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RESEARCH MEMORANDUMCHANGED

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ANALYSIS OF THE VERTICAL-TAIL LOADS MEASURED DURING

A FLIGHT INVESTIGATION AT TRANSONIC SPEEDS

OF THE DOUGLAS X-3 RESEARCH AIRPLANE

By WilliamL.Marcy,HarrietJ.Stephenson,andThomas V.Cooney

High-SpeedFlightStationEdwards,Calif.

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CMSSIFIELIDCCUMENT

Thismaterfnlcontains informationaffecthg* NationalDeierseof* United States witldn b meaningof the espionage laws, Title 18,U.S.C.,Swcs.793& 794,thelrvmnlssionorrevelationofwhichinw-r toanunautimrlzedpersonisprohibitedby18w.

NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTNovember

.?7,1956

\

NACA RM H56H08

NATIONALADVISORYCOMMITTEEFORAER~_!

.

ANALYSIS OF THE

RESEARCH MEMORANDUM

VERTICAL-TAILLOADSMEASUREDDURING

A FLIGET INVESTIGATIONATTRANSONICSI’IEDS

OFTHEDOUGLASX-3RESEARCEKIRPIANH

ByWilliam L. Marcy, Harriet J. Stephenson,and Thomas V. Cooney

suMMARY

Results are presented of an snalys is of the strain-gage measure--—ments of vertical-tail loads experienced in rudder pulses, graduallyincreasing sideslips, and rudder-fixed aileron rolls at transonicspeeds with the Douglas X-3research airplane. A Mach number rangefrom approximately 0.7to 1.2 at an altitude of about 30,000 feetwas covered during this investigation.

The lift-curve slope of the vertical tail increased with increasingMach number from a value of 0.038per degree at a Mach number of 0.7toa msximum of 0.048per degree at a Mach nuniberof 0.94,followed by areduction to 0.041 at supersonic Mach numbers. A comparison with avail-able methods of estimating this psrameter indicated good agreement withflight results. The effectiveness of the rudder (lift-curve slope ofthe vertical tail due to deflecting the rudder) decreased from approxi-mately 0.020 per degree at subsonic Mach numbers to 0.013 per degree atsupersonic Mach numbers.

In several violent roll maneuvers, sideslip angles of as much as21° were reached without stalling of the vertical tail, althoughvertical-tail effectivenes~ was reduced at sideslip angles above about12°.

At sideslip angles below about 6°the center of pressure of the loadon the vertical tail in sideslip was practically unchanged with Machnumber, remaining at about 55 percent span and 30 percent chord. Atsideslip angles above this value a more rearward center of pressure wasindicated. The center of load due to displacing the rudder was fartherinbosrd, 45percent span, and farther rearward, 63percent chord, thanin sideslip, snd moved rearwsrd to 85 percentnumbers.

coNFmIAL

chord at supersonic Mach

@gq

—-—-—_

2 CONFIDENTIAL NACA RM H56H08

The variation of airplane yawing-moment coefficient with sideslip,as determined from vertical-tail-loads measurements, increased from0.0023 per degree at a Mach number of 0.7 to a maximum of 0.0032 at aMach number of 0.94, and decreased to 0.0023 at supersonic Mach numbers. .

INTRODUCTION..

The configuration of the Douglas X-3several features of interest to designers

research airplane embodiesat the present timej such as

a long slender fuselage extending well forwsrd of the airplane centerof gravity, low-aspect-ratio unswept wings} and twin jet enginesexhausting nesr the fuselage. Because of general interest in thistype configuration and particular interest in the vertical-tall loadsassociated with maneuvering such aircraf%, a flight investigation wasundertaken at the NACA High-Speed Flight Station at Edwards, Calif.,to obtain information on the loads encountered by the vertical tail inthe transonic speed range.

Flight measurements were obtained in rudder-pulse and steady side-slip maneuvers and in aileron rolls in the Mach number range from aboutM= O.7toM= 1.2 at an altitude of about 30,000 feet. An analysisof the measured data was made in order to isolate such parameters asvertical-tail lift-curve slope, rudder effectiveness, and airplaneyawing-moment coefficient due to sideslip. The results of the analysisare presented and comparisons with existing data are made where possible.

This investigation is a continuation of the Air Force-Navy-NACAjoint effort, using special research aircraft as flight test vehicles,to investigate loads, performance, and stability at trsmsonic and super-sonic speeds. Results of several phases of the overall investigationwith the X-3 airplane have been reported in references 1 to 6.

