JOURNA AIRCRAFF LO T Vol. 31, No. 3, May-June 1994

Lift Enhancement of an Airfoil Using a Gurney Flap and Vortex Generators

Bruc . StormsL e * Sterling Software, Moffett Field, California 94035 and Cor . JangyS t California Polytechnic State University, LuisSan Obispo, California 93407

Experimental measurements of surface pressure distributions and wake profiles were obtained for an NACA 4412 airfoil to determine the lift, drag, and pitching-moment coefficients for various configurations. The addition oGurnea f y flap increase maximue dth m lift coefficient from 1.96 o decreaset 1.4d p an 9,u drae dth g neae rth maximum lift condition. There was, however, a drag increment at low-to-moderate lift coefficients. Additional nose-down pitching moment was also generated by increasing the Gurney flap height. Good correlation was observed betwee e experimennth Navier-Stoked an t s computation e airfoith f o sl wit ha Gurne y flapo Tw . deploy able configurations were also tested with the hinge line forward of the trailing edge by one and 1.5 flap heights, respectively. These configurations provided performance comparable to that of the Gurney flap. The applicatio vortef no x generator baseline th o st e airfoil delayed boundary-layer separatio yielded nan increasn da e in the maximum lift coefficient of 0.34. In addition, there was a significant drag penalty associated with the vortex generators, which suggests that they should be placed where they will be concealed during cruise. The two devices were also show woro nt k wel concertn i l .

Nomenclature desig f mechanicallno y simpler high-lift systems wit dego hn - radatio performancen ni . However maintaio t , e hignth h lift Ctl = section drag coefficient, dlqc C, = section lift coefficient, IIqc coefficients require r approacdfo landingd han technolw ne , - 2 og s needei y o providt d e lift enhancemen d separatioan t n sectio= Cm n pitching-moment coefficient, m/qc control. Cp = pressure coefficient, (p — p-^lq c = reference airfoil chord, ft One candidate technology is the Gurney flap which consists d = section drag, Ibf osmala f l ordee plateairfoie th th f 1-2 n ro f o , o l% chor n di h = flap height, ft height, locate trailine th t da g edge perpendicula prese th -o t r LI lift-to-dra= D g ratio sur eairfoie sidth f eo l (Fig . Thi1) . s devic originalls ewa y used = sectio/ n lift, Ibf on the airfoils of performance race cars to increase the down m = section pitching moment, ft-lbf force for the lateral traction required during high-velocity 1 stati= pc pressurei ps , turns. Liebeck tested a 1.25% chord Gurney flap on a New- dynami— q c pressure, ipVi 2ps , V = freestream velocity, ft/s x = axial distance from airfoil leading edge, ft a = angle of attack, deg d = boundary-layer thickness, in. = densitp f airyo , Ibm/ft3

Subscripts

Downloaded by UNIVERSITY OF CAMBRIDGE on October 16, 2014 | http://arc.aiaa.org DOI: 10.2514/3.46528 ma = maximux m value sc = freestream value

Introduction payloaE H d rangan df subsoni o e c transport e dicar s- T tated d oftean , n limitede performancth y b , f theio e r high-lift systems. These systems are generally quite complex, chor% 2 Figd 1 .Gurne y 441 a fla n p2o airfoil. consisting of a leading-edge slat and two or three trailing-edge flaps. The high maintenance and weight penalty associated with such configurations have provide n impetue da th r fo s

Presented as Paper 93-0647 at the AIAA 31st Aerospace Sciences Meeting, Reno , JanNV ., 11-14, 1993; received Feb, 199322 - . re ; vision received Apri , 199319 l ; accepte r publicatiofo d n Apri, 19 l 1993. Copyright © 1993 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. * Aerospace Engineer, Fixed Wing Aerodynamics Branch, MS 247- 2, NASA Ames Research Center. Member AIAA. tGraduate Student, Aeronautical Engineering Department. Stu- dent Member AIAA. Fig 2 . Submerged vortex generators detaile Refn di . 9 . 542 STORMS AND JANG: LIFT ENHANCEMENT 543

