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RESEARCH MEMORANDU~
COLLECTION AND SUMMARY OF
FLAP-TYPE-AILERON ROLLING-EFFECTIVENESS DATA AT ZERO
LIFT AS DETERMINED BY ROCKET-POWERED MODEL TESTS AT MACE
NUMBERS BETWEEN 0.6 AND 1.6
By H. Kurt Strass, Emily W. Stephens, E. M. Fields,and Eugene D. Schult
Langley Aeronautical LaboratoryLangley Field, Va.
NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS
● WASHINGTONSeptember 2, 1955
L_____ . -.. . –_ ---— .-..— --- .- .= .. . .. -, —..!— - --- - .— -- .— ----
IT
,&
NACARIL55F14
NATIONAL ADVISORY COMMITTEE FOR
RESEARCHMEMORANDUM
COLLECTION AND SUMMARY
TECHLIBRARYKAFB,NM
Illlllllu!lu!llulullfllllAERONAUTICS
OF
I?LW-TYPE-AIIJ3RONROILJNG-III?17T(XNINWISDATAAT ‘ZERO
LII?J2AS D~ BY RO=-POWERED MODELTESTSAT MACH
NUMBERSBETKEEN0.6AND 1.6
By H. Kurt Strass, lhily W. Stephens, E. M. Fields,and Eugene D. Schult
A collection and summaryeffectiveness data which have
have been made of the wing-aileron roU.ing-been obtained as a part of a general inves-
tigation of lateral control being conducted by the Langley PilotlessAircraft Research Division qtilizing rocket-powered test vehicles in freeflight over a range of Mach number from 0.6 to I-.6.Some effects oftrailing-edge angle, aileron-chord ratio, aileron span and location,aspect ratio, wing sweepback, and wing-tail interference are presented.Rough design charts have been prepared to show some effects of ailerontrailing-edge angle at two sweepback angles, aileron-chord ratio, wingaspect ratio, and spanwise extent and location of aileron. These roughdesign charts have been prepsred for use in the preliminary design stage,and esthates from these charts were in fair agreement with measuredrocket-mmiel data for several configurations simulating existing or pro-posed aircraft wing-aileron combinations.
INTRODUCTION
A general investigation of lateral control at transonic and super-sonic speeds is being conducted by the Langley Pilotless Aircraft ResearchDivision utilizihg rocket-powered test vehicles in free flight. The firstsuccessful roll test was achieved in May 1946and since that time a largenumber of successful test vehicles, comprising a variety of wing-controlconfigurations,have been flown at thp [email protected] Aircraft ResearchStation at Wallops Islandj Va. It is the purpose of this report to collectand swmar ize the rigid-wing flap-type aileron data obtained from thesetests under one cover in order to aid in the design of aircraft intended
- —-- .---—- -—. –- -. - --— -. --— ----- —-— -—..——- ._ —-..— .—-— . . ——_ .—..--..
. .— —
2 NACARML55F14
to be flown at transonic and supersonic speeds. A major portion of these .data has been reported in references 1 to 33. The data for a few models(see refs. 1 to 4 and 6 to 10) were obtained during the earlier phasesof the development of the testing techniqpe, and although they werebelieved to be reliable as to trends and magnitude, they do not meet thestandards of accuracy of the present report and are not included. Datafor only the solid-steel-wingmcdel (ref. 19) have been tdcluded sincethe other models had such large aeroelastic corrections as to make thecalculated rigid-wing results questionable.
It should be noted that the data in the present paper represe~t arigid-wing tailless (exceptwhere noted) configuration in essentiallysteady-state roll at zero lift and zero yaw, with each aileron differ-entially deflected 3° to P. No data concerning the variation of rollingeffectiveness with aileron deflection or wing stiffness are shown.
.
U the following sections an attempt has been made to separate the“effectsof the major geometric variables of the wing and aileron onrolling effectiveness. Except for the effects of trailing-edge angle,the data obtained by varying a major gecnnetricparameter are first pre-sented as rolling effectivenessplotted against Mach number and thencross-plotted against the major gecmetric variable for sevend Mach num-bers. These cross plots maybe considered rough design charts shotingsome effects on rolling effectiveness of aileron trailing-edge angle attwo wing sweepback angles, hileron-chord ratio, wing aspect ratio, andaileron spanwise extent and location. AU the aforementiotid cross plotsunavoidably contain sane effects of trailing-edge single.
,
The design charts have been used to esttite the rolling effectivenessof an assortment of wing-aileron combinations simulating the wing-aileroncombinations (without fixed tails) of some existing or proposed aircraft.These estimates have been compared with measured data fram rocket-modeltests to indicate the applicability of the design charts for prelhlnarydesign purposes.
The geometry, pertfient wing-control Parameters, ~d sources of pub-lished data for each of the test vehicles are listed in table I.
.-.,An index to-the figures which shows the effects of the major geometric
variables on-the rolling effectiveness.has
SYMBOIS
A aspect ratio, b2/S
b diameter of circle swept by wing
k.~
been included as ta~le II.
tips, ft
—— —-—— —
NACARML551?I-4 3
“
rolling-mment coefficient
c
z mean
h
chord, parallel-to.model
exposed wing
thickness at
()pb/2v5KA = A
()
pb/2v
5A=3.7
KC=
%=
L
M
P
(),pb/2Vb
Ca c
()pb/ZVb
ICac=1
()pb/2V6part-span
center line, ft
cr+~chord, ,ft
2
trailing edge, ft
aileron
I —1\ b /m-spm aileron
( 1)cl 5part-span control
( /)Cz 6full-span control
le@”h of model fuselage, 4.58 ft
Mach number
rolling velocity,radians/see
-c pressure”,
Reynolds number
positive when right wing is moving downward,
lb/sq ft.’
—..— —— . .—. .—- — ——. -—— ~ ——— --. — ...— . . ..— . --———-—— —---- — —-——
—
4
s
t
v
x
Y
pb/2V
@/2v5
area of two wings to model center Line, sq ft
wing local InaMmum thickness, ft
velocity, ft/sec
distance from model
distance frcm mdel
nose to quarter-chord
center line, measumddicular to maiel center line,
rolling-effectivenessparameter
rolling-effectivenesspara&ter
l.b
(wing-tip
NAC!AM L55F1.4
point of ‘5,ft
in a plane perpen-
helix angle), radians
(not b(pb/2V)
a5 ), radians/deg
average aileron deflection for one wing in a plane parallelto model center line, positive for trailing edge up on right*, deg
Y_—n- b/2
A angle of.sweepback of quarter-chord line, except where noted,deg
A ratio of tip chord to chord at model center line
9 trailing-edge angle, defined as the angle measumd in a planeparallel to model center line between straight lines drawnbetween 0.97 chord and 1.00 chord on upper and lowersurfaces, deg
Subscripts:
a aileron
i inboard
o outboard
R rigid-wing rolling effectiveness
r wing-fuseWge intersection
t wing tip
.
.
— —-. . .— —
NACA FM L55F14 5
MODELS
A general arrangement of typical test vehicles is illustrated infigure 1. The geometry, pertinent wing-control parameters, and sourcesof the published data for each of the test vehicles are Listed in table I.More detailed information regarding most of the models can be obtainedfrom the appropriate references. The airfoil sections for all models,except where specifically noted in table I, were taken in a directionparallel to the model center line. All the models had constant percent-chord ailerons. It may be noted in table I that in some cases thetrailing-edge angle varies between models having the same airfoil section.This variation generally arises from the hand-finishing operations nearthe extreme traillng edge during model construction.
A few of the models (2, 4, 6, 8, 83, 85, 89, and 93) having a sire- .plified construction were used in order to provide additional data onthe effects of wing aspect ratio and aileron-chord ratio. A typicalmodel of this series (model 89) is shown in figure 2. The wings of thisseries were made of l/2-inch-thick aluminum alloy and had wedge-shapedleading edges exteting to 0.20 chord. The dark spanwise strip near thetrailing edge is the filled-in slot along the aileron hinge line, theslot having been cut to alluw bending the aileron to the desired deflec-tion. Welded fittings were used to attach the wings to the rocket-motorcase which also served as the fuselage. A tape~d sleeve was used as afairing behind the standm.d spinsonde head.
In the investigation of the effects of wing location and number ofwings on rolling effectiveness, the wings were located at a more forwardposition on the body and sane of these models had a free-to-roll tailfor directional stability. Each free-to-roll tail consisted of four finsand two ball-bearing asseniblieslubricated with a special tide-temperature-range silicone grease. The photographs presented in figure 3 are typicalof the sweptback configurations.
TEST TECHNIQUE
The flight tests were made at the Iangley Pilotless Aircraft ResearchStation at Wallops Island, Va. “Thetest vehicles were propelled by a two-stage rocket-propulsion system to a ~um Mach number of approximately1.6 to I-.8in about 3 seconds. During the following 10 to 20 seconds ofcoasting flight, time histories of the rolling velocity measured in zero-lift flight were obtained with special radio equipment (designatedspinsonde; see ref. 34),the flight-path velocity was obtained throughthe use of CllDoppler radar, and the model space coordinates were obtained
.-\.
— -— ..—.—--—-— .— ..—. ——_— _ ——— - — .. ——— — —..——
6 NAM RML55F14
through the use of modified SCR @l tracking radar. These data, in con- “junction with atmospheric data obtained with radiosondes, petit theevaluation of the rolling effectiveness in terms of the parameter pb/2Vas a function of lhch number. The variation of Reynolds numbed and-c pressure With Mach number is presented in
ACCURACY AND CORRECTIONS
General Discussion
During the course of a general investigationwhich has been conducted by the Ian@ey Pilotless
figure4.
of rolling effectivenessAircraft Research
Division for the past 8 years, maq-changes have been made in the designand construction of the test vehicles and in the testing technique inorder to @rove the accuracy and reliability of the data. As a result,the data have been corrected to a standardized set of conditions to allowdirect comparison of the data obtained at various stages in the evolution -of the present techniqpe.
The rolling-effectivenessdata were obtained under essentially zero-lift conditions and have been corrected to rigid-wing values and are pre-
sented in terms of the param&ter ‘b/a, where pb/2V results from the8
b(pb/2V)aileron deflection 8 and should not & confused with
ab “
Accuracy
The following factors must be considered in the assessment of the ‘overall probable accuracy of the data presented here.
(1) The accuracy of measurement of pb/2V at a given Mach nudberfor any given test model is dependent upon the following values:
M. . . . . . . . . . . . . . . . . . . . . . . . . .p,radians/sec . . . . . . . . . . . . . . . . . . .V,ft/sec. . . . . . . . . . . . . 1. . . . . . . .b/2. . . . . . . . . . . . . . . . . . . . . ... . .