SYMBOIS

All derivatives, inertias, and airplanebody system of axes.

b wing span, ft

bv vertical-tail-panel

Cbv vertical-tail-panel

SpS,Il(outboard

bending-moment

CONFIDENTIAL

motions me referred to the

of strain-gage station), ft

Mvcoefficient ~svbv

NACA RM H56H08 CONFIIIENTIAL 3

.

dCyv

CL

Cnp

c%

Cn%

CTV

dCTv

q

CYV

e

g

1X

%

lZ

1=

&

lV

M

Mv

spanwise center of load, percent bv

lift-curve slope, per deg

yawing-moment

yawing-momentpanel

yawing-moment

coefficient due to sideslip of complete airplane

coefficient due to sideslip of vertical-tail

coefficient due to sideslip of airplane lessvertical tail

vertical-tail-panel

chordwise center of

vertical-tail-panel

Tvtorque coefficient, —

qS.@

load, percent C

kside-force coefficient, ~q+

mean aerodynamic chord of vertical-tail panel, ft

acceleration due to gravity, ft/sec2

moment of inertia about X-axis, slug-ft2

moment of inertia about Y-sxis, slug-ft2

moment of inertia about Z-axis, slug-ft2

product of inertia, slug-ft2

aerodynamic load on vertical-tail panel, lb

distance between the airplsne center of gravity ud thequarter chord of the vertical-tail panel 5, f%

free-stream Mach nwiber

vertical-tail panel aerodynamic bending moment about strain-gage station, ft-lb

CONFIDENT

NACA FM H56H08coInumNTm

rolling angular velocity, radians/see

free-stream dynamic pressure, lb/sq ft

yawing angular velocity, radians/see

yawing angular acceleration, radians/sec2

wing area, sq ft

vertical-tail-panel srea, sq ft

vertical-tail-panel aerodymmic torque about the quarter chordof the panel 5, ft-lb

time, sec

airplane angle of attack, deg

angle of sideslip, deg

total aileron deflection (sum of right and left), deg

rudder deflection, deg

DESCRIPTION OF AIRPLANE

The Douglas X-3 is a single-place research airplane with 4.5-percent-thick modified hexagonal section low-aspect-ratio wing, and a longpointed fuselage. The wing is unswept at the 75-percent chord. Thevertical-tail surface has a 4.5-percent-thick modified hexagonal sectionwith quarter-chord line swept back 40°. The horizontal tail is locateda distance of approximately half the depth of the fuselage below thevertical tail. All controls are actuated by an irreversible hydraulicpower system incorporating artificial feel devices. The airplane ispowered by two jet engines with afterburners that exhaust near the fuse-lage. Figure 1 Is a photograph of the test airplane and a three-viewdrawing is shown in figure 2. Table I presents pertinent physical char-acteristics of the airplane.

CONFIDENTIAL

INSTRUMENTATIONAND ACCURACY

NACA RM H56H08 coIwllmTm

Standard NACA recording instruments were used to measure the fol-lowing quantities pertinent to this investigation:

Airspeed and altitudeAngles of attack and sideslipNormal, lateral, and longitudinal accelerationsat the airplane center of gravity

Pitching, yawing, and rolling angular velocitiesat the airplane center of gravity

Control-surfacepositions

!5

The error in Mach number based on a flight calibration of the air-speed system in the test airplane is estimated to be within ti.01. Angleof attack and angle of sideslip, measured by vanes located on the noseboom, are considered for the test conditions encountered to be accurateto within *0.3°. Increments in angle of sideslip, however, as used inthe subsequent analysis of the measured data are considered to be accurateto @lo. Lateral acceleration measurements pertinent to the presentinvestigation are estimated to be accurate to i9.Olg; ~gular velocitiesin yaw, *0.01 radian/see; and in roll, +0.03 radian/sec. Control-surfacedeflections were measured at the control surface and are estimated tobe accurate to +9.1°.

In addition to the standard instruments, strain gages were installednear the root of the vertical tail (see fig. 2) and were calibrated toindicate shear, bending moment, and torque. Based on the results of acalibration of the strain gages, and considering the accuracy with whichthe flight records could be read and the accuracy of inertia correctionswhich were applied to the structural loads, the estimated accuracy ofthe strain-gage results is *1OO pounds aerodynamic shear, and +2,000 inch-pounds aerodynamic bending-moment and torque. When expressed in termsof load coefficients, these shesr and moment accuracies result in anaccuracy of approximately W.02 for ~v and *0.01 for ~v and ~v

based on the dynamic pressure value corresponding to the lowest testMach number.

