NATIONALADVISORYCOMMI’ITEEFOR AERONAUTICS

TECHNICALNOTE 3859

COMPARISONOF FLIGHT AND WIND-TUNNEL~

OF HIGH-SPEED-AIRPIANESTABILITYAND

CONTROLCHA.RACTERISTICS1

By Walter C. Willisms,HubertM. Drake, and Jack Fischel

SUMMARY

Comparisonsof wind-tunneland flight-measuredvalues of stability~d controlcharacteristicsare of considerableinterestto the designer,sincethe ~nd-tunnel method of testing is one of the prime sourcesuponwhichestimatesof the characteristicsof a new configurationare based.IIIthispaper comparisonsare made of some of the more importantstabilitymd controlcharacteristicsof three swept-wingairplanesas measured inflightand in wind tunnels. Wind-tunneldata from high-speedclosed-throattunnels,a slotted-throattransonictunnel, and a supersonictun-nelare used.

The comparisonsshow that, generallyspeaking,the wind tunnelspredictall trends of characteristicsreasonablywell. There are, how-ever)differencesin exact values of parameters,which couldbe attrib-utedsomewhatto differencesin the model causedby the method of support.Thesmall size of the models msy have some effect on measurementsof flapeffectiveness.When nonlinearitiesin derivativesoccur duringwind-tunneltests, additionaldata shouldbe obtainedin the region of thenonlinearitiesin order to predictmore accuratelythe flight character-istics. Also, nonlinearitiesin static derivativesmust be analyzedonthebasis of dynsmicmotions of the airplane. Aeroelasticcorrectionsmustbe msde to the wind-tunneldata for models of airplaneswhich havethinsurfacesand are to be flown at high dynsmicpressures. lkleteffectscan exert an influenceon the characteristics,dependingupon airrequirementsof the engine snd locationof the inlets.

.%e informationin this reportwas also containedin a paper by

thesame authorsentitled: “Some Correlationsof Flight-MeasuredandWind-TunnelMeasuredStabilityand ControlCharacteristicsof High-SpeedAirplanes.”The latterwas presentedto the Wind Tunnel and Model TestingPanelof the NA!IOAdvisory Group for AeronauticalResearch and Developmentatthe meeting in Brussels,Belgium,August 27-31, I-956.

.-

2 NACA TN 3859

INTRODUCTION

One of the principaltools of the aircraftdesignerin predictingthe stabilityand controlcharacteristicsof a new airplaneis the USe

of models tested in wind tunnels. There is, of course,the questionwhether the model results accuratelypredict characteristicsof the air-plane in free flight or, in other words, the questionof the de~ee ofcorrelationbetween the two results. This problem has receivedconsid-erable attention. Most of this work, reference1 for example,has beenperformedat subsonicspeeds and indicatesthat, in general,good come.lation can be obtainedwhen the model accuratelyrepresentsthe actualaircraft,and the tests,both flight and wind tunnel, are carefullyperformed.

Some work has been reportedon the correlationbetween the wind-tunnel and flight-measuredstabilitycharacteristicsin the transonicspeed regime (ref. 2). Correlationsof transonicand supersonicresultsare currentlyof particularinterestin view of the availabilityof ~~tunnels capableof testingthrough the transonicspeed range. Problemsof correlationsin this speed range sre complicatedby the compromisesimposedon the model by the mounting system;for example,sting supportsrequire that the rear end of the fuselagebe altered. It is also neces-s~ in high-speedtunnelsto utilizemuch smallermodels-than were possi-ble in the low-speedtunnels. The purpose of this paper is to presentsome correlationsof severalof the more importantflight-measuredandwind-tunnel-measuredstabilityand controlcharacteristicsof high-speedairplanes.

