NASA-TP-1865-1981 Design and experimental results for a flapped natural-laminar-flow airfoil for general aviation applications《一般航空应用摆动自然层流翼设计和实验性结果》.pdf

《NASA-TP-1865-1981 Design and experimental results for a flapped natural-laminar-flow airfoil for general aviation applications《一般航空应用摆动自然层流翼设计和实验性结果》.pdf》由会员分享，可在线阅读，更多相关《NASA-TP-1865-1981 Design and experimental results for a flapped natural-laminar-flow airfoil for general aviation applications《一般航空应用摆动自然层流翼设计和实验性结果》.pdf（130页珍藏版）》请在麦多课文档分享上搜索。

1、(NASA-TP-1865) DESIGN AN_ EXPERIMENTAl.RESULTS FOR A FLAPPED NATUIAL-LAMINAR-r_.O_AIRFOIL FOR GENERAL AVIATION APPLICATIuiS- (_ASA) 125 p HC AO6/_F A01 CS_ uIA Unclas-j-. G3/02 12o30 _ _: . / -. -. .-: . :. . - _83-3038_ I-I_PRODUCEDBYNATIONAL TECHNICALINFORMATION SERVICEU_,DEPARIME_TOF COMMRCESPRIN

2、GFIELD,VA, 22161 “IIII IiIi IiIIIi II iiI liI_k i i if“Ii_ _ _Provided by IHS Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INASA Technical Paper 1865Design and Expe

3、rimental Results for aFlapped Natural-Laminar-Flow Airfoilfor General Aviation Applications_Dan M. SomersLangley Research CenterHampton, Virginia I5iiN/LSANational Aeronauticsand Space AdministrationScientific and TechnicalInformation Branch1981FYILProvided by IHSNot for ResaleNo reproduction or net

4、working permitted without license from IHS-,-,-u_ t_d_a_amlmlProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-ORIGINAL PAGE ISOF POOR QUALITYINTRODUCTIONResearch on advanced-technology airfoils for general aviation applicationshas received considerabl

5、e attention over the past decade at the NASA LangleyResearch Center. The initial emphasis in this research program was on thedesign and testing of turbulent-flow airfoils with the basic objective of pro-ducing a series of airfoils which could achieve higher maximum lift coefficientsthan the airfoils

6、 in use on general aviation airplanes at that time. For thisseries of airfoils, it was assumed that the flow over the entire airfoil wouldbe turbulent, primarily because of the construction techniques in use (mostlyriveted sheet metal). A summary of this work is presented in reference .While these n

7、ew NASA low-speed airfoils did achieve higher maximum lift coeffi-cients, the cruise drag coefficients were essentially no lower than the earlierNACA four- and five-digit airfoils. Accordingly, the emphasis in the researchprogram has been shifted toward natural-laminar-flow (NLF) airfoils in anattem

8、pt to obtain lower cruise drag coefficients while retaining the high maxi-mum lift coefficients of the new NASA airfoils. In this context, the term“natural-laminar-flow airfoil“ refers to an airfoil which can achieve signif-icant extents of laminar flow (_30-percent chord) solely through favorable p

9、res-sure gradients (no boundary-layer suction or cooling).Research on natural-laminar-flow airfoils dates back to the 930s at theNational Advisory Committee for Aeronautics (NACA). (See ref. 2. ) The workat NACA was culminated with the 6-series airfoils (ref. 3). The 6-series air-foils were not gene

10、rally successful as low-drag airfoils, however, because ofthe construction techniques available at the time.The advent of composite structures has led to a resurgence in NLF research.The initial applications were sailplanes, but recently, a number of poweredgeneral aviation airplanes have been const

11、ructed of composites - most notably,the Bellanca Skyrocket II (ref. 4) and the Windecker Eagle (ref. 5). In Europe,powered composite airplanes have also been produced. One such aircraft, theLFU 205, used an NLF airfoil specifically tailored for its mission (ref. 6).Thus, the introduction of composit

12、e construction has allowed aerodynamiciststo design NLF airfoils which achieve, in flight, the low-drag characteristicsmeasured in the wind tunnel (ref. 7). The goal of the present research on NLFairfoils at Langley Research Center is to combine the high maximum lift capabil-ity of the NASA low-spee

13、d airfoils with the low-drag characteristics of the NACA6-series airfoils.As part of the present research, an NLF airfoil, the NLF()-046, wasdesigned using the method of reference 8 and verified experimentally (ref. 9)in the Langley Low-Turbulence Pressure Tunnel (LTPT) (ref. 0). Based upon thesucce

14、ss of this airfoil and the excellent agreement between the theoretical pre-dictions and the experimental results, a second, more advanced, airfoil wasdesigned using the method of reference 8. An experimental investigation wasthen conducted in the Low-Turbulence Pressure Tunnel to obtain the basic lo

15、w-KY.k.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-OE POOR QUALITYspeed, two-dimensional aerodynamic characteristics of the airfoil. The resultshave been compared with the predictions from the method of reference 8.Use of trade names or names of