TESTS

Maneuvers pertinent to this investigation of the X-3 airplane con-sisted of rudder pulses, gradually increasing sideslips, and rudder-fixed aileron rolls. The maneuvers were performed at an altitude of

coNFIDmIAL

6 cowmmuu NACARM H56H08

about 30,000 feet and covered a Mach number range from about 0.70 to 1.2.Reynolds nuribersfor these tests, based on vertical-tail-panel mean

aerodynamic chord, varied from 8.4 x 106 to 13.2 x 106.

The control inputs and the subsequent airplane motions and measured.

loads for each type maneuver performed are illustrated by the time his-tories shown in figures 3 to 5.

RESULTS AND DISCUSSION

Vertical-Tail Loads

The rudder-pulse maneuvers performed during this investigation, asshown in figure 3 for the maneuver made at M = 0.93, consisted of anabrupt displacement of the rudder to peak deflection before appreciablesideslip had developed, followed by an abrupt return of the rudder toneutral or near neutral. Following this t~e of control input, alateral-directional oscillation resulted with rudder held constant.Variations of vertical-tail side-force coefficient, bending-moment coef-ficient, and torque coefficient associated with rudder deflections wereobtained from the initial portion of the maneuver where sideslip was con-stant. These variations are shown in figure 6 for Mach numbers of 0.93and 1.14. ‘l’hevariation of vertical-tail side-force coefficient with

‘WV was determrudder position~

“ned in this manner for each test Mach

number. During the oscillations which followed the rudder pulses, vari-ations with sideslip angle of vertical-tail side-force, bending-moment,and torque coefficients were determined and are shown in figure 7 for themaneuvers performed at M = 0.93and 1.14. The variation of vertical-

dCyvtail side-force coefficient with sideslip angle — during ~he oscil-

d~

lations was considered to be the vertical-tail lift-curve slope.

Another type maneuver, the rudder-fixed aileron roll, performedduring this investigation also resulted in sideslip response of the air-plane. Variations of vertical-tail side-force, bending-moment, andtorque coefficients with sideslip angle during an aileron roll atM . 0.90 are shown in figure 8. In this figure the variation ofvertical-tail load with sideslip angle, which is typical of the dataobtained in the roll maneuvers, is essentially linear) indicat~~ thatthe effects on the vertical-tail loads of the rolling and yawing motionsencountered during these maneuvers are of a secondary nature. There-fore, data from several roll maneuvers throughout the test Mach numberrange were also used to determine the variation of vertical-tall side-force coefficient with sideslip angle.

NACARMH56H08

The vmiation with Mach

CONFIDENTIAL 7

d% dCyvnumber of ~and— obtained from

dbr d~

both the rudder-pulse and aileron roll maneuvers is shown in figure 9.A reduction in effectiveness of the rudder to produce side force on the

d%vertical tail from a value of --# = 0.020 per degree at subsonic Mach

rnumbers to 0.013 at Mach numbers above 1.0 is shown in figure 9. The lif*-curve slope of the vertical tail increases with Mach number from a valueof 0.038 at low Mach numbers to a peak of 0.048 at M = 0.94, followed bya reduction to 0.041 at supersonic Mach numbers.

The dashed curve in figure 9 (labeled “estimated”) is the vertical-tail lift-curve slope obtained by using the charts of reference 7.Before using the charts, however, a value of effective aspect ratio wasdetermined. The vertical-tail panel (portion of the vertical tail out-board of the strain gages) has a geometric aspect ratio of 1.18. Thisvalue of geometric aspect ratio was increased by 55 percent (ref. 8) toaccount for the end-plate e~fect of the horizontal tail and fuselage,res~ting in an effective aspect ratio of 1.83. The curves of refer-ence 7 for an aspect ratio of 1.83 and sweep of the quarter chord of 400yield a value of C% = 0.038 for the lift-curve slope at M . 0. The

modified Prandtl-Glauert factor was employed to account for compress-ibility effects and resulted in the increase in lift-curve slope withMach number shown by the dashed line in figure 9. Good agreement isindicated between the flight measurements and the estimated vertical-tail lift-curve slope.

