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THE

NATIONALADVISORY COMMITTEEFOR AERONAUTICS

TECHNICAL NOTE 2229

EFFECT OF END

By John M. Riebe

ON

and

SWEPT WINGS AT LOW SPEE:

James M. Watson

Langley Aeronautical LaboratoryAir Force Base, Va

Washington

November 1950

w- -~-r”- -n!, ,-,1,, ,L. l.-

. . .

. ..J l—= .-> [, :

. . . . . . . . . . . . . .. . .------- -------- . .. . ... . . .,. : ----- ”----..--”- --------... ..:. . ... ., .. ... .

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IllmllmlllulllulloI?ATI(INALADVISORY COMMITTEE FOR AERONAUTICS

0Clb50bl

mCHNICALNOTE2229

THE D?FECTOF END PLATES ON SWEPJ!WINGS AT IOW SPEED

By John M. Riebe and Jsmes

SUMMARY

M. Watson

An investigationwas made in the Langley 300 MPH 7. by lo-foot ‘tunnel to determine the effects of various sizes end shapes of end plateson the aileron characteristics and on the aerodynamic characteristics inpitch end yaw of a tig of aspect ratio 2 with no taper and a sweepbackof 450 aud of a wing of aspect ratio 4, taper ratio 0.6, and sweepbackof 46. p. Free-roll characteristicswere obtained with two end-plateconfigurations on a wing of aspect ratio 3, taper ratio 0.6, and asweepback of 350 h order to determine the effect of end plates on‘wing -damping in roll.

The addition of the end plates to the swept @rigs increased thelift-curve slope, reduced the maximum lift-drag ratio, generallydecreased the maximum lift coefficient, end increased the longitudinalstability slightly in the low lift coefficient range.

The variation of wing effective diQedral with lift ,coefficientwas.

reduced by increase in end-plate size. The effective dihedral at zero “lift could be changed from positive tb negative by lowering the endplates. The directional!.stability of the swept.wings was ticreased withticrease in end-plate area aud with rearward movement of the end plates.

The flap-type aileron end spoiler-aileron effectiveness increased ‘with the addition of end plates to the swept tigs; however, the increaseof the wing damping in roll may reduce the ro.11.inneffectiveness for someend-platejconfigurateions. In addition, end plates located bekw the wingchord line reduced the adverse yaw of flq-type ailerons.

Theoretical.and experimental.tivestigati&s on unswept wings endtail surfaces (for example, references 1 to 3) have indicated that theaddition of md plates will.generally improve the wing aerodynamicefficiency. The use of end plates which acted as a -barrierto the span-wise flow along the outboard portion of the spsn and around the tips ofairfoils resulted in increased lift-curve slopes, less induced drag,end higher maximnn lift coefficients.

.

.- ,.- .-— -- --- -— -- --- .-—---- —--.-—— -—.-. ———y------.— --: .—. . . . —.. —-. —________.. . . .. .>, .-.,

2 i NACATN 2229 “

The application of end plates to sweptback wings has been consideredas a possible mesns of overccmdng some of the lateral-stability diffi-culties (such aa large changes b effective &lhedral with lift coefficient)and other adverse effects (such as reduced lift-curve slope, ~tiUIU lfit~end aileron effectiveness) that result through the use of sweepback hwings.

The present p~er presents the results of a low-speed investigationmade tn the Langley 300 MPH 7- by lo-foot tunnel to deterndne the effectsof end plates having various sizes and shapes on the stability end con-trol Characteristicsof several swept w5ngs. For the most part, theresults are for end plates on a’wing of aspect ratio 2 with no taper anda sweepback of 45° with -ted results for two other swkpt wtngs ofaspect ratios 3 and 4.

The results, in general.,include the longitudinal stability,lateral stability, and lateral control characteristics as affected byend-plate size, shape, end location. The lateral control characteristicsinclude results for both flap snd spoiler ailerons, and reml.ts of afree-roll investigation of two end-plate configurations on a wing ofaspect ratio 3, tap= ratio 0.6, and 35° sweepback to deterxlne theeffect of end plates on wing damping in roll.

SYMBOLS

The forces and moments measured on the wings (fig. 1) are presentedabout the stabili@ axes, which intersect at the center-of-momentposi-tions shown in figures 2,”6, and 9. The Z-axis is in the plane of-try and Perpentitiar to the relative wind, the X-axis is in theplane of symmetry and perpendicul= to the Z-axis, and the Y-sxis ismutually perpendicular to the X-axis and Z-axis (fig. 1).

CL.

%

Cy

Cz

cm

Cn

The symbols used are aa follows:

lti coefficient (L/qS)

drag coefficient (D/qS)

lateral-force inefficient

rolling-mment coefficient

.

(Y/qs)

(Lt/qSb)

t coefficient (M’/qSb)pitching—mallen

moment coefficient (N/@b)Y=Q3-

--— -. . .. ---- . . . ----- .-—..”:-..!... -~.”. Yy———---------”. .. . ..,.. --” .’.,..- .’-. . . .,. -.. . .

NACATN2229 3“

.

.

L ltit of model, pounds (-z) .

D drag of model, pounds (%X when $ = 0°)

Y force along Y-axis, pO~dS.

x force &long X-axisj pounds

z force along Z-axis, pounds

LI rolling moment about X-axis, foot-pounds

M! pitdhing moment about Y-sxis, foot-pounds

N yawing moment about

L/D lift-drag ratio,

~ free-streem dynsmic

foot()1 ~2~P

z-axis, foot-pounds

pressure, pounds per sqyare

P rate of roll, radiaus per second.

s wing srea (6.oo square feet on wing model of aspectratio 2, 3.17 square feet on whlg model of aspectratio 3 snd 2.25 square feet on wing model of aspectratio kj

.lateral a?ea of both end plates, square feet

wing mean aerodynamic chord (1.73 feet ofitig modelof aspect ratio 2, 1.05 feet on wing model of aspect

se

E

ratio 3, and 0.765 foot onratio 4)

c wing tip chord

b wing span (3.46 feet on wing3.09 feet onw3ng model of—

Wing modelof aspect,

model of aspect ratio 2,aspect rati~ 3, and 3.06 feet

on wing model of aspect ratio 4)

effective height of end plate (Se/2c), feet

ah velocity, feet per second

ma8s density of *, slugs per cubic foot

.- .:..- -------- --- . —.——,— ._=_ —. . _ -...— —— ~. —.—. .. .——. — —--- -. . ,,.. .

4

A

M

R

a

M

A

A’

h

A

pb/2v

NACATN2229.

