AgI27 t 7 w

The elevon had sufficient pitching-mnt effectiveness to provide longitudkal balance at all test Mach numbers for all positive lift cmfficients at which the model had static longitudinal stability. !L'he positive lift coefficient at which the loss of static longitudinal stability occurred (about 0.5) was reduced with increasIng negative elevon deflection at a Mach number of 0.20, and generally increased with negative elevon deflectfon greater than -50 at higher Mach atiers. A slight forward movement of the aerodynamic center at zero lift was noted as the elevon was deflected negatively, and the movement became larger at the higher Mach numbers.

The change of eleven hinge moment with angle of attack was nearly uniform between angles of .attack of -lo and +8o at a Mach number of 0.20 and between -lo and +6O for all other test Mach numbers. The variation of hinge-moment coefficfent with angle of attack became considerably larger at angles of attack beyond these ranges. The sharply defined change of slope of the hingemment curves occurred coincidentally with the re arward movelnent of the aerodynamic center noted In the pitching- Munsnt data.

The variations of lift coefficient, pitching+nomnt coefficient, andhinge -moment coefficient with elevon deflection are presented in figure 9 for constant angles of attack at several Mach nurihers. !T!he pitching-momsnt effectiveness of the elevens was generally malntatied throughout the entire range of elevon deflection.

The effect of Mach nuniher on the pitching-mome nt effectiveness of the elevons and on the lift coefficient for longitudinal balance is shown in figure 10. The pitching-nmme nt effectiveness was nearly inde- pendent of Mach number at lift coefficients below 0.20 over the test range of Mach numbers. The effectiveness -Cm* increased wfth increasing Mach number at lift coefficients greater than 0.20. The lift coefficient for longitudinal balance was essentially unaffected by compressibility up to a Mach number of 0.80 for negative elevon deflection of loo or less, and it is indicated that for negative deflections of 50 or less the lift coefficient for longitudinal balance was little affected by compressUzility throughout the entire test range of Mach nuuibers.

Liftilrag ratio.- Figure II presents the variation of lift-drag ratio with lift coefficient for various elevon deflections at several Mach nuribers. The highest maximum liftilrag ratio occurred with an elevon deflection of -5O, which suggests that Increasing the wing twist would result in a higher maximum lW%-drag ratio for the wing tith the elevons undeflected.

Lateral Control

Eleven effectiveness and hinge moments.- Rolling+uo~nt coefficients due to elevon deflection are presented in figure 12 as a function

8 NACA RM AgI27

cf angle of attack for differential eleven deflections of f loo, f 20°, and f 30' at Mach nun&era ranging from 0.20 to 0.93. Also presented In figure 12 are elevonainge 4bsment coefficients for the left elevon only (the deflection of which was posftive) over the same range of elevon deflections and Mach numbers. These data indicate that the effectiveness of the elevens in producing rolling momxt was maintained throughout the test range of angle of attack and Mach nuriber. The effectiveness was nearly constant at angles of attack between--lo and +80 for a Mach number of 0.20, and between -lo and +6O for the higher Mach numbers. The sxlgles of attack at which the rolling- nt effectiveness of the elevons began to decrease rapidly coincide with those at which the rearwar d movement of the aerodynamic center is noted in the pitching- ntdata. The variation of elevotiinge+noms nt coefficient with angle of attack remained fairly uniform over the same angle-of-attack range for which the msxlmum rollinginament effectiveness was maintained. At angles of attack just beyond these ranges the variation of hinge-moment coefficient with angle of attack became considerably greater, and at the larger positive angles of attack became erratic.

The variation of rolling- nt coefficfent with total eleven deflection (the arithmetic sum of the positive and negative deflections) was

' moth to the largest deflectioqas may be seen in figure 13. Increasing the Mach number from 0.20 to 0.93 reduced the effectiveness byrougbly 10 percent for an angle of attack of 6O and by about 25 percent for en qle of attack of 10' at the largest eleven deflection au = f 30°. The effect of Mach number on the rolling-& effectiveness of the elevons is summsr ized in figure 14 for angles of attack of O" and ho. The rollfng moment produced by a given elevon deflection was generally reduced slightly with increasing Mach nu&er, the effect becoming greater with increasing deflections.

