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EXPLORATORY INVESmGATION OF LEADING-EDGE CHORD-EXTENSIONS TO IMPROVE THE LONGlTTuDINAL STABILITY CHARACTERlsTICS OF TWO 52' SWEPTBACK WINGS t By G. Chester Furlong Langley Aeronautical Laboratory Langley Air Force Base, Va.

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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON .E March 10, 1950 NACA RM L50A30

NATIONAL ADVISORY COMMITTEZ FOR AEROHAmICS

By G. Chester Furlong

Results are presented of explorgtory tests obtainedwith leading- edge wing chord-extensions on two 52 meptback wings. Both wings exhibit a vortex flaw similar to that whichhas been observed OIL tri- angular wfnga. One wing incorporated circular-arc airfoil sections, and the otherwing incorporated IVACA 641-m airfoil sectians. The w-as -85 wlng aspect ratio approximate- 2 for each . The tests were 6con- ducted at a Reynolds nmber and Mach numberin the viclnityof 6.0 x 10 and 0 .E, respectively. - The results constitute preliminary observationsand indicate tht, on the wing incorporating circular-arc airfoil sections,sharpnose ’ chord-extensions reducedthe aerodynamic-center shiftobtained over the entire lift rangewith the plaFn wing from 37 to 5 percent of the mean aerodynamic chord. The chord-extensions coveredthe outer 0.25 semispan and had chords0.147 of the mean aemmcchord. With a reduced chord or span, the sta3ilizing effectivanessof the chord-extensionswas re- duced inthe vicinity of maximum lift.

The applicationof round-nose chord-extensionsover approximately the outer 0.43 semispan m’ the whg fncorporating NACA 641-~airfoil sections substantially improved the lmgitudinal stability.

An analysis of the flow would indicate that the dimion, at the plan-form discontinuity pmvided by the chord-extensione, of the vortex flow emanating fram the wing apex, and for sqconditima the. formation of a secondary vortex flow over the chord-extensions, wa8 responsible for the hprovement in the longitudinal stabilityof the two wings L tested.

Y

UNCLASStFIED 2 K4CA RM L50A30

The longitudinal stability obtained KJth chord-extenaims w&e ae good as that obtained with extensible leu-edge flaps.

INTRODUCTION

' The presence of a vortex type of flow an Elweptback win@ of certain aspect ratios and thickness rattos can producg large undeeirable changes in langltudinal stability at lift coefficients well below maximum lift (reference I). The vortex flow referred to is similar to that aich hae been observed on thin or @harp-edge triangular winga and which is quali- tatively described Fn referenae 2.

The data presented in references 1, 2, and 3 Fndicate that the presence of the vortex flow over the tip sections caused an incream In lift at these sections which wae reepmsible for the initial atabilizlng aerodynamic-center shift. This initial ahift in the aerodynamic center, furthermore, constitutes a large portion of the aemdynemic-canter travel over the entire lift range. It was assumed that the ellmhatian or dissipation of this vortex flow over the tip sectians would remit in an appreciable rsductian fn the over-all aerodynamic-center travel. Eane reduction was obtained in the tests reported in reference 1 through the applicaticmof extensiB1e leadfng-edge flaps. Subsequent teste on the same wing indicated that the greatest reducticm in aerodynamic-center trahel that could be obtained was critically dependent an the spnwiee location of the plan-form discorit-lpuity that occurred at the Inboard end of the 1eading"edepe flape. The dependence of the improvement ln stability on the plan-form dAscmtfnuity suggested the possibility that malJ"span chord-extensicme could provide the discontinuity necessary to upset the effect of the vortex flow. Such a device could be extended when required, or if the high-epeed characteristics proved satisfactory, it could be fixed. In either case, the chord-extanaims would appear to be hSS caplicated than the staJ-1-control devices currently being used on Bwept- back wings which exhibit the vortex flaw.

The present papr preeants the resulte of exploratory tests obtained with outboard wing chord-extensiane on two 52* aweptback rhgs xhich exhibit the vortex flow. A camprism le made of the effects of extensible leading-edge flaps with the effects of the chord-extmaione on the lift, drag, and pitching-mmnt characteristica of each wing. One wing fncolporated _circular-am aieoil ~c.tf~,and the other .wing incorporated NACA 641-112 airfoil eectfcme. The aspect ratio was approxl- mately 2 "6 for each wing. The term "chord-ertaneion" ie med herein to designate a device which extend8 the normal wing chord over an outboard portion of the leading - d NACA RM L50A30 t-I 3 edge. The purpose of this device is to improve the 1ongltudFnal stability characteristics of meptback wings wbfch exhibit a vortex flow uing ~11 the upper surface. ..