.angle of attackof chord ldne at root of model, degrees

angle of yaw, degrees ‘

aileron deflection, measured h a plane perpendictiar tothe hinge axis, degrees

ticrement in coefficient due to end plates

Ma& number (V/a) -

Re~lds nmiber (p~/v)

speed of sound, feet per second .

coefficient of Absolute viscosity, slugs per foot-second

* aspect ratio ( 1)b2 S

efiective -g aspect.ratio with end plates

taper ratio (TiT chordfioot chord)

sweep angle of

wing-tip helix

coefficient of

o

gyarter-chord ltie

angle, radians

()%.demping in roll —(9a~

.

.

, .

.

.

. .-. . -,- ~-..:.... . . . -——~-=. -“-—— ‘“” -- “—”-”— ‘-~,. ......-””.. - . ...’”’..’. . .. . .-.,-.-’:. .’..

I

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NACATN2229 5

.

acnc%= ~

.

SUbscrlpt: . .

max nlsxhnum

CORRE&l!IONS

Wing of aspect ratio 2.- The sngle-of-attack snd the drag data havebeen corrected for jet-boundary effects according to the methods out-lined iu reference 4 for unswept wings; as csn be seen from reference 7,there is little effect of sweep on the.jet-boundsry effects. Blockagecorrections were applied to the test dsta by the method of reference 6.The data have been corrected for the effects of the mo&l support strutby the use of tare corrections deterndned for the tig without end plate.

Wing of aspect ratio 3.- Blockage corrections were applied ti thetest data by the method of reference 6. A small tare correction beca~eof bearing friction has been applied to the free-roll results in theform of an increment of dsmping-in-rol.lcoefficient equal to -0.005.

Wing of aspect ratio 4.- The angle-of-attack and the drag data havebeen corrected for jet-boundary effects according to the methods out-lined in reference 4.

MODEL AND API?-S.“,

Wing of aspect ratio 2.- The 45° sweptback ~ model of aspectratio 2 (fig. 2) was mounted horizontally on a single strut in the

_eY 300 ~ 7- by lo-foot tunnel (fig. 3). !l?h~untapered wing hadNACA @lAOIO airfoil sections normal to the wing leading edge and hadneither twist nor dfhedral. The ‘wing,which was “constructedof wood,hadfor

rounded tips which were removed forward of the aileron hinge lin~the investigationwith end plates.

,

.“.. —.. - ____ ._..__ ____ .. . . .. ——.. —.—— .. --— ._. —.—.—

.’.,.- ... . .. ...” . . .

6

\

NACATN2229

The end plates investigatedwere constructed of –-inch sheeti

duralumin with rounded edges to the MmensionE shown in figure 4. Acutout was made-in the trailing em of each end plate to,allow fordeflection of the outboard, ha13-semispan, O.25-chord, P1*J se~edaileron.

The stepped spoiler ailerons investigated were constructed of “

~- inch aluudnum angles which were fastened to the wing upper surface

as shown in figure 5 end projetted-8 p-cent of the vdng chord. Thisconfiguration corresponded to one of the more promising stepped-spoilerconfigurationsfor this fig plan form (unptiLLshed data).

.

wm of aspect ratio 3.- The .35°sweptback ~ model of aspectratio 3 snd taper ratio O~6 used for the free-roll.investigation ismown in figure 6. The wing was supported by a sting extending forwsrdinto the test section from a vertical strut. A schematic drawing of .the support system and rolltig appsratus is shown in figure 7. Thesngle of attack of the model waa changed by varying the angle ofincidence of the wing relative to the sting. Rolling—nunent data wereobtained by an electrical strain gage with-the sting restrained in roll.When the model was permitted to roll freely under the moment created bythe deflected ed.1.eron,the rate of roll was recorded electrically.

The ordinates of the symmetrical, 12-percent-thick airfoil sectionof the 37 sweptback wing are given in table I. The model was con-

.strutted of steel end the * end pities were constructed of k -inch

8sluminwn sheet with rouuded edges. The model was equipped with au oti-board flap-type ed.leronwith seeled gap.

wing of aspect ratio 4.- The 46. P sweptback wing of aspect ratio 4and taper ratio 0.6 was tested on a sting-mounted electrical strain-gagebalance (fig. 8). The sting was attached to a single strut which variedthe engle of attack and sngle of yaw of the model. The wing r~ed @the center of the test section at various angles of yaw but was dis-placed vertically at various angles of attack.

Dimensions of the wing, which had NACA 65A206 &oil sections, endof the end plates Qvestigated on this wing are given in figures 9 and 10,—

respectively.

durelumin with

1 tich or+The end plates -e constructed of ~- -inchrou+d or beveled edges.

._.-_. — _— ___... . .. — . . ... -,.. . . -.. . .. . .’. .

.,

I,!

“~

.

!EmJ!s

The conditkm end types of tests made on the three nings h the Langley 330 MPH T-”by

lo-foot tunnel are as follows:

conaitlQn or wing

type of testAspect ratio 2 Aspect ratio 3 Aspect ratio 4

q 100 145 no ‘

H 0.27 0.32 0.28

R 3.2 X 106 2.3 x lc$ 1,84 x 106

Longitudinal .91xMlity a . -P to stalJ. ------------------------ -- ~ .40 ~ 240

Lateral stability

{

+’ P ma, -? :-------------------—---- 9.@aud50a . -6° to atau -------------------“-----a. -40 to stall

Directional stabflity

{

+=!P.and-p ---” ------------------ ---q ----------------

a=-~ta stall -------------------------- ----------------

Lateral. control a.-fP to stall a. 6.50 ---------------Flapdqpe ailerons 6a . -lOoJ Oo, ~, ad 10° ~~ = -lF to 9.40 --—----- -----Spoiler ailerons 4.08 ~ chord projection ---------------------- ------ --. —---- -------

{

“e

Free rolla = 0.30, 3.F, auti6.50---------.----.----.----”- ---------------~a . -150 -&J9*40 ----------------

8 NACATN2229 ‘

REsums .

...

The r~sults are presented

Aerodynamic characteristics inLateral-stability parameters .

in the foUow3ng figures:

pitch . . . . . . . . . . . . . . ..11. . . . . . . ...0. .. 0.00 ● E

Flap-type-aileron characteristics . . . . . . . . . . . . . . . . . 13Stepped-spoiler-aileronchsmcteristics . . . . . . . . . . . . . . 14Vari;tion-ofVariation of

Variation of

Variation of

Variation of

Variation of

size . . .Vsriation of

size . . .

AC~tith end-plate size. . . . . . . . . . . ...15

AC&with end-plate size. . . . . . . . . . ...16

ACDwith lift coefficient.. . . . . . . . . . ...17

()

& withend-plate size . . . . . . . . . . ...18D-

(%)c c@ I

and acz bcL tith end-plate size ..19 and 20w

%end CyV at CL = 0.5 with end-plate

. . . . . . . 9****** ● *=9*** .** 21and22

%klsnd Czaa at a . 00 tiw end-plate ,.

