NASA-TP-1919-1981 Wind-tunnel results for a modified 17-percent-thick low-speed airfoil section《改良型17%厚度的低速翼剖面风洞结果》.pdf

《NASA-TP-1919-1981 Wind-tunnel results for a modified 17-percent-thick low-speed airfoil section《改良型17%厚度的低速翼剖面风洞结果》.pdf》由会员分享，可在线阅读，更多相关《NASA-TP-1919-1981 Wind-tunnel results for a modified 17-percent-thick low-speed airfoil section《改良型17%厚度的低速翼剖面风洞结果》.pdf（86页珍藏版）》请在麦多课文档分享上搜索。

1、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NASA Technical Paper 1919!ii:,f Wind-Tunnel Results for aModified 17-Percent-ThickLow-Speed Airfoil SectionRobert J. Md3hee and William D. Beasley Ix, agh:F Research Center _Har,tpton. Virgima1National

2、AeronautiCsand SOacL, Admiriistration .Scientific and Technical ,_InfOrmation Branch198i !L _. L_. _ - a I _,t . i . I II _ rt .- . - I - - ii I I II I I .Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-t:ii./P_ SUMMM_iHAn investigation was conducted

3、 in the Lanqly Low-Turbulence Pressure Tunnel toevaluate the effects on p0rforinanOe of modifying a 17_perce,lt-thiek low-speed air-.ilL foil. lheairfoil contour was altered to _educe the pitching-moment coefficient byincreasing the fo_ard loadillg and to increase the climb llft-drag ratio by decrea

4、s-ing the aft uppel _ Surface pressure gradient. Tl%e tests were conducted over a Machnumber raDge.from 0.07 to 0.32, a chord Reynolds:number range from 1.0 106 toDI_.0 x 10 , and an angle-of-attack range from about -i0 _ to 20“. +The results of the i_%vestig, Subscripts:i_ A local point-on airfoili

5、_ max maximum,“ S SeparatiOn,+ free-stream conditionsi_ Airfoil designations:LS(I)-0417 low speed (first series); 0.4 desigD lift coefficient, 17 percent thick Mod modifledAIRFOIL MODIFICATIONThe airfoil contour was changed with two objectives in mind: to reduce thepitching-moment coefficient by inc

6、reasing the forward loading and to increase the +_iclimb lift-drag ratio (c% = 1.0) by decreasing the aft upper surface pressure gradi- +ent. The maximum thickness ratio, trailing-edge thickness, and design llft coeffi- cient (cI = 0.40) of the original airfoil were retained.The upper+surface modifi

7、cation to the original airfoil was accomplished byusing the computer code of reference 2. This inverse method calculates inviscidcoordinates of an airfoil, from a prescribed velocity distribution. A boundary-layercorrection is made to allow for viscous effects by computing the displacement thick-nes

8、s of the turbulent boundary layer, which is subtracted from the inviscid coordi-nates. The inviscid velocity distributions for both airfoils areshown in figure+l,and figure 2 illustrates the change in airfoil shape. The design conditions for theairfoil were a lift coefficient of 0.40, a ReynOlds num

9、ber of 4.0 106, and a Machnumber of0.15. Figure 3 compares the mean thickness distributions and camber linesfor the two airfoils. Coordinates for the modified airfoil are presented in table I.Theoretical chordwise pressure distributions (_ef. 2) for both airfoils areshown in figure 4 for a Reynolds

10、number of 4.0 10_. +Boundary-layer transition wasspecified at x/c = 0.03 for the calculations to ensure a turbulent boundary-layerdevelopment on the airfoils. A reduction in the pitching-moment coefficient atdesign lift of about 28 percent is indicated by the theoretical calculations. Notethat a fla

11、t pressure distribution or reduced pressure-gradient region extends +forabout 0.20c prior to the start of the aft upper Surface pressure recovery for themodified airfoil. This reduced pressure-gradient region with the “corner“ locatedat x/c = 0.60 iS Considered to be an important feature of the airf

