-4 mN#EmuQL co,,!? RM LfbJ26 .. -,.--.-———— ..-–-–: - _m__.7z-.-. “. :.... .__-= ., -=-. .- ——-. -. - - ~ ._J... ..== RESEARCH MEMORANDUM- COMPARATIVE TES!13 OF THE ROLLING EFFECTIVENESS OF CONSTANT-C HORD, FULL-DELTA, AND HALF-DELTA DELTA WINGS AT TRANSO?$IIC AND SUPERSONIC By Carl A. Sandahl and H. Kurt Strass Langley Aeronautical Laboratory Langley Air Force Base, Va. , f w“ ~! 1 !>/ “’ NATIONAL AILERONS ON SPEEIX3 FOR AERONAUTIC= WASHINGTON December 12, 1949 .-. . . ... . --.—- . . -= -, --- , _ ~ ,.= — https://ntrs.nasa.gov/search.jsp?R=19930086137 2020-06-22T00:32:12+00:00Z
Comparative tests of the rolling power of plain constant-chord,full-delta, and half-delta ailerons on delta wings having 45°and 69° leading-edge sweepback have been made utilizing rocket-prope31edtest vehicles in free flight. The rolling power of the constant-chordailerons was reduced abruptly in the Mach nwiber range from 0.9 to 1.0.For a given ratio of slleron erea to wing area, the full-delta aileronshad a higher level of effactiveness at superscmic speeds then theconstant-chord or helf-delta ailerons. The half-delta ailerons had thesmallest variation of effactiveness with Mach nmiber end were about aseffectiye as the full-delta ailerons at M = 1.9. The wing-aileronrolling effectiveness of the helf- and full-delta ailerons can beaccurately predicted by calculations based on the Uneerized flow“equations. Similar calculations for the constsnt-chord ailerons yieldvelues which are considerably lsrger than the measured values.
INTRODUCTION
Of the numerous wing plen forms which have been proposed for flightat trensonic end supersonic speeds, the delta plan form afforalscertainaero&mni c end structural admntages. W en approach to the problemof providing such wings with adequate aerodynamic control surfaces,comparative tests of the rolling power of several delta-wing aileronconfigurations have been made. Phln constant-chord, full-delta, andhalf-delta ailerons were tested with delta wings having 450 snd 600leading-edge sweepback. A hslf-delta conf@uration identicel to thatused in the investigations reported in reference 1 was elso tested. Thetests, which were made in ftreeflight with rocket-propelled test vehiclesby means of the technique described in reference 2, permit the evaluationof the rolling power of wing-aileron configurateions centinuously over the
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IWC!ARM L9iT26
Mach num%er range fran about 0.7 to as high as,1.9. The flight testswere conducted at the Lsmgley Pilotless Aircr~ ReseeJxJj Station,WaU.ops Island, Va. The present yaper Includes and extends the workreported in reference 3.
SYM130LS
pb/2v
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wing-tip helix angle, radims
rolling velocity
diameter of circle swept by wing tips
flight-path velocity
Mach number
(Ii)4aspect ratio
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aileron area
expmed wing area
sweepback of wing leading edge
aileron deflection measured inchord plane and psrallel todegrees
plane normal to wingmodel center line,
semivertex angle of wing (CQU@aXUent of Am)
DESCRIPTION OF TECHNIQUE.-.
Only a brief description of the technique will be given in thispaper; a more complete Uscussion is contained in reference 2.
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Test Vehicles
3
The generel arrangement of the test vehicles is shown in figure 1.The gecmetric details of the wing-aileron configurations tested aregiven in figure 2 end further yertinent information is given in table I.The constant-chord and the full-delta ailerons were formed by deflectingthe wing chord plane at the required hinge-line locations. Half-deltaaileron configurations 5 and 6 were tested with the gap due to ailerondeflection o~en and sealed. Configuration 7waa tested only with thegap sealed by the fence as shown in figure 2. Photographs of sane ofthe test vehicles are shown in figure 3.
The bodies of the test vehicles were constructed of lalsa exceptat the wing attachment where spruce was employed.(configurations1,
Same of the wings2, 3, and 6) were constructed with a leminated spruce
core to which a steel skin was cycle-welded to provide the requiredrigidity. The remaining configurationswere of solid dural.min. Forsll configurations, the exposed wing ereawas 1.563 square feet.