SYMBOLS

pitching-moment coefficient about 0.25E

- =aC approximate aerodynamic-center location, percent c

a angle of attack, degrees

X Reynolds nrrmber

¶ stream dynamic pressure, pounds per quare foot

M Mach number

S wFng area (basic wing) b

E

A aspect ratio h taper ratio

C local chord

Y spanwise ordinate

rate of chane of pitchm-mcnnent coefficient with lift coefficient 4 NACA RM L50-0

MODELS, TESTS, p;[QD CORRECTIOIVS

Model.- The delplan fom, together with pertinent geometric dimensians, are shown In figure 1. meither model had dihedral or twist. The wing incorporatln@; symmetrical circular-arc airfoil secttone normal to the 0 .?O-chord line (line of maximm thichneee) had a 9 .€!-percent- chord thicknees at the root & a 6.1-percent-chord thickness at the tip. The other wing incorporated NACA 641-~airfoil sectim normal to the 0.282-chord line.

The extensible leading-edge nape and xLng chord-ertansiani teated on each wing model are &own in figure 2.

The 52O meptback KLng havtng an aspect ratio of 2.84 and incorpo- rating circular-arc airfoil. sections is referred to as the "circular-arc" wing, and the 52O eweptback wing having an aspect ratio of 2.88 and .. incorporating RACA 641-112 airfoil sections is referred to a6 the "64-series" wing.

The 64-series wing is ahown mounted an the two-support system of the Langley 19-foot pressure tunnel Fn figure 3.

Tests.- The tests were conducted in the Langley 19-foot pressure tunnel with the air cmpreeeed to an absolute preseure 0-f 33 pound6 per I 9q-e inch.

Data were obtained an the circular-arc wing and 6bseriee wing at the conditione listed in the fallowing table:

Configuration .. plain wing 6.0 x 166 0.12 46 Circula" King All othere' 5.5 .11 , 40 Plain whg 6.0 .I2 46 64-series wing All others 6.0 .12 46

The Reynolds numbem were based an the respective mean aerodynamfc chards. ~ACARM ~50~30 5

Lift, drag, and pitching-moment data wgre obtainedthrou& an angle-of-attack rangefrom approximately -4 to an angle beyond maximum lift . Correctims .- The lift, drag, and pitching mament have been corrected for support tareand strut interferenceas determined by tare tests. The angles of attackand drag data have been corrected for Jet- boundary effects by the method presentedin reference 4. The pitching- moment data have been corrected for jet-boundary effectsan extension by of the method presented in reference 4. ~naddition, the angles of attack have been correctedfor air-stream misalinement.

Force Characteristics

The lift, drag, and pitching-mament cha;racteristics obtainedfor the circular-arcwing equipped with chord-extensions of variousspans and chords and an extensible leadfng-edge flap arepresented Fn figures 4 to 7- Similar data are presented for the 64-series whg in t figures 8 to Il . Cmpari~smsof the longitudinal ekbility obtained with these devices have been inmade figures 12 and 13 where variations. of instantaneous aerodynamic centerare plotted against lift coefficient * for the circulgr-arcand 64-series a,respectively. The values of aerodynamic centerin the high-lift rangeare of an appmrlte nature, inaemuch asthe drag has not been taken into account in the calculations.