Helix angle.- On the basis of the methods presented in reference 10, helix angles generated by the wing tip in a steady roll have been calculated utilizing the data of fQure 12. For the purposes of the calculations no torsional deflection and O" of sIdeslip were assumed. Values of the damping-mo43n.t coefficient Cz calculatedbyther&hod of reference 11, varied from 4.226 at a Macgnuziber of 0.20 to 4.231 at a Mach number of 0.93.

The variation of the predicted wing-tip helix angle with tot&, elevon deflection &r is presented in figure 15 for various Mach numbers at a lift coefficient of 0.20. As anticipated from the decrease in rolling effectiveness above an sngle of attack of 8', calculations of pb/2V at a lift coefficient of 0.40 indicated a considerable decrease from its value at a lift coefficient of 0.20. No such calculations are presented herein, however, since above a Mach nuziher of 0.20 the test angle-of-attack range was insufficient to evaluate corrections to the , rolling- nt coefficient in roll. The variation of pb/2V with- BT was fairly linear throughout the range of elevon deflections considered.

NACA PM A9127

Increasing Mach number generally reduced the helix angle. While the predicted wing-tip helix angle is large enough to insure high rolling velocities, it must-be emphasized that the present calculations are for a rigid wing and that deflection of the wing could cause serfous reduc- tions in the magnitude. of the rolling velocity.

Longitudinal Control of a Hypothetical Airplane

Data from the tests have been used in the calculation of the stability, maneuverability, elevon hinge moments, and power-off sinking speed of a hypothetical tailless airplane, geometrically similar to the model tested. Dimensions of the airplane were assumed to be as follows:

Wing span, feet . . . . . . . . :. . . . 50 Wing area, square feet . . . . . . . . . 714.3 Total elevon area, square feet . . . . . 89.14

The center of gravity was assumed to be at 25 percent of the mean aerodynamic chord, and a wing loading of 40 pounds per square foot was assmd.

Figure 16 presents elevon hinge moment, elevon deflectian, and lift coefficient as functions of Mach nuniber calculated for the airplane in level flight and as affected by normal acceleration at en altitude of 25,000 feet. The variation of elevon deflection withMach number and ' with normal acceleration factor was smooth and uniform. A very large variation of hinge moment with Mach number is noted for normal acceleration factors greater than 1.0. For unaccelerated flight (n = 1.0) increasing Mach nuniber would require a gradually increasirg push force up to a Mach number of 0.90. For a norm& acceleration factor of 2.0, increasing Mach nuziber is accompanied by a gradually decreasing push force. For con&an-peed maneuvers with varying normal acceleration there are large and erratic changes In the hfnge moment.

Power-ff sinkin@; speed, elevon deflection for balance, elevon hinge moment, and angle of attack are presented in figure 17 as functions of

'power-ff gliding speed for sea-level operation.(Data at a Mach number of 0.20 were used 3.n calculating the performance parameters shown in this figure.) The minimum power-3ff sinking speed is 22 feet per second and it occurs at a forward speed of approximately 215 miles per hour. The variation of elevon deflection required for longitudinal balance with gliding speed was stable for gliding speeds greater than 180 miles per hour. No computations are shown for gliding speeds less'.than l&J miles per hour, since the data indicated that the airplane would be longitudinally unstable at the required lift coefficients.

10

SaMMARY OF REm

BACA RM A9127

Tests have been made of a cambered and twFsted wing with the lead- iqg edge swept back 63O Fp codination with a slender fuselage. The wing was equfpped with constant-chord elevens extending over the outer 50 percent of the span. The tests were conducted at a Reynolds n&e& of 2.0 million and at Mach nmibers ranglug from 0.20 to 0.93. The following results were obtaFned:

1. At low speed (M = 0.20) negative eleven deflections reduced the lift coefficient at which the loss of static longitudinal stability occurred, while at higher Mach nu&ers this lift coefficient generally increased with negative eleven deflections greater than -5'. (With the elevens mdeflected the loss of static longitudinal stability generally occurred at a lift coefficfent of-about 0.5.)