b

c2p

c%%

SYM1301S

wing span, ft

damping-in-rollcoefficient,per radian

rolling-momentcoefficientper degree ailerondeflection

pitching-momentcoefficient

staticmargin,percentmean aerodynamicchord

(%JW- (%JF

RACA TN 3859

“q

3

%-jt

%+%

CN

cNa

c~P

-E

Iy

it

tit

M

ma

q

s

v

Xa

a

6

..e

T

pitching-momentcoefficientper degree stabilizerdeflection

dsmping coefficientin pitch

normal-forcecoefficient

normal-force-curveslope,per deg

directionalstabilitypsrsmeter,per radian

wing mean aerodynamicchord, ft

airplanemoment of inertiain pitch, slug-ftz

stabilizerangle,negativewhen stabilizerleadingedgedown, deg

()‘t ~ - (it)F

Mach number

mass rate of air intake,

wing-tiphelix angle per

slugs/see

degree ailerondeflection,radians/deg

dynamicpressure, lb/sq ft

airplanewing area, sq “ft

true airspeed,ft/sec

distancefrom airplanecenter of gravity to air intake ofjet engine,ft

angle of attack,deg

pitchingvelocity,radians/see

pitching acceleration,radians/sec2

relativeelevator-stabilizereffectiveness

4 NACA TN 3859

Subscripts:

1 initial conditionat start of msneuver

F flight

WT wind tunnel

AIRPLANESAND TESTS

Three swept-wingairplanesare consideredin this study. All aresingle engine,fighter-typeairplaneswith a sweep range from 35° to 600.Much of the flight data were obtainedat an altitudeof 40,~0 feet withsome of the supersonicdata extendingto altitudesas high as 60jO00 feet.The overallReynold.snumber variationwas from 8 million to 19.5 million.The flight data were obtainedwith power on, involvingfor the most partbetween 90 percent and 100 percent availablethrust.

The wind-tunneltests for these airplaneswere performedin thefollowingNACA wind tunnels:

Langley8-foot transonictunnelIangley8-foot high-speedtunnelLangleyhigh-speed7- by 10-foottunnelLangley#- by #--footsupersonicpressuretunnel

All models were sting supportedand the forceswere measuredbyinternallymounted strain-gagebalances. The Reynoldsnumber ranges ofthe tests varied from 1.9 million to 3.6 million. The model tests weremade with no power simulationand the inletswere faired, except for air-plane A which employedan open duct. There were differencesbetween themodels and the actual airplanesin most cases. These differencesandthe model scales are as follows:

AirplaneA (l/n-scale model)8-foot transonictunnelHigh-speed7- by 10-foottunnel

(1) The wind-tunnelmodel incorporatedan enlargementat therear end of the fuselageto accamodatethe sting support.

(2) The wind-tunnelmodel exposed-horizontal-tailarea wasmaintained,and an increasedtail span thereforeresulted.

The plan form differencesfor airplaneA are shown in figure‘1.

NACA TN 3859 5

AirplaneB (1/16-scalemodel)8-foot high-speedtunnel (closedthroat)8-foot transonictunnel7- by 10-foothigh-speedtunnel (closedthroat)4- by 4-foot supersonicpressuretunnel

(1)

(2)

The wind-tunnelmodel incorporatedan enlargementat therear end of the fuselageto accommodatethe sting support.

The wind-tunnelmodel incorporatedconstsnt-percentage-chord wing sections,whereas the airplanewing incorporatedsimilsrroot sectionsbut thickertip sectionsthan thewind-tunnelmcdel. In addition,during tests in the 8-foothigh-speed(closedthroat)tunnel and the 4- by A-footsupersonicpressuretunnel,the model was testedwithouta cockpitcanopy.