23. . ...00 .. 0009. ● .**.*. ● ***=*” . .cl duetoadleron deflection . . . . . . . . .. . . .9 . . ...24Variation of pb/2V with aileron deflection . . . . . . . . . . . . 25Variationof (pb/2v)ba witha . . . . . ... . . . . . . . ...26

The slopes presented h the figures were taken over a lift-coefficient’range of about *O.1, an sngle-of-atiack rsnge of *2°,’en aileron-deflection renge of *10° or 0° to 1oo, =d a yaw rauge of 0° to 5°.(For the wing of aspect ratio 4, the values of AC& (presented in

fig. 15)”were determined from lateral-stabilitytests made through anEU@e-of -attack range at 5° angle of yaw; however,.a few tests * at0° angle of yssishowed that small angles of yaw had little effect onthe incremental values of Cl&.)

DISCUSSION

Aerodynamic.

Characteristics b Pitch

Lift-curve slope.- The addition of end plates to either the wingof aspect ratio 2 or aspect ratio 4 increased the lift-curve slope inthe low lift-coefficientrauge (figs. U and 15). A compwison of theticrease in lift-curve slope for the mptba *gs ~fi end plates

...

—..—. -.—. ..- -.—- .,- =.. ...=— -,—.. -,’ ..-; ,, .., . .. .., . . .,. .,

#

,.

IucATN 2229

obtained fromobtained from

.

9

the exper~ntsl- data of this investigation and thatunswept-w5ng end-plate theory (reference 7) with the use “

of reference 8 shows good agreement (fig. 15). It can be seen tl&the increments in C% due to end plates decrease as the wing aspect .

“ ratio is increased for & given value of h‘/b ,but are relatively inde-pendent of end-plate shape. This is as would be expected, since thetheory of reference 7 tidicates that regardless of wing.aspect ratiothe increase in A’/A is dependent on end-plate-height - wing-spanratio end eqy,al.increases h A’/A result in smaller ticreases in C&at the higher aspect ratios (rtierence 8).

.

The increases in lift-curve slope (fig. 15] represent increases ineffective aspect ratios of about 1.8 for the wing of aspect ratio 2 and

3.6 for tie ~ of aspect ratio 4 at ~ues of ~ = 0.5.

A comparison of the results of this investigation of swept wingswith the results of unswept wQgs (reference 3) indicates similar chsngesh L!C~ end Ai/A with h’/b for values of h’/b less then 0.5; for

values ~eater them 0.5, however, the unswept-Mng data indicate thatthe chsnges tn AC% end A’/A are less thsn those predicted by the

theory of reference 7. It would be expected, therefore, that furtherticreases in end-plate size on swept wings would give further increasesh C%, but that the fncrease6 would proba@ly be less than those pre.

dieted from the theory of reference 7. Wing taper ratio might beexpected to have an effect on A% for a given value of h ‘/b for ,

wings having fairly him tap~. The results of this tivestigationinticate that the effects of ts&r, if any, are very small for the rmgeof mqdel geometry used.

The results of this investigation indicate that the lift-curveslope of a swept ~ wifh end plate cau be satisfactorily predictedfrom the”theory for end plates on unswept wings up to a vshe of h’/%of 0.5.

Maximum lift coefficient.- The maximum lift coefficient C*

,yae generally decreased by the addition of{end plates to the 45 swept- ‘

back wtng of aspect ratio 2 (figs. 11 end 16). T%is @crease is o~ositeto the effect found on unswept wings (fig. 16) where increases in C-generally resulted from installation of end plates.

End plates located below the wing chord plane were found to haveless adverse effect on the velues of C~ fi= the smhe end plates

located ~ove the chord plane. In fact, three of the end plates testedbelow the chord plane (a triangular eqd plate, a semicircular end plate,

.

--- . . . . .- - ... ——--- .,- ———— —.—..._ --—— —.. — —- ——. —.. .i..,. -, ,...,. ..- ,.. , .. “. .,. ,-.

10 NACATN 2229“

and a 45° sweptback end plate, all exten~ c/2 below the chordplsne) had a negligible effect on CL. F

Drag coefficient.- The drag coefficients of the 45° sweptback wing .of aspect ratio 2 were generally increased by the addition of the variousend plates to the wtng, except h the titermediate lU%- coefficient rsngeof about 0.4 to 0.8 Were the reduction h induced drag resulting fromthe increase h wtng effective aspect ratio exceeded the drag of the end’ylates (fig. U). C&paring the incremental drag coefficients estimatedfor unswept wings by the method of reference 1 with the experimentalvalues for the swept wings of aspect ratios 2 and 4 (fig. 17) *OWS veryBimilsx trends. For the cskulations, a skin-friction drag coefficientof 0.011 was assumed for the end plates. The estimations show that verylittle, H any, drag reduction can be expected below ltit coefficientsof about 0.4. The comparison is limited to lift coefficients up todbout 0.6 since the lift curves (fig. U) indicated a nonlinear varia-tion of lift coefficient with angle of attack at higher values of CL.Above values of CL = 0.8 the data, in general, show increases in dragcoefficient when the end plate is added to the w3ng.

Tuft studies of some of the end plates of the present investigationshowed that there was unsteady flow on the surface of the end plate.This unsteady flow developed at intermediate angles of attack andgradually becsme more unstea~- as tie angle of attack was inaeased.The disturbed flow was generally more prevalent on the outboard surfacesof the end plate. It is believed that more careful design of the airfoilsection of the en! plate could result @ more favorsble drag character-istics at the higher 1~ meff icients. .

The reduction in drag coefficient was less for the swept wing ofaspect ratio 4 th= for the swept wing of aspect ratio 2 (fig. 17).,These results sre consistent with the trend indicated in reference 1and correspond to the previously noted conditioiiwherefi larger incre-ments of I.Ht-curve slope were obtained on the wing of aspect rtiio 2then on the ~ of aspect ratio 4 with the addition of end plates ofa given area ratio. (See fig. 15.)

The change h the values of (L/D)E for the wing of aspect

ratio 2 generally had some scatt& with end-plate size end shape “(fig. 18); however, the values of (L/D)E generally decreased with

increases in end-pkte area ratio. inspection of figure U. shows that(L/D)E OCCUrS at lift coefficients less then *out 0.3. me resflts

(fig. 17) indicate that unless the end-plate drag is very mall noappreciable gains in (L/D)E can be expected since there is ~

increase in dreg coefficients due to the end plate for lift coefficientsbelow 0.4.

. .

●

.

.

——. — --- ,. . . . .. --- .-: . ----- .- .——--– -. ....” ..- . .. .., .. . . .-.....--—. .

.. .

NACATN 2229 IL

.