Tests
The launching of the test vehicles was acccqilished at the WallopsIsland test facility of the Langley Pilotless Aircref’tResearch Division.The test vehicles were propelled by a two-stage rocket system to a Machnumber of about 1.9. During a 12-second period of flight followingrocket-engine burnout, in which time the test vehicles coasted to a Machnumber of about 0.7, measurements of the rolling velocity Poduced hythe ailerons (obtainedtith speciel radio eq.ui~nt designated spinsonde)end the flight-path velocity (obtainedwith Ihppler rader) were made.These data, in conjunction with atmospheric data obtained with radiosondesj_tted the evtiuatlon of the rolling effectiveness of the particularwing-aileron configuration under investigation in terms of the ..——
parameterpb
m78 asa function of “theflight Mach nuniber. The scale of
the tests is &dicated by the curve of Reynolds number against Machnumber in figure k.
Accuracy
The accuracy is estimated to be within the following Mmits:
pb~75 (due to limitations cmmodel constructional accuracy) . . +0.000~
~~a (due to limitations on instrmnentation)
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It should be noted, as pointed out in reference 2, “~hat,owing to
the relatively mnall.roUing mament of inertia the values of~2v/8
obtained during the larger part of.the flzght are substantially steady-state values even though the test vehicles are experiencing an almostcontinuous rolling acceleration or deceleration. Except for the Machnumber range from 0.9 to 1.1, in which range a@upt changes in rollingvelocity mey occur, the deviaticm fl?omsteady-state conditions isestimated to be within 3 percent. tiasmuoh as it is no$ now possibleto estimate the dmnphg in roll of these models with suitable-accuracyin the Mach number range fiwn about 0.9 to 1.1,”en accurate calculationof the deviation from steady-state conditions cemnot be made. However,for an extreme exmrpl.ehaving a rolling acceleration of 1.00radiansper second per secand and assmning & damping-in-roll derivative of 0.2,the maximum deviation would be about.10 percent”.
RESULTS
The results of the present
AND-DISCUSSION
investigation are shown in figure 5 as
curves of the wing-e31eron
functions of Mach nuniber.summsrlzed in figure 6 andlinearized supersonic-flowcalculations, which in all
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The rolling-effectivenessresults areere compared with c4culatio~ based on theequations in figures 7 to 9. The theoreticalcases applied to isolated wings only, were
based on the wing plan form defitid in figure 100 The damping in rollwas obtained &cm reference 5 for ell cases. The results presented arefw essentially infinitely rigid wings.
Effect of leadim-edae Sweepback.- The effect of leading-edgesweepback on the rolling effectiveness of the Configurations tested canbe noted in figure 5. For the constsnt-chord ~lerons (fig. 5(a)),increasing the,sweqpback &am 45° to 60° incre%ed the subsonic ‘effectiveness, decreased the abruptness of the loss of rolling power inthe Mach number range flrcm0.9 to 1.0, and increased the rolling powerat moderate supersonic Mach rumhers. At the hi&hest Mach numbers investig-ated, the rolling power is independent of lee.Ung-edge sweep for thevalues tested.
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For the full-delta ailerons (fig. 5(b)), increasing the leading-edgesweep &an k50 to 600 decreased the rolling power about 20 percent at Machnumbers less then 1.0. Above a Mach number of 1.0, increasing theleading-edge sweep had little effect on the rolling effectiveness.
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NACA RM L9J26 5
For the half-delta ailerons (fig. 5(c)), the main effect ofincreasing the leading-edge sweep was to increase slightly the rollingpower at moderate supersonic s~eeds. The effect of gap seal for theseconfigurations is negligible.
Wmu3ariaon of rolllna effectiveneas.- The rolling effectiveness ofthe configurations tested is ccmpared in figure 6. In figure 6 theresults were averaged for those configurations for which results wereobtained f?xnntwo models. Of the configurations tested, the constant-chord ailerons exhibited the lmgest variation of effectiveness overthe Mach number range and the lowest effectiveness at supersonicspeeds. The full-delta ailerons have the hfghest effectiveness atsupersonic speeds. For the same ratio of control area to wing area, theeffectiveness of the half-delta ailerons is less than that of the full-delta ailerons and, of the configurations tested, has the mal.lestvariation over the speed range.
Ccmmrisonwith theor.v.- h figures 7 end 8 the expertiental resultsobtained are capered with results calculated using methods based on thelinearized equations of supersonic flow. The results fpr the constant-chord ailerons’were calculated according to reference 5 with a correctionfor the effects of finite trailing-edge angle given in reference 6. Thecalculated results for the full-delta ailercm were obtained &cmreference 7 for the case of the Mach lines behind the leading edge andfrom reference 8 for the case of the Mach lines ahead of the leadingedge. The calculated results for the half-delta ailerons for the caseof the Mach lines ahead of end behhd the leadlng edges were obtainedftmxnreferences 8 end 5, respectively.