AirplaneC (l/lA-scalemodel)8-foot transonictunnel4- by k-foot supersonicpressure tunnel

RESULTSAND DISCUSSION

One of the prime considerationsin the measurementof airplanecharacteristicsis the lift-curveslope of the airplane. A comparisonof the variationof normal-forcecoefficientwith angle of attack forairplaneA, as measured in flight and in the 8-foot transonictunnel atMach numbersof 0.76 and 0.9X, is shown in the upper part of figure 2.The data are for trinmed conditions. As can be seen in this figure, thecorrelationis reasonablygood in the linearrange. At angles of attackabove peak lift or above the break in the curve that are indicativeofseparatedflow, there are discrepancies. The lowerpart of this figureshows the variationwith Mach number of the ratio of f13ght-determinedto wind-tunnel-determinednormal-force-coefficientslope for airplanesAand B. These slopeswere taken at a normal-forcecoefficientclose tothe value for level flight. As can be seen, the resultsare within10 percent of each other,with the flight-measuredvaluesbeing generallyhigher. The transonicdata up to M = 1.15 were obtainedfrom the 8-foottransonictunnel,the data at M = 1.2 from the 8-foothigh-speedtunnel,and the higherMach number data were obtainedfrom the 4.-by A-foot super-sonicpressuretunnel.

Determinationof the staticmargin is importantin establishingthenecessg center-of-gravityposition for a configuration. The variationof staticmargin ~ with Mach number is shown in figure 3 for air-

Lplane A, as measured in the 8-foot transonictunnel,and as me&sured in

6

flight from pulsecenter-of-gravity

NACA ~ 3859

disturbances. The data are referencedto the sameposition. This figure shows that similarvariations

of staticmargin with Mach number are exhibitedin the two sets of data.The flight data, however,show a value of staticmargin consistentlyhigherby about 3 percent. It is believedthat differencesbetween themodel and the airplemeat the rear of the fuselageand horizontal-tailcon-figurations(fig. 1) could accountfor these discrepancies. The lowerpartof figure 3 shows the incrementaldifferencein staticmargin

%~between the data from the two test mediums for airplanesA and B atnormal-forcecoefficientsfor level flight. As statedpreviously,thedata for airplaneA exhibita constantdifferenceof about 3 percent.The flightvalues for airplaneB sre about 5 percent higher than the datafrom the closed-throattunnelup to a Mach number of about 0.85. Abovethis Mach number the differencedecreases,and at a Mach number of about0.95 the wind-tunneldata show about 5 percent greaterstaticmarginthan that shown by the flight tests. This variationbetweenMach numbersof 0.85 and 0.95 is believed to be causedby chokingeffects in theclosed-throattunnel. The results from the transonictunnel (slottedthroat) are similarto those from the closed-throattunnelup to a Machnumber of 0.813. Above this Mach number the differencein staticmarginvsries scznewhat,but throughoutthe Mach number range of this test theflight data show higher staticmargins by 1 to 5 percent. The highersupersonicdata for airplaneB show similarincrementsin staticmargin.

In additionto checkingthe levels of longitudinalstability,itis importantwith high-speedconfigurationsto establishthe variationsof stabilitywith angle of attack in order to explorefor the existenceof nonlinearitieswhich may lead to an undesirablecharacteristic,suchas pitch-up. l’ypicalvariationsof pitchingmcxnentwith angle of attackfor airplaneA, as measured in flight and in the 8-foot transonictunnelat Mach numbers of 0.76 and 0.91, are shown in figure 4. The flight datafor the wing-fuselagepitching-mamentcoefficient(tail off) were obtainedfrom measurementsof horizontal-tailloads. It shouldbe noted that thetail loads were measuredby strain gages mounted at the roots of the hori-zontal tail and representonly the panel loadingwithout carry-overto.the fuselage. Thesemeasurementsare in error, the~efore,by the unknownsmount of thecarry-over. The overall airplanepitchingman@ wasobtainedprimarilyfran flightmeasurementof the variationof stabilizeranglewith angle of attack in acceleratedmaneuvers,turns, and pull-upsmade at constantMach number. These variationsof stabilizerangle withangle of attackwere correctedfor pitching accelerationby the expression

IF

()TF

Ait6=rit

NACA TN 3859 7

The correcteddata were convertedto pitching-momentcoefficientby thefollowingsmlified expression,which includesthe effectsof pitchdamping