12、oil design.Research reported in reference 3 for a modlfied 13-percent-thick airfoil clearly R = 4.0 x 06 . 18Variation of maximum lift coefficient with ReynOlds number forE LS(I)-0417 and L8(I)-0417 Mod airfoils; M _ 0.15 . , 19*: LS(I)-0417 a d LS( )-0417 Mod airfoils roughness on;R = 6.0 x 10 . 20

13、Comparison of maximum lift coefficients of LS(I )-0417 Mod airfoilwith those of NACA airfoils; models smooth; M = 0._5 . 21i Variation of drag_coefficient with Reynolds number for LS(I )-041“IMod airfoil; M _ 0.15; cI = 0.40 22Variation of lift-drag ratio with lift coefficient for LS(I)-0417and LS(I

14、)-0417 Mod airfoils; roughness on; M = 0.15 23DISCUSSION OF RESULTSSection CharacteristicsLif_,t.-Thelift-curve slope for the 17-percent-thick modified airfoil in asmooth condition (roughness off) was about 0.12 per degree for the Reynolds numbersI investigated (M = 0.15) as indicated by figure 10(a

15、). The angle of attack for zerolift coefficientwas about -3.5 . Maximum lift coefficients in_reased from a_outp 1.70 to 2.10 as the Reynolds number was increased from 1.0 x 10_ to 12.0 x 10_.(See fig. 19.) The largest effect of Reynolds number on maximum lift coefficientI occurred for Reynolds numbe

16、rs below 6.0 10b. The stall characteristics of theairfoil _re of the trailinq-edge type as shown by the pressure data of -figure 15_However, the nature of the stall was abrupt for Reynolds numbers greater than2.0 x 106. (See fig. 7.) Abrupt stall characteriStiCs were not expected for thisr airfoil,

17、and a detailed discussion is included _n a subsequent-sectionentitled“_ressure Distributions.“ iThe addition of a narrow roughness strip at 0.075c (fig. 9) reculted in theexpected decambering effect for thick airfoils because of the increase in boundary-layer thickness. The lift coefficient at a = 0

18、o decreased about 0.03 at the lowerReynolds numbers, but only small changes occurred at the higher Reynolds numbers. !The roughness strip decreased the Cl,ma x performance of the airfoil as much as 0.04for the test Reynolds number range. (See fig. 19.) iThe effects o_ Mach number on the airfoil llft

19、 charaCteristicS at a Reynoldsnumber of 6.0 x 10 with roughness located at 0.075c are shown in figure 12(a).Increasing the Mach number from 0.10 to 0.32 resulted in the expected increase inlift-curve slope, and the Stall angle of attack was decreased about 2.2 . However,there were only small changes

20、 in Cl,ma x due to Mach number effects. (See fig. 20.)The lift data for the original and modified 17-percent-thick air_oils arecompared in figure 13 for Reynolds numbers from 2.0 x 10 to 6.0 x 10with fixedtransition at 0.075c, The data indicate that the linearity of the llft Curve isextended to high

21、er angles of attack and that C_,ma x is increased for the modifiedairfoil. This result is attributed to reduced upper Surface boundary-layer separa- I6Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-,L_:,/i tion for the modified airfoil, as illustrat

22、ed by the pressure-data comparison of figure 18(c). Note, however, that _he nature of _he stall iS m_re abrupt for the_ modified airfoil for Reynolds numbers of 4.0 x.10 and 6._ x I0“. The variation of_i C_,ma x with Reynolds number and Mach number_is compared for both airfoils in fig-ri ures 19 and

23、 20, respectively. In the low Reynolds number range (R _ 4.0 x I06), ani=i_ increase in C%,ma x of ab0ut 10 percent if shown for the modified airfoil. HOwever,I for ReynOlds numbers _reater than 9.0 X 10_, both airfoils .develop about the same_ C%,ma x . Increasing the Mach number results in a decre