2. There was little effect of compressibility on the pitchlng- moment effectiveness of the elevens at lift coefficients of 0.20 or less. At higher lift coefficients the effectiveness increased with increasing Mach nmiber.

3. The effectiveness of the elevens in producFng rolling moment was reduced slightly with Increasing Mach nmiber. The effectiveness was nearly constant at angles of attack between -lo and +8O for a Mach number of 0.20 and between -lo and +6O for the higher Mach nmibers.

Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Moffett Field, California.

1. HopkFns, Edward J:: Aerodynamic Stuq of aWing4uselage Corcibina- tion Employing a Wing Swept Back 63 .-Effects of Split Flaps, Elevens, and Leading-Edge Devices at Low Speed. RACA RM ~9~21, 1949.

2. Reynolds, Robert M., and smith, Donald W.: Aerodynamic Study of a Wing+Fuselags Combination Employing a Wing Swept Back 630.- Subsonic Mach and Reynolds Number Effects on the Characteristics of the Wing and on the Effectiveness of an Eleven. 1948.

NACA RM A6D20,

3. McCormack, Gerald M., and Walling, Walter C.: Aerodynamic Study of a Wing-Fuselage Co&in&ion Employing a W-Lng Swept Back 63O.k Investigation of a Large-Scale Model at Low Speed. NACA RM Am02, 1949. *

-

NACA RM A9127 - ll .

4. Madden, Robert T.: Aerodynamic Study of a Wing+usslage Conibination Employing a Wing Swept Back 63O.- Characteristics at a Mach Number of 1.53 Including Effect of Small Variations of Sweep. NACA RM A8J04, 1949.

5. Madden, Robert T.: Aerodynamic Study of aWing+uselage Combination Employing a Wing Swept Back 63O.- investigation at a Mach Number of 1.53 to Determine the Effects of Cambering snd Twisting the Wing for Uniform Load at a Lift Coefficient of 0.25. NACA RM A9CO7, 1949.

6. Jones, J. Lloyd, and Demele, Fred A.: Aerodynamic Study of a Wing- Fuselage Cofiination Employing a Wing Swept Back 63O.- Chsracter- istics Throughout the Subsonic Speed Range with the Wing Cambered and Twisted for a Uniform Load at a Lift Coefficient of 0.25. NACA RM AgD25, 1949.

7. Jones, Robert T.: Estimated Liftarag Ratios at Supersonic Speed. NACA TN 1350, 1947.

8. Glauert, Ii.: Wind Tunnel Interference on Wings, Bodies and Aircrews. R. & M. No. 1566, British A. R. C., 1933.

9. Herriot, John G.: Blockage Corrections for Threeqimensional-Flow Closed-oat Wind Tunnels with Consideration of the Effect of Compressibility. NACA RM ~7~28, 1947.

10. Swanson, Robert S., and Priddy, E. LaVerne: LiftingSurface~ory Values of the Damping in Roll and of the Parameter Used in Esti- mating Aileron Stick Forces. NACA ARR L5F23, 1945.

ll. Polhamus, Edward C.: A Simple Method of Estimating the Subsonic Lift and Damping inRoll of Sweptback Wings. NACA TN 1862, 1949.

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NACA RM A9127 13

(a) Rear view.

(b) Plan view.

Figure l.- Model of the cambered and twisted wing with the leading edge swept back 63O in coz&ination with a fuselage.

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Eguotion for fuseloge ordinoies: --

f Mate: Ail &ennians givan h &et unfass oMerwk90 qoecmf.

/=6.3X5 L

Ffgure .?.- Dimensions of wing and fuselage.

M&: All dimensions g&n in ket un/ssb

7)pkal sect/on paralW fo plane of symmetry

AU sections have NACA a=/.0 mean- comber lines and 64A005 thickness

Spanwise camber disiribution

Lj 60 .0/f 7

0 80 .Ol/5

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An@ of tif, at, deg

figure 3.- Plan form of right half of wing showing spanwise variation of camber

and twist and location of sections for which coordinates have been calculated.