• Instrumented center section ' Boundary layer fences > 7ft. Tip Row pressurf so e taps Tip extension extension

Boundary Supports layer fence Support <-1.75ft-3 6.5 ft. • -1.75 ft>

Fig. 3 NACA 4412 in the NASA Ames 7- by 10-ft Wind Tunnel test section.

man airfoil which resulted in an increase in lift and a slight reduction in drag. Larger lift increments were observed for greater flap heights drae th gt increasebu , d noticeably beyond height f approximatelo s chord% 2 y . Experimental investi- gations incorporating a Gurney flap on the trailing element of multielement race-car airfoils were later conducte Katdby z and Largman2 and Katz and Dykstra.3 Both investigations reported a substantial increase in lift (as much as 50% using chor% 5 a d Gurne e ycorrespondin th flap) t s bu , wa D gLI reduced over the design angle-of-attack range. A computational study of an NACA 4412 airfoil with a Gurney conductes flapwa Jany db t al.gincompressible n 4A e -2.2 Navier-Stokes code was used to compute the flowfield about the airfoil with Gurney flap heights ranging from 0.5 to 3.0% chord. With the addition of a Gurney flap, the computations Fig 4 . Span wise pressure coefficient variatio t 0.25na variour cfo s predicte dsignificana t lift increment that increased with flap angle f attackso . size, althoug t linearlyhno e computeTh . d pressure distribu- tions indicated increased loading alon entire gth e airfoil when primare Th y objectiv presene th f eo tprovid o t stud s ywa e compared wit baselinee hth , particularl suctioe th t ya n peak an experimental dat aperformance basth n eo Gurnee th f eo y and near the trailing edge. The computations also predicted flap on a single-element airfoil. Several flap sizes were tested draa g penalt t low-to-moderatya e lift coefficients. in order to determine the effect of Gurney flap height and to validat e Navier-Stoketh e s computations conducted previ- Similar trailing-edge devices, such as the wedge flap and 4 the divergent trailing edge, have been designed to increase ousl Jany yb t al.g e secondar A y objectiv determino t s ewa e airfoil efficienc t cruisa y e conditions e wedgTh . e flapa s i 5 the effectiveness of vortex generators in delaying airfoil stall. ramp located on the airfoil lower surface at the trailing edge with a height between 0.5-1.5% of the chord. At transonic Experimental Setup velocities, the wedge flap lowers the angle of attack for a currene Th t experimen 10-f y conductes b - ttwa 7 Win e th dn di given lift coefficient and reduces the drag for lift coefficients Tunne NASe th t a lA Ames Research Center. This facilits yi above 0.52. Divergent trailing-edge (DTE) airfoils6 have also a closed-circuit wind tunnel incorporating a test section 15 ft shown significant improvement in transonic performance when long with a constant height of 7 ft and a width of 10 ft with applie supercriticaa o dt l airfoil. This modification incorpo- divergence% 1 a . turbulence-reducino Thern e ear g screens rates strongly divergent upper and lower surfaces to produce in the circuit and the test section turbulence intensity level is a blunt trailing edge. Relativ baselina o et e supercritical air- ft/sl datobtaine0.770 s Al .10 a wa %t chora a t da d Reynolds Downloaded by UNIVERSITY OF CAMBRIDGE on October 16, 2014 | http://arc.aiaa.org DOI: 10.2514/3.46528 airfoiE DT foille exhibitth , increasn sa t loadinaf n ei wels ga l numbe f approximatelo r milliono ytw . as a decrease in compressibility drag for a given lift coefficient. modee Th l consiste NACn a f do A 4412 wing wit chorha d Vortex generators have a long history of successful appli- of 36 in. spanning the 10 ft width of the wind tunnel (Fig. 3). cation to aerodynamic surfaces to prevent flow separation and Boundary-layer trip strip chor% s wer10 d d ean place5 2. t da increase efficiency. By generating streamwise vortices, these uppee o nth lowed ran r surfaces, respectively airfoie s Th . wa l devices serve to energize the boundary layer which tends to instrumented with a total of 200 pressure taps which composed dela floe yth w separation commonly encountered unde- ad r one spanwis thred ean e chordwise rows e spaTh .n wise taps verse pressure gradients. First introduced by Taylor,7 the vane- were locate t one-quarteda r chord, whil chordwise eth e rows type vortex generator mose th e t ubiquitoussar , consistinf go were located at midspan and one-half chord on either side. a flat plate mounted normal to the surface at a small angle d pitching-momene lifan t Th t coefficients were determined of incidence to the local flow. Because they typically extend by an integration of the centerline pressure distribution, while well beyond the boundary layer, this type of vortex generator the spanwise and additional chordwise rows of pressure taps produces considerable parasitic drag.8 Submerged vortex gen- served to monitor the two-dimensionality of the flow. In ad- erators, however, derive their name from the fact that they dition, boundary-layer fences were mounte airfoie th 1 n 2 dlo measure on the order of the boundary-layer thickness in height in. from e wall eacth promoto f t sho e two-dimensional flow and are, therefore, mostly "submerged e boundaryth n i " - on the instrumented center section. layer flow e submergeTh . d vortex generators investigaten di The drag coefficient was determined by an integration of this study were the Wheeler wishbone9 type consisting of V- the stati d totaan c l pressures measured wit a wakh e rake shaped ramps pointing in the downstream direction (Fig. 2). situated 0.7 chord downstream of the airfoil trailing edge. The Each device generate counter-rotatino stw g vorticesf of e on , rakcomposes estatiwa 9 tota 1 d 9 cf an l do pressur e probes each edge, that grow with downstream distance. distribute d. wit ovein h6 3 rclusterin g nea centerlinee rth A . 544 STORMS AND JANG: LIFT ENHANCEMENT