The maximm probable error in pb/2V from these
. . . . . . *o. 01
. . . . . . *1.5
. . . . . . *5.ONegligible error
sources is estimatedto be *0.0020 at subsonic speeds and M.OO1O at supersonic speeds.
(2) The systematic errors caused by deviations from the desiredmodel geometv, a result of constructional.tolerances which can alter
f . —. ——_ ._
NACARML55F14 7
the roll effectiveness, normally are limited to variations in ailerondeflection and wing alinement. The method of model measumne nt used iscapable of measuring the angular deviation to within approximately
* 0“’3 degrees per foot of ~ or aileron chord.chord
The accuracy of meas-
urement for a typical model (model 28) can be illustrated as follows:
5 (average of models 28a, b, and c) = 5.22°* O.OvO
iu (not published, average of models 28a, b, and c) .=0.02° * 0.0180 ;
where im is the average angle of wing misalinement (differential’inci-
dence), positive when tending to roll the model in a clockwise directionas seen from the rear, and ‘isbased upon distance from’leading edge toaileron hinge line (O.472 foot).
Corrections
Incidence.- The data were,corrected for deviations in wing incidencefrom the nominal value of ~ = 0° by use of the following equation
which was derived by using very simple aerodynamic assumptions:
Pb _ Ziu) 1-I-2A—_—m 57.31 + 3A
The validity of this correction was demonstrated in reference 21 whereinit is shown that this simple formula provides good estimates of pb/2Vresulting from differential incidence for a wide range of wing plan forms.It is estimated from the data published in reference 21 and additionalunpublished data that the probable accuracy of prediction of this formulais within *15 percent for most configurations. (Relativelythick unsweptwings, NACA 65Ao09,show an abrupt discontinuity in the variation of pb/2Vwith M at M = 0.92 which is not predicted by the theory.)
Aileron deflection.- Corrections for deviations in 5 were madep’b/2v
simply by presenting the data as —.5
Pb/2v for m~el 28 restitbgAs an example, the probable errors in —8
from the previously mentioned limitations are tabulated as follows:
.—. .— . —.— .. ..— — —— .— —.---. — —--
8
pb/2VSources of probable error in —
5pll/2v
M 56
. ‘@ ‘gnominal ma
$;Total
Measuremen~ Calculated correction(a)
).8 0.02040 w. 00036 to.00028 to.00009 *O.(jQtmI *O.00074
..40.00510w. 00018*o.00007 w. 00009 HI.Oooo1 m. 00035
aEstimated at *15 percent of theoretical correction.
Aeroelasticity corrections.- It was necesssryto correct all thedata for the effects of aeroelasticity, and the Mge number of mdelswhich were tested precluded the use of very refined methods of aeroelasticanalysis. For this reason, a special engineering methcd was developedand is presented in reference 25. The probable errors in the values
of pb/2V
6resulting from t% application of this method are very difficult
to assess and are dependent upon a large nmber of ~ia~~s. u~ess
otherwise specified, it is believed that errors from this source are neg-ligible as the test wings inmost instances were very stiff and neededa relative~ small.correction.
Effect of model roll inertia.- For one-degree-of-freedomconfigura-tions such as the rolling-effectivenessmodels of this report, the equa-tion governing their behavior is
% dpc%%
+cz6b=— —qS’b dt
or
()pb “
()
= pb Ix ~p—
2V steady state G mess - CzpqS’b ~
.)
,,
.
,
.
. —.— .——— — _— — —
2TNACARML55F14
where
9
Rolling moment
qS‘b
dCz%p = -d(pb/2V)
dC~c1 =%5
1- measured model moment of inertia about ro~’ axis, slug-ft2
A
s’ area of three wings to
t time, sec
See the section on “SYMBOLS” for
model center line, ft2
additional defWtions .
The data for a“model with lxmge rolling accelerations (model 2i’a)have been corrected for roll inertia effects and the results are shown in
figure 5. Physical constsnts for the model were Ix = 0.0697 slug-ft2,
b= 2.18ft, ~d s’ = 1.93ft2; and CZ values for this configurationP
were obtained fran reference 35. The midnmn rollinn acceleration for
the model was 175 raifians/sec2.
Figure 5 shows that the clifferences between measured rolling effac-tiveness d steady-state rolling effectiveness are negligible for thismodel despite the very large values of rolling acceleration. Very fewof the models in this report have rolling-accelerationvalues evenapproaching those for model 27a, and no inertia corrections have beenmade to any of the data presented h this report.
DISCUSSION OF RM3ULTS
Effect of Trailing-Edge Angle
General ccmmwnts.- Aileron rolling effectiveness is affected by thecontour of the entire wing, and particularly the contour over the aileron.
—. . -.. ..— —..—— .— .. . . . -—-. —.-. _ -— —.- —.— — — .. . ——— ——-—.-
10 NACAI@l L51?14
Thus, two ailerons cliffering widely in contour but having the sametrailing-edge angles may have different values of rolling effectiveness.
It was shown in reference 20 that thq rolling effectiveness ofuntapered wings tith 0° and 45° sweepback and employing full-exposed-spanailerons could be correlated as a function of the trailing-edge angle fora wide rsmge of airfoil profiles and thiclmesses. b general, the corre-lation was good but the scatter of the data indicated that the traiUng-edge angle was not the only variable; however, it was obvious that itwas a very imporlqnt.factor. Since that initial effort, several attemptshave been made to improve the correlation by making-use of the transoticsimilarity laws. The pb/2V values were plotted against parameters con-taining various combinations of airfoil thickness ratio and traiMng-edge angle at constant Mach nuniberand constant parameters containingMach number, airfoil thiclmess ratio, and trailing-edge angle in com-bination. None of these attempts provided any marked hprovement overthe original correlation against trailing-edge angle in reference 20 and,in addition, they were much more complicated to use.
Figure 6 presents the rolling-effectivenessdata correlated againsttrail$ng-edge angle and includes some additional data not available atthe the of publication of reference 20. In addition, an improved methodof correcting for the effects of aeroelasticity was used (ref. 25) whichprimarily affected the data for the sweptback wings.
Wswept wimgs.- For unswept wings it is apparent that trailing-edgesngle has the ~eatest effect in the speed rage between M s 0.8 toM = 1.2. At M = 0.7 _ M = 1.2 the rolling effectiveness does notVarY ~hu with q. In the trsnsonic range there is considerablescatter and the only clear indication is that a small trailing-edge angle(appro-tely 7’Oor less) maintains rolling effactiveness throughout thespeed range and that larger traillng-edge angles exhibit varying smountsof rolling-effectivenessloss or even reversal of control. There is anindication, although not shown as such in figure 6, that” the 6-percent-thick wings appe~ to have gener~ higher ro~llinneffectiveness thanthicker wings of about the same traiUmg-edge angle. “b view”of thescatter of the data and the relatively small nuniberof thiclmess ratiosfor comparison at a given traiUng-edge angle, it is not clear whethera clifferentiation between various thiclmess ratios should be made, andso the trailing-edge angle has been the sole variable considered in fairingthe data points.
Sweptback wings.- For wings swept back 45°,,there was.generaddy lessscatter than for the unswept wings and the effects of trail&g-edge anglewere evident at a lower subsonic Mach number. No direct data are availablefor very low trailing-edge angles on a wing having the aspect ratio of thewings used in this correlation (A = 3.71), but flat-plate data (q = 0°)for aspect ratios of 2.31 and 8.00 (models 6 and 89) have been interpolated
.
,,
NACA RM L55F14 u
to provide an estimated end point toto q = OO. The faired curve in theas a dashed line to indicate that itthe data.
aid in fairingregion nesr qis essentially
the measured data= 0° is presentedan extrapolation of
Etfect of airfoil thicbess ratio.- Although it is evident thattrailing-edge angle is a major factor in determining the level of roltigeffectiveness for an aileron, it is also of interest to note the effectsof changing the thickness ratio for a given family of airfoils. Such achsnge necessarily involves a corresponding change in trailing-edge angleso that it is not possible to determine the effects of thickness ratiodivorced from the effects of trailing-edge angle without altering”thebasic profile of the family. Figure 7 presents a comparison of therolling-effectivenessdata for wings having several thickness ratios,aspect ratios, and pbll forms.
Concluding comment.- The preceding discussion indicates that thetrailing-edge angle of an aileron is a parameter of prime importance andmost of the undesirable characteristics of increased trajhg-edge angle,such as abrupt changes in pb/2V with M and unusually large losses hrolling effectiveness, can be avoided by employing ailerons with trailing-edge snglesof q~w for A= 00 wings and qS12° for A=45° wings.In actuality, most of the high-speed aircraft which are designed to flyat the speeds where large trailing-edge angles are to be avoided shouldexperience little or no trouble from this soqrce-because drag consider-ations preclude tm use of thick wings (which would normilly have largetrailing-edge angles).
Effect of Wing Sweepback
Figure 8 presents some effects of wing sweepback on aileron rolllingeffectiveness for a variety of test configurations employing full-exposed-span ailerons.
At subsonic speeds, increased sweepback generaldy resulted indecreased rollinn effectiveness.
At transonic speeds, the most significant effect of increased sweep-
Pb/2v ~th l&ch num-back was the smoothing effect on the variation of ~
ber. The rolling-effectiveness“bucket,” characteristic of unswept wingsat transonic speeds with moderate traikhg-edge @es, was virtuallyeliminated as the sweepback was increased to 45° and disappeared can-pletely at 60° sweepback.
. . —..-—..-. ..__. _—.—. —.— -—.-. . —— -- . ...— — .. .. -:-.——. ..— ——
.
12 NACARML55F14
.At supersonic speeds, increased sweepback did not cause a consistent
‘b/2v with change in sweepback.angle; however, the generalvariation of —5
tendency was a decreased rolling effectiveness with increased sweepback “angle.
Effect of Aileron Chord Ratio
Same effects of aileron chord upon rolling effectiveness axe shownin figure 9 for a wide range of wing plan forms and airfoil sections.Unless otherwise specified, the following discussion pertains to full-exposed-span ailerons.
NACA 65AO09 ahfoil secti.ons.- Figure 9(a) presents the effect of
aileron chord as measured on unswept wings. The variation of pb/2v6
with Mach number for all the aileron configurations is characterizedbyan abrupt dip near M s 0.9. This is a wing dropping phenanenon whichon the basis of past experience is restricted primsrily to unswept wingsemplo@ng trailing-edge angles greater than q ~ P and thickness ratios~eater than t/c s 0.06. (See refs. 20 and 36.) Aerodynamic controlreversal.was measured for the O.1-chord ailerons in this”region. Itshould be noted that a simil&r reversal of rolling effectiveness has beenobtained for full-choti ailerons at a very small angle of deflection.
(See ref. 23..) The VELLuesof ‘~ presented in figure 9 were obtained
from mcdels on which the ailerons were deflected approximately 5°. Other“tests have shown that aerodynamic reversal.maybe eliminated by increasingthe aileron deflection. (See ref. 10, for example.)