24、ase i_ Cl,max for the_, original airfoil (fig. 20); however, only small. Mach number effects on C_,ma x are:ii shown for the modified airfoil. Comparisons of the values of C_,max for the modi-.fied airfoil with the NACA 4- and 5-digit airfoils and 65_se_ies airfoils are _hownin figure 21 for Reynold

25、s numbers from 3.0 x 10 to 9.0 x 10 . Substantial improve-ments in Cl,ma x throughout the Reynolds number range are indicated for the modifiedlow-speed airfoil. For example, at a Reynolds number of 3.0 106 a 35-percentIi improvement in C_,ma x is shown for the mc_dified airfoil compared with the NAC

26、A23018 airfoil.Pitching moment.- The pitching-moment-coefficient data of figures 9, I0(c), and11(c) illustrate the expected positive increments in cm dueto decreasing theReynolds number or adding roughness at a constant Reynolds number. _lis result istypical of the decambering effect associate_ with

27、 boundary-layer thickening for thickairfoils. At a Reynolds number of 6.0 10 , increasing the Mach number from 0.10 to0.32- (fig. 12(c) shows small effects on the pitching-moment data to about _ = I0.At the higher angles of attack, a positive increment in cm is shown.The pitching-moment data for the

28、 original and modified airfoils are compared infigure 13. The design objective:of reducing cm by increasing the forward loadingof the airfoil was accomplished. A reduction in the magnitude of cm of about20 percent at cI = 0.40 (cruise condition) is indicated for the modified airfoil.The theory indic

29、ated a reduction in cm of about 28 percent. This result is impor-tant because of the expected reduced trim penalties for the modified airfoil atcruise:_onditions. _Dra and lift-draq ratio.- Natural transition usually occurs near the leadingedge of wings in actual flight conditions of general aviatio

30、n aircraft because of theroughness of construction or insect remains gathered in flight, Therefore, the dis-cussion of the drag data is limited to data obtained with fixed transition at 0.075c.The profile-drag c_efficient at design lift (ci = 0.40! decreased from about0.01.15 at R = 2.0 x 10v to abo

31、ut 0.0085 at R = 12.0 x 106. (See figs. 11(b)and 22.) This reduction in cd is due to the decrease in the skin-friction coeffi-cient of theturbulent boundary layer at the higher Reynolds numbers. Thereare onlysmall effects of Mach number on cd (fig. 12(b) over a Mach number range from 0.10to 0.32.The

32、 drag data for the original and modifie_ .airfoils are compared in figure 13for Reynolds numbers from 2.0 x I0b to 6.0 x 10 with fixed transition at 0.075c.The design o_Jective of reduced Cd, at typical climb conditions (ct = 1.0;R = 4.0 x 10 ) by decreasing the aft upper surface pressure gradient w

33、as accomplished(cd decreased about 0.0012). An increase in climb lift-drag ratio of about i0 per-cent was measured for the modified airfoil. (See fig. 23.)7Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-F _ _ _ _-,-_.i. -i -.r._ ._,_ _ _,_ , _,=_ _,

34、.,_._ _ !?! Pressure Distributions _ iiThe chordwise pressure data Of figure 15 illustrate the effects of angle ofh attack for Several Reynolds n_mbers for the modified airfoil. As the angle of attackis increased, upper surface traillng-edge separation is first indicated by the_!_ approximate COnsta

35、nt-pressure region on theairfoil. At a Reynolds number of2.0 10 , separation is indicated at about u _ 12o, (See fig15(a).) Additionalincreases in angle of attack result in this constant-pressure region moving forward:_ along the airfoil, and the stall characteristics are docile. The li_t Curve is w

36、ellrounded at stall as shown by figure 9(a) However, at R = 4._0 x 10_ (fig. 15(b)Ii! the trailing-edge separation point moves rapidly f_rward for an increase in angle ofattack from about 16“ to 17o, and the-stall characteristics are abrupt, The liftdata of figure 9(b) illustrate the abrupt decrease

37、 in llft at stall. As discussed inill a previous section entitled “Airfoil Modification,“ docile stall behavior was antici- pated for this airfoil because of the flat pressure distribution (reduced pressure-i gradient region) prior to the start of the aft upper surface pressure recovery at thedesign

