1、RESEARCH MEMORANDUM LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A SERIES OF SWEPT WINGS HAVING NACA 65A006 AIRFOIL SECTIONS “ (Revised) NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5016 .” LOW-SPEED
2、 AEEiOmWAMIC CHIIRACTERISTICSOF A SERIES OF SWEPT WINGS HAT7ING NACA 6s006 AIRFOIL SECTIONS* (Revised) By Jones F. Cahill and Stanley H. Gottlieb SUMMARY . An investigation was made to determine the effect of Bweep, taper ratio, and aspect rat.io on the aerodynamfc characteristics of nlne semispan w
3、ings of MACA 65006 airfoil section with and without split flaps. Lift, drag, pitchi-nt , and -oat tje-ment characteristics were measured through a range of Reynolda nmibers from 1.5 x 10 to 12.0 x lo6. One of these wbgs was tested with a hinged lea-dge flap of vazioua spans and deflections to determ
4、ine the effect of this type of flap on longitudinal stability near maximum lift. For wings of aspect ratio 4, increases in 8treep angle increased the maximum lift coefficient of the plain wings but decreased the maximun lift coefficient of the wings wi$h half”span sglit flaps. Rather abrupt unstable
5、 change8 in pitching mment occurred at lift coefficients ke11 below maximum for newly all of the swept wings tested. Increases in sweep angle or aspect ratio reduced the lift coefficient at which these unstable changes occurred. . Increases in lift-curve slope asd stable changes in pitching moment o
6、ccurring at Lar to moderate. llft coefficfents for the aweptback wings were incresed in magnitude bg increase in taper ratio and decrease in aspect ratio. Reynolds nt-Reynolds nmibers as near as possible to those at which the wings are expected to be used. A number of investigations of the character
7、istics of swept wings at high Reynolds numbers have been made, but as yet there does not exist nd span on loqgitudinal stability at the I t I I I 1 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA RM 5016 staJ.1. Fence8 were tested on the wing
8、with one leadiwdge-flap configuration in an attempt to delay spaawiee flows. Drag coefficients and angles of attack were corrected for jet- boundary effects by means of boundary-induced upwash corrections calculated by the method of reference 5. The highest Mach number attained during these tests wa
9、s approxi- mately 0.20. PIiESENTATION OF DATA The aerodynamic characteristics of the wing8 tested with and without split flap and roughness 8378 presented in figures 4 to 12. Figure8 13 to 17 present data showirq the effect of leadingedge-flap deflection and span on the aerodynamic characteristics o
10、f the 454-0.6 wing with and without splft flaps and fences. These data are presented as plots of angle of attack, root bendin whereas changes in taper ratio have little effect. The addition of the “pan split flap increases the lift coefficient at which the pitching moments breek -unstable for all of
11、 the wings except the 45O sweptforward. The effects of variation in Reynolds number on either maxlmma lift coefficient or on the lift coefficient for the unstable pitching-nt break were dl in U. casea except for the 30 sweptback wing (fig. 6) which showed a higher madmum lift coefficient at a Regnol
12、ds number of 6 X 10 than for Regnolds nmibers either above or below this value. This phenomena ut be associated with some peculias scale effect on the laminar flow mound the lea- edge since the addition of leadingedge roughness decreased. the mx3mum lift to approxiinately the value obtained at other
13、 Reynolds nmibere. hide from this isolated instance, leading- edge roughness has very little effect on the lift charecteristics of my of the wings tested. 6 The effects of changes in airfoil section on the lift coefficient at which the unstable pitching-momsnt break occurs may be deduced from a comg
14、arison of these data with other previously published data. Data . for a wlng having plan+form pmameters (4M .625) roughly similar to the 4M.6 King teated in this investigation but with circular-src sections are presente,d in reference 9. These data show that the lift coefficient at which the unstabl
15、e pitchi-nt break occurs i8 tipproximately v the same for the circul- wing 88 for the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 MACA RpiI 5016 Gpercentdhick low-drag wing. Data in references 10 and 11 for two wings (40-4-0.625 and 52.9-0.625