Ih the presentinwstigatian the inboard-end lmatians of the chord- extensims were selectedti the basisof moreertemive tests (unpuklished data) with extensible leading-edge flaps. Circular-arc w%.- The plain wing exhibited large undesirable shifts in aemdyndc center throughout the lift range. The ehift of the aero- Qnamic cmter from its mostforward. to most rearward positicm between zero andmaxhum lift amounted. to 37 percent of themean aerodynmic chord (figs. 4 and =(a) ) . Outboard chord-extensions having 6-inch chords (14.7 percent E) and located over approximately the 25outer per- cent of the semispas reducedthe aerodynamic-center travelof the plain wing to a ?-percent mean-aerodynazaic-chord shiftbetween zero and nmximm lift (figs. 4 and 12(a)). The spans of the 6-inch chord-extensions were reamed to o .13 and o .06 of the semispan while the inbard ends of the chord-extensicme were fixedat the same spanwise position. The pitching- moment characteristics obtainedwith the 6-inch chord-extensionsof reduced spans are presented in figure 4. The variatiane in aerodynamic center with lift coefficient obtained withthe chord-extensions of reduced span (fig. E(a)) indicate increasesin aemdynemic-center travel 6 NACA RM ~50~30 in the high-lift range over that obtained with 0.2553/2 chord-extaneione . Ln both cams, however, the reduced-span chord-extarmione eliminated the initial aerodynamic-center dift and hence materially reduced the mer- all aerdynamlc-center travel of the plain wlng. With the chord of the chord-exteneiamreduced to 3 inches (7.4 percent e), the aerodynamic- center travel was caparable to that obtained xfth the 6-inch chord- extanelms of reducedspans; that is, the initial aemdynamic-center shift wae eliminated but meaeurable shlf'te occurred in the hi&-lift range (fig. 12(b)). .Reductions in the spans of the 3-lnch chord- extenaims allowed aerodpmnlc-center ahifts in the high-lift range which may be comidered objecticmable (figs. 6 and U(c)). The application of chord-erteneiaus had, in general, a straightening effect on the lift curves (figs. 4 to 7). It ie interesting to note that amall Increases in maximunn lift were obtained that were eamswhat greater than the increaee in area pmvided by the chord-ezteneims. The drag data (figs. 4 to. 7) indicate that chord-ertecneims had practically a negligible effect on the drag throughout the lift range although a Blight decrease is obtained in the maximum-lift range. The data preeented in figure6 6 and U(d) indicate that chorb- exteneiane are RH effective in reducing the Etercdynmic-center travel of the plain wing as the extensible leading-edge flaps. The extensible leading-edge flaps, it Bhould be pointedout, caueed an increme in drag over most of the lift range.

64-eeriee wing.- The &ift of the aemdynmlc center from ita most forward to most rearward position between zero and maximum lift for the plain wing amounted to an 87"percent man-aerodynemic-chord travel (figs. 9 and 13(a)) . Outboard chord-exteneione having 6-inch chord8 (eimilar to thoee ueed on the circular-arc wing) and located Over approximately the outer 0.43 percent of the eemiepan reduced the large aerodynamic-center trawl of the plain wing appreciable between zero an& mimum lfft (figs. 9 and 13(a)). To indicate the effect of the mae ehape of the chord-exteneians on a wlng havhg subeanic airfoil sectlom, tests were aleo made xith round-nose chord-exteneton (fig. 2). The reeulteobtained (fige. 8, 9, l3(a), and 13(b)) indicate that the bitid aerodynamic-cat- shift is delayed to a hlgher lift coefficient and the magnitude of the shift is reduced with either the round- or eharp-nom chord-erteneione. The data do Indicate, however, that the round-nose chord-exteneims me superior to the sharp-noee chord-extenelma in regard to langitudinal stability, max- lift, and drag.

The results obtatned on the 64-series wing, even with round-noee chord-extensions, are not eo- favorable ae thoee obtained with chord- ertensions QD. the circular-arc wing. This fact ie eapecfdly true from cmsideratiom of the increaee in drag fn the moderate-lift range and to NACA m ~50~30 ” 7

-some extent the lcmgitudinal stability just at maximum lift. The increase in drag was greatly reduced by using round-nose rather than sharp-nose chord-extensions on the &-series wing. InaErmuch as the round nose used in these tests was arbitrarily selected, the possibility exists that ft is not the optimum nose ahape.

The effects of variaticms in length of chord and span of the chord- extensians are aimflar to those obtained on the circular-arc WLng .(figs. 8, 9, 10, and 13(c)). The 6-hch round-nose chord-extensions provide a greater stabilizing effect than the extmsible leading-edge flaps. The 3-Fnch round-nose chord-extensions wereabout equally as stabilizing aB the ext;ensible leading-edge flaps.