S3milar data for 45° sweptback wings are presented in figure 9(b).
‘b’*v with Ca c is similar to that experienced byThe variation of —5 /
the unswept wings.
mm 65Ao06 airfoi3_ sections.- The effect of aileron chord as meas-ured on tapered, sweptback wings is presented in figures 9(c) to 9(f) foroutboard ailerons of various spanwise etients. With the exception of thefull-exposed-span ailerons, the data for the full-chord ailerons (figs. 9(d),9(e), and 9(f)) were obtained for ailerons of different spanwise extentsthan the psrtial-chord ailerons.
The 0.15-chord ailerons were approximately 60 percent as effectiveas the 0.30-chord ailerons for the two aileron configurations of greaterspan (figs. 9(c) and 9(d)) but became relatively more effective for thesmaller spans (figs. 9(e) md 9(f)).
.
r.
.-——— .—.
NACA RM L751?14 13
It may be of interest to note that, with the exception of the abruptdip in effectiveness at MsO.9 for the unswept wings (fig. 9(a)), mthe full-chord ailerons exhibited very little vsxiation of effectivenesswith Jkch number.
Flat plate airfoil sections.- Figures 9(g) to 9(j) present some datawhich show the effects of aileron chord upon rolling effectiveness forseveral wings all of which employed flat-plate airfoil sections (q = OO).The effect of aileron chord was approximately the ssme for all the modelsin that the 0.2-chord ailerons were appro-tely 75 to 85 percent aseffective as the 0.4-chord ailerons at sfisonic speeds with the relativeeffectiveness decreasing with increasing Mach number until at M = 1.6,the 0.2-chord ailerons were approxhnately 50 percent as effective as the0.4-chord ailerons.
Correlation of data.- Where data me available for both part-chordand full-chord controls of the same span (figs. 9(a) to 9(f)), the part-
chord ~ ~vebeen di~idedb yt~f~-chord ~ and plotteds 6
as ~ at the appropriate ca c/
value in figure 10 to illustrate non-
dimensionaUy the effects of aileron-chord ratio on rolUiqg effectiveness.Because of the scatter and the relatively small nuder of tests available,only the theoretical two-dimensional curves for thin plates are shownfor comparison with the test ~oints.
The O.1-chord ailerons on the unswept wings exhibited control reversal.at M = 0.9 and zero effectiveness at M = 0.93. At M = 0.96, no con-trol reversal was observed but all of the unswept wing-aileron cotiigu-rations were appreciably less effective than the comparable swept-wingmodels. Figure 6 shows that the effectiveness of 0.2-chord ailerons canbe ~eatly increased in this speed rsmge by recourse to smaller trailinn-edge angles. Although no direct evidence is available for O.1-chord con-trols, there is no reason to suspect that a similar improvement could notbe achieved on these controls by the use of smaH. trailing-edge angles.It shouldbe pointed out that swept wings with part-chord ailerons exhibit
low ‘w values for large trailing-edge angles in this ssme speed range.
(See re;. 20.)
Effect of Aspect Ratio
The effect of wing aspect ratio upon the rolling effectiveness offull-span ailerons is illustrated in figure II for untapered wings havingsweepback angles of 0°, 45°, and 600 dnd several airfoil sections. Ond.ytwo configurations (fig. n(a)) have been tested which show the effectof aspect ratio upon 600 sweptback wings; these data do not conform with
,,
.—.—. ..—.— ——. . ..-. .— ——z. — —.
—
14
the trend shown by the wings of lower sweep but
ratio caused an increase in - at subsonic5
decrease at supersonic speeds.
NACA RM L55F14
show that increased aspect -
speeds and a very slight
Correction of data.- lh order to illustrate better the effects ofaspect ratio, the data of figure 11 have been nomalized as a fraction ofthe A = 3.7 values, and the resulting values of KAJ the normald.zed
aspect-ratio factor, are cross-plotted agatit aspect ratio in fi~e 1-2.It is obvious that, although the general trend of the data is for therolling effectiveness to decrease with increasing aspect ratio, consid-erable variation exists in the rate of change of the variation as isevidenced by comparison of the faired curves of the various testconfigurations.
At Mach nun?hersof 0.9 and greater, the effect of aspect ratiodepends upon the configuration. For exsmple, at M = 0.9, the rollingeffectiveness-of the unswept wings decreased rapidly with increasingaspect ratio, whereas the 45° sweptback wings showed little effect ofaspect ratio. The aileron traihg-edge angle had considerable effectin that
changes
No
because
for a given sweepback angle there was weater sensitivity toin aspect ratio for the larger trailing-edge angles.
plots are shown for the region between M - 0.92 and M = 0.98pb/2V
of the rapid variation of — with Mach number in this region.8
-plots of KA against aspect ratio at these speeds would at best have.
doubtful value.
General discussion.- It is evident from the foregoing discussionthat the variation of rolling effectiveness with aspect ratio is very com-plex and is dependent upon a nmber of variables of which Mach number andtrailing-edge angle are of ~eat hportance. However, it is possible togeneralize to some extent. With certain exceptions, increased aspectratio apparently causes a decrease in rollinn effectiveness. At transonicspeeds the effect is very variable and is influenced to a large extent byth trailJng-edge sngle.- Jncreased aspect ratio tends to decrease rollingeffectiveness more at supersodc than at subsonic speeds.
Effect of Aileron IOcation
.
Some effects of aileron location and spmwise extent upon aileronrollinn effectiveness are presented in figure 13 for a variety of wing-aileron configurations.
—
NACA RM L55F14 15
NACA 65AO06airfoil sections.- Figures 13(b), 13(c), and 13(d) showthe rolling effectiveness of various ailerons (includingfull-chordailerons) on sweptback wings having NACA 65AO06 airfoil sections. Gen-erally, the outboard half-exposed-span ailerons were from one-half totwo-thirds as effective as the full-exposed-span ailerons. An inbosrdaileron was generally more effective than an outboard aileron of thessme span at subsonic and transonic speeds, but the results were mixedat supersonic speeds. Although all the part-chord ailerons showeddecreased effectiveness as the speed increased, the full-chord aileronshad essentiu constant effectiveness throughout the speed range tested.
NACA 65Ao09 airfofi sections.- Figure 13(a) shows that on the unsweptwing the outboard half-exposed-span control was more effective than theinboard half-exposed-span control, whereas on the sweptback wing thereverse is true. The outboard control was generally more than one-halfas effective as the fall-exposed-span control for the unswept wing, butabout one-half as effective for the sweptback wing.
NACA 651A012 airfoil sections.- Figure 13(e) shows that the outboard
half-exposed-spareailerons were more effective than the inbo&d half-exposed-span ailerons at subsonic speeds for both taper ratios. At speedsgreater than M = 1.0, all the configurationshad poor control character-istics, particularly near M = 1.2 where the controls were either com-pletely ineffective or contr@ reversal was obsened.
Miscellaneous airfoil sections.- Several miscellaneous wing-aileronconfigurationshaving full-choti ailerons are presented ti figure 13(f).~ese data, except for model 36, show a relative~ constant ro~ effec-tiveness throughout the speed range tested.
Comparison between experimental and estimated values.- Reference 37presents a method for estimating the effect of aileron spsnwise locationon rollinn effectiveness for unswept wings having various aspect ratiosand taper ratios at 1~ mibsonic speeds, and reference 38 compares theresults shown in reference 37 with experimental data for swept wings andpresents a design chart based on the comparison. It is shown that aspectratio has little effect on the spanwise variation of rolling-moment coef-ficient and that the effect of taper ratio is not of major hportance.Figure 14 shows the design chart of reference x in nondimensional.formwhich was obtained by dividing the value of Cz/Az (equivalentto C2/8
in the notation of the present report) at any span stationby the full-Spau value.
The qpantity of rocket-model data in figures 13(a), 13(b), 13(c),13(d), and 13(f) is not considered to.be sufficient to.establish a gen-eral plot of the variation of rolling effectiveness with control spanwlselocation, so the data from figure 13 are ccmpared in figure 14 with the
‘.
-..—— .—— -- .—---
16 I?ACA& L55F14
nomalized design c= taken frcm reference 38. @ comparisons infigure 14 have been made at two representativeMach numbers and are inreasonably god a~eement; the data of figure 13(e) (A,= 8.o) have notbeen included at M = 1.4 because of the poor control characteristicsat this speed.
‘I!hedata of figure 13 were normalized according to the procedure
11.0
“ pb/2V
r
.0 ~ qi
‘%Vr 1
1.0pb/2V
8% .
where%
appearing on the right-hand side of the equation is the value
for a fuJ1-exposed-span control fram figure 14. The Vdllf2S of ~ thus
(obtained for outboard ailerons 70)
= 1.0 are plotted as data points on
figure 14 at the appropriate vaiue of vi. The experkntal ~ valuesfor ~oard ailerons are not shown in figure 14, but they agree fairlywell tith estimated values from figure 14.
lWfects of Wing-Tail titerference, Wing
Location, and N.miberof Wings.
General discussion.- All the data thus far discussed were obtainedby the me of three-winged test ve~cles which do not resemble typicalairplane configurations. Some uncertainties @st regarding the appli-cation of these data to conventional aircraft. In order to partiallyclsrify this situation, a lJmited investigation has been conducted todetermine some effects upon aileron rolling effactiveness of wing-tailinterference,wing “locationon the fuselage, and the number of wings.
Wing-tail interference.- In order to determine some effects of wing-tail interference upon the aileron rolling effactiveness at zero lift, aninvesti~ation emlofi five t-t mmls W- conducted. ~ test mode~
were co&Aructed- with two wings, instead ofappro@nate an airplane-type configurateion.
=6
the usual three, in order to(See fig. 3(b).) Two of
\
.
.
3T
NACA RM L55F14 17
these models employed free-to-roll tail assedil.ies(models 54 and 58)which provided directional stability without introducing rolling moments.Bench tests of these free-to-roll.tails, under simulated drag loads sev-eral magnitudes greater than those estimated for flight conditions, showedfriction forces corresponding to a rolling mment of 0.17 ft-lb which isnegligible when compared with the kO to 50 ft-lb dsmp~ moment of thetest wings. A more complete description of therolJ tails is given in reference 33. Figure 15with these models.
Ihbosrd ailerons: Figure 15(a) shows thatailerons resulted in a considerable decrease in
test models tith free-to-shows the results obtained
fixing the tail for @oardrollbg effectiveness.
The chmge in rolling effectiveness due to the fixed tail was approxi-mately constant throughout the speed range tested and was enough to causea slight roll reversal.at speeds geater than M = 1.3. Whether thiscondition etists for other aileron deflections and tail fore-smd-sftlocations is not known at this time. Z* the location of the fixedtail from the plane of the wings to 0.18c above the plane of the wingshad little effect on the fixed-tail rolling effectiveness.