midspan — B lef- t --A--right

0.4 0.6 0.8 1 c x/ ) b c x/ a)

Fig 5 . Pressure distribution cleae th nn so airfoi deg5 1 l showin. = a charactee ) b gth floe d th wan f rg neao beyon d de ran 2 1 d = C /maxa ) :a

100- ^ 0.5%c Gurney flap Baseline ^^ l-0%c flap 80- ""

0.5%c Gumey flap 60- 1.0%c Gumey flap L/D 40- l.5%c flap 2.0%c flap 20-

0 0.5 0.75 1 5 1. 1.21.755

Fig .6 Effec Gurnef to y flap heigh chordwisn to e pressure distribution Fig. 8 Effect of Gurney flap height on lift-to-drag ratio. deg9 = . aa t

2- 0.5 %c Gurney flap 2.0%c Gurney flap—— 1.5%c Gumey flap-rr* V ,-;S,'- / ^^L 1.5- 1.5- &•,'•* --*,' /

1- Baseline 1- 2.0%c A 0.5%c Gurney flap Baseline 1.5%c / 1.0%c 0.5- 0.5- -0.25 -0.2 -0.15 -0.05 10 15 C a, degrees Fig 9 . Effec f Gurneo t y flap heigh pitching-momenn o t t coefficient. Downloaded by UNIVERSITY OF CAMBRIDGE on October 16, 2014 | http://arc.aiaa.org DOI: 10.2514/3.46528

0.06

Fig. 7 Effect of Gurney flap height on the lift and drag coefficients. Fig. 10 Deployable miniature split-flap configuration. motor-driven traverse was used to center the rake vertically on the airfoil wake at the midspan location. Both the wake 2.0% chord. Wishbone vortex generators with a height of rake pressure surfacd san e pressures were measured witn ha 0.5% chord were then tested with and without a Gurney flap. electronically scanned pressure syste r rapimfo d data acqui- e vorteTh x generators were mounted acros entire sth e span sition. All aerodynamic coefficients are reported in the win dchor% 12 d t a fro leadine mth g edge werd an , e evenly spaced axis system. with a distance of six vortex generator heights between cen- o lift-enhancintw e Th g devices were studied both inde- ters. At this location, their 0.5% chord height corresponds to pendently and in concert. The effect of Gurney flap size was approximately three to four boundary-layer thicknesses and first investigated by testing flap heights of 0.5, 1.0, 1.5, and they are, therefore, only partially submerged. STORMS AND JANG: LIFT ENHANCEMENT 545