The data points presented in figure 9 were determined from measure-ments made at sideslip angles less than about 6°. lh several roll maneu-vers made at two Mach numbers, 0.92 and 1.05, however, large sideslipangles were encountered. A time history of the vertical-tail side-forcecoefficients and other pertinent data is reproduced in figure 10 for amaneuver where a sideslip angle of 21° was reached at M = 1.05 at about25,000 feet. The dashed line in figure 10 (labeled “calculated”) was

d% d%obtained by using the value of ~ of 0.042 and J of 0.013 from

d~ &r

figure 9 corresponding to M = 1.05 and multiplying these coefficientsby the measured sideslip angles and rudder positions. The calculationsdeviate from the experimental data when a sideslip angle of approxi-mately 12° is exceeded, indicating that the vertical tail suffers a lossin effectiveness above this angle of sideslip. The discrepancy betweenmeasured vertical-tail side-force coefficients and calculations basedon the coefficients derived from small sideslip angle data was found tooccur at about 12° wherever this sideslip angle was exceeded.

CONFIDENTIAL

8 Comn)mm NACARM H56H08

In an attempt to account for this indicated reduction in effective-ness, wing-span loadings from unpublished pressure measurements obtained .

simultaneously with the vertical-tail strain-gage data were examined forindications of abrupt changes in wing-span loading near the inboard por-tion of the wings. Inspection of the wing-span loadings, however, did

-.

not show evidence of any changes which were consistent with changes insideslip angle and which might be correlated with the loss in effective-ness of the vertical tail as indicated by the reduction of side-forcecoefficient with sideslip. Therefore, the effectiveness is apparentlymodified by flow conditions at the tail, which sre the result of inter-ference effects (rear fuselage and engine wake, for example) ratherthan effects which emanate from the wings.

Center of FYessure

Variations of the bending-moment and torque coefficients with side-force coefficient during the initial portion of a rudder-pulse maneuverand during the sideslip oscillations following the pulse are shown infigure 11 for a maneuver at M = 0.93. Similsxvariations are shownin figure 12 for a roll maneuver at M . 0.90. Slopes of the linesfaired through these data yield the center of pressure of the load out-

‘c% ~dboard of the strain-gage measurement station —‘%v

chord of the vertical-tail-panelmean aerodynamic chord

d% d%~ obtained from the pulse portion of

‘f # ‘d d%v v

where sideslip was constant gave the center of pressurethe rudder. The center of pressure due to sideslip wasthe oscillations in sideslip with rudder held steady in

about the quarter

dC~“Lv Values

q“v

the maneuver

due to displacingdetermined fromboth the pulse—

and roll maneuvers. Centers of pressure were obtained in a similar man-ner from data at each test Mach number and the variations with Mach num-ber of the spanwise and chordwise centers of pressure are presented infigure 13. The circular symbols of figure 13 are the data from theoscillations in sideslip following the rudder pulses and in the aileronroll maneuvers; the square synibolsare the data from the rudder-deflectedportion of the pulse maneuver with sideslip constant. In sideslip withrudder constant, little change with Mach number is noted for the span-wise center of load which is about 55 percent of the panel span. Thechordwise center of pressure remained essentially constant with Machnumber at about 30 percent panel mean aerodynamic chord except at thehighest test Mach number. These data were obtained in maneuvers wheremaximum sideslip angles reached were about 6°. During roll maneuversmade at M . 0.92 and M = 1.05, sideslip angles in excess of 6° were

CONFIDENTIAL

.

NACARM H56H08 cowmmnm 9

encountered. The variations of%

and CTV ~th CyV for these

maneuvers are presented in figure 14. It can be seen from this figurethat extension to the higher sideslip angles shows that the variation of~v with Cyv became nonlinear in such a way that a more rearward chord-

wise center of pressure is indicated.

The center of load associated with deflecting the rudder at constantsideslip is farther inboard, 45 percent of the panel span, than the span-wise center of load due to sideslip. Chordwise, the rudder-deflectedload center is 63 percent of the chord from the leading edge and movesrearwsrd to about 85 percent of the chord at supersonic Mach nunibers.