Outboard ailerons: Figure 15(b) shows that fixing the tail did notcause my appreciable cha@e in the rolling effectiveness of the outboardaileron at all speeds for which data are available.
.General.discussion: It is-apparent from the limited data presentedh&e that the rol.lb effectiveness of inboard ailerons is markedlyaffectedly the fixed-tail assembly. Reference 33 shows that the addi-tional damping in roll due to the addition of the fixed-tail assemblycaused a decrease of about 15 percent in the rolling effectiveness atall speeds, a value which is outweighedby the large losses resultingwhen the aileron-generated downwash strikes the fixed tail. Reference 33shows that the effects of the downwash could be estimated with fairlygood accuracy by simple theoretical means.
Effect of wing location and nudber of wings.- l?i.gure16 shows theeffects on rolling effectiveness of locating the wings at a more forwsxdposition on the fuselage and reducing the nmiber of wings from three totwo. ~ the wings were untapered, had NACA 65Ao09 airfoil sections,and 0.20-chord ailerons. The effects of wing forward location and num-ber of wings were not appreciable except for the umwept,wings at sub-sonic and transonic speeds (see fig. 16(a)) where any physical modifi-cation to the standard model (three wings aft) gener~ caused a decreasein rolling e~fectiveness.
.._
. .—. .—.. ——— ..—— —. —— —— — . . . _— —.— ...- ———.—.. ..- . .
— —— ----- ___ ___—
18
datatype
ADDITIONKG TESTS
NACA ti L55F14
References 1 to 33 contain additional aileron rolling-effe.ctivenessnot considered appropriate for inclusion in the present report. Theof investigation and corresponding references are listed as follows:
!&pe of investigation
Aeroelastic effects . . . . . . . . . . . . . .
Delta wings . . . . . . . . . . . . . . . . . .Leading-edge and tra.ng-edgeailerons inccmtd-nation... . . . . . . . .
Effects of aileron chord extension . . . . . .Effects of airfoil nose shape . . . . . . . . .Horn-balanced ailerons . . . . . . . . . . . .Wing-tip ailerons . . . . . . . . . . . . . . .Effects of wing leading-edge roughness . . . .Interferencebetween ailerons andtip stores . . . . . . . . . . . . ...’..
Bellows-actuated ailerons . . . . . . . . . . .Mfects of Imllt-in wing twist . . . . . . . .
Reference
1.6 18, 19, 22, 23,24, 25, 26, 28, 32
5, 1.2,24
6101314
17, 3027
303132
SPECIFIC WING-JIIIJZRONCONFIGURATIONS
-ComparisonBetween Measured and Esthated Values
A num%er of models simulating etisting or proposed airplane wing-
.
aileron configurationswithout fixed tails did not fit handily into thebasic data plots showing the effects of aspect ratio, trailing-edge-, ~ so forthj ~d t~se speci~ models ~ve been ~cluded infigure 17. It was felt that these special mmiels would provide some.indication of the applicability of the preliminary deSig?l ChUtS (figs. 6,10, 12, and 14).
W estimating the rolling effectiveness of a given configuration, itwas assumed that the rolling effectiveness could be approximated by con-sidering only six major variables: (1) aileron trailing-edge angle, and(2) sweepback at the hinge line, figure 6; (3) ratio of aileron chord towing chord, figure 10; (4) wing aspect ratio, figure I-2;and (5) spanwiseextent, and (6) location of ~eron, figure 14. No consideration wasgiven to the effects of wing location on the body, number of wings, and
.
..
.
—
NACA RM L55F14
SO forth. No considerationwas given to theother than the use of the hinge line for thewhen using figure 6.
19
effects of wing taper ratio, .reference sweepback angle
As shown h figure 17, the agreement between the measured rollingeffectiveness smd the estimated rolling effectiveness (from figs. 6, 10,U, and 14) is fair on the whoh.
Methml Used in Estimation
For the purpose of estimating the rolling effectiveness, it isassmned that the rolling effectiveness for any configurationmay beexpressed as the rolling effectiveness for some arbitrary reference con-figuration with the proper corrections applied to account for deviationsfrcm the geometry of the reference configuration. b the present case,the reference configuration geometry is given in the legend of figure 6.
h order to indicate theeffectiveness frcm figures 6,a ssmpl.ecalculation is shown
pb/2V _5
procedure used10, U, amd 14as follows for
A = 2.8
in estimating the rollingfor a given configuration,model 99 at M = 1.00:
A= 38.5° (hingeline,calculated)
q =U.60 ‘
Ti L 0.41
~. = 0.75
>= 0.23c
()Kc
pb/2V8
ref aK’l Kcref
()pb/2Vwhere —5
is the rolliqg effectiveness for reference cotiigu-
ref
ration having 38.5° sweepback at the hinge line and ~.6° traiu-edgeangle, using straight-line interpolationfor effect of hinge-line sweepback
.- _
.. —..-. . .. . ——. -.——— _ .-— -—-. --— —-...-. -- ——— ——.. —.——
—
20..- e -- ---
L-—-—–-
()Pb/~6 = 0.01L6ref
NACA RM L55F14
(interpolatedfrom fig. 6(f))
= 1.15 (from M >1.0 theory curve, fig. 10(f))
KA = 1.12 (from A = 45° curve, fig. 12)
.
~ = 0.76-0.24
= 0.52 - (frm Am = 51.3° curve, fig. 14)
therefore,
= 0.0078
CONCLUDING REMARKS
. A collection and summary is presented of the wing-aileron rolling-effectiveness data which have been obtained with each aileron differ-entially deflected 3° to 70 during a general investigation of lateralcontrol being ccmducted by the Langley Pilotless Aircraft Research Division.Some effects of trailing-e&e angle,,fileronchord ratio, spanwise extentand location of aileron, aspect ratio, wing sweepback, sad wing-tail inter-feren~e are presented. The quantity of data which has been obtained ineach of these categories varies considerably.
———...—.—— .-——. .— —
NACA RM L55F14 a
It is felt that the effects of some psmmeters, such as ailerontrailing-edge sagle, are fairly well defined, whereas others, such aswing-tail interference, are not satisfactori~ deffied.
Rough design charts have been prepared to show 6me effects ofaileron trailing-edge angle at two sweepback amgl.es,aileron chord ratio,aileron span and spsmise locationy ~ ~ wet ratio.design charts have been prepared for use in the prelimimryand esthates have been made frcm these charts for thirteenconfigurations (without fixed tails) simulating existing orcraft wing-aileron caibinations. The esttited values werement with measmed rocket-model data.
These roughdesign stage,wing-aileronproposed air-in fair agree-
These design charts should be used with caution when the configurationhas a fixed tail assembly, since the Limited data available indicate thatin s-, cases the wing-tail interfe~nce appreciably affects the aileronrolling power.
Iamgley Aeronautical Laboratory,National Adtisory Camnittie for Aeronautics,
Ian@ey Field,.Va., June 8, 1955.
—-— ..— .—— .— .-—
—
22 I?ACARML55F14
lmFEKENcEs
1. Sands& Carl A., and Marine, Alfred A.: Free-Flight ?kr.wsti@tionof Control Effectiveness of Full-span O.2-Chord Plain Ailerons atHigh Subsonic, Transonic, and Supersonic Speeds to Determine SomeEffects of Section Thickness and Wing Sweepback. NACA RM L7D02, 1947.
2. Ssz&hl, Carl A.: Free-l?l.ightErvestigation of Control El?factivenessof Full-Span, O.2-Chord Plain Ailerons at High Subsonic, Transonic,and Supersonic Speeds To Detemine Some Ei?fectsof Wing Sweepback,Taper, Aspect Ratio, and section ThiClnlessRatio. NACA RM L7F30, 1947.
3. Sanddl, Ml A., and %.rass, H. Kurt: Additional Results in Free-Flight Investigation of Control Effactiveness of lRdl-Span, O.2-ChordPlain Ailerons at High SUbsonic, Transonic, and Supersonic Speeds ToDetezmine Some Effects of Wing Sweepback, Aqect Ratio, Taper, andSection Thickness Ratio. NACA RM L~l, 1948. “
.4. ~, Ml A.: pre~ l?ree-l?lightInvestigation of the Effect
of Airfoil Section on Aileron RoX1.innEffectiveness at Transonic andSupersonic Speeds. I?ACAm L8B26, 1948.
5. Sandabl,CarlA.: I?ree-FliLghtlhvestigation of the Rolling ECfactive-ness of Several Delta Wing-Aileron Configurations at Transonic andSupersonic E@eeds. NACA w L8D16, 1948.
6. stI’ELSS, H. Kurt: Free-Flight Investigation of the Rolling Effactive-ness at H%gh Subsonic, Transonic,‘andSupersonic Speeds of bading-Edge and Trailing-Edge Ailerons in Conjunction With Tapered andUntap-eredPlan Forms. NACA RM L8Elo,1948.
7. Stmdabl,CarlA.: Free-Flight hvestigation at Transonic and Super-sonic Speeds of the Rolling Effactiveness of a 42.7° Sweptback WingHaving Partial-span Ailerons. NACA RM @25, 1948.
8. SanddiL,CarlA.: lEree-Flighthvestigation at Transonic and Super-sonic Speeds of a Wing-Aileron Configuration Simulating theD-558-2 AirPlme . NACA RM L8E28, 1948.
9. Sand&d, CarlA.: F&ee-l?lighttivestigation at Transonic and Super-sonic Speeds of the Rolling lZYfectivenessof a Thin, Unswept WingHaving PartisJ--SpanAilerons. NACA RM L&204, 1948.
10. SandaW, Carl A.: Free-Flight Investigation at Transonic and Super-sonic Speeds of the Rolling Effectiveness of Several Aileron Con-figurations on a Tapered Wing Having 42.70 Sweepback. NACA RM L8K23,1949.
— —
NACA~ L55F14 23
U. sandabl, Carl A., Bland, Wimm M., Jr., - fi~ss, H. K~: ~fectsof Same Airfoil-Section Variations on Wing-Aihron RolJing Effective-ness and Drag as Determined in ?&ee Flight at Transonic and Super-sonic Speeds. NACA FM L9D12, 1949.
12. sadahl, Carl A., and fimss, E. K~: cq~atlve Tests Of theRo~ing Effactiveness of Constant-Chord,Fdl.-lklta, and Half-DeltaAilerons on Delta Wings at Transonic and Supersonic Speeds. NACARM L9J26, 1949.
13. Strass, H. Kurt, and Fields, Edison M.: Flight Investigation of theEffect of Thickening the Aileron Trailing Edge on Control 12ffective-ness for Sweptback Tapered Wings Having Sharp- and Round-Nose Sec-tions. NACARM L9H9, 1950.