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Angle of ottac& , CT, &g .16 .I2 .a!3 .04 0 704 708

Pltc#ng-moment coeff/cient , C;, (al 4 V-S a, CL vs Cm.

F&w 4.- T7re efiect of eleven deflection on the aeIo@mImrC characterfstlcs of the wing-f&age comb/i&ion and on the eleven hinge-mount coefflcents at a Mach number d 0.20.

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18 NACA FM A9127

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F@un? d- The effwt of 6Won d&ct.&w on tile avodyncrmic characteristtlcs of the wing-fuselage combb&ion and an the efetm ftbtge-moment coefffc/enls at a Maoh number of 0.60.

20 NACA RM A9127

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Figure 6.- ?71e effect of ekwvn deflect/on on the aetvdynamk characterlsl/cs af the wing-fuselage combfeatfon Ond on the eleven hinge-moment coefffcfents at a bfach number of 0.80.

22 ITACA RM A9127

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Figum Z- The effect of elewn deftecth on the M/c choructeristics of the whg-fuse/a@ combinaffon and on the eiewn hinge-moment coeffM&nts at a At& number of 0.89. II!2

24 - NACA RM ~9127

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26 NACA RM A9I27

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I I I\I I I I

-8 -4 0 4 8 12

Angie of Muck, u o&g

&I CL vs Go, G/J vs a.

Rgure 8. - Gcmct’uded.

.

.-

-

.-

.

.

NACA RM A9127 27

.6

./6

-28 -24 -20 -/6 -12 -8 -4 0 4

E/even deflect/on, 8 deg , T

- /a) M, 0.20. f@UT8 9. - 728 vadufion of’ //icI, pfchhg-momen and hhge-

mOt778t7f Coef%i8nfS Wifh 8/eVOn def/CfiOn for VffhOUS Ung/eS Of OffUCk Uf S8V8fd hi&h numb8a

,

28 NACA RM A9127

0 ./6

-.4 I I1

: 2 0 ./2

A 4 3

-.08

I I I I I I I I I I I I I I I -28 -24 -20 -/6 -/2 -8 -4 0 4

E/even deflscfion, S , .deg Pj@i7 --*I

I

(b) M, 0.60. figure S.- Gonfthued. .

. -

N A C A R M A 9 1 2 7 2 9

.

.4

Q --y .2 * ra

0

,.2

./6 ,.4

./2

.0 8

e .O B

p ’ -9 2 II- .-? % -I I I I I I I I l - -O 8

B e k -.0 4

-8 -.0 8 I I I I ! l d l I I I I I I I I I 1 7 1 7 I I

I I I I I I I I I I II I II I I. -2 8 -2 4 -2 0 -/6 -1 2 -8 -4 0 4

& W O n d 8 f/8 C ti M , 8 , d 8 g v

(c ) M , 0 .8 0 .

fi g u re 9 . - C o n fh u e o !

NACA RM A9127

.6

-l<! !-5t i I I I

P = -.08

tO8

-28 -24 -20 -/6 -/2 -8 -4 0 4

Ef8VC.W def/eCt/on, 8 , Cf8g v

Figure 9. - Cont/nued.

(4’/ M’, 0.89.

CONFIDENTIAL

.

NACA RM A9127 .B 31

I I I I I

I I hI

I tYJa d--i

x? 8 .04

8

6 O

c , -.04

.& ’ -.08

I I I l\hI I I

-28 -24 -20 -16 -/2 -8 -4 0 4

Ef8VOn def/eCf/On, 8 , d8g

(8) hf, 0.93. F&f8 9. - Concf’uded.

./6

32 ?XACA RM A9127

-008 .> .

---- - - -- -/50/*---Y-- +- \ -- -/0*-t---------'--- .---a_

I I I I I I

./ .2 .3 .4 5 .6 .I .8 .9 l.0 MQch mm&w, M

--- : 4

Figure IO.- The effect of Moth number on C&,..* and on the lift coefficient for /ongtWd~no/ bahnce.

EIACA RM A9127 33

24

/6

8

4

0 0 .2 .4 .6 .8 I.0 fur &=O”

Lift coefficienf , CL

ngure I/.- The varht~on of lff-drag rat/o With Yff CO8ff/c/enf for various 8f8VOn def/ecf/ons at severa/ Mach numbers.