1.5-

Baseline i- - B Gurne-- y Flap --A---Split flap Ax/h=1.0 -—*-—Split flap Ax/h=1.5 0.5- 0 5 10 15 0 0.02 0.04 0.06 , degreea s Cd Fig 1 Compariso1 . Gurnee th f no y flasplid pan t flap configuration = 1.25%c) h ( s .

I——————T 0.5 1 1.5 2 Gurney flap height, %c Fig 2 Compariso1 . f experimentano computed an l d pressure distri- Fig 5 Variatio1 . f lifno t coefficient increment with Gurney flap height. butionbaseline th r deg9 sfo e = .airfoi a t a l Results and Discussion e extenTh f two-dimensionao t l flogooa s wi d measurf eo o experiment date th a qualit two-dimensionan yi l airfoil testing e spanTh . - computations wise pressure variation for a representative case is presented in Fig . Thes4 . e distributions indicat- e es tha e flos th tw wa -3- sentially two-dimensional at the lower angles of attack, while some three-dimensionality is apparent near maximum lift -2- (a = 12 deg). At angles of incidence beyond maximum lift, ( however, strong three-dimensional effect evidente ar s . These -1- effects can also be observed in a comparison of the three SCDo, chordwise pressure distributions (Fig ) wher5 . e good corre- lation is observed at the maximum lift condition, while three- 0 ^.-.npnnOoOOonnOOO CftO dimensional flow is indicated for the stalled condition. Airfoil pressure distributions and wake surveys were first 1 I i obtaine e baselinth r dfo e clean configuratio d ther an nfo n 0 0.2 0.4 0.6 0.8 Gurney flap height chor% 2 so dt fro withou5 m0. t vortex x/c generators. A comparison of the pressure distributions for Fig. 13 Comparison of experimental and computed pressure distri- various Gurney flap heights is presented in Fig. 6. The pres- Downloaded by UNIVERSITY OF CAMBRIDGE on October 16, 2014 | http://arc.aiaa.org DOI: 10.2514/3.46528 butions with a 1.0 %c Gurney flap at a — 9 deg. ence of the Gurney flap considerably increased the aft loading als s airfoili e ot ofi th t notebu , d that muclife tth incremenf ho t is derived fro ma genera l increas loadinn i ea highe d gan r suction peak. As the Gurney flap height is increased, higher 1.0% Gurney flap loading is noted along the entire airfoil. experiment The lift and drag coefficients are presented in Fig. 7 for 1.6- computations angles of attack from 0 to 15 deg. It can be seen that the additio e Gurneth f no y flap produce significansa t lift incre- ment compared with the baseline configuration. The smallest Gurney flap tested (0.5%c) e yieldeth n i increas n da % 13 f eo maximum lift coefficient, while the larger flaps produced suc- Baseline cessively larger increments, althoug t proportionallyhno p u , experiment 0.8- to 32% for the 2%c flap. Also shown in Fig. 8 is the drag computations e samth pola r e fo rconfigurations t low-to-moderatA . e lift coefficients, there is a drag penalty associated with the Gurney flap which increases with flap height. At higher lift coeffi- 0.4- cients, however, the drag is significantly reduced. As a result, -2 6 10 14 18 the effect on the maximum lift-to-drag ratio is small, but the a, degrees lift coefficient for a given lift-to-drag ratio is significantly in- Fig 4 Compariso1 . f experimentano computed an l d lift coefficient crease s alsi d t oI (Fig note. 8) . d thae maximuth t m lift-to- witwithoud han Gurnea t y flap. drag ratio is reduced for Gurney flap heights greater than 1% 546 STORMS AND JANG: LIFT ENHANCEMENT

2- Vortex generators and 1.25%cGurneyflap ">

1.6-

1.2-

0.8-

0.4- 12 16 20 0.06 a, degrees Fig 6 Effec1 . f vorteo t x generator dra d life an tth g coefficientn so s witwithoud han Gurnea t y flap.