14. Strass, H. Kurt: The Effect of SpanWise Aileron Location on theRolling Effectiveness of Wings With 0° and 45° Sweep at Subsonic,Transotic, and Superso~c Speeds. NACA RM L50A27, 1950.
15. strass, H. Kurt, Fields, E. M., and Schult, E. D.: lZree-Flightinve-stigationat Transonic and Supersonic Speeds of the Rolling Effec-tiveness of a Partial.-SpanAileron on an hversely Tapered Sweptbackwing. NACARM L50B08, ,1950.
16. Sandabl,CarlA., and Strass, H. Kurt: Rolling Effectiveness of aThin Tapered Wing l%ving Partial-Span Ailerons As IkterminedbyRocket-Powered Test Vehicles. NACARML50D17, 1950.
17. Sandshl, tilA., Strass, H. Kurt, and Pil.and,Robe- O.: The RollinnEffectiveness of Wing-Tip Ailerons As Determinedly Rocket-PoweredTest Vehicles and Linear Supersonic Theory. NACA RM L50F21-,1950.
18. Strass, H.’Kurt, Fields, E. M., E@ Purser, Paul.E.: ExperimentalDetermination of Effect of Structural Rigidity on Rolling lKfective-ness of Some Straight and Swept Wings at hkch Mmnbers From 0.7 to 1.7.NACAFlIL50G14b, 1950.
19. Strass, H. Kurt, Fields, E. M.,”and Schult, Eugene D.: Some El?fectsof Spanwise Aileron Ucation and Wing Structural Rigidity on theRolling Effectiveness of 0.3-Chord Flap-Type Ailerons on a Taperedwing Having 63° Sweepback at the Leading Edge =d NACA 64AO05 Air-foil sections. NACARM L51Dl&, 1951.
20. Fields, E. M., and fi~SS, H. K@: Free-Flight Measuramnts at I@chNumbers Frmn 0.7 to 1.6 of Some Effects of Airfoil-ThiclmessDis-tribution and ’Trailing-EdgeAngle on Aileron Rolling Effectivenessand Drag for Wings With 0° and 45° Sweepback. NACARML51G27, 1951.
——-—.— .—.— . ...— —. . ..—. — . . .— —— —- ———
— —
24 CQ NAC4RML55F14
21..S&ass, H. Kurt, and ~ley, EdwardAU-Movable Wings at Emall AnglesIYOm 0.6 to 1.6. NACA RM L51H03,
22. Schult, Eugene D., Strass, H. Kurt,
T.:‘ Rolling EYfectiveness of .
of Incidence at Mach Numbers1951.
and Fields, E. M.:Meas&ments of-Some IMfects of Aileron span; Chord,and of Wing Flexibility on the Rolling EPfactivenessSweptback Wings at Mach Nmbers Between 0.8 to 1.6.1952.
,,
l@ee-Flightand Deflectionof Ailerons onNACA RM L5w6,
23. Strass, H. Kurt: Summary of Some ~fective Aerodynamic Twisting-Moment Coefficients of Various Wing-Control Configurations at ~chNumbers Fran 0.6 to 1.7 As Determined IEromRocket-Powered Models.NACA RM L511C20,1952.
24. Marley, Edward T., smd ~sh, Roland D.: Some Effects of Aeroel.as-ticity at Mach Nmnbers From 0.7 to 1.6 on the Rolling EY?ectivenessof Thin Flat-Plate Iklta Wings Having 45° Swept leading Edges sndFull-span Consts@ -Chord Ailerons. NACA RM L511L15,1952. .
25. Strass,H. Kurt, and Stephens, lMly W.: An Engineering Method forthe Determination of Aeroelastic Effects Upon the Rolling Effective-ness of Ailerons on Swept Wings. NACA RM L53H14, 1953.
26. ~sh, Roland D.: Sane lWfects of Aeroelasticity and Sweepback onthe Rolling Effectiveness and Drag of a l/H-Scale Model of theBell X-5 Airplane Wing at Wch Nmibers Fran 0.6 to 1.5. NACARM L53118b, 1953.
27. ~lish, Roland D.: Some EYfects of Leading-Edge Roughness on the ,’
Aileron EPfactiveness and Drag of a Thin Rectangular Wing Enployinga Full-span Plain Aileron at lhch Nwnbers From 0.6 to 1.5. NACARM L53125, 1953.
28. Strass,H. -: ~ of SOIIII?Rocket-Model hvestigations ofEffects of Wing Aspect Ratio and Thickness on Aileron Ro12ing Effec-tiveness Including Some Effects of Spanwise Aileron Location forsweptb~ck Wings With Aspect Ratio of 8.o. NACA RM L53L12.,1953.
.
29. ~sh, Rohnd D.: Flight Investigation of an Aileron and a Spoileron a Wing of the X-3 Airplane Plan Form at Mach Numbers From 0.5to 1.6. mm RM L54D26a, 1954.
30. l@lish, Roland D.: Some Effects of External Wing Tip Stores on the .Rolling:Effectiveness and Drag of Plain and Half-Deltd Tip Aileronson a 4-Percent-Thick,Tapered, Unswept Wing. NACA RM L5klZ29a,1954. ,
4TNACAFOfL5’jF14
31. Schult, Eugenetiveness andFlap Aileron
.
25
D- IYee-Flight Measurements of the Ro~g Effec-O&rating Characteristics of a Bellows-Actuated SpEt-Control on a 600 Delta Wing at Mach Nmibers Between 0.8
and-1.8. NACARML54rn7, 1954.
32. Strass, H. Kurt, and Tucker, Warren A.: Some EPfects of Aileron Span,Aileron Chord, and Wing Twist on Rolling Effectiveness As Determinedby Rocket-Powered Mcdel Tests and Theoretical Estimates. NACARML54G13, 1954.
33. Ih@ish, Roland D.: lWee-Flight Investigation To Determine SomeEffects of Tail Damping and Wing-Tail Interference on the RolLingEffectiveness of Inboard and Outboard Ailerons on an UntaperedSweptback Wing. NACARML54L17a, 1955.
34. Harris, Orville R.: Determination of the Rate of Roll of PilotlessAircraft Research Models by Means of Polarized Radio Waves. NACATN 2023, 1950.
35. Bland, WilJ-ismM., Jr., and Dietz, Albert E.: Some Effects of Fuse--e hterference, Wing bterference, and Sweepback on the Dampingin Roll of Untapered Wings As lletermhedby Techniques EnployingRocket-Propelled Vehicles. NACA RM L51D25, 1951.
36. Stone, David G;: Wing-Dropping Characteristics of Some Straight andSwept Wings at Transonic Speeds As Determined With Rocket-PoweredModels. NACARML5-OCO1, 1950. “
37. Weick, FredE., and Jones, Robert T.: R&ti tiAIX@SiS of N.A.C.A..,~.
Lateral Control Research. NACA Rep. 605, 1937.
38. Lowry, John G., and Schneiter~ Leslie E.: Estimation of Effectivenessof Flap-Type Controls on Sweptback Wings. NACA TN 1674, 1948.
F-.—-..--3
-....--—.... -..——. - ..———-———- .-—— —-—
TABLE I
WING CONTROL CONFIGURATIONS
.
NACARML55F14
n) Section is taken freestream unless noted otherwise.
.
.
J
.
.,
NACA RM L75F14
TABLE I.- CONTINUED
WING CONTROL CONFIGURATIONS
27
a) Sectian is taken freestream unless noted otherwise.
NACA 65AO06
~Wedge to Q4c
Con.sttiickiOT.EJt/q!Q06
~ ircdar arc to 0.4c
Ccqy&-6EJ
fi&cu!-3r arc to Q4.c
Flot toTE(8=O)~Iwt=omit/c=ao6
“rcular orc to C14c
VC=O.06
Crcuior arct/c=C106
GmsLtliclq r
Circdor m.LEfo05c06ctoT.E.
VC=O.06Const Ihicq i-
QKa&x Orci-E.to O.&a7c toTE.
“ t/c= 0.06‘Ccms.t.ti-icbt~
C2r’cah aTcl_E.to 0.5CO.sciom.
tlc =Q06
~
VC*O.09
~
tlc =0.09
NACA I 6-009
NACA 65AO09
.
. . . . . . . . . .--. — ____ ___ ___ -——— .. —.— —. -—. .—
28
;onfigurotion
I1L(a
-(b
1-A?’.
——
—
...
TABLE I.- CONTINUED ~
WING CONTROL CONFIGURATIONS
Wing Parameters
Airfoil section(o)
NACA 65ACX39
INACA 651A012
NMA 65AO09
INACA 65AO09
NACA 65AO09
s
9
mm RM L55F14
NACA 65AO06
~
VC=O.06
H-& thickq 1-
Cirwlar 9rcl-EbJ 05c06ct0TE
tlc =O.ols
1Irst.ttlic+ j--
3cub arclx.ta0.5cQ7c taTE
tk=O.06
IrLsthick-q-mmmmuL2cubr arcL.Et005c
O-SCtoTEt/c=OJ36 I
(a) Section is taken freestreom unless noted otherwise.
(b) 3 wings, free-to-rot I tail.
T
1.00 —
20,2
20,2
20
f
20
20
20
r
.
CE2EHZ-2
NACA RM L751?14
onfiguration
_z——
(b]
(cl
z?—.—(c)
(d)
-AZ———(c]
(d]
(e]
E——
L
TABLE I.- CONTINUED
WING CONTROL CONFIGURATIONS,
al+HH-l-l +++++1
F-H+ltlt
+1+1+1+1 : 1 I 1~49
50
51 NACA 651A012
52: NACA 65AO09
c
53
54
55
a
56b
.1 1 I Ill57
58
59
601+1+1+1+1+1+
61
29
1 I I I 1 1
62 t i t t
—- —-— I I ! !
~1”1111/
I I I I 1111’’+rt--t
-&z’ ybpqq”qNACA 65AO09 .82 —
11.5_—. .. . .
(a) Section is taken freestream unless noted otherwise.
(b) 3 wings, f ree-to-rol I tail.
(c) 2 wings, free-to-roll taiL
(d) 2 wings, fixed-tail, horizontal tail in wing chord plane.
(e) 2 wings, fixed-tail, horizontal tail 0.1 ~ above wing chdrd plane.
m!!F!Bm-.--—...—.————-.—————-.——-——-——.—._ — ...—. ___..———
.-
TABLE I.- CONTINUED
WING CONTROL CONFIGURATIONS
mcfl RML55F14
\
(o) Section is token free.sheam unless nated otherwise.
.
TABLE I.- CONTINUED
WING CONTROL CONFIGURATIONS
31
.
— _.. .. ----
o) Section is taken freestream unlessnoted otherwise.
—.. .—..~— z___ ——. .— -. ——- ———.