Longitudinal stability.- The addition of end plates to the swept-back wing of aspect ratio 2 resulted in an increase in the longitudinalstability of the wing .inthe 0.1 to 0.65 lift-coefficient range (fig. U).The shtit in ~odynamic center varied almost linearly vith end-platearea, the aerodynamic center moving back ~out 5 percent mean aerodynamic

chord as the end plate was increased to ~ = 1. Variations in red-plate

shape and location had only small effects on the longitudinal st~ility.The increase in longitudinal stability of the sweptback wing probably isdue to a shift of the center of pressure outbosrd as a result of restraintof flow about the wing tip with the end plate in place.

Data obtained with the sweptback wing of aspect ratio 4 (not pre-sented hereiu).showed similar results.

Lateral Stabili~

Effective d3hedral.- The rate of chsnge of effective dihedral with

lift coefficient at low lift coefficients ac~/bCL was reduced withincrease in end-plate area on both the sweptback tigs of aspect ‘,ratios 2 end 4 (figs. 12, 19, and 20). The reduction in &@/&L

was generally independent of end-plate shape. The small end plateslocated ahead of the tip chord (fig. 20) generally appeared to be moreeffective in reduchg @V/&L than end plates located farther back.

The reduction in the values of ~ZV/&L with increase in end-

plate srea can be partly attributed to a side force on the end plates. -As the wing angle of attack incYeases the end plates move downwardrelative to the moment axis, and the side force acting on these endplates produces a rolling moment o~osite to that produced by the Mng. ,If the value of CyV (fig. 21) obtained in the investigation and thegeometric properties of the @ of aspect ratio 2 are used, the com-puted reduction h W2~laCL due to the end plates is only sboti 1/3 the

reduction shown h figure ig. The remahing reductions in ~@CLmay have resulted f!comthe end plates caudng separation and loss oflift on @e lea&ng wing and reducing the tip losses on the trailing

W*

The value of the wing effective-~edral parameter cz~ at zero

lift was dependent upon the end-plate area and upon the distribution ofend-plate area above and below the wing-tip chord line (figs. 12, 19,snd 20). Positive increments in Cz$ resulted frm placing the end-

plate area above the tig chord line and negative increments wereobtained when the end-plate area was added below the chord We. This

,

-. .. ----- .. --— —.z—--- ——-—..-— -7’--~ T .~-—— —— ----- .— .--—.—-,-.—---- -— -

,.. , . . ,. . ...’ .-” -,. .“..

.

12 NACATN 2229

change in CZV with end-plate position results from the side force of

the end plate (figs. 21 end 22) acting above snd below the chord lineend smounted to en increment of Czv of about 0.003 when an end plate

Seof — = 0.5 was placed either above or below the chord ‘lineon the

wing ‘of aspect ratio 2.

The maximum effective dihedral Czv was obtained on both the wQg of

aspect ratio 2 (fig. U?) and aspect ratio 4 at moderate lift coefficientssnd the values of CzWM exhibited the same trends with ficrease in .end-

plate area as were exhibited at zero lift (figs. 19 and 20). Unpublisheddata indicate that an increase in Reynolds nuuiber,‘tovslues correspondingto flight would increase the ~um effective ~edral by extending ther-e of ltiear veriation of C1* with CL to higher lift coefficients

and would *US delay aud possibly decrease the reversal tendenciesof CZ* exhibited by most of the wing and wing-end-plate combinationsat the low Reynolds nunibersof the present investigation.

wing-moment coefficient.- The plain wings of aspect ratios 2 and ~had a~>oxhately neutral tiectionsl stability (C% x O) over thelift-coefficientrsnge (figs. 12 and 21 . An ticrease h ~ectionalStdXU.ity (Cn

1

becoming more negative1 occurred with increase h end-

plate area; th s effect was reasonably tidependent of end-plate locationmove end/or below the wing chord line. This fact is indicated by thedata of fQures 21 and 22, which are for a lfit coefficient of 0.5 andalso generally applied to the variation of Cn* with end-plate srea

throughout the lift-coefficient range (fig. ld). The data of figure 21also show en effect of forward and rearward location of a given end-plate mea on the values of Cn*; this effect results from a change h

the moment erm between the wing center-of-momentposition end the centerof pressure developed on the end plate at sn#.es of yaw.

Side-force coefficient.- The variation of CyW with CL was

negligible throughout the lift-coefficient range for all end-plate con-figurations on the wing of aspect ratio ? (fig. I-2). Increase in end-plate size resulted in larger positive changes of CY* for the swept

wings of both aspectpendent of end-plate

.

ratios 2 and 4; this effect appe&ed fafily @de-shape (figs. 21 end 22).

.

.

Aileron Characteristics

Flap-type ailerons.- For most”of the ~ end-plate configurationson the swept wing of aspect ratio 2 the rolling-moment coefficientsproduced by the ‘kileronwere lexgest at low angles of attack (fig. 13).

.

\

.

-—— ... .. ...- ------ ---. —---.- . ..- --.: -. -—--- --,—,.- . . ,.-’,... , .,, ... .“. . ... .-” -,----

NACATN 2229 13

The aileron-effectivenessparsmeter c%a .increased when the area of end

plates was increased (fig. 23) and Clbe generaUy became more effective

when the end-plate area was concentrate~nesr thewhg trailiug edge.One of the most promising end plates was the trisngular~shaped end platewhich had a value of Cz5a at-0° angle of attack (0.0013) sJmost equal

to that of the ma?dmum obtained with any of the end plates in theinvestigation audyethad a relatively small end-plate area (fig. 23).

The wing with end plates having area only above the wing Chord linehad positive values of Cl%a> whereas the wing with end plates hatig

‘area only below the wing chord line had negative values of C~a (fig. 23). 0

For a given end-plate shape and position, increasing the end-plate axearesulted in an increase h the magnitude of the values of C~a. One

of the largest posittve -d one of the largest negative values of Cwa

were produced by the wing with t+e relatively smaU triangular end platelocated move end below the wing chord line, respectively.

Analysis of yawing-moment data obtained from aileron tests madethrough an sngle-of-attack range (but which sre not presented herein)has indicated that the reversal in C~a as the end plate was shifted

from above the -g chord line to below the ~ chord ltie resultedprimsrily from a change iu lateral force on the end plate. With endplates located above the wing chord ~e, down deflections of the ri@taileron resulted in side force in the negative direction on the rightend plate because of the increased negative pressure above the wing.Positive yawhg moment on the wing resulted because the center of pres-sure of the end plate was behind the center of moments of the wing.With an end plate loc,atedbelow the wing chord Me, down deflectionsof the right aileron resulted in negative yam mment because of theincreased positive pressure below the w5ng and the positive side force.With ailerons located on both wings the effects mentioned above wouldbe additive. For example, with end plates located below the wing chordline, a negative yawing moment would be produced from down deflection ofthe right aileron and also from up deflection of the left aileron.