16、) having airfoil sections 9.6 and 7.8 percent thick, respectively, (and, therefore, larger leading-edge radii than the 6percent“thick wings) show that the pitchi-nt break occurs at appreciably higher lift coefficients than would be indicated by the data of the present investigation for wings of roug
17、hly similar plan form and 6-percent-thick sections. It seem .likely therefore that, for wings swept back apKoxWtely 45O, the changes in leading-edge radius corresponding to decreases in thickness ratio below 6 percent will have lltt-le effect on the lift coefficient at which the pitching moments bre
18、ak unstable but that this lift coefficient mey be raised substanttally by relatively small increases in thic-ess. Pitchiw moments.- An examination of the pitching4ument data in figures 4 to 12 shows that abrupt variations in the slope of the pitching- moment curve occur at lift coefficients well bel
19、ow maximum lift for nearly all of the wings teated. In all caaw except the 01c-o.6, the 3U.6, and the 452-0.6 wings, unstable variations occurred. These changes in pitchimwnt- characteristics am in agreement with the bo-ezy curve for stability at high lift coefficients presented in reference 12. The
20、 unstable changes in the pitching-mament cmes occw at the same lift coefficient as the shift8 in the spamrlse center of pressure (fig. 20), which would indicate that thfs instability c The data ahow that increasing any of the three plan-form pazameters does came a decrease in the negative pitchinwme
21、nt increment caused by flap deflection an Wing-root bendiw? momenta .- The data for these vlngs show that the -root bending maments are roughly linem up to about the lif% coef- ficient- at which. the pitchkg+mwnt break occurs. The spanwlae centers of pressure shown in figure 20 help to show the chai
22、ges which take place in the loading on the wings. The spaarise centers of presure are generally constant at noderate lift coefficients and move rather rapidly inboard for the sweptback wings and outboard for the Bweptforward wing at high lift coefficiente . Values of the spaswise center of the addit
23、ional load distribution, . indicated by the slope of the bendirtg+nomsnt cme through zero lift, are shown plotted against the varipus plan-form parameters in figure 22. I I I I I I t . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 NACA RM 5016
24、 These data show good agreement with calculated values except for the wing with aspect ratio 6 and the wiw with taper ratio 0.3. In both of these cases, the spanwise center of pressure is farther inboard than indicated by the computations. Characteristice of 4w.6 Wing with HinLf”spm leadiwdge flap d
25、eflected 30 provided a atable va;ri+ tion in pitch- moments at high lift coefficients although the over-all pitchhg+noment variation.could not be considered desirable. Use of the leadiwdge flap also provided appreciable reductions in drag at hi- coefficients. b Y In an attempt to delay spanwise flaw
26、s and, therefore, improve the pitchinmment vrniation at high lift coefficients, chordwise fences were installed on the wing with the 0.7% +eadiq+edge flap deflected 30. Use of the fences at the positions tested increased the maxirnum lifi coefficient and the lift coefficient for the pitching- moment
27、 break slightly but caused no Fmprovement; in the direction of the pitchlnmmnt break at ;coefficient range. A hinged leadiw whereas chamgea in taper ratio came relatively ch8nge8. Deflection of a haU+pan eplit flap increases the lift coefficient for the unstable break for all of the wings tested exc
28、ept the 45O sweptforwmd wing. (4) -creases in sweep came increases in maximum lift coef- ficient of the &lapped wings for either positive or negative sweep. increased and is actually negative for positive sweep angles greater than 45. Increasing aspect ratio of a 45O sweptback wing hecreases maxim=
29、lift, but changes in taper ratio have little effect. -Flap effectiveness on eum lift decreases as the sweep angle is (5) Deflection of a full-span hinged leadiwdge flap 30 provides the largest increase in the lift coeffictent for the unstable . -. . . . . . - Provided by IHSNot for ResaleNo reproduc
30、tion or networking permitted without license from IHS-,-,-14 UCA RM 5016 . pitchinvnt break of any of the ccadbinationer of leading-edge-flap span and deflection tested but did not prcduce a stable pitching-moment variation at the eta. (6) Scale effectl on aerodynamic characteristics were confined t