%=-(% +%)6’@Ji~ - iq) - ~v

h these calculationsthe pitching-momentcoefficientdue to sta-bilizerdeflection ~.lt was assumedconstantover the angle-of-attack

range. The data of figure4 show that the pitching+nomentcurves fromthe two sourcesare generallysimilar. At both Mach numberstinecom-parisonbetween flight and wind-tunnelresultsyielded a differenceinthe angle of attack for trim. At a Mach number of 0.76, however,thenonlinearitiesoccur in the tunnel data at lower angles of attack andthe data do not exhibitthe large dip in the curve that is shown for theflight results. This differencecouldpossiblybe accountedfor by thelack of sufficientwind-tunneltest points to define such a variation,since there is no wind-tunneltest point between an angle of attack of10° and 12°, where such a dip might be expectedto appesr if it existedin the wind-tunnelresults. The data at a Mach number of 0.91 are con-sideredto be reasonablysimilar,both with tail off snd tail on. Itshouldbe pointed out that inspectionof the shape of the-pitching-momentcurvesis not sufficientto determinewhether or not a pitch-upproblemexists. It has been found that pitch-up can .bea problem even with air-planeshaving neutral stabilityor even slightlypositive stabilityinthe nonlinesrregion. The degree of stabilityabove the pitch-up isalso important. In order to evaluatepitch-up,it is necessary-tomakecalculationsof the motions of the airplanein dynsmicmaneuversby usingassumed=bitrary pilot controlinputs (ref. 3). It is believed thatthesewind-tunneldata representthe flight case closelyenough for suchcalculationsto be of value in predictingthe maneuveringcharacteristicsof the airplane.

Another importantlongitudinalcharacteristics the variationwithMach number of the longitudinalcontroldeflectionrequired,forlevelflight. Data of this type are shown in figure 5. The upper portion ofthe figure shows the variationwith Mach number of the stabilizerdeflec-tion for trim for airplaneA as measured in flight and in the 8-foottransonictunnel. As can be seen, the variationsare generallysimilarfor the two tests,with flight-measureddata showinga larger change instabilizerdeflectionrequiredabove a Mach number of 0.90 than shown bythe wind-tunneldata. In the lowerportion of the figurewhere the dif-ferencesbetween flight and wind-tunnelmeasurementare shown for air-planesA and B, it can be seen that the differencebetween flight andwind-tunneltrim values exceeds 1° of stabilizertravel only at a Machnumberof 0..98for airplaneA. Over most of the range there is lessthan 0.50 differencein stabilizerdeflectionrequiredfor trim.

8 NACA TN 3859

Although elevatorcontrolon high-speedairplanesis being rePhcedby all-movableor one-piecehorizontaltails, it appearsthat flap-typerudders and aileronsmsy continueto be used. Some comparisonsof meas-ured values of relativeelevator-stabilizereffectivenessare shown infigure 6. The upper portion of this figure comparesthe variationsofrelativeelevator-stabilizereffectiveness T with Mach number as meas-ured in flight and in the wind tunnel. This figure shows that there is ~appreciabledifferencebetween the flight @ wind-tunneldata) partic-ularly above a Mach number of 0.9 where a much largerdecreasein rela-tive elevator-stabilizereffectivenesswas measured in flight than inthe wind tunnel. Data are shown in the lower psrt of figure 6 on thebasis of the ratio of flight-measuredto wind-tunnel-measuredvalues ofT for airplanesA and B. Although the values of T fr~ the two sourcesare within 10 percent of one anotherbelow a Mach number of 0=8j the dif-ferencesbetween flight and wind-tunnelvalues at transonicspeeds sreas high as *25 percent. Somewhatbetter agreementis shown for thesupersonicdata than for the transonicdata. At a Mach number of 1.6 thedata for airplaneB are in perfect agreement,which msy be fortuitous.The small size of elevatorsused on wind-tunnelmodels such as thesemake the measurementdifficult.