Spoiler ail=ons. - At angles of attack below approximately 160, theaddition of the circular end plate to the swept wing of aspect ratio 2increased the spoiler-aileroneffactiveness substantially (approximately ‘ ‘a 75-percent increase at- a = Oo). (See fig. 14.) The ticreasedeffactiveness of the spoiler aileron probably resulted from the tncreasedeffective aspect ratio and correspondinglyhigher lift developed by t&ewing-end-plate combination. move 16° sm@e of attack, where the liftof the tig with end plates was less thau that of the plain ~, the

.— -. —— ~—. —. —.-—.-.. . . . ,... . .. ,.- . .

,#

a NACATN2229

effectiveness of the s@ler aileron on the plain ~ was greater thanon the wing with end plates (fig. 14).

Above 2° angle of attack, ti the region where the end plates hada favorsble effe~t on Cl, the value of Cn produced by the SPOfler

qilerons was reduced when the end plates were -addedto the wing. Thew- moments were unfavorable for the wbg with the cficulsr endplates (fig. 14) *OVS sn sngle of attack of 1P.

Rolling Characteristics#

The addition of the end plates to the 450 sweptback wtng of aspectratio 2 resulted in ticreases h Czaa; however, the wtng dsmping-in-roll

coefficient Czp may increase at a greater rate with the addition of

the end plates than did C~5a snd thus result in lower values of pb/2V.

In order to investigate this effect, a .few static-roll and free-to-rolltests were made cona roll rig (fig. 7) of a 350 sweptback wing of aspectratio 3 (fig. 6). One of the end-plate configurationstested on the .35° sweptback wing was similsr to that tested on the ’45°sweptback wingof aspect ratio 2 with which the lsrgest value of %~a was obtained.

For the other end-plate configuration investigated on this wing, tieupper half of the aforementioned end plate was removed. The flap-typeaileron on the 350 sweptback wing did not extend to the wing tip;therefore, no CUtOUt was made h the-end p~te to P-t tie ~eron todeflect - as was necessary on the 45° sweptback wing of aspect ratio ~.The plain-wing data for the 350 &eptback wing presented herein wereobtained by extrapolating some unpubli~ed data obtained in the Lsmgleyhigh-speed 7- by 10-foot tunnel at high subsonic Mach numbers. Thesedata obtatned at high Mach numbers were readily extrapolated to the Machnuniberof the present investigationbecause they varied linearly withMach nmber at-all.except the highest Mach numbers.

Results of the static-roll tests of the 35° sweptbacku = 6.50 with and without the two aforementioned end-plateshow some nonlinearity of Cl with aileron deflection forend plates (fig. 24). The value of Czaa determined from

●

wing at ,configurationsthe wing withfigure 24

* resulted in s- aileron-effactiveness trends with end plates as thedata obtained on the 45° sweptback wing of aspect ratio 2 over the ssmedeflection rsnge (figs. U =d 23).

The effect of the two end-plate configurations on the variationof pb/2V with aileron deflection was detemined from free-ro13.testsat angles of attack of 0.30, 3.5°, and 6.50. The variation of pb/~

y --- ~:.. .... - .,..—.- .. .. ,..-=— .-—..”., -“—=. . . .. .. . . .. . . .. . ~ _. _.. -:. ,4.,.. . . . . . . . . . .. ’,,.. . . . . . .

NACA

overeter

TN 2229 15

the aileron-deflectionrange is linear (fig.’2~). The parsm-(pb/2V)ba was reduced by the addition of the end plates through-

out the sngle-of-attackrange tested; this reduction in (pb/2V)~a

varied little with end-plate area (figs. 25 and 26) for these two end-plate configurations. ~wo@ cz~a increased with end-plate area,

the damping-in-roll coefficient Czp ticreased at about the ssme rske,

as shown by values of Czp computed from the data of figures 24 to 26,

(-O. 305 for the plain wing, -0.365 for the wiug with the sweptba~ endplate located below the wing chord line, and -0.436 for the sweptbackend plate located ~ove and below the wing chord Me). The Czp VSIL1.les

were determined from the relationship

C2 c2~aC2P =-— -pb/2V= (pb/2V)~a

The values of Czp were computed for an angle of attack of 6.5° inasmu@

as no static rolling-moment data were obtained at any other angle ofattack. The increase in the values of Cz

Pwas proportional.to end-

plate area; that is, dodd.iug the end-plate area about dcnibledtheficrease in Czp.

.

As noted previously, the triangular-shaped end plate of smaller areathan end plates of.other shapes may be utilized to obtain a given increase

fi c2~a (fig. 23). Because the dsmping.inrolJ.is shown to vary with J

end-plate area, the triangular end plate should result in a smsller incre. ‘ment of Czp. It may therefore be possible for low-aspect-ratio swept-

back wings with end “platesof this type to have *out the ssme or.largervalues of (pb/2V)~a as those of plain wings.

Unpublished data from a free-roll investigationmade in the Langley300 MPH 7- by lo-foot tunnel actually did show a very slight increase~ (pb/~)8a ~~ a - ad plate was attached to a sweptback-wing

model with the end-plate area mncentrated near the aileron. The wingreceiving benefits from the addition of end plates - such as increasedlift-curve slope or reduced variation of effective dihedral with liftcoefficient - would thus not be penalized by reduced rolling power.

..

.. .>-. ---- ----. -,-. - .... . ...- —:. — ..-. — .. —-— -— —-—-. -------., ., .,-—- —— --— . . .——-- -... - . .. .. . .

NACATN 2229

CONCLUDING REMARKS

An investigationwas made in the Langley 300 MPH 7- by lo-foottunnel to detezmdne the effects on aileron characteristics and on wingcharacteristicsin pitch and yaw of various sizes snd shapes of endplates on several sweptback.wimgs.

The addition of end “platesto sweptback tigs increased the l’fit-curve slope ti the low-lift-coefficientrange. This increase in lift-curve slope tended to increase with end-plate size and could be pre-dicted from unswept-wing end-plate theory. The end plates also generallydecreased the maximum lift coefficient, decreased the maximum lift-dragratio, end slightly increased the longitudinal stability in the low-lift-meff iciat range.

The vsriation of wing effective dihedral with lift coefficient wasappreciably reduced by ticrease in end-plate size. The effectived3hedrsJ.at zero lift could be changed from positive to negative bylowering the end plates. The directional stability of the swept wingswas increased with increase ti end-plate area mid with rearwsrd move-

,

ment of tie end plates. 1.

Although the end plates increased the j?lap-typeaileron and ~iler-aileron &?ectiveness, free-roll tests showed that end plates 4s0increased the dsmping in roll and may result ‘ina reduction in rollingeffactiveness for some end-plate configurations. b addition, end plateslocated below the wing chord lfie reduced the ~verse yaw of flap-typeailerons.