— . — —
TABLE I.- CONTINUED
WING CONTROL CONFIGURATIONS
.
NACA RM L551?14
.
—
Re
1Ct cft if
o I.c
~
I
Pa
1.51 1.0 .6
.25 20 .1I
- {. ,7
I .40 J]
.252 01 7
-i-- t-- t
202 OJ 3
Jx
J 3
5 7
J3
.40 .09
?62 OJ3
ontrol]meters ;, $ $ I
o ted (8.6 4m
I.0 .s
6.7 498
1.0 0 4.6s a
Lo I 1.8 4.68
1.0 0 5.07
I 1.84 96. *
‘“0 I 1.94 .93
-t I5.74 94 {
.0 IQo + 67. 7(
.0 04 .1I
15.75 24.0
I6249 9
.0 E. 6 5.06
57 15.64 SC
.0 0 45 7 f
.015. 74-9 0 .7(
Wing Parameters;anfiguratian hod
HAirfail(:ection b;
NAcA 65AO06 1.2A 62 12
—
—ccmtJdckrOK
1.2
VC=O063
63 ).5
-.
I
-.
—
).5-t
—
39
84 NACA 65AO09 II
II42?L2?f
6!3
====+
—0=s-t
—3s
————28
B7 NACA 651A012 1+
NACA63AO09 1,.5!36
36
.30(— ‘m 6’lAO’2 I I 28
L——_/2f—-—-.A2!!17
31 28
32
—
33
28
we .9b a2C
Qhz&zq% Iconsttickro-cE.
tJt*Q108
—
—
28
—
!7
—
—
50
—
—
Is
—
NACA 65[A012 1~$
I) Section is i— I
n unless nated otherwi[ken f
.
5TNACAR4 L55F14 33
TABLE I.- CONTINUED
WING CONTROL CONFIGURATIONS
Wing Parameters:anfiguration Mod&-
A A Airfoil (:ection b/2A (d@ (ft~
L “—.——- .
/ : I4;‘i ‘AcA’r“i—.——-— HexaguldA 97 28 1’ 39 wedw L+. fOo.3c 1.02
0.7C to-cEt/c=o.045
——98. 30 16 .40 t L04
E.
NACAOO06.6-116
99 ~ 45 z 38/1.14rnodifiedMr ootNACA 0006.4-L16 1.42
36/1.14 (mcdfled)at tip.—_ —-—
_zz Republic R-4,40- 17[OX101 3.1 35 1.63 (modified) normol to !2u- .8s
persent chordIina——
A ’02 55 ‘1 25 ‘Am’’5A00’ 11’—— —-.
d0 NACA 0010-64
103 S6 38 .45 nwmol to 44.8 1. I
b percent chord line
—-——I I I I
Same0s 103except104 oileron is modified:
h=+ (t)~
0 105 4.0 0 .60 NACA 69AO04 ,., ~
[a) Section is taken freestream unless noted otherwise.
Ct Cr 3‘ft) (ft ) ‘ft
27 50 L26
‘1 I I41.87 L4E
39 .86 L4Z
.69 123 28
.64 L17 27
.80 .56 1.2
.26 .8 8’ 1.4
.39 0.7613
I I 1’
.43 .67 13
]
?5
25—
23
—
25
—
27
—
30
—
.17
--
I
-
L
.15
lefs
28
28
16
29
—
—
19
—
13
13
30
— .—-.. . . — --- -— . .. . ——.— .—-— ———-
Configuration
..—
A ((
47 (c
A
MO
—
10(
101
10[
TABLE I.- CONCLUDED
WING CONTROL CONFIGURATIONS
Wing Parameters
$=A Airfoil section
a
~ 10percent circular arcnormal to cf41ine
WA 64(lo)AOil nom
to 38-percent-chard
he atrootW 6~@ AO08Z3
normal to 38-percenl
chard line at tip ( f 1
ii(.!
1.I
—
1.?
(<
—
0
(9—
81
[g:—
—
7~
3!
—
3(
—
4(
—
51
—
—
7
(!
.7
—
.6
—
81
—
.0
—
—
NACARML55Flk
10
—
?5,2
—
5,2
1) Section is taker reestream unless nated otherwise.
:) 2 wings, free-to-roll tail.(f) Variable wing sweepback configuration. Wing pivots about axis normal towing chord
plane at intersection of fuselage and 38-percent-chord line.
(g) To tip at 38-percent-chord line.
.
.
.
NACA RM L55F14
TABLEII
35
INDEX To !?lGuREs
Basic data Cross plots
V=iableFigure Figure
rrailing-edgeangle . . . . . . . . . . . -- 6
L&l?oil thickness ratio . . . . ● . . . . 7 ---
Sweepback . . . . . . . . . . . . . . . . 8 (a)
Aileron chord ratio ..’... . . . . . . 9 10
JAspect ratio . . . . . . . . . . . . . . U 12
Aileron span d spanwise location . . . 13 14
Wing-tail interference . . . . . . . . . 15 ---
Wing location @ number of wings . . . . 16 ---
Comparison between measured andestimated values . . . . . . . . . . . 17 ---
%railing-edge angle plots show data at two wing sweepbackangles.
.
. —-- —.. --.—— -— .— —. —. —— ——— -——— —— — — ---
36. ... .
k NACA RM L5mk
3 wings spacedof 120° around a“’;rva’sw7
3.25 aircraft rocket _
,.~ _________ _.––_–--–’L_! I 1)———-——_.
1~-1 I-—.—
L–_~__–__–~_-__–j_ –__––”_– ---~ .
Spinsonde 5.00 diam.
1. If.
I13.1
.— —— ——— .______
——”1
1
I
—— —. ———— . —— —— ..—— . .
L—
Figure l.-
.- —
(a)
General
Model 28. A = 3.7; A = 00;
aztranganent of typical test
are in inches.
.
.\
\.—,.—— -
.. . . . ——’
A = 1.0. L-67857 ●1
vehicles. All dimensions
.
,
@ —--
‘-_-]T= -.-, - .= -.’-..-.-..———.L———-- .-. —-+
“-. . ..M
-
.—
NACAlW L55F14 37
3 wings spacedof 120° around
at intervals
bed, ~’~
3.25 aircraftrocket
~ Spinsonde .1 // ““r
———. ———-
——— ——- —-
5.00 diam.
1
— —.——
-L1 ..–. ——— . .——
(b) Model 71. A = 4.0; A = 45°;
Figure 1.- Continued.
A = 0.60. L-T3023el
.-. ——. — -.-..—— ..—. . ..— ---- _——. —
38. NACAIOf L551?14
3 wingsspaced at intervals
of 120° around body
!
18.62
3.25 aircraftrocket{
~–_––– –-L____ __–____ —----- ------ -1
L–~––––-;____–~–
5.00 diam.~
55.00 ~
—I
I
I
I
/
-./ .
,->,.:
! @
‘ :%.,:”
...>
; ,.y1
/“ ‘/‘/
—— ....— . .)
p~
+.- — – -“-. .
.- .. .- - ..- ----1
1
I
\
‘[[I ._._2r...i.-~
(c) Mode188. A=8.0; A= 450; A=l.O.
R@re 1.- Conclud&i.
.
.
L-69353.1.
..”— — .—. —— ..:<2. .—
NACA RM L5~4 39
3 wingsof 120°
spacedaround
at intervalsbody
_.—__— ——
.— !?!.-...—Figure 2.- Photographs showing typical construction
8, 83, 85, 89, @ 93= Au ~=iow ‘e
L-82789 .1for modeh 2, 4, 6,in inches.
——.——..-_____ - .. —.—- — - — — -——_.__ ——..——
40 I’UCARlf L55F14
3 wings spaced at intervalsof 120° around body
3.25 aircraftrocket
+~—. ––. __ L______ –––_–– –______
L__–__– –__
5.00 diam. t
.
.- -.
‘2/
/
,/’
L-71131.l(a) Three wings with outboard half-span aileron for investigation of
effects of wing location. Mcilel53.
Figure 3.- ~ical test vehicles for investigating the effects of winglocation and wing-tail interference. All dimensions are in inches.
6TNACARML~5Flk
~——————.—— —— —_______ ___
——— -—— ——. —.— --- -
rocket
H 55.00 *
--1
L-75895.1(b) TWO wings with outbosz’dhalf-span aileron with fixed tail for inves-
tigation of effects of wing-tail interference. Model 55.
Figure3.- Concluded.
-’.—— — ———. — .—. — . .—. ——— ___ -— —— —._
42 NACA RM L55F14
12XI06
8 —
Rec
4-
L
0 I I I I I I I I 1 I I I.6 .8 -Lo
(a) For all tests, the ReynoldsShaded band. Theminfmum E
4XI03t
3
2
I
!
1.2 1.4 [.6 1.8
M
number per foot R/E fell within thetested was approxhately 0.4 foot.
01 I I I I I I I I I I I I.6
- (b) For all
Figure 4.-
.8 Lo 1.2
M
tests, th,edynsmic pressure
Variation of Reynolds numbernumb= for all test
1.4 1.6 1.8
q ‘fell within the shaded ban~.
end @w.nic pressure with Macti;vehicles.
f,!:I,!,,J,.
.
,
.
.
<
NACA RM L55F14 43
.02
.0 I
(pb:2v)R
o
-.0 I
A= 00’
for roll
.6 .8
Figure 5.- Effect of model
Lo 1.2
M
roll bertia on rollingModel 27a.
1.4 1.6
effectiveness.
‘\
..
,
—.—— . —— —— .— —--——— .= —- .-. .. . . . . . . .
—. -. —
44
(pbi2V
8 )R
()pbJ2V
t3~
.02
.0 I
o
I’UCARML55F14
o 10 20 30 40
.
.
.
0 10 20 30 40
+,deg
(a) M = 0.70.
Figure 6.- Variation of rolling-effectivenessparameter m ~~5
trsiling-edge a&le ~. Unless otherwise indicated, average valuesare plott~ where two or more nminally identical modeti were tested.A= 3.7; A = 1.0; Ca/C = o .20; full-exposed-span ailerons ~ =
(0.19);
fjmsotop. NumlIers in symbols denote model nunikrs.
.
-.
NACA RM ‘L75F14
.02
()pb/2Vs~
.0 I
o
45
0 10 20 30 40
#,deg
y.,
, ..
.02
(pb:2v)R.0 I
o
0 10 20 30 40
#bde9
(b) M = 0.80.
Figure 6.- Continued.
.— ._—. .— - —.- —-—
46
.02
(P’:2V)R.0I
1
0
-.010
.02
()pb/2V
8 R.
.01
0
10 20
+,deg
ILICAIU4L551CL4
30 40
o 10 20 30 40
+,deg.
(c) M = 0.90.
Figure 6.- Continued.
—
NACA FM L55F14
.02
(W*”)
8R
.0 I
o
-.010 10 20 30 40
+,deg
47
0
0 , 10 20 30 40
+.deg
(d) M = 0.93.