Flow Chamcteristics

Althow pressure-distribution teats would be required to explain fully the action of the chord-extensions, sone caments can be made on the baeis of tuft probing and &e force data obtained. -

Visual flow observations, by means of a tuft attached to a wooden probe, indicated that the vortex flaw lies alang.th0 leading age from the wing apex to the plan-f om discontinuity. Although the probe observations then became indefinite,, it appeared that the vortex flow trailed off the wing at this statim. It was observed that in the case of the sharp-nose chord-&ensions a secondary vortex formed over the chord-extensime. No tuft obeemations were made on the 64-series wing with the round-nom chord-extensians . The ellmlnation of the initial aerodgnamic-center shifts which occurred on the plain wings when the vortex flow formed, by even the smallest-chord and malleat-span chord-extensions, Indicate that the discontinuiQ in plan form evidently diffused the vortex flow emanating from the Xing apex in such a way as to prevent any additional lift over the tip sections.

The strength of the secondary vortex which forme over the chord- ertansims is dependent on the span of the chord-exteneions and angle of attack. This fact is Fndicated by the force data where the large-span chord-extensions provided slightly more negative pitching mcHnant near maximum lift than the ahorter-span chord-extensions and, hence, were sli&tly more effective in reducing the aerodynamic-canter travel of the wings. Ih addition, the secondary vortexprobably accounts for %he - increase in maximum lift notaccounted for by thearea increase. 8

Ln both cases the dditiai of chord-extensions produced straighter lift curves althou& not in the same manner. In the cam of the circular- arc wing, chord-ertfmeions caused reductions in lift coefficient in the moderate-lift range that can be attributed to the prevention of the increases Fn lift over the tip sections which normally occur when the vortex flow is formed. In the caseof the plain 64-series wing, the inflection in the lift curve is rather abrupt and occurs at a moderately high angle of attack. The lift cme is stral&tened with chord- extensions by an increase in lift throughout the low- and mcderate-lift range, and this additional lift is due to the increme in wing area in the cam of the round-nose chord-extensiona and to -the increase in area and eecmdary vortex in the cam of the eharp-nose chord-extension6 (fig. 9). *

The drag due to the chord-extensim on the circular-arc wing wae negligible. The increase to be expected fran the increased wing area was probably compensated for by the diffusion of the vortex flov ema- nating fram the wing apex. . In the cam of .the .&-eerie8 wing, the aharp- nose chord-extensions formed a vortex flow at much lower angles of attack than the me at which iC normally occurred on the plain wing and, hence, caused a drag increase in additian to that obtalne3 with the increase in wing area. The increase In drag due to the round-nose chord-extensions was much lees became of the fact that a aecandary vortex flow ISnot c precipitated . The results obtained on the circular-am and 64-series vings Fndlcate that altering the effects of the vortex flow by meam of chord-extensions extending over approximately the outer 25-pement and 43-percent semispan, respectively, produced satisfactorylangltudfnal stability. Thus, two wings having the me sweep meand aspect ratio require greatig Uf- ferent spans of chord-extension, which is attributed to the fact that the formation, strength, and position of the vortex flow is, amang other things, dependent 0x1 leading-edgeradius. The two wings caneideredhere differ greatly fn thie respect.

The results obtained with leading-edge wing chord-extensions an two 52O exeptback wings which have aapect ratios of approxIm8mRtel.y 2 -85 indicate that:

On the wing incorporatine circular-- airfoil sections, aharp-nose chord-extensions reduced the Eterodynmic-center &if% obtained over .the entire lift range with the plain wing from 37 to 5 ‘prcent of the mean aerodynamic chord. The chord-extensionscovered the outer 0.25 semispan and had chorda 0.147 of the mean aem~cchord. With a reduced chord 9 or span, the eta%ilizing effectiveness of the chord-extensicme w&8 reduced in the vicinity of marimmu lift.'

The application of round-noee chord-extenaim over approxlmately the outer 0.43 semispan on the WLng incorporating NACA 641-112 airfou secticms substantially improved the Imgitudinal stability.

An analysis of the flow would fndicate that the diffusion, at the plan-form discontinuity provided by the chord-exbensiane, of the vortex flow emanating from the wing apex, and for ecme cditims the formation of a secondary vortex flow over the chord-extensicme, was responsible for the impmvement in lmgitudimd stabili- of the two winga tested.

The longitudinal stability obtained xith chord-exteneione was as good as.that obta*ed with extensible lenAfnn-edge flaps.