Variation of Yawing-Moment Coefficient With Sideslip

In these maneuvers, the load on the vertical-tail surface acting adistance lV rearward of the center of gravity of the airplane, may be

approximated by a component associated with yawing acceleration and acomponent which balances the yawing moment of the airplane less verticaltail. This relationship can be expressed as

During a steady sideslip maneuver with zero yawing EUIgularacceleration,the vertical-tail load is a measure of the wing-fuselage yawing-momentcoefficient in sideslip. In the sideslip maneuvers performed duringthis investigation, steady sideslip maneuvers were approximated sincethe sideslip angle was increased gradually and the resulting yawingaccelerations were small. Representative data are shown in figure 15for sideslips made at Mach numbers of approximately 0.75 and 0.92 wherethe rudder required to balance the airplane in sideslipping flight andthe resulting vertical-tail side-force, bending-moment, and torque coeffi-cients are shown. Variations of vertical-tail side-force coefficientwith sideslip angle as shown in figure 15 were obtained from all sideslip

d(+maneuvers, and variations of ~ with Mach number me presented in

d~

fimu-e 16. The variation of the yawing-momentfu=elage with sideslip was then evaluated

CONFIDENTIAL

from

lV

%-

coefficient ofthe expression

the wing-

10 CONFIDENT NACARM H56H08

and the variation with Mach nuxxiberof C‘Pwf

is shown in figure 17. It

cam be seen that the wing-fuselage becomes slightly more unstable asM = 1.0 is approached, with a reversal of this trend at supersonic speeds. . .The sideslip maneuvers from which the data of figure 17 were obtained wereperformed with variable aileron deflections, and Cn

P~presented in the

figure contains the yawtig-moment contribution from the ailerons. The data,however, have not been modified”to take into account the effect of theailerons, since available data for the test airplane configuration indi-cate that yawing moments contributed by the ailerons are small and canbe neglected for the flight conditions of the present investigation.Also shown in figure 17 is the variation of the yawing-moment coefficientwith sideslip contributed by the vertical tail as determined by substi-

d~tuting values of ~, obtained from the pulse oscillations and roll

d~

maneuvers and given in figure I-2,in the expression

This parameter increases with increasing ~ch number to M . 0.95, andexperiences a reduction thereafter similar to the data of figure 9.

The sum of C‘Pv

plane. This value is

c%vary from 0.0023

and Cn results in Cl%f ‘P

of the complete air-

shown in figure 17by the solid curve. Values ofper degree at M= 0.7 to 0.0032 at M= 0.94,

then drop off again to a value of 0.0023.

The rudder,displacements required to balance the airplane in side-slipping flight have been determined for each test Mach number from thevariations of rudder position with sideslip angle as indicated in fig-Ue 15 for M= 0.75 ando.92. The results for these and the other testMach numbers are summarized in figure 16 in which the variation with

Mach nurriberof the sideslip angle associated with rudder position ~dbr

iS shown. A progressive loss in rudder effectiveness as Mach nmber isincreased is indicated in figure 16.

coNFImNTIAL

.

CONCLUSIONS

NACA RM H56H08 coImDmIAL 11

Analysis of the measurements of vertical-tail loads made during aflight investigation of the Douglas X-3 research airplane has indicated:

1. The effective vertical-tail lift-curve slope determined fromsideslip oscillations following rudder-pulse maneuvers and in rudder-fixed rolls increased with increasing Mach number from a value of 0.038at low Mach numbers to a peak of 0.048 at a Mach number of 0.94, followedby a reduction to 0.041 at supersonic Mach nunibers. Estimates of thisparameter based on published span loading data and estimates ofhorizontal-tail and fuselage end-plate effects indicate good agreementwith the flight results.

2. In several violent roll maneuvers, large sideslip angles werereached (210) without stalling of the vertical tail, although the lifteffectiveness was reduced above about 12° sideslip.

d%3.Rudder effectiveness parameter .& as determined from rudder-

rpulse data, decreased from approximately 0.020 per degree at subsonicMach numibersto 0.013 per degree at supersonic Mach numbers.

4. The center of pressure of load on the vertical tail in sideslipwas practically unchanged with Mach number, remaining at about 55 percentspan and 30 percent chord at sideslip angles below about 6°. At side-slip angles above this value the variation of the vertical-tail-paneltorque coefficient ~v with the vertical-tail-panel side-force coeffi-

cient -Cyv became nonlinear in such a way that a more rearward chord-

wise center of pressure was indicated. The center of pressure due todisplacing the rudder was farther inboard, 45 percent span, and fartherrearward, 65 percent chord, than in sideslip and moved rearward to 85 per-cent chord at supersonic Mach numbers.