Additionalflap-effectivenessdata are shown in figure 7 in whichsome aileroneffectivenessinformationfor airplaneB is shown. In theupper part of this figure the ratio of flight-measuredto wind-tunnel-

1‘b b is shown as a functionof Mach number. The

‘easured‘alues ‘f mflight-measuredvalues”are generallylower than the wind-tunnelvalues,reachingonly 70 percent of the wind-tunnelvalues at Mach numbers aboveO.go. This differenceis understandablewhen it is consideredthat thewind-tunneldata for rolling-momentcoefficientwere obtainedunder staticconditionsand the aileroneffectivenesswas calculated,on the assumptionof freedomonly in roll, by the followingexpression

I c%p&=-2V

c%

lh addition,it shouldbe noted that the outboardwing sectionsof theairplanewere thickerthan those of the wind-tunnelmodel, as discussedpreviously. Moreover,inasmuchas the damping-in-rollcoefficient

c%was not measured for this model, values of C~ wed in the present

calculationswere based on thosemeasured for almost caparable wingconfigurations. Better correlationwould probablybe obtainedif theeffectivenesswere calculatedby assumingfreedomin roll, yaw, and side-Slip. In some cases it may be necessaryto includefreedom in pitchand angle of attack as well. The usual testingtechniqueis to obtainthe flight data in rudder-fixedaileronrolls where the airplaneexperi-ences motions about all axes. Aeroelasticityis not believedto be an

NACA TN 3879 9

tiportantfactor in the differencebetween results,because the flight-iest resultsdid not show a significanteffect of dynamicpressurewithinthe r~e tested. The lowerpart of figure 7 shows the variationofaileroneffectivenesswith Mach number for the two tests with the datambitrarily normalizedto the value of effectivenessexistingat M = 0.6.The data show close agreementbetween the flight- and wind-tunnel-measuredvariationof aileroneffectivenesswith Mach number. It appears,there-fore,that if the level of aileroneffectivenesscould be determinedaccuratelyfrom model tests at low speeds,it might be expectedthat thewind-tunneltests could accuratelypredict the decreasein effectivenesswith increasingMach number.

Staticdirectionalstabilityof a new configurationis of importanceto the designersince it is one of the more importantparametersused indeterminingairplanebehaviorunder dynamic as well as static lateralconditions. It has been found that many of the high-speedconfigurationsexhibitlarge changesin directionalstabilitywith angle of attack.Typicaldata for airplaneA sxe shown in the upper portion of figure 8,where the staticdirectionalstabilityderivative Cn

Pis plotted as a

functionof angle of attack. These data were obtainedin the 7- by 10-foottunnelat a Mach number of 0.70. There are no comparableflight data forthis case because of the difficultyof measurementin flight. As can beseen in this figure,the directionalstabilityparameterbecomes zero atan ule of attack of about 180. From data such as these, the variationwith Mach number of the angle of attack at which C

mined.

~ is zero was deter-

This boundary is plotted on the lowerpart of this figure. Alsoshown are pointswhich representthe combinationsof angle of attack andMach number at which directionaldivergenceshave occurredin flight. Itshouldbe noted that, for any given Mach number,divergencesoccurredatanglesof attackboth less than and greaterthan that requiredfor zerodirectionalstability. It appe~s that, as in the case of pitch-up,dynamicanalysisof the airplanemotions is requiredin order to assessthe problem.