90- vortex generators e vortee effecth Th f . o xt generators wa s Baselii investigated both witwithoud han Gurnea t y flap shows A . n Vortex generators and additioe in th vorteFige , th 16 .f no x generator baseline th o st e 70- 1.25 %c Gurney flap configuration delaye e flodth w separation fro n anglma f o e attack of 12-19 deg, and increased the maximum lift coeffi- cient by 23%. However, it is apparent from the drag polar L/D 50- that ther significana s ei t dra moderatd g an penaltw lo et ya lift coefficients causee parasitith y db c drag associated with 30- the vortex generators. Figure 16 also presents the results from the synergism of the Wheeler vortex generators and a 1.25%c Gurney flap which yielded a 36% increase in the maximum 10- lift coefficien n angla t f attacdeg7 a o e corret1 Th f . o k - 0.4 0.8 1.2 1.6 sponding lift-to-drag ratio (Fig. 17) was significantly reduced for both configuration large th eo t increas e sdu dragn ei . Thus, Fig 7 Effec1 . f vorteo t x generator lift-to-drae th n so g ratio witd han e resultinth g reductio cruisn ni e performance greatly offsets withou Gurnea t y flap. high-life th t benefits obtained from these vortex generators. However, it has been proposed in Ref. 11 that vortex gen- chord. Finally nose-dowe th , n pitching moment about 0.25c erators located nea leadine rth g edg flaa f peo coul stowee db d (Fig. 9) is shown to increase with Gurney flap height although, flae inth p e mai covth f neo element during cruise withoua t again, not proportionally. drag penalty. Similarly, these devices may be placed near the Because of the drag increment at lower lift coefficients leading edge of the main element where they would be con- causeGurnee th y db y flap t wouli , desirable db stoo et e wth cealed by the slat in the cruise configuration. flap during cruis n aircrafi e t applications sharA . p trailing edge, however t conducivno s i , a hing o t e e lind othean e r Conclusions hardware necessary for deployment. This prompted the con- Two lift-enhancing devices were experimentally investi- sideration of a miniature split flap with a 90-deg deflection gated on a two-dimensional single-element airfoil and the (Fig. 10). Two configurations were tested for a 1.25%c split following conclusions were drawn: flap with a hinge line located forward of the trailing edge by 1) The Gurney flap can significantly increase the lift of a I.0 and 1.5 flap heights, respectively. As illustrated in Fig. single-element airfoil with a small increase in drag at low-to- II, these configurations yielded essentially the same results moderate lift coefficients. 1.25%e asth c Gurney flap wit degradatioo hn flan ni p effec- 2) To avoid a drag penalty during cruise, the miniature split tiveness. flap configuration can be employed with a hinge line forward The measured effect of the Gurney flap is in general agree- of the trailing edge, providing performance comparable with ment wit Navier-Stokee hth s computations presente Refn di . e Gurneth y flap. 4. Figur present2 1 e comparisosa measuree th f no cald dan - Vorte) 3 x generator delan sca y flow separatio increasd nan e Downloaded by UNIVERSITY OF CAMBRIDGE on October 16, 2014 | http://arc.aiaa.org DOI: 10.2514/3.46528 culated pressure distribution on the airfoil in the baseline maximue th m lift coefficien airfoin a n o tl that suffers trailing- clean configuration at an angle of attack of 9 deg. The lower edge stall. However, the large drag increment associated with pressures measured experimentall attributee b n yca walo dt l these devices suggests that they must be incorporated into a effects sinc e wine wallth eth f do s tunnel wer t modeleeno d high-lift system by placing them near the leading edge of the e computationsith n . Larger discrepancie e evidenar s a n i t flaps and/or the main element where they will be concealed similar comparison shown in Fig. 13 which incorporated a during cruise. 1.0%c Gurney flap. Nevertheless, the computations effec- 4) The Gurney flap and vortex generators can be employed tively predicted the higher suction peak and the increased aft in concert to generate greater lift enhancement than either loading causeGurnee th y db y flap comparisoA . come th f -no device individually. pute experimentad dan l lift coefficient e baselinth r d fo s ean \%c Gurney flap case presentes si Fig. n Goodi 14 . d corre- latio observes ni d between computatio experimend nan r fo t References the baseline airfoil except near the maximum lift coefficient. computationse Th , however, appea underpredico t r e lifth tt 'Liebeck , "DesigH. . R f ,Subsoni o n c Airfoil r Higfo s h Lift," Gurneincremene th o t ye flapdu t . This discrepanc notes ywa d Journal f Aircraft,o , 19789 . . 547-561 pp ,No Vol , 15 . . Largmand 2Katzan , J. , , R., "EffecDegre0 9 f o t e th Fla n po for each Gurney flap height tested. As illustrated in Fig. 15, Aerodynamic Two-Elemena f o s t Airfoil," Journal f Fluidso Engi- the computed lift increment for flap heights between 0.5- neering, Vol. Ill, March 1989, pp. 93, 94. 2.0% consistentls cwa comparew ylo d with experiment. 3Katz, J., and Dykstra, L., "Study of an Open-Wheel Racing Cafs Previous experimental and computational results4-10 indi- Rear-Wing Aerodynamics," Society of Automotive Engineers Inter- cate that the NACA 4412 suffers a trailing-edge stall pro- national Congress and Exposition, SAE Paper 890600, Detroit, MI, gression, a behavior especially suited for the application of Feb.-March 1989. STORMS AND JANG: LIFT ENHANCEMENT 547