38、 lift coefficient. However, as previously discussed, the rearward extent ofthis flat pressure distribution must be determined from experimental tests becausethe present airfoil theories are inadequate when large regions of separation arepresent. Thus, more desirable stall characteristics for this ai

39、rfoil at the higherReynolds number would be expected by reshaping the airfoil with a more rearwardextent of this flat pressure distribution.CompariSons of the pressure data for the original and modified airfoils 8reShown in figure -18 at a Mach number of 0.15 and a Reynolds number of 4.0 10 . Thepre

40、ssure _ata at _ = 0 (fig. 18(a)illustrate the increase in forward loading and ithe decrease in the aft upper surface pressure gradient for the modified airfoil. 1This reduced pressure gradient has a favorable effect on theairfoil boundary-layer_development (reduced thicknesS)and results in a decreas

41、e in cd at typical climb Ilift coefficients. Comparisons of the airfoil pressure data at _ = 12 :_(fig. 18(c) indicate thatthe modified airfoil exhibits about 0.20c less separationthan the original airfoil.CONCLUDING REMARKS.Wind-tunnel tests have been conducted in the Langley Low-Turbulence 9ressur

42、eTunnel to evaluate the effects on performance of modifying a 17-percent-thick_low- iispeed airfoil. The _irfoil contour was altered to reduce the pitching-moment coeffi- icient by increasing the forward loading and to increase the climb lift-drag ratio bydecreasing the aft upper surface pressure gr

43、adient. The tests were conducted atfree-stream Mach numbers from 0.07 to 0.32. _le chord Reynolds number was variedfrom about 1.0 106 to 12.0 x 106 .The results show that the modification to the airfoil contour produced thedesign objectives of reduced pitching-moment coefficient _t cruise and increa

44、sedlift-drag ratio at climb. The magnitude of the pitching-moment coefficient wasreduced about 20 percent at the design lift coefficient of 0,40, and the llft-dragratio was increased about 10 percent at the climb llft coefficient of 1.0. The maxi-Provided by IHSNot for ResaleNo reproduction or netwo

45、rking permitted without license from IHS-,-,-j:mum lift6coefficient WSs also increased about 10 peruent at Reynolds numbers of2.0 x10,and 4,0 x 10 . However, the stall charaCterlstiCs _f the modified airfoilwere less desirable at Reynolds numbers greater-titan 2.0 x 10 beaausa. Of_u rapidk_. forward

46、 movement of the trailing-edge separation point.ir!ip._? Langley Research CenterNational Aeronautics and SpaCe Administration_. HamptOn, VA 23665_ August 18, 1981Ii1t9Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REFERENCESi,i,. I. MCGhee, Robert J

47、._ and Baasley, William D.: Low-Speed Aerodynamic Characterist_Qsof a 17-Percent-Thi0k Airfoil Soctlon Designed for General Aviation AppliCa-il tlons. NASA TN D-7428, 1973.P:;_ 2. Bauer., Frances_ Garabedian, Paull and Korn, Davidl Superoritical Wing Sec-tions III. Volume 150 of Lecture Notes in Ee_

48、nomies and MathematicalSystems, Springer-Verlag,1977._ 3_ McGhee, Robert J. 1 and Beasley, William D Low-Speed Wind-Tunnel Results for aModified 13-Percent-Thick Airfoil. NASA TM X-74018, !977.4. Von Doenhoffs Albert E.; and Abbott, Frank T., Jr.: The Langley Two-DimensionalLow-Turbulence Pressure T

49、unnel. NACA TN 1283, 1947.5. Braslow, Albert L.; and Knox, Eugene C.; Simplified Method for Determination OfCritical Height of Distributed Roughness Particles for Boundary-Layer Transitionat Mach Numbers From 0 teL5. NACA TN 4363, 1958.6. Pankhurst, R. C.; and Holder, D. W.: Wind-Tunnel Technique. Sir issac Pitman and Harper, John J.: Low-Speed Wi