Another variationof directionalstabilityof concernto designersis that which occurswith changesin Mach number. Figure 9 relates thevariationof Cn

Pwith Mach number as measured in the wind tunnel to

thatmeasured in flight for airplaneC. As can be seen, there are largediscrepanciesamountingto as much as ‘jOpercentdifferencebetween thebasic wind-tunneldata and the flight-measured-values.In the previouscases shown, relativelythick airfoilsectionswere used on the empennageand the dynamicpressurefor the tests was relativelylow, less than400 pounds per square foot. In the present case the vertical-tailthick-ness was about half that of the other airplanes,and the maximum dynamicpressureexperiencedwas of the order of 8X pounds per square foot.Aeroelasticeffectswere found to be of importance. When the wind-tunneldata were correctedfor aeroelasticeffects,primarilybending and

10 NACA TN 3859

twistingof the verticaltail, the agreementbetween the data from thetwo sourceswas improvedbut differencesas high as 20 percent stillremained. Because airplaneC has a large jet engine and a nose inlet,the wind-tunneldata were then correctedfor inlet effectsby theexpression

As can be seen,when this correctionwas made, the wind-tunneltestsgave values of the directionalstabilityparsmeterthat were within10 percent of the flightvalues throughoutthe Mach number range.

CONCLUDINGREMARKS

Comparisonsof wind-tunneland flight-measuredstabilityand controlcharacteristicsshowedthat the wind-tunneldata predictedall trends ofcharacteristicsreasonablywell. Discrepancieswere found in exact values,which may be attributedto differencesin the models causedby mountingconsiderationsand, in the case of controleffectivenesses,to the smallsize of the models. Where nonlinesritiesin derivativesoccur duringwind-tunneltesting,it may be necessaryto obtain additionaldata pointsin the region of the nonlinesritiesin order to predict”moreaccuratelythe flight characteristics.Nonlinesritiesin staticderivativesshouldbe analyzedunder dynamic conditions. Aeroelasticitymust be consideredin evaluatingdata dealingwith thin airfoilsand high dynamicpressures.Inlet effectscan be important,dependingon the size of the engine andthe locationof the inlets.

High-SpeedFlight Station,NationalAdvisoryCommitteefor Aeronautics,

lliwsrds,Calif.,August 21, 1956.

NACA ‘TN3859 11

13EFEmNcEs

I. Kayten, Gerald G., and Koven, William: Comparisonof Wind-TunnelandFlightMeasurementsof Stabilityand ControlCharacteristicsof aDouglasA-26 Airplane. NACA Rep. 816, 1945. (FormerlyNACAARR L5Hlla.)

.2. Duddy,R. R.: Some ComparisonsBetweenWind TunnelModel and FlightTest Results on Aircraft at High Angles of Attack. Papers Presentedat the Joint Sessionof the Flight Test Techniquesand Wind Tunneland Model TestingPanels, Ottawa, Canada. AG18/P8,AGARD, NorthAtlanticTreaty Organization,June 1955,pp. 107-125.

j. Stone,Ralph W., Jr.: Interpretationof Wind-TunnelData in Terms ofDynsmicBehaviorof Aircraft at EUgh Angles of Attack. Papers Pre-sented at the Joint Session of the Flight Test Techniquesand WindTunnel and Model TestingPanels, Ottawa,Canada. AG18/P8,AGARD,North Atlantic TreatyOrganization,June 1955,pp. 24-50.

12 NACA TN 3839

—MODEL —

—AIRPLANE -----

Figure 1.- Plan form of airplaneA. T

r M=O.91

% :L k:..o 10 20 30 0 10 20 30

a, DEG a, DEG --1.5

l!l,

(%Ja)F ‘“0 ~ 0 0 0

(%&T .5--- AIRPLANEo— AIRPLANE

0 .6 1.0 1.4

Figure 2.- Correlationcharacteristics

M

of flight andfor airplanes

1.8

:

-wind-tunnelliftA and B.

NACA TN 3859 13

-.4

[

AIRPLANE AFLIGHT

CmcL -.2

———WIND TUNNEL

/“——— ——— ——— —.— —

L, I 1 I I I 1 I I IO .60 .70 .80 .90 1.00

M

AIRPLANE‘—— A (SLOTTED THROAT W.T.)

o } B (CLOSED THROAT W. T.)

.2

[

------ B (SLOTTED THROAT W.T.)