Figure 6.- Continued.
.02
()pb12V
8R
.0 I
.——— ._. _.— — . -- .—. — .——— —— -— ——-—
.-
NACA RM L55F144a.
.02
.0 I
o
.
0 10 20 30 40
$,deg
.02
.0 I
o
(pb:2v)R
o 10 20 30 40
+,deg
(e) M = 0.96.
Figure 6.- Continued.
7TNACA RM L55FIL4
.
.02
(pb~2v)R.0 I
o
49
0 10 20 30 40
$,deg
.02
.0 I
o0 10 20 30 40
+,deg
(f) M = 1.00.
~gure .6.- continua.
. . _————— ——..—— —-___ .—— .—-
m
50
.02
()
pb~2V
s~
.0I
.
0
t A=O” h##diI Iiiii iii
o 10 -
NACARML55F14
20 30 40
+,deg
)R
.02
.0 I
o
(
0 10 20 30 40
+,deg
(g) M=l.20.
~gure 6.- continued. .
NACA RM L55F14 51
t
/pb/W )R
.02
01
00 10 20 30 40
#,deg
.02
()
pb/2V
s~
.0I
o
/
o 10 20 30 40
+,deg
(h) M % 1.40.
l?igure6.- continued.
-..-——-—— .— — _—— ..— —-— — ———..-—. —
52
.02
(pb’2v)TR””
.0 I
o
‘.:.L2!!!EE!!E NACARML55F14
.
.
0’ 10 20 30 40 .
$,deg”
-.
.02
()
pb/2V
a~
.0 i
o
0’10”20 30 40
#,deg
(i) M = 1.60.
~gure 6.- Concluda.
G.RR@g?
—-——...——.—..-.——————___ .——..—.-—___-.—....
NACAIU4L55F14 53
()pb12V
8R
.02
.0 t
o
.6 .8 LoM
1;2 L4 1.6
Figure 7.- liYfectness. Munibersfoil sections.
.8 Lo 1.2 1.4 1.6
M
(a) A= 3.71.
of airfoil thicbess ratio on aileron rolling effective-in parentheses denote model nunibers. NACA 65A~ air-Ca~C = 0.20; A = 1.0; 5 =30 to 70.
-. . . . —-.——— -—- ——-. -—.—— .— —. ——.—. ———.———
54
.02
(P’’2V)—.8 R .01
0
,.NACARML55F14
.6 .8
11111111
I 1111 I Ill [11
1.0
M
(h) A = 5.0.
.
.02
o
.6
1.2 1.4 1.6
65Ao09, # =10.00 (88)651AOW, @=16.oO (90)
.8 Lo 12,.
M
(C) A = 8.0.
Figure 7.- Concludei.
1.4 1.6
~
.—— ———. .——. ..— -——-- ———— —.. — ——. ——--— .—... —-——-
NACAm L55F14 55
(pb/2V3 “)R
(a)
.02
.0I
o
A=
.02
;0I
o
.6 .8 Lo 12 1.4 L6
M
3.7; A = 1.0; NAcfl6LjA006airfoil sections; Ca/C = 0.20.
1.8
.6 .8 Lo 12 1.4 1.6 18
M
(b) A = 4.0; X = 0.6; I?ACA65Ao06 airfoiI sections; Ca/C = 0.30.
Figure 8.- Sconeeffects of wing sweepback on rolling effectiveness forfull-exposed-span ailerons* Numbers ti parentheses denote model num-bers. ~ U30 to To.
— —-. . .——— . .. --—. -. — .—. . .—
56_. .--—
e“=!z!? NACA RM L5@14
(P’:2V)R.02
.0I
0‘.6 “. 8 Lo
M
L2
= 00, j = 11.50 (3)=300, #.= 10.6 (5)
=450, d=u.oo (7)s(joo, fj=~~*~O (g)
(C) A= 2.3; A = 1.0; WA 65Ao@ ~tioiI
.02
()pb12V8~
.0I
o
14 1.6 1.8
sections; cJc = 0.20.
.6 .8 LO . 12 1.4
M
(d) A= 2.9;A= 1.0;NAcA 6SAOOgairfoil sections;.
~gure 8.- continued.
1.6 1,8
c~jc = 0.20.
—.-. — -— ——.———. — ——.—— —. — —-——.—
8T
NACA FM L55F14
t-
—---------— ~ ---—=,
57
.02
()
Pb12V
8R
.0I
.fl
(-e)
“.6 .8
~ = 397; h = 1.0;NACA
.02
(pb:2v)R.0I
o
Lo 1.2 1:4 ‘ 16
M
6~Aoog airfoil sections; ca/c = 0.20.
1.8
.6 .8 Lo L2 1.4 1.6 1.8
M’
(f) A = 5.0; A = 1.0; NACA 6~AO@ sirfoil sections; ca/c = 0.20.
~gure 8.- Continued.
-—-— -——-. .-— — ———— -- --- —-—.. —._ ______ —— ..- ——— _._—. _ . .
.
.02
.0I
(P’{2V)Ro
-.01
—
NACA RM L55F14
.6 .8 Lo 12 L4 16 1.8
M
(g) + = 3*7;~ = 1.0; NACA 651AO12 airfoil sections; ca/c = 0.20.
.02
.0I
(P’;2V)Ro
-.0I. .6
(h) A= 3.7;
.8 Lo 12 14
M
A = 1.o; ~CA 16-o(Y3airfoilsections;
Figure 8.- Continued.
16 L8
Ca/C = 0.20.
.
.
.
—.
NACARML55F14 59
.
(pb12V8
.- .—.
.03
.02
o.6
(i) A = 2.5;
(3)
.03
.02
“)R
.0I
o
liii! !!-.:—--..--.+
“..>.*. >. . . . . .=.
.8
h = 1.0;
Lo
flat-plate
1.2
M
airfoil
1.4
sectioni3;
1.6 1,8
ca/c = 0.20.
.6
A= 2.3;
.8 Lo 1.2
M
A = l.O; flat-plate airfoil
1.4 1.6
BeCtiOIIS; Ca]C = 0.40.
Figure8.- Conchuied.
. . .. . . -——- ---- —.—— —
1.8
60
.03
.02
()pb12V8R
.0I
o
-.01
. .. . ———._
L -@)————————————./
-n_IUCA RML55F14
. .
.6 .8 Lo -
M
(a), A = OO; A = 3.7; h = 1.O;’NACAexposed-span
Figure 9.- Effectof aileron~ S 3° *O 70, ~cept for
model numbers.
12 1.!4
6yLoogairfoilaileron.
L6
sections;
L8
full-
chord ratio talc on rolling effectiveness.
Ca/C = 1.0. Numbers in parentheses denote
------ ...-
.
—— .
NAC!ARML55F14 61
.
.-.U3
.
.02
()pb12Va~
.0I
o.6 .8
(b) A =45°; A = 3.7;
Lo L2 L4 L6 1.8
M
A= 1.0; NACA 65AOOg airfoil sections; full-
exposed-span aileron.
Figure 9.- Continual.
.— _- —— -- . ———. —— —
62 NACARML55F14
.—
.
.03
.02
(7).
.0I
(3
.6
(c) A = 45°;
.8 Lo 1.2 14
M
A =4.0; A“= 0.60; NACA 65Ao06 tifoilexposed-span aileron.
Figure 9.- Continued.
L6 1,8
sections; full-
,,
—
NM% R&lL55F14 63
//,
/
/
.03
.02
(P’:2V)R
c1
.0I
.6 .8 Lo L2 L4 L6 1.8
M
(d) A = 45°; A = 4.o; x = 0.60; Wm 6~006 airfoil sections; out~otithree-qyarter-span aileron.
Figure 9.- Continued.,.
4iiallh.-. ——. . .____ ___ ___ —.— ._ _ .—— _— ..__ ..—. . ___
64 MACA RM L7jl?14
= 0957
..-—.03
.02
()pb12V3~
.0I
o
Oa— = 1.00, (interpolated)I I I I II I o
o~—=0.30, $=7.4° (73)o08~ = 0.15, # = 6.8° (68)
.6 .8 Lo 12
M
(e) A = 45°; A = 4.0; A = 0.60; NACA 65Acx16. .half-span
Figure 9.-
aileron.
Continued.
14
airfoil
L6 1.8
sedio~j outbod
.
.,
—
9T
NACA101L571?14
.
65
//
.03
.
.02
(“’2”)8R
.0I
o
.—
0.78
.6 .8 Lo 1.2 1.4 1.6 1.8
M
(f) A =45°; A =4..0; A = 0.60; NACA 65AQ06 airfoil sections; ~tboaraone-quarter-span aileron.
Figure 9.- Continued.
. —--— —.— ... ..—
66
.03
.02
(P’:2V)R.0I
o
(g) A = OO; A
NACA R?4L55F14
..
.
.6 .8 Lo 1.2 14
M
= 2.3; A = 1.O; flat-plate airfoil sections
Figure 9.- Continued.
1.6 1.8
(t/c = o.o’j4).
-!
.
.. —.—. —— -——.-—
NAC!ARML55F14
.
67
-IL.—
.03
.
.02
(%3,.0I
o.6 .8 Lo L2 1.4 L6 1.8
M
(h) A=O°; A=5.0; A= I.O; fkt-phte airfoil sections (t/c = 0.083).
FiWe 9.- Continued.
.
. ..— ——. _—__ . .. . . _._ —___ __ ______ . . -—-—. — —.— — —
68
.03
.
.02
()pb/2va~
.0I
o
—
.6
(i) A = 45°;A = 2.5;
.8 Lo” 12 1.4
M
A= 1.0; flat-plate airfoil sections
Figure 9.- Continued.
1.6 1,8
(t/c = o.@).
NACARML55F14
.
69
.03
.02
()pb/2V8R
.0I
o.6 .8 Lo L2 L4 L6 L8
M
(j) A=45°; A = 8.o;A = 1.0;flat-pkte airfoil sections (t/c =o.1o8).
Figure 9.- Concluded.
.
—— . —-. ..— —.— -.——— .———
70
0 .8-“
am- NAC!ARM L55F14
.
Ii
X*
o0
.2 .4 .6
co7
.6 1.0
?
(a) M = 0.70.
,.
Figure 10.- Variation ofratio . ~.30~p
model numbers.
.2 .4
‘0T
(b) M = 0.80.
the effectiveness factor
except for Ca/C = 1.0.
.6 .6 1.0
~ with aileron chord
Numbers in symbols denote
.
-.—— — ——.
NACAIM L55F14
.
-
II
4?
.
.
.8
.4
0
.8
.4
0
0 .2 .4 .6 8 1.0
‘a-E-
“ (c) M = 0.90.
0 .2 .4
Ca
.6 .6 Lo
c
(d) M =0.93.