Langley Aeronautical LaboratoryNational Advisory Committee for Aeronautics

Langley W Force Base, Va., August 24, 1950

.

,

,. .-= .- . ..T —.. ,.—.— .— -.,. .--— .., ..’--”. ..a’v ::”--:~ -.”---- ..,. . ..” ;.”.-

●

✎✎✎✎ ✎✎

NACATN 2229

.

.

1.

2.

39

4.

5.

6.

7.

8.

REFERENCES

1“”

Hemke, Paul E.: Drag of Wings with End Plates.. NACA Rep. 267, 1927.,

17

Msngler, W.: The Lift Distribution of Wtigs with End Plates. NACATM 856, 1938.

Bates, Willism R.: Collection snd Analysis of’Wind-Tunnel Data onthe Characteristics of Isolated Tail Surfaces with and withoutEnd Plates. l?ACATN 2291, 1947.

Gillis, Clarence L., Polhamus, Edward C., and Gray, JoSe@ L., Jr.:Charts for Determirdmg Jet-Boundary Corrections for Complete Modelsh 7- by 10-Foot Closed Rectangular Wtid Tunuels. NACA ARR L5G31,1945.

po~us, Edward C.: Jet-Boundsry-hduced-Upwash Velocities forSwept Reflection-PlaneModels M@mted Vertically in 7- by 10-Foot,Closed, Rectangular Wind Tunnels. NACA TN 1752, 194.8.

Herriot, Johu G.: Blockage Corrections for Three-Dtiensional-FlowClosed-Throat Wind Tunnels, with Consideration of the Effect ofCompressibility. NACA RM A7B28, 1947. .

Von K&m&n, Th., and Burg~s, J. M. : General Aerodynamic Theory -Perfect Fluids. Airfo~ snd Airfoil Systems of F@ite Span.Vol. II of Aerodynamic Theory, div. E, ch. IV, sec. 19,W. F. Duraud, ed., Julius Springer (Berlin), 1935, pp. 211-212.

po~smus, Edwd C.: A Simple Method of Estimating the %bsonic “Lift and Dsmping iu Roll of Sweptback Wings. NACA TN 1862,1949.

.

.

------ -— .. . . ------ . .. —---- ---. —----- ----- ——--——-.. --—.—— --— --. .--— ...-

18

.

—WA TN 2229

TABLE I

AIRFOJZ SECTION ORDINA!CESOF 350 SWEPI’BACKWING

r@l dimensions in percent“ofwing chord parallelto plane of symmetry of wind

.’

Station

o● *7.880

1.4662.9265.8308.71011.5(3317.21522.77328.24133.62038.9U44.I.I.649.23454.26559.21264.07468.95973.54678.15882.68887.3.3791.50495.792100.000

Ordinate

o1.0961.323L 6692.2602.9983.4923.8634.3934.7524.9955.149‘5.232

%%4.9094.5514.0783.5322.9552.382I.840L 338.876.441.021.

..

.

... . . . . -- ----.,. :.., --- ,. . -. .. . .y. i.-, ~-- —,- , . .- --- —---... . . ——-. ..,. . . .... . . . . .. ,. .-, - . . ..-, ,’,. ..

.

. ‘\Relative wind .

.

.

I

1.

I

Figure 1.- System of stability axes. Positive values of forces, moments,ad angles are indicated by arruws.

. . . . . . -- .- -—---,------- —.-- - ..——— --.—- —.. —-—— -. -. . —-— .. -.-—-—— -- -. . . . . ,.... :’. ., ..” - ,-.. . .

.

.-

.

1

TQ25 chord fine

Haih=n7ng-tI@, outline

“iK,+,..N:.!1/ I

PY.gure 2.- ffeomtric characteristics of the 45° swsptback wing of aspwtratio 2 and taper ratio 1 with tious end plates. S.6~fiet.

, , .

I

,.

,,

Ik@u’e 3.- The 45° sweptback wing of aepect mtto 2 with triangular endplates mounted q single strut in the Langley 300”MPH 7- by lo-foottunnel. t3a = 10°.

,:.,

.

.-

.

.

●

. .

.

.

.

.

0

. .

.,. . —

.; I ,“J”---’ .-.-.--— -.....-.—,------- ---

..,..;.. ,.-. . . .,. .: .-”.. . . . . . . -, -

.

se/s= 0.230 $/s.0.236 S#S=O.378

-mLz!22.$&;lL2k .@~.5#s.a479 Se/3=C42W Sef!!=(2236 se/s =0235

R=aoPc

.&~T*w~w-gG-

5#s=0366 q#s=o.46# Sef.s=am @W!466 q@-a??6

&/SSO. 756 s@as58 q?/s=o.459 q#3=Q459

Figure k- Geometric character18ticB of the varioua end plates

on the 45° swept’back wing of aspect ratio 2.Investigated

24 NACA ~ 2229

/

.

IHgure 5.- Stepped spoiler ailercms investigatedof aspect ratio 2.

,

‘!1

O.o+ ,\ ,w/hg

Ickw’Spoiler deh2il

on the 45° swe@back wing

.

.

,

. . . . . . . —.. ..= —.- . ;.;. .- “..: . - ~ -~.-.=. -- -, ---- -...- .-, - .,, . . -“”7.~”- ..” .. ’..’.. .....’.’. .

I I

~s~fz” +)?igure6.- Geometric characteristics of the 35° sweptback W@ of aspect

ratio 3 and taper ratio 0.6Utith two end-plate configwatlcma.

s . 3.17 equare feet.

fine

.

.1

,.,:,

~.,.,,1,:

.!,..,“. ,

.i!

Strut falrlng

,.iSting fa~@7~

Tunnel center the

1.

Sting su~ti

.— [

‘Test wing

‘Block fw chonghg =6rusttmgfe of attd beorlng

.,

EY.gure -(.- Schematic dxawing of the free-rolllng sting mounted in the testsection of the Langley 300 MPH i’-by lo-foot tumel. “

.—

‘. 1

, IiAcfi m 2229 27

,.

RLgure 8.- The 46.70 sweptback wing of aspect ratio 4 with end platesmounted on sting-balance system in the Langley 300 MPH 7- by 10-foottunnel.

,

- . ...... . . . . . . . . .__. _._. .— ____ .. -.. ..- .-— .— --- -—- -Q. — . ——. - —- - .,. :.. .

.●

\

I

..

#

.

.

.

.

.— .. ..—- +.., .=. .”—. .

... . . . . .-.-—-—- —-

.“: . . .. . .: . ...’ .. . .... ,.,.

s

I

‘1

End

1/ ~f’$$ L‘i127b50.oCenter of moments

0,25chro’he

plate ~

.