AcmcL o .—~-=-.0 u

-.2 - 1 I t 1 I 1 1 1 I0 .4 .8 1.2 1.6

M

=2?=Figure 3.- Comparisonof staticmargin determinedin flight and

wind-tunneltests for airplanesA and B.

MN 0.76 M-O.91

:[+%=+[+

~~

%h

-.04 -‘Q

Cm ‘~ ‘k

‘b h-.08 - Y+”-%,

\\

\-.12 - \

\—FLIGHT \

‘+- WIND TUNNEL ‘b-.16- 1

~~0 16 240 16 24CZ,DEG Q, DEG

(a) M = 0.76. (b) M = 0.91. v

Figure 4.- Flight and wind-tunnelpitching-momentcharacteristicsforairplaneA with and without horizontaltail.

14 NACA TN 3859

-6

-4

1~

--~

it, DEG-. --- .

AiRPLANE A-- /“

-2---

— FLIGHT

L– – – WIND TUNNEL1 1 1 I I

O .60 .70 .80 .90 1.00

-2

Ait, DEG O

2

M

–-– AIRPLANE A

o } AIRPLANE B

1 I 1 1 1 I 1 1 I.4 .8 1.2 1.6 2.0

M

=s=Figure 5.- Trim characteristicsdeterminedin flight and wind-tunnel

tests for airplanesA and B.

.6

.4

T [

--- —-- -- —--— ------

\

-.. ----

.2

[111111111

— FLIGHT--- WIND TUNNEL

O .60 .70 .80 .90 1.00M

1.5 -

4

/--— --- 1

(T)F I“o -0

\\ o

‘7)WT .5 ---–AIRPLANE Ao— AIRPLANE B

I I I 1 Io .4 .8 1.2 1.6

M

TFigure 6.- Relativeelevator-stabilizereffectivenessdetermined

in flight and wind-tunneltests for airplanesA and B.

NACATN 3859 15

1.5

1.0

.5

1

1.

1.

,5-

0 -

5 -

1 1 1 1 1 1D

1 I

—--— —___ __

— FLIGHT- ‘-– WIND TUNNELI I 1 1 I t 1 t

O .60I

.70 .60 .90 1.00

Figure 7.- Mleronin flight

M

=5=effectivenesscharacteristicsdeterminedand wind tunnel for airplaneB.

.3

h

WIND-TUNNEL DATA,.2 M = 0.70

Cn , PER RADIAN.1B

_.1 L

~o 24

a, DEG30 r — Cng =O, WIND-TUNNEL

1“KG;* ‘N’T’ALD’VoL4~

LoM

Figure 8.. Comparisonof flight and wind-tunnelcharactersties. Airplane

vdirectional-instabilityA.

16 NACA TN 3859

180

[

—--——.——_

160 ____

Cnpv 140

PERCENT120

1

-----

WIND-TUNNEL DATABASICCORRECTEDFOR FLEXIBILITYCORRECTEDFOR FLEXIBILITY AND AIR INTAKE FLOW

/,//----

/ “-’w

h/0 ‘i-.+ -—-—

100 ‘-./”---

./ .- /-FLIGHT DATA--\ -- —-- /“ 1

-’ ---

80L I I 1 1 I 1 1 I

0 .8 1.0 u L4M

vFigure 9.- Coxqparisonof flight and wind-tunneldirectional-stability

characteristics.AirplaneC.

NACA - LangleyField, va.

NATIONALADVISORYCOMM17TEEFOR AERONAUTICS

TECHNICAL NOTE 3859

COMPARDON OF FLIGHT AND WIND-TUNNEL MEASUREMENTS

OF HIGH-SPEED-AIRPLANE STABIUI’Y AND

CONTROL CHARACTERISTICS

By Walter C. Williams, Hubert M. Drake, and Jack Fischel

High-Speed Flight StationEdwards, California

.Wash&tonAugust 1956

—