5.The variation of airplane yawing-moment coefficient with side-slip cn~ as determined from vertical-tail-loads measurements, increased

L? coNFIDmIAL NACA RM H56H08

from 0.0023 per degree at a Mach number of 0.70 to a -- of 0.0032at a Mach nuniberof O.~ and decreased to 0.0023 at supersonic Machnumbers.

High-Speed Flight Station,National Advisory Committee for Aeronautics,

Edwards, Calif., July 30, 1956.

NACA RM H56H08

REFERENCES

13

1. Bellman, Donald R., and Mu@Y, Edwud D.: Lift and Drag @racter-istics of the Douglas X-3 Research Airplane Obtained During Demon-stration Flights to a Mach Number of 1.20. WCA RM H54U7, 1954.

2. Day, Richard E., and Fischel~ Jack: stabilitYandcontrolcwacter-istics Obtained During Demonstration of the Douglas X-3 ResearchAirplane. NACARMH55E16, 1955.

3. NACA High-Speed Flight Station: Flight Experience With Two High-Speed Airplanes @ving Violent Lateral.-LongitudlnalCoupling inAileron Rolls. NACARMH5~3, 1955.

4. Stephenson, Harriet J.: Flight Measurements of Horizontal-Tail Loadson the Douglas X-3 Resesrch Airplane. NACARMH56A23, 1956.

~. Jordan, Gareth H., and Hutchins, C. Ke~eth, Jr.: preli~nary F~@t-Determined Pressure Distributions Over the Wir& of the Douglas X-3Research Airplane at Subsonic and Transonic Mach Numbers. NACARMH55A10, 1955o

6. Keener, EarlR., and Jordan, Gareth H.: Wing Loads and Load Distri-butions Throughout the Lift Range of the Douglas X-3 Research Air-plane at Transonic Speeds. NACARMH56G13, 1956.

7.Deyoung,John, and Harper, Charles W.: Theoretical Symmetric SPan

Loading at Subsonic Speeds for Wings Having Arbitrary Plan Form.NACA Rep. 921, 1948.

8.Pass, H. R.: Analysis of Wind-Tunnel Data on Directional Stabilityand Control. NACA TN 775, 1940.

coxwmmm

14 CONFIDENT NACA RM H56H08

WI

PsYsIcALmARACTmmrcs oFTsEDxJCtAS x-3 AnPIAsE

wing:Airf0118ectim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .AirfOilthickne13sr8t10,percmt chord . . . . . . . . . . . . . . . . . . . . . . . . . . .Total.srea,nq ft... . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Span,ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Maneerodmumic chord,ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Taperratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Aspectl’atio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .m0f75~mttiu=, deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . .bcidence,deg . , . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Dihedral,deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .GSCiO#XiC_Bt,d W . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

!kdifled hexagon. . . .. . . . 1.66:$. . . . 22.69. . . . 7.04. . . . 0.39. . . . 3.09. . . . 0. . . . 0. . . . 0. . . . 0

Hm-izmt.alt.dl.:Airfoilsection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . McdifledhexssonMrfoilthickness mtioatmcrt chord,percentchord . . . . . . . . . . . . . . . . . . . . . . . . 8.01Airfoilthictiasratio outbosrdofstation26, perceotcbor d.... . . . . . . . . . . . . . . . .Tots.larea,sql%, . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4;:%m=,ft . . . . . - . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1; .7JMemmmdymlicctl ord,ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Taperratio. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.io5Aspectratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.38sweepatlee.dingedge,deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21.14-wat-~w-,~s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0Dihe&al, deg. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 0Tmvel,l.eadingedgeUP,deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6Travel,leadingedge dmm,deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .~We-~nel@ati~, =centrmtch~. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46.i!

Verticaltail:Airfoilsect.ion.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . l.biifiedhexaganAirfoilthicknessratio,percentchord . . . . : . . . . . . . . . . . . . . . . . . . . . . . . . .Area, sqft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .sw,ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Meanaeradynamiccherd, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Rootchord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Tipchord, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Tapermtio. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Aspectratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .sveepof lead.l~edge,deg . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Sweepoftrailingedge, deg. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Rudder:Area, rearwardofhinge lim,sqft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .smat~=e~,fi . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Rootchmd, f% . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Tipchord, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Tmvel,deg. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Vertical-tailpanel (outboardof strain.~ referencestation):A1’S3,eqft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Sw,ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Meanaercdylunl cchard,ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . ... . . . . ..Rcmtchord, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .!flpchmd,ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Taperratio. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Aapectratio . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Fuselage:Lengthincludingbom,ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Maximmwidth, ft. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .lkximmhei@t, ft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Basearea, sqft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . , . .

Airpl.meweight,lb:Safiic(withoutfuel,oil,wat.er,pilot) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Total(fullfuel, oil,wster,nopilot) . . . . . . . . . . . . . . . . . . . . . ., . . . . . . . .

Center.of-grmltyIceation,percentman aerodym.mtcchord.Es8icveight,geardmm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Totalveight,@ar down . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Totalweight,gear up..... . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Momentsof inertia,slug-ft2(estimted fcm tot,dweight):

Ix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Iy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Iz . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

%2. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . .

colmcDENTm

4.523.73

W6.738

1.930.2921.315

9.;;

3.443.5351.98logMD

17.05b.b8kAl5.681.93

0.3411.18

66.756.084.817.*

16,12021,900

2.634.593.91

Jl,mti,m65,m

4,200

NACA RM H56H08 CONFIDENTIAL 15

.

.

g

Ejal

.sJ

!ijInalLi

1

l-i

CONFIDENTIAL

NACA RM H56H08

I

I

I

(

L:’El

-.

74

NACARM H56H08 CONFIDENTIAL

2 .4

r, radians/see ,1

0p,radians/see

L Ko

d “ L/

L -4

L

Cyv

o / \1 /~

k

-. 2

tc~

v

o / /

x

-1

Right

~, deg

I

CTV

o

-. I

4

0 / \ /\

Left I O

$, degP

o /

-102 4 6t, sec

8 10 12 14

Figure 3.-Time history of the various quantities measured during a

rudder-pulse msneuver at M = 0.93 at 30,000 feet.

CONFIDENTIAL

17

NACA RM H56H08

CYV

Cbv

CTV

o \

-.2’

0

\

-.04

-.08

Right 10

~, deg

Right I

~r, deg

-1 3 4 5 6 7 [t, sec

Figure 4.- Time history of the various quantities measured during asteady sideslip maneuver at M = 0.93 at 30,000 feet.

.

NACARM H56H08

.

CONFIDENTIAL

1.0

M

.8 -

4 .4

2-- — -

___ — ---- .2p, radkmsAec / \ P- 7

rfl r, radions/seco ~

\ \ti

\ /’. 0

-2 -.2

CYV

Cbv

c Tv

Right

~, cieg

Right

~a, Sr, deg

.4

—-

\

o— ‘/

\

-.4

.2

. ~ - \

o —. -- ‘ \

\.

-.2

1-

0 f —

-1

10

0 r \/

Y / ‘“

-10

2(3

10 1. Sr. o

0 ‘.

-I% I 2 3 4 5tg sec

Figure 5.- Time history of the various quantities measured during arudder-fixed aileron roll maneuver at M = 0.90 at 30,000 feet.

CONFIDENTIAL

19

20 I?ACARM H56H08

.2

.1

Cyv

o

-.1

.1

0

Cbv

-.1

CTV

8 4 08r,deg Right ~r, deg Right

(a) M= O.93. (b) M=l.14.

coefficientsrudder-pulse

Figure 6.-Variation of vertical-tail side-force, bending-moment, andtorque with rudder position (sideslip angle constant)during maneuvers at M = 0.93 @ 1.14 at 30,000 feet.

CONFIDENTIAL

NACA RM H56H08

.1

CYV

_.21__l__l

CONFIDENTIAL 21

.1

cbv

o

-.I

.1

0 -~ -@@@m

CTV

1.1

2 0 2@,deg

Right

2 0 2

~, degRight

(a) M= O.93. (b) M=l.14.

Figure 7.- Variation of vertical-tail side-force, bending-moment, andtorque coefficients with sideslip angle (rudder angle constant)during rudder-pulse nianeuversat M = 0.93 and 1.14 at 30,000 feet.

22

CYV

●

Cbv

CONFIDENTIAL NACA

.4

0 ‘

-.4”

.2‘

o “

-.2

RM H56H08

.1

8 4 0 4 8~,deg

Right

Figure 8.- Variation of vertical-tail side-force, bending-moment, andtorque coefficients with sideslip angle (rudder angle constant)during a roll maneuver at M = 0.90 at 30,000 feet.