/# 1/

●

\

\’\ \

\

~

‘1’- t: /800” —3600”Figure 9.- Geometixlc characteristics of the h6. ~ swaptback wing of aspect

ratio 4 and taper ratio 0.60. S = 2.23 Bquare feet.

7“ 6.75”\

1

\“

=w=

.

I

.,

,. .,,’

. .t

.’

I

4

. .

’8’

\ n ;*C0.36 c++ ‘e& = o~2’8~0.s7; ‘+s = 0.248 !_ 0.37C \ \ _L

.—

Figure 10. -Ea

Geometric ckxacteristics of the various end plates investigated

on the 46. P meptback wing of aspect ratio 4. - “IIw

. & ,

NA.CATN 2229 3-

.

●

.

.

.

.,

J

E0

32

24 - — — — — — — — — — — —

~ 16“ — — — — — — — — — —%8’8 _ _ _ _ _ _ _

Q ‘ wT

p

--4-2’ 02’ 4.6.8 LO~

G

,

Figure 11.- Aerodynamic characteristics in pitch of the 4-5° syeptback wingof aspect ratio 2 with and without various end plates.

. . . ...-— —.——-——.—- --~ .. .,.—-----. - .- - — ----------- .— -----, —-- —-— -— —-. ---. . . . ,.. .,. . . .- .,- .

.

.

32 s NACA TN 2229

m

h Q.

1

, 0

3

G2’

./

o

-x? I 11(1b)

Figure 11.-”Continued.

,

c

.

. ..7 ..- .,, — ..- -—- ------ ----- ., . . . . ~.:-. ---J .------ —-.. -. -... . . ...’.... .. . . . ., ’., :. - --,. . . .. . “,

●

.

NACA TN 2229 33

.

.‘ q.

Figure 11.- Continued.

.. . . . ... .. . . ------- -—- . .,.-.

-.— —. .-. ,_—_ . ..—,. -.-Y—- -=. .-T-. —-- — ————. - .—- ;.

.. .

:. .-. . . . . . . . -... , . .

34

.

M.ACATM

J \

c. o Qf% %+ @ @- ~ - e %&* * *. - - -. . . . . .,

.-J

End plate.5

0 Uain wing0

,a -~—

4

~ .+

e

e— —3

CD

32 ~

24

b /6%

$8

0

e

.Q .

Figure 11.- Continued.

.. .... . . . . .=-— - . ---- . . .--:, ,. . . . .- . .-.- “.’”..-..- .... . .. . . . .. . . .. . . . . .-, ,

.

2229

,

.

.

NACA TN 2229 35

,.

.

.

c

. . . . . . ... .—--

End plate

‘-+-

l?tgure11.- Concluded.

. ..-_ ..-— — —~- -- -.--y-—————- ..--— --, . :- —~---- “-=-7-- —-,. .,- ---- . . ..-. ” .. . . . ,.

36 NACA TN 2229

c,/?P

4

?-

, -Plain wing/

o —/

~ ._ /“.

—- --- .. / ~

im?

:004

-.006

.-

1 I .l’”l J-

.

t II I II I---

- _

,.

.002/ ‘— -’ \

// .

/ d-- -x. \

I I I It 1 1 I I I 1 I 11~o/o _@ I I

.02

-$ 72 0 .2 “$ .6 .8 LOLift mefilicient, CL

CY’o

mgure 12.- Lateral-stability parameters in pitch of the 45° sweptback wingof aspect ratio 2 with and without various end plates.

I

,

.

.,. . . ., .-. .-, .- .-... . .-... --.:: .. .( . ,. !., -.: ...’....” ---- -,;-.. ,..

,

.

wwmv 2229 37

c%$

.

0

:002

:004

7006

.006

.004

.002

0

:002

loin wing ,/ p/

— — — — - — \ .— /“.- - -- .~ -- - -- - I

xlI l-l \l I I I I I I I

‘1--d .La.I

II

. 1

‘

.

.-$ ,.2 0 .2- 4 .6 .8 10” .Lift @?ffkh?nt,cL

Figure 12.- Continued..

------- . . . . . ..-= - —-- . .. .. -,—---- —--—-. —.—.“ -- Y--—--- —.—. .- —.-. .-. . . . . . . . .,

. . ..-. . ..” ‘,.

38’. WA TN 2229.

.

-4? ~20.24.6.8-i.0Lift coefficient ~

Figure 12.- Continued.

.

. . . ..— ---- >--- ——— ..— —-.. . . . - . . - ..- -.>“. ..”..” . ... . ... . ‘.’

NACA TN 2229 39

.02 Cyv

o

I

-$ ;2 O .2 $ .6 .8 /0

Lift coefficient,CL

l?igure12.- Ccmttied.

.-, --—- -— -—-- —, ——— -— ,-----. —. -. ..— — ---- - .—. . ... .. . . . . . . .,. .,- --”.,,:-,. -,, . . . ,,”.

40

.04

.02 Cyy

o

FiRure 12.- Concluded.

\— —------ - . ..——.—,- . . . ..--7.,=--,--—. - --- . . .., . . .---’”: ,..

0

.

;

m

NACA TN 2229

.

.

.02’ ~ -/u —

.0/

Ct o

-01

I1 I 1 I I I ( 1 1 I I I I

%&znixII I I I I I I I — I\- /[

.UG

-8

.gure 13. -

-4 “O 4 8 /2 16 20 - 24 28m, deg

41

l?lap-t~e-aileron characteristics of the 45° sweptback wing ofaspect ratio 2 with and without various end plates.

\

----- .-. .—-. ---- --—.. .— _— —— .- ..——i.— ---.-=----- — . .- —.. -—— —.. ---— ..-. . . .. .

42 WA ~ 2229

.

.0/. . ..- .

-.0/

~ /0El -10 RtzYhwing13 /0Q –/0 w.a /0

4 -10 m

.02

.

.0/, ~.

‘2 .0

-.02 ‘

-8 + O 4 8 I’2 /6 20 24 28~, deg

Figure 13. - Continued. ‘

#-

. . . . .. . -.

Mm m 2229 43

.0/

% o

-.0/

.02

.0/

-.0/

-.02

.

. - .

,

.-.

W&37 (c)

-8 + O 4 8 /2 /6 20 24 28cc, deg

Figure 13. - Continued.

--- - --- . . - -...--—— .. —.--— —.—— —._!,. .

,. - —--— — — --.—— .. —.— — —-. .-. — .

. . . . ..“

44 N+A TN 2229

C*

o

-.0/

.0/

-.0/

-.C22

A d

I.

Cz

-8-40 ”48 /2/6 20 24 28tz, deg

Figure 13. - Continued.

.

‘.

.

*

—--, . . . .. ..-— — . ... - . . . . -—,----- -~..-. . ..- .,.-. . ,. .:.. . .. . .. . :. ,’.,. . . .