34

i6

8

NACA RM A9127

.i,X

.- --- .

-.-

fd M, 0.80.

8 --

4

.2 .4 .6 .8 /.O

Lift coefflc/ent , 15,

fel M, 0.93.

for a# * o”

.

i

ffgur 8 if. - Conclutf6w!

XACA RM A9127

0 .08

35

.04 0 - IO /o

0

-8 -4 0 4. 8 /2 16

Angle of otfack, + ~ deg

1.) iw, 0.20. Hgure I2.- The vuriai~‘on of robing-moment and hinge-momenf

coefficienfs with ung/e of otttxk for vadous e/won cief/ections of severuf Much numbers.

36 NACA RM A9127

0 .a9

I &p,* 8Lur

deg deg -

-8 -4 0 4 8 I2 16

Angle of oftack, au , deg

Figutv fP.- Continued. fbl M, 0.60.

.

.-

.

37

.

/ . c

.04

.03

.02

.O/

0

.08

-.08

-./2

-./6

-8 -4 0 4 8 /Y

Ang/e of oftck, q, , deg

(cl M, 0.80.

figure /2.- Confhued.

.

I

38 NACA RM A9127

‘“n-l-l-r t I I I

.02

.O/

Ottttttt

1-72 \f,-\I I

t/6

-4 0 4 8 Angfe of affack,eu,d8g

.

.

. 39 NACA RM Agi27

.05

B F” .02

0 I

8 .o/

$

0 .08 111

-.08

-./2 I I I I I 1 I

-8 -4 0 4 8 /2 Angle of af tuck, U, , deg

Fgure 12.- Conchded. fe/ fW, 0.93.

7

NACA RM A9127

.o/

0

g .04 k P, e

.03 2 !$ .02 9 8 .oi

$ 0

.05

.04

-03

.02

. .o/

n

I i i i i i

0 30 50 60 10. .eO 49 .- -q&g7

-0 10 20 30 40 50 60 Tofd e/even deflection, Sr , deg

Figure /3- The vt?rhtf/on of ro/l/g-momenf coeff/c/Bnf w/th toto/ e/even det%&?n at vatfous angles of otfuck for severvl Mach numbers.

QU, deg 0 0 0 6 Q 8 A /o

NACA RM A9127

-05 -05

.04 .04

.03 .03

.02 .02

.O/ .O/

0 0 .f .f .2 .2 .3 .3 .4 .4 .5 .5 .6 .6 .7 .7 .8 .8 .9 .9

Mu& number, M

Figure /4.- The effect of Mach number on the ro//fng-moment effectiveness of the e/evens.

42 RACA RM A9127 *

10 PO ,a30 40 50 60

Totd efewm def/ecHon, &, deg

.

Fifgwe fS.- The wuffofh of fhe wing-ffp /refix ong/e wffh fofof efevon def/ecft/on for severof A/o& numbers of 0 ffff coefJ/c/bnt af 0.20.

NACA FM A9127 43

-800

.5 .6 .7 .8 .9 10

M&h number, M v

Figure /6.- The VOrhfiOn with Mach number of Mnge moment, of eleven deflect/on, and of //ff coeff/c/ent for severui normal acce/eration fuctors of a taMess arp/one at 25,000 feet a//l/ude. Whg /oao!?ng, 40 pounds per square foot; wing area, 714.3 square feef; cenfer of grovffy af 0.255.

44 m==mh lUCA RM A9127

u, I I I I I I.1 I I I I I I

-24

-8

16

8

0

I I I I I f I I I I I I

, f60 200 240 280 320 360 400

Gl/dhg speed, Ve , mph

Figure 11. - The vutiiat~on with gl/ding speed of &king speed, hh?ge moment, Devon d8f/eCf/On, and cmg/e of oftuck for 0 fuh’/ess ofi@/cme ut S8Q /eve/. Wfng /ood/ng, 40 pounds per squure foof; cet?ter of grov/y ot 0.25C

.

.

-. I Ill -..

-. . .’

.

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