16、manufacturers in this report does not con-stitute an official endorsement of such products or manufacturers, eitherexpressed or implied, by the National Aeronautics and Space Administration.mSYMBOLSValues are given in both SI and U.S. Customary Units.calculations were made in U.S. Customary Units.PZ

17、 - P_Cp pressure coefficient,q_Measurements andc airfoil chord, cm (in.)ccCdcd ,cZcmcnhMPqRtsection chord-force coefficient, fCp d(_ 1section profile-drag coefficient, f C d d(!)Wakepoint drag coefficient (ref. I)section lift coefficient, cn cos _ - cc sinsection pitching-moment coefficient about qu

18、arter-chord point,C xvertical height in wake profile, cm (in.)free-stream Mach numberstatic pressure, Pa (ibf/ft 2)dynamic pressure, Pa (ibf/ft 2)Reynolds number based on free-stream conditions and airfoil chordairfoil thickness, cm (in.)“I11 l 1 1 li Jl 1 I 1 1 It I I Jl 1 Jl Jl I Ir_Provided by IH

19、SNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-xz_fairfoil abscissa, cm (in.)airfoil ordinate, cm (in.)angle of attack relative to chord line, degflap deflection, positive downward, degSubscripts:Z local point on airfoilmax maximummin minimumfree-stream condition

20、sAbbreviations:1sLTPTNLFuslower surfaceLangley Low-Turbulence Pressure Tunnelnatural laminar flowupper surfaceOF PO0_ Q_.,A_,AIRFOIL DESIGNOBJECTIVES AND CONSTRAINTSThe target application for this airfoil is a high-performance, single-engine, general aviation airplane. This application requires low

21、sectionprofile-drag coefficients cd at a Reynolds number R of about 9 x 0 6 forthe cruise section lift coefficient (c z = 0.2) as well as for the climb sectionlift coefficients (c z = 0.5 to .0).Two primary objectives were identified for this airfoil. The first objec-tive was to design an airfoil wh

22、ich would produce a maximum lift coefficientCZ,ma x at R = 3 x 0 6 comparable to those of the NASA low-speed series air-foils. (See ref. .) A requirement related to the first objective was thatCz,ma x not decrease with transition fixed near the leading edge on both sur-faces. This means that the max

23、imum lift coefficient cannot depend on theachievement of laminar flow. Thus, if the leading edge of the wing is con-taminated by insect remains, etc., the CZ,ma x should not decrease. Thisrequirement is set by safety considerations relating to stall and, therefore,to landing speeds. The second objec

24、tive was to obtain low profile-drag coeffi-cients cd from the cruise lift coefficient cz of 0.2 to about .YU 1 11 1I tl 11 U _ 1t _ 11 _ _ 11 _ 1 il IProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-OF POOR O_;A:-,YThree constraints were placed on thi

25、s airfoil design in order to make itcompatible with existing aircraft designs. First, the airfoil thickness t/cmust be 5 percent. Second, the pitching-moment coefficient cm should be nomore negative than -0.05 at the cruise lift coefficient (c Z - 0.2). Third, theairfoil must incorporate a simple fl

26、ap having a chord equal to 25 percent ofthe airfoil chord c.PHILOSOPHYGiven the previously discussed objectives and constraints, certain charac-teristics of the design are evident. The following sketch illustrates thedesired c z - cd curve, which meets the goals for this design:C 11.8 m CC dII:I:FSk

27、etch rThe desired airfoil shape can be related to the pressure distributions which occurat the various lift coefficients shown in the sketch. Point A is the cruise con-dition (c z = 0.2, R = 9 06). The value of cd for this point is determinedby the extents of laminar flow on the upper and lower surf

28、aces. There is littleaerodynamic advantage in achieving low drag below c z = 0.2. This is especiallyimportant if high maximum lift must be obtained (point C). However, in anattempt to insure a low-drag coefficient at the cruise lift coefficient(c Z = 0.2) despite contour deviations due to constructi

29、on tolerances, the lowerlimit of the low-drag range was extended downward to c z = 0. . Notice that thedrag at point B (c z = .0) is not quite as low as at point A (c Z = 0.2). Thisfeature is quite important because it shows that the transition point on theupper surface moves slowly and steadily tow

30、ard the leading edge with increasingc Z, as opposed to the sudden forward jump characteristic of the NACA 6-seriesairfoils. This feature leads to an airfoil with a relatively blunt leading edgewhich, in turn, should produce a high maximum lift coefficient as well as goodflap effectiveness. L_This ou

31、tline of the desired section characteristics is not sufficient todesign the airfoil, however, primarily because of the variable introduced byM ti 11 11 11 11 1t 1t 1I ti II, 1t 11 11 11 11 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Of _ the flap

32、 and the unconstrained extents of laminar flow on the upper and lowersurfaces. In order to evaluate the effects of these variables, it is helpfulto examine the design goals with respect to overall aircraft performance. Forthis airfoil, the primary goal is a reduction in wing parasite drag. This goal