Figure 10.- Continued.
.—— —. .-— —— —-——— .— —.— —.—
72
0--
o-
.-
lilCA RML55F14
.8
.4
n
“o .2 .4
CaT
(e) M = 0.96.
.6 1.0
.8
.4
00 .2 .4 .6
‘a7
(f) M = 1.00.
Figure 10.- Continued.
●
..
Lo
—. .—
10T
NACA RM L55F14
.
o
1- 73
.2 .4 “ .6
(g) M = 1.20.
x
Lo
0 .2 .4 .6 .6 1.0
CaT
(h) M = 1.40.
Figurq 10.- Continued.
—.————— —-
74 NACA RM L55F14
o-“
810
x-
.8
.4
00 .2 .4 .6 B 1!0
CaT
(i) M = 1.60.
Figure 10.- Concluded..
.
—-- — —
NACA RM L55F14 75
.02
.0I
n‘.6 .8 Lo L2
M.L4 L6 L8
.
0.6 .8 Lo 1.2 L4 L6 1.8
M.02
.0I
o.6 .8 Lo.
M
1.2 1.4 1.6 1.8
Figure 11.- Effect of aspect ratio on rolling effectiveness for full-spanailerons on untapamlwings. 8 -3° to 70. Numbers inyrenthesesdenote model nuders.
—— ——_—...
—.. —
76
.02
.0I
(pb12V8
-.01
t--=
-.*.— NACA RM L55F14
)Ro
.6 .8 Lo L2
M
(b) NACA 651A012 -Oil
Figure 11.-
,
sections.
Continued.
.
14 L6 1,8
ca/c = 0.20.
.
77
.03
.02
(,,/2”)
8R
.0I
o.6 .8 Lo L2
M1.4 1.6 1.8
.
0.6
(c) Modified
,8 Lo L2 1.4 L6 1,8
M
flat-plate airfoil sections. c~lc = 0.20.
Figure il.- Continued.
-....——— .. -—- .—. .. —-— _—. —.——...
78 IiMA RML55F14
.
.03
.02
.0I
n
‘.6 .8 Lo 12 L4 1.6 L8
.03
()pb/2V~“R
.0I
o
(d)
M
.6 .8 LO 12 1.4 1.6
M
Modified flat-plate airfoil sections. ca/c = 0.40.
Figure 11.- Concluded.
1.8
.
.
NAC!AIU4L55F14 79
.
n
ax
o A.oo
❑ A=45”o A=60°
2
I
II
n
u
ax
Figure 12.-6%
“2
Variation
30 to 70.
3 4’ 5 6A
of the effectiveness factor ~
Numbers in symbols denote model
u
7 8
with aspect ratio.
numbers.
.. ..— ———.. —.——..———— —— —.— --- .— ——— — —
-.
80
Iymn
u
X4
II
i?
II
ax
o
8
A=wA = 45°A = 60°
A
.
I
o-2 ‘3 4 5 6 7 8
A
Figure 12.- Concluded.
llTNACARML55F14 81
.
(“’2”)8R
.0I
:6 .8 Lo ‘1.2 1.4 1.6 1.8M
(“/2”8“ )R
.6
(a) NACA 6qAocg
.8
airfoil
Lo 1.2 1.4 1.6 1.8
M
sections. A = 3.7; h = 1.0; CJC ‘0.20.
Fimxre 13. - Variation of rolling effectivenesswith Mach mmiber for con-—stant percent-chofi ailerons located at various spanuise positions.8 - 3° to 70 except for Ca/C = 1.0. NMbers in parentheses denotemodel nunibers.
.- — -. .- —.——.———. —— _——. —____ -—— ..—. — -——— .4
82 I?ACAFtML55F14
()Pb12V8R
.02
.0I
o.6 .8 Lo 12 L4 1.6 1,8
(b-)NACA 65Ao06 -oil sections. A = 4.0; x = 0.60;A = 35°; Ca c = 0.50./
.
.02
()Pb/2V8R
.0I
f)
.02
.0I
o
.8 Lo 1.2 1.4 1.6 1.8
‘.6 .8 Lo L2 1.4 L6 1.8
M
(c)NACA 65AOC6airfoilsections. A = 4.o; x = 0.60; A = 450.
Figure 13.- Continued.
NACA RM L55F1.4
(“’’2”)8R
.6
(d) WA 6~A006 *fOil
83
.8 1.0
sections. A
Figure 13.-
L2 L4 L6 1,8
M
= 4.0; ~ = 0.60;A = 45°; Ca/C = 1.0.
Continued.
...— —— —— ----- —— _.--_—_—
—
84
(Pb/2V8 )R
()Pb/2V8R
.02
.0I
o
‘.0[
(e) NACA
NACA BM L5+3F14
.
.6 .8 Lo 1.2 L4 L6 1,8
M
.02
.0I
o
-.0I.6
651Aoi2
.8 Lo L2 L4 L6 L8
M
airfoil sections. A = 8.0; A=k5°; C~/C =0.20.
Figure ‘13.- Continued.
.
NACA FM L55iF14
.
.04
.03
()
pb/2V
8R
.02
.0I
o
------ -.—-——--——— -—— --—- -.
1.00
1.00
1.001.001.001.001.00
(36)(62)(81)(82)(lo)(11)(11
I
.6 .8 Lo 12 1.4 1.6 1.8
M
(f) Mist.dlaneous COnfigl.lmticm. Ca/C = 1.0.
Figure 13.- Concluded.
.——— .- ———-
86
1.0
.8
.6
x? .4
.2
n
NACA RM L55F14
“o .2 .4 .6 B 1.0
(a) M = 0.80.
Figure 14.- Variation of the effactiveness fxtor ~ @th tileron SPan.
N@bers in symbo~ denote model numbers.
NACA RM L55F14
1.0-
.8
.6
.4
.2
00
87
.2 .4 .6
qi
(b) M = 1.4.
Figure 14.- Concluded.
.6 1.0
..— —c _—. ———. —.—.
— — —
88
(pb/2V--T- -)R
()Pb12V8R
-.:- --
*.. * NACA FM L55F14
?dodel58 F’ree-to-rplltall
.— -—.
Model 59
__&%%zA ‘:;-to-rol’ ~
-G??dL-----------
d wired +.51
.0I
o
lFlxedtall, @ = =.OO (59)’#M#
.6 .8 Lo L2 1.4 1.6
M
(a) Inboard ailerons.
.0I
n
:6 .8 Lo L2 1.4 L6
M
(b) Outboard aikI?OIM.
Figure 15. - 8cme effects of wing-tail interference on rolling effective-ness for inboard and outbosrd ailerons. A = 3*7; A = 1.0; A = 450;NACA 65AOOg airfoil sections; ca/c = 0.20; 8 = 3° to 7°. Numbers in
parentheses denote mcdel nunikrs.
.
.
.
—
12TNACA RM L5@?14
=2H!J!L‘=’’”70~“’’’”A%dl-4-JsdL4 ”’”
.02
.0I
(?)Ro
Figure 16. -
A = 1.0;
3 wings,XL
{*
3 wings,Xfi
3 wingsplusfree~to-yell~ = 0.63
2 wingsplusfree-to-rollxfi = 0.63
-.01.6 .8 Lo 1.2 L4
M
(a) l?ul.1-expos~-spanaileron. A = OO.
16
89
= 0.89
= 0.63
tail,
tall,
sane effectsof wing location and number of wings. A = 3.7;Ca/C = 0.20; 5 -3° to 70; NACA 6~AOOg airfoil sections.
-———. ——.. — ———. . —
90_-—
$ = 11.7°
g = 13.6°—-. — —
g = 10.9”
.02
()pb/2V
8R
.0I
o
NACAFM L55F14
3 wings, Xfi
3 wings, x/L
3 wings, plusfree-to-rollx/Z = 0.63
2 wings plusfree-to-rollxfi = 0.63
= 0.80
= 0.63
.6 .8 Lo 12 L4 1.6
M
(b) Full-qmed-spm dkrOIIS. A = 45°.
~igure I-6. - Continued..
tail,
tail,
.
.
.
.
.
.
NACA RM L55F14
Mociel / —/ /
1 -— .— - 1
()pb/2V8R
3 W*8, x/z
3 wings, x/L
2 wings plusfree-to-roll
XL = 0.63
.0I
o‘.6 .8 Lo 1.2 1.4 L6
M
(c) outboard half-exposed-span ailerons . A = 45°.
J%&4zlv”303wtigs8x
()pb/2V-rR
-————.-.. .-. —
.0i
o‘.6
(d) Inboard
.8 Lo L2
——. .._—
M
half-exposed-spin ailerons.
Tigure 16.- Concluded.
—.. — .— —___ —
1.4 1.6
91
= 0.80
= 0.63
tall,
z 0.80
= 0.63
tail,
.
. ——— ..— — ——.——. ..
92
I
(pb12V8
NACARML55F14
(pb12V8
(pb12V8
.02
)R
.0I
o
.8 Lo
M
12 1.4 “16 1.8
.
‘.6 .8 LO - 12
M
.02
)R
.0I
o
L4 1.6
.8 Lo 12 1.4 16 ..
M
Figure 17.: Messured and estimated rolling effectiveness for severalmodeh sbmlating edsting or proposed airplane wing-aileron configura-tions (without fixed tails). Estfmated values were obtained from fig-ures 6, 10, M, smii14. Numbers in parentheses denote model numbers.b = 30 to 70 except for models 99, 107, and 108.
kk?w~vziiqziq
— ——— —— ~—
NACARML55F14
.02
()pb/2V
8-R
.0 I
93
0.
..’6 .8 - LO 1.2 L4 1.6
M
.02
()
pb/2V
8R
.0 i
o.6 3“ Lo 1.2 L4 1.6
M
.02
()
pb/2V
8R
.0 I
o.6 .8 Lo 1.2 L4 1.6
M
Figure 17.- Continued.
‘aiiimwi. . . . . . ..———. ..—. — —. —.. ..-— -. —____
—
94
.02
.0I
o
–.01
(pb12V8
.—–r -
,-- -- !--
6-,....-.>-.--‘---NACA BM L5~4
.6 .8 Lo 12 14 1.6
M
.02
‘)R.0I
n
‘.6
()pb/2V8R
.0I
o.6 .8 Lo L2 14
M
Figure 17. - Continued.
—
.
.8 Lo
M
L2 1.4 L6
1.6.
NACA FM L55F14 95
(pb12V
8
‘.6
.02
R
.0 I
o.8 Lo
M
12 L4 L6
.02
(pb/2V
8
.0I
o‘.6 .8 Lo 1.2 [.4
M
Figure 17. - Conclud@.
L6
NACA - I.an@y Fiek4 V&
-. —.——— .—— — -— —— -.— —-— -— -— — .--— —-——— -———