CONFIDENTIAL

NACARMH56H08 23

.06

Estimate!7

.04–——~ *——7

“.02

0

dCyII

1 1 {

O Sides lip oscillations

•1 Aileron rolls

I

‘M

Figure 9.- Variation with Mach number of thecoefficient due to sideslip (rudder angletail side-force coefficient due to rudderconstant).

coNmDEmmL

1.0 1,1 1.2

vertical-tail side-forceconstant) and the vertical-position (sideslip angle

NACA RM H56H08

P,

M

radians Aec

CY”

o

-.4

-.8\

o

Cbv -.2 \

-. 4 \

.2

/

CTV O — ~

-.2

“’de’2ELEIEIIRight 20

~r, ~ , deg

o

(-.2% I 2t, sec

3 4:

Figure 10.- Time history of the various quantities measuredaileron roll maneuver at M = l.0~ at 25,000 feet whereslip angles were attained.

during anlarge side-

.

NACA RM H56H08

Cbv

CTV

.1

“o

-.I

.1”

0

–.1-.1 0 .1

Cyv

CONFIDENTIAL 25

-.2 -J o .1 .2Cyv

(a) Constant sideslip angle. (b) Constsat rudder position.

Figure 11.- Vsriations with vertical-tail side-force coefficient of thevertical-tail bending-moment and torque coefficients during a rudder-pulse maneuver at M = 0.93 at 30,000 feet.

CONFIDENTIAL

26 NACA RM H56H08

cbv

.2

.1

c

—. I

-.2

.1

0

CTV

-. I-.2 –.1 o .1

c y“

—

.2 .3 4

Figure 12.- Vsriation with vertical-tail side-force coefficient of thevertical-tail bending-mommt and torque coefficients during anaileron roll maneuver at M = 0.90 at 30,000 feet.

.

NACA RM H56H08 CONFIDENTIAL 27

dCbv

dCyv ‘

percent

0 Constant rudder position

•! “Constant sideslip angle

dCTv

dCyv ‘

percent

00”

la80 ❑

El #

El

60 ❑ — El..Flm n /

40 o–0 0 0 -( mm cm o al)

20 ‘

o.6 .7 .8 .9 I.0 1.1 I.2

M

.—

Figure 13. - Vsriation with Mach number of the center of pressure of thevertical-tail load due to changing sideslip with rudder constant(additional load) and due to displacing the rudder at constantsideslip.

CONFIDENTIAL

CONFIDENTIAL NACA RM H56H08

Cbv

.2

.1 ‘,

o

–.I

-.2

–.3

cq-V

—

1’

0

1

-.6—.I -.4 -.2 0 .2 .4

CYV

(a) M=o.921

Figure 14.- Variation with vertical-tail side-force coefficient of thevertical-tail, bending-moment, and torque coefficients during aileronroll maneuvers.

CONFIDENTIAL

NACA RM H56H08

Cbv

CTV

.1

0

-.1

-.2

–.3

-.4

-.5

F.1

o

I

‘.8 –.6

CONFIDENTIAL

/

3

I

‘Yv

(b) M=l.0~.

Figure 14.- Concluded.

CONFIDENTIAL

o .2

29

CONFIDENTIAL NACA RM H56H08

CYV

.10 M=O.75i7M 0.76=

o

-.1 ‘

\

-.2

I\

oCbv

\--.1-

.1

CT0

v–.1

RightI

Sr, deg

I

o

0

n“lo 10

P,;eg Righi

\

\

10 0 10~, deg Right

Figure 15.- Variation with sideslip angle of the vertical-tail side-force, bending-moment, and torque coefficients and rudder positionin steady sideslip maneuvers at M = 0.75 and 0.92 at 30,000 feet.

CONFIDENTIAL

NACA RM H56H08 31

0

-.02

-.04

‘6 .8 I.0 1.2 1.4

Figure 16.- Variationwith Mach number of the vertical-tail side-forcecoefficient per unit sideslip angle and the sideslip angle for aunit rudder deflection in steady sideslips.

32. ;.; ,,, ~~ , .

I . . . ..— . . . --—.

Q-00

CONFIDENTIAL

d-001. 1“

%-W

bQ

.

Q

u=

NACA RM H56H08

ii

NACA - LangleyField,Va.

.

.