33、can be achieved in a number of ways, two of which are discussed in this report.First, if a high maximum lift coefficient Cz,ma x can be realized, the wingarea can be reduced relative to a wing with a lower Cz,ma x. This conclusion isbased on the assumption that both aircraft must achieve the same mi

34、nim_ speed.Second, if the amount of laminar flow on one or both surfaces can be extended,the minimum profile-drag coefficient Cd,mi n will be reduced. Further analysisindicates that by maximizing Cz,max/Cd,mi n, the wing parasite drag is minimized.Unfortunately, a reduction in Cd, mi n through the e

35、xtension of the amount oflaminar flow on the upper surface generally results in a reduction in CZ,ma x.(See ref. 3.) By trial and error, it was determined that c Z,max/cd,min wouldbe maximized for this application if the extent of laminar flow was about 0.4con the upper surface and about 0.6c on the

36、 lower surface.The effect of the flap on the design can be evaluated by examining the con-straint on the pitching-moment coefficient (Cm, cruise -0.05). The objectiveof a high maximum lift coefficient is in conflict with the pitching-moment con-straint. For this design, the flap can be used to allev

37、iate this conflict byemploying negative (up) flap deflections. This concept allows an airfoil to bedesigned which has a fairly large amount of camber (conducive to a high Cz,max)but retains the ability to achieve a low pitching-moment coefficient at thecruise lift coefficient (c Z = 0.2). This conce

38、pt has the added advantage that,by deflecting the flap up or down, the low-drag range can be shifted to lower orhigher lift coefficients, respectively. (See ref. 2.) Based upon experiencewith other airfoils, the negative flap deflection 6f was limited to -10 .From the preceding discussion, the press

39、ure distributions along thec Z - cd curve from points A to B in sketch can be deduced. The pressure dis-tribution for a flap deflection of 0 at a lift coefficient of about 0.7 (i.e.,between points A and B) should probably resemble sketch 2.CP+0usI I0.4 0.6X/CSketch 2I1.OFR5r-LProvided by IHSNot for

40、ResaleNo reproduction or networking permitted without license from IHS-,-,-ORI_NAL PP,_ ISOF POOR QUALITYFor the reasons previously stated, a favorable pressure gradient on the uppersurface is desirable up to x/c = 0.4. Aft of 0.4c on the upper surface, ashort region of slightly adverse pressure gra

41、dient is desirable to promote theefficient transition from laminar to turbulent flow (ref. 3). Thus, the ini-tial slope of the pressure recovery is relatively shallow. This short region isfollowed by a steeper concave pressure recovery which produces lower drag and ,has less tendency to separate tha

42、n the corresponding linear or convex pressurerecovery (ref. 13). The proposed pressure recovery, although concave, does notapproach the extreme shape of a Stratford recovery (ref. 4). The Stratfordrecovery is not appropriate for an airfoil which must operate over a range oflift coefficients and Reyn

43、olds numbers (ref. 15).For the reasons previously stated, a favorable pressure gradient on thelower surface is desirable up to x/c = 0.6. A rather abrupt and very steep con-cave pressure recovery is introduced aft of 0.6c, which results in a largeamount of aft camber. This camber, although limited b

44、y the pitching-moment con-straint (Cm, cruise -0.05 with _ = -0), helps produce a high maximum liftcoefficient.For point A (c I = 0.2) in sketch , the pressure distribution should resem-ble sketch 3. For this lift coefficient, the flap is deflected up 0 . AlongCP+ I I I0 0.4 0.6 1.0x/cSketch 3the lo

45、wer surface, the pressure gradient is initially adverse, then zero, andthen increasingly favorable. Basically, this concept is to transition as theStratford pressure recovery (ref. 4) is to separation. The concept was sug-gested by Richard Eppler of the University of Stuttgart, Stuttgart, West Germa

46、ny.For point B (c I = 1.0) in sketch l, the pressure distribution should resem-ble sketch 4. For this lift coefficient, the flap is deflected down 0 o. Afavorable pressure gradient on the upper surface to x/c = 0.4 and a zero pres-sure gradient on the lower surface to x/c = 0.6 is expected to result

47、 in lowdrag, albeit at the lower limit of the low-drag range for this flap deflection.i11 II 11 11 11 11 1t 11 11 11 1t 1l 1l 1Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CPusOF PDOR QUALITYI I I0.4 0.6 1.0x/cSketch 4It should be noted that the c

48、ruise-flap concept may not be optimum for allapplications. If the construction tolerances at the flap hinge are not suf-ficiently tight, lift and drag penalties due to a disturbance to the turbulentboundary layer may be sufficient to offset the advantages of this concept. (Seeref. 6. )EXECUTIONGiven the pressure distributions for c Z = 0.2, c Z = 0.7, and c Z = 1.0,the design of the airfoil is reduced to the inverse problem of transformingthe pressure distributions into an airfoil shape. The method of reference 8 wasused because it is ideal for handl