Langley Aeronautical Laboratory National Advisory Ccmrmittee for Aeronautice Langley Air ForceBaae, Ta. 10 RACA RM ~50~30

1. Fitzpatrick, James E ., and Foster, Gerald V. : Statfc Longitudinal A6rOdynamic Characteristics of a 52O Eweptback Wing of Aepect Ratio 2.88 at Reynolds Numbers fran 2,000,000 to il,OOO,OOO. NACA liM L- ,’ 1948. 2. Uileon,Herbert A., Jr., and Lovell, J. Calvin: Full-scale Inveati- gation of the Maximum Lift and Flow Characterietice of an Airplane Having Approximately Triangular Plan Form. NACA €M L6K20, 1947.

3. Lange, Roy E., Whittle, Edvard F., Jr., and Fink, Marpin P.: Inveeti- gation at Large Scale of the Preseure Diatributian and Flow Phenomena over a Wing with the Leading Edge Swept Back 47.5O Having Circular-Arc Airfoil Sectims end Equipped with Drooped- Nose and Plain Flaps. NACA RM L9Cl5, 1949. 4. Eiaonetadt, Bertram J.: Boundary-Induced Upwaah for Yawed and Swept- Back Wings in Cloeed Circular Wind TUIXI~~E.XACA TN 1265, 1947. . .. .-

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Figure 1.- Ceometry of 52' sweptback wings having circular-arcand NACA 641-112 airfoil sectlona. All dimensions are in inches. r ....

A B Le. fk7p Le. chord- exbnsmn

Secfiian K-X (mtmpd) of cimhr-arc wing. AkuiLMness I V

Figure 2.- Details of typical leading-edge devices for circular-arc and 6bseries xings. All dimension8 are in inches unless otherwise specified.

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wr Figure 3.- NACA 64-ser1es 52' sweptback wing mounted in the Langley 19-foot presaure tunnel.

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Fimre 4.- The effect Qf a 6-inch leading-edge chord-extenaion of several spans on the aWOwic characteristics of a 52O sweptback circular- arc wing. 16 RACA

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Figure 5.- The effect of a 3-inch and 6-.inch leading-edge chord-extension on the aerodynamic characteristics of a 52' sweptback circular-arc wing; 0.23/2 span. 18 WARM L5QA30 ,

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Figure 5. - Concluded. WA RM ~~30

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Figure 6.- The effect of a 3-inch leading-edge chord-extension .of several spans on the aerodynamic characteristice of a 520 sweptback circular- arc wing. 20 NACA RM ~50~30

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Figure 6.- Concluaed. WARM LWQO 21

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Figure 7.- A comparison of the effects of leading-edge flaps and 3-inch leading-edgechord-extensions on the aerodynamic characteristics of a 52O sweptback circular-arc wing; 0.256/2 span. 22 NACA RM L5OA30

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Figure 7.- Conciuded. ~CARM ~5a~30

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-4 s2 0 -2 .4 .6 .8 1.0 /.2 1.4 - CL Figure 8. - The effect of a round and sharp 6-inch leading-edge chord- - extension on the aerodynamic characteristics of an NACA 64-series 52O sweptback wing; 0.43b/2 span; R = 6,000,000. 24 ZACA RM ~5~30

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Figure 8.- Concluded. WA RM ~50~30

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Figure 9. - The. effect..of a round and sharp 3-inch leading-edge chord- extension on the aerodynamic characteristics of a.n NACA 64-series 52O iweptback wing; 0.43b/2 span; R = 6,000,000. 26 EIACA RM L5OA30

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Figure 9. - Concluded. NACA RM ~5~~30

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Figure 10.- The effect of a round 6-inch leading-edge chord-extension of several apana on the aerodynamic characteristics of am NACA 64-series 52O sweptback wing; R = 6,OOO,OOO. . 28 WA . .60

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Figure 10. - Concluded. NAGA m ~50.~30

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Figure 11.- A comparison of the effects of leading-edge f-kps and round 3-inch leading-edge chord-extension on the aerodynamic characteristics of an NllCA 64-series 52O sweptback wing; 0.43b/2 span; R = 6,000,000. .60

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-PO 0 .e .4 .8 1.0 1.2 0 .2 .4 .6 .8 i,O 1.2 - CL Figure 12.- Aerodg?lamic-ctnter variation with lift coefficientof a 52' sweptback circular-arc wing.

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