.

,-

NACA TN 2229

,

45

.

.0/

80(deg) E~~@ute

-.0/~ /0 /?/uh wing

0/0 B

:;Z3’ – – - – –

-.02 ‘“-8 -4 0 4 8 12 16 20 24 28

CC, u’eg.

Figure 13. - Continued.

.

. .

#

..

. . . . .. .... ---— —,—.--- .-—- 7—. —- ----- .— — —.. -— —-- . . _— —-— — -- —.- .—. ——-. -—-.

.,.,

“f.

NACA TN 2229

\

.-.

Go

-=0/

o

-.02

..-

)

.

-8 =4 O 4 8 /2 /6 20 24 28m, deg “

Figure 13. -

.

Continued.

.

.,-. .---— --n- --~ --- ., -----,.....-.,, .-. .-. . ... . .. . . .

mm m 2229 47 “

.0/

Cn o

-0/

oq

-42/

-.02-8 ‘4 O 4 8 f2 /6 20 24 28

C, u’eg

~gure 13. - Concluded.

.. .--— ------- . . . . . ... -— ------ .— - -—-— ---,— -- —-..—— —--- -—___

48 WA TN 2229

.02

D/

-D/

.02

.0/

o8 + O 4 8 ./’2 16 20 24 28

cc,&g

Figure 14. - Effect of an eqd plate on the rolling~ and yawing-moment “coefficients of the 45° sweptback wing of aspect ratio 2 with steppedspoiler ailerons.

.

NACA TN 2229. .

b

●02

FL

J *Z ,3 .4 ,5AJ

I

b“

(a) A = 2; A = 45°; ~ = 1.0..

.— — Thei9reticdl (%7?enience

Ac~d

o

.0/

Q .

0

Figure 150- The incrementaland 4 with

, / .2 ,3 ,4 ,5h’7

=s=

(b) A= 4; A= 46.7°; 1= 0.6.

lift-curve slopes ofvarious end plates.

the wings of aspect ratios 2CL = O.

.

-.. —---- --- -—. -—.= —- ....-. — - -,._. —--—— .-—. -e . ..—. . . . . . . .. . .

,.

..-. —

.“. ,. . .. . .- .,. - ..

,

I

.!

.

,,

A

. ,,

o ./ 2!

Figure 16. - The ticremental values of maxb.uu MN coefficient of the

45° aweptback wing or aspect ratio 2.

I

mm TM 2229c

51

.01“

(2a&o -

ACD

-.O/it

‘=0.//9

:02“o 2

CL CL

(a) A = 2; A = 45°; A =1.0..

.0/ACD

o .2 ,4 .6CL

— — - Esthxzted(refereflce 1>

.n Experimental

(b) A = 4; A = 46. 7°;

Figure 17. - The incremental drag coefficientsplates.

h = 0.6.

of the wings with various end “

. .. ..—.. ---- ...-.. .———.-— , .. -.—.———-— ——-——----- . — .- .- —————-— ..-—— .-——-- —, .. . . . .. . . . . .

.

ulN

,,

. “i,.,4

Figure 18. - The maximum lift-drag ratios”aspect ratio

or the 45° aweptback wing of2.

.

.

.

.

L1

. .

.!IIti

‘;,:I

.1. .

.1.j

:,/. . .:, !

,..

-i:

‘{

I -,i /

se ~

/ Flgura 19.-()

Lateral-atabilitypmmnetera Cl* cL=O /

ad %* % “f “

I45° t3we@back wing of aspect ratio 2 with various end.plates. ‘

,

.

.0//?

-,008 ‘ . .’ .----------..- . . ..———.

.

,,!

.,

i

“ji

‘.

‘.,1

..,i

‘,/

““1‘1.1,,

i

1

I

I

I

I

o

.0(22

.004

. -ax?

-.00

412

.04

cY# .02

J-11-H,,,--0 _- 1 I I I I I I I I I I I I I0 .1.2,34.5.6 ,r-ea.gl~

s.#s “

Figure21. - Lateral-8 tabllity parameter Cn+ * %) of the 45° Ewept-

bati WIX Of aepect ratio 2 with various end plates. CL . O.~.ulUI

56

cYp

NACA TM 2229.

.

.002am\ \\

o\

\-.002‘— — — — — — ~ — — — — —

\\ ‘\ .

\\-.W4 \\\

\

.04’ .

.02 -,

// /,--/=s=

“O ./ .2 .3 .4 .5 .6

/se s#

..

,

d

.

IHgure 22. - Lateral-stability parameters C~

and ~+

of,the 46.70 swept-

back wing of”aspect ratio 4 with various en~ p@tes. CL = 0.5.

. . . ,. ...7- .---— . .,“. .-:.,. . ... .. . “.’. .,, .“

. .

I

,:(I1

.i

t.i

I.{,,1

.1‘/

f?,a

.0m4

o

-.0004

‘.00/6

-.00/2

70008

-0004

0’

0.J.2.3, ”# .5.6 .id3.9k0

S*/s

Fip 23. - Parameters $a and Clb of the 45° m?eptback wing of.a

aspect ratio 2 with flap-t~ ailerons and with variom end plates. .

a E OO.ul-1

1 ,.

I

,,

‘1...

: .’

i

“1.,.1

‘1,, -/ 6 -12

Figure d+. - Effect of’

aileron deflection

-8 -4 0

$0

two end-plate configumtions

on the 35° mwsptback wing of

4 8 12

on c1 resulting from

aspect ratio 3. w 6.5°.

.%’

.

\

NACA TN 2229 59 ‘

.

08

.04

/@2v o

.04

‘ -.08~

P%”

late .08

7-- -––– Plain hg

-“I— I&l -n” r .(J4a= 3’5°

0.

// .. . .

.08— — — — — — — — — — — — — — -08.

.6.04

0

-.04

/, 0-

= 0.3 0P-

.

.

.

/pb2v

\

.

-16 -12 -8 -4 0 4 8 12

Figuw 25. - Effect of two end-plate configurations on the variation of pb/2Vwith aileron deflectionof the 35° sweptback wing of aspect ratio 3.

.. . . .. ----- .. ..------ ~.. — .— ----- —-.-,-.— ~ . ..—— —... — --.—-.—— -— —-—- . .. ..- . ..” . .. ,,.

60

.

,008

.0Q6

.004

NACA TN 2229

.

.06!!

o0

~Ptain wing

-- -.—~-—

/,

/ ev

!

z 4 6 8

IHV 26.- Effect of two end-plate configurations

35° -Ptback wing of aspect ratio 3 in the low angle-of-attack

.

NAcA-Langl~-11-X-W-1000

--.- —---------., ..-.-..-”.:-....... ......... ‘--- .“- “ -’.-T--”—.. -.””.. ----“..