Except for the constant-chord ailerons, the agreement betweentheory end experiment is good: This ~eement, however, is probablyfortuitous because other tests (reference 1) have shown that thetheoretical values of both the aileron rolling mcment sndwing dampingin roll are higher then experimental values by roughly the same factor.The experimental values for the constant-chord ailerons are considerablylower than the theoretical values, probably due to the adverse effectsof wing end fuselage boundary layer which would be larger for theconstant-chord eUerons then for the other ailerons tested. lh order toestablish if there was agreement between the shapes of the theoreticaland experimental curves, the results for the constant-chord ailerons infigures 7(a) end 8(a) were plotted as relative values in figure 9. IYOmfigure 9 it appears that the theory can, at least, predict the shape ofthe effectiveness curves for constant-chord ailerons.
,
QSO shown in figure 8(c) is the rolling effectiveness at a ~chnumber of 1.9 obtained in the wind-tunnel tests of configuration 7reported in reference 1. Good agreement is obtained between the wind-tunnel test point and theory. The theoretical curve also agrees well withthe results of the present tests.
—CONCLUSIONS
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The folluwing conclusions regarding the rollQ effectiveness ofdelta wing-aileron configurations are basal on the tests reported herein:
1. The rolling power of constant-chwd p.l@n ailerons on deltawings having a 45° leading-edge sweep is reduced abruptly in the Machnumber range from 0.9 to 1.0. Increasing the leading-edge sweep to 600decreased the abruptness of the loss of effectiveness at trsnsonicspeeds end increased the effectiveness at moderate supersonic speeds,but had little effect at the limits of the Mach number range investi-gated (M=. O.8tol.9).
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2. For a given ratio of control area to wing srea, full-deltaaileron configurationshave a higher level of yelling power at super-sonic speeds than constant-chord or half-delta ailerons. —
~. Mf-delta ailerons have the smalle”stvariation of effectivenesswith Mach number of the configurations tewted. At M= 1.9 the rollingeffectiveness is comparable to that of tho full-delta control.
4-.The wing-aileron rolling effectiveness of hsd.f-deltaend full- ●..delta aileron configurations can be accurately predictedby calculationsbased on the linearized flow equations. WnIJ.er calculations forconstant-chordplain ailerons predict accurately the variation ofeffectiveness with Mach ntmiber,but yield absolute values of effectiveness
.,.
* which are considerably higher than the measured values.—
Lengley Aeronautical LaboratoryNational Advisory Ccmmittee for Aeronautics
Langley Air Face Base, Va.
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REmRENcEs
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1. Conner, D. Williem, and Msy, Ellery B., Jr.: Control EffectivenessLoad end Eh.ge-McnnentCharacteristics at a Tip Control Surface ona Delta Wing at a Mach Number of 1.9. NACARM L9H05, 1949.
2. Sendahl, Cerl A., and Msrino, Alfred A.: l?&ee-Fli.ghthvestigationof Control Effectiveness of Full-SIem 0.2-Chmd Plain Ailerons atIkLghSubsonic, Transcmic, and.Supersonic Speeds to Determine SaneEffects of Section Thickness and Wing Sweepback. NACARML7D02,1947. .
3. Sendahl, Carl A.: I@ee-llight limestigation of the Rolling Effective-ness of Seversl Delta Wing-Aileron Configurations at Transonic sndSupersonic Speeds. NACA RM L8D16, 1948.
4. Brownj Clinton E., end Adems, Mac C.: DQing in Pitch and Roll ofTriangular Wings at Supersonic Speeds● NACA Rep. 892, 1948.(Formerly NACA TN 1566.)
5. Tucker, Werren A., and Nelson, Robert L.: Clsracteristics of
1
Triangular Wings with Constant-Chord Partial-Span Control Sur es& Supersonic Speeds. NACATN 1660, 1948.
I
6. Tuc r, W&ren A., and Nelsm, Robert L.: Theoretical Cheracteris icsSupersonic Flow of Constant-Chord Partial-Sp3n Control Surf esRectangular Wings Having Finite Tb3ckness. NACA TN 1708, 1 8.
7. Tucker, Warren A.: Characteristics of Thin Triangular Wings with -Trisnguler~ip Control Surfaces at Supersonic Speeds with Mach Linesbehind the Leading Edge. lUCA TN 1600, 1948=
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8. Lagerstrcnn,P. A., and Grahsm, Martha E.: Linearized Theory ofSupersmic Control Surfaces. Jour. Aero. Sci., vol. 16, no. 1,Jen. 1949, p~. 31-34.
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Figure 2.- Details of configuration tested.
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Constant-chord ailerons
Configuration I
Full-delta ailerons
Configuration 3
= 45°.
Half-delta ailerons
Configuration 5
Constant-chord ailerons
Configuration 2
Full-delta ailerons
Configuration 4
Half-delta ailerons
Configuration 6
Half-delta ailerons
Configuration 7
(b) A~oEe =&)”.
Fi&nrre3.- Photo=anhs of confimrations tested.—“-— - .*-
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