4Jang, C. S., Ross, J. C., and Cummings, R. M., "Computational 1947. Evaluatio n Airfoia f o n l wit a hGurne y Flap," Proceedingse th f o 8Rao, D. M., and Kariya, T. T., "Boundary-Layer Submerged 10th AlAA Applied Aerodynamics Conference (Palo Alto, CA), AIAA, Vortex Generators for Separation Control—An Exploratory Study," Washington, DC, 1992, pp. 801-809 (AIAA Paper 92-2708). Proceedings t National1s e th f o Fluid Dynamics Conference (Cincin- 5Boyd . A.J , , "Trailing Edge Devic Airfoil,n a r efo " U.S. Patent nati, OH), AIAA, Washington , 1988 . 839-84DC , pp , 6 (AIA- APa 4542868, 1985. per 88-3546). 6Gregg, R. D., Hoch, R. W., and Henne, P. A., "Application of 9Wheeler , "LoO. . wG , Drag Vortex Generators," U.S. Patent Divergent Trailing-Edge Airfoil Technolog- Designe De th a o yf t o 5058837, 1991. rivative Wing," Society of Automotive Engineers Aerospace Tech- l('Wadcock, A. J., "Investigation of Low-Speed Turbulent Sepa- nology Conf Expositiond an . PapeE SA , r 892288, Anaheim, CA , rated Flow Around Airfoils," NASA CR 177450, Aug. 1987. Sept. 1989. nLin, J. C., Robinson, S. K., McGhee, R. J., and Valarezo, W. 7Taylor . D.,H , "The Eliminatio Diffusef no r Separatio Vortey nb x O., "Separation Control on High Reynolds Number Multi-Element Generators," United Aircraft Corp., R-4012-3, East Hartford, CT, Airfoils," AIAA Paper 92-2636, June 1992. Downloaded by UNIVERSITY OF CAMBRIDGE on October 16, 2014 | http://arc.aiaa.org DOI: 10.2514/3.46528