31、o the more highly 8vep-t wings and consisted principalLy of a delay in the lift coefficient at which the increase in lift-cme slope occus. Langley Aeronautical Laboratory National Advisory Comrmittee for Aeronautics Langley Air Force Base, Va. I Provided by IHSNot for ResaleNo reproduction or networ
32、king permitted without license from IHS-,-,-WA RM 5016 15 . 1. Sweberg, Hiwold E., and kge, Roy H. : Sumary of Available Data Relating to Repolds Nmiber Effects on the Maximum Lif% Coef- ficients of Swept-Back Wings. NACA RM 6120a, 1947. 2. Wilson, Herbert A., Jr., and Lovell, J. Calvin: FuXL4caJ-e
33、Dme8tI- gation of the Maxbmm LiFt and Flow Characteristics of as Airplane Having .Approx&w.tely Triangular Plan Form. RACA RM -0, 1947. 3. Schuldenfrei, Marvin, Cdsarow, Paul, and Goodson, Kenneth W.: Stability and Control Characteristics of an Airplane Model Having a 45 .lo Swept-Back Wlng with Asp
34、ect Ratio 2.50 esd Taper Ratio 0.42 and a 42.8 Swept-Back Horizontal Tail with Aspect Ratio 3.87 and Taper Ratio 0 .kg. NACA RM L7B25, 1947. 4. Cahill, Jones F.: Comparison of Semispas Data Obtained in the Langley !Iko4imemional Lowbulence Pressure Tunnel and Full- Span Data Obtained iq the Langley
35、Ig-Foot Pressure Tunnel for a Wing with 4.0 Sweepback. of the 0.27-Chord Line. NACA RM Lp25a, 1949 5. Ilat zoff, S., and Eannah, Margery E. : Calculation of Tunnel-In&uced Tfpwash Veloclties for Swept and Yawed Wings. NACA TN 1748, lgh.8. 6. DeYoung, John: Theoretical Additional Span Loading Charact
36、eristics of Wings with Arbitrary Sweep, Aspect Ratio, asd Taper Ratio. NACA TN lh-91, 1947. 7. Lange, Rog H., Whittle, Edward F., Jr., and Fink, “Tin P. : Emestigation at Large Scale of the Pressure Distribution and Flow Phenomena mer a Wing with the Leading Edge Swep.t Back 47.5O Having Circular-Ar
37、c Airfoil Sectfons and Equipped wlth Drooped- Nose and Plain lEps. IIACA RM LgG15, 1949. 8. Anderson, Adrien E. : Chordwise and Spanwise Loadbqy Measured at Low Speed, on Wge Triangular Wings. RACA RM AgB17, 1949. 9. Meely, Robert H. and Kown, William: Low-Epeed Characteri-stics in Pitch of a 42 6 S
38、weptback Whg with Aapect Ratio 3.9 and Circulm Arc Airfoil Sections. mACA RM L7E23, 1947. 10. Graham, Robert R., and Comer, D. Wllliam: InveAigation of Eigh- Lift and Stall-Control Devices on an NllCA 6LrSeries 42 Swept- back Wing with and without Fuselage. WCA-RM LiGO9, 1947. I I Provided by IHSNot
39、 for ResaleNo reproduction or networking permitted without license from IHS-,-,-16 NACA RM 50n6 11. Foeter, Gerald. V., and Fitzpatrick, James E. : . Longitudinal- Stability Investigation of High-Lift and Stall-Control Devices on a 52 Sweptback Wing with and wlthout Fuselage and Horizontal . Tail at
40、 a Reynolde Number of 6.8 x 10 . NACA RM 8108, 1948. 6 12. Shortal, Jo8eph A., and Maggin, Bernard: Effect of Sweepback and Aspect Ratio on Longitudinal Stability Charmteristics of Winge at Low Speeds. MclCA TN 1093, 1946. 13. Koven, William, asd Graham, Robert R.: Wind-humel Investigation of High-U
41、ft and StU-Control Devices on a 370 Sweptback Wing of Aspect Ratio 6 at High Reynolds Ikutibers. NACA RM L8D29, 1948. 14. Guryanslrg, Eugene R., and Lipson, Stanley: Effect of High-Lift Devices on the L0ngitudlm.l and Eteral Characteristics of a 45O Sweptback Wing with Symmetrical Circultur-Arc Sect
42、ions. NclCA RM 8x16, 1948. Y Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-i “1 r . -. . . . . . . . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . . . . . . . . . . 45-4-0.6 -1.0 Far
43、iatian in taper ratio t Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MACA RE4 5016. 20.00 ? 1 811 Mmensions in inches 19 -L I .Figure 2 .- Leading-edge flap anB fences OBI 45 Bweptback wing. * I f Provided by IHSNot for ResaleNo reproduction or ne
44、tworking permitted without license from IHS-,-,-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. “ . “ I I t Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. Provided by IHSNot for ResaleNo reprod
45、uction or networking permitted without license from IHS-,-,-I c I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-r * . Provided by IHSNot for ResaleN
46、o reproduction or networking permitted without license from IHS-,-,- . . . . . . . . . . . . . . . . . . -. . . . (b) Wing with 0.5G It flap, smooth. Figure 4.- Ccoltinusd. 2 . . . . . . . . . R) o . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
47、I I w (c) Wing with leading-edge rovghness. R = 3 -0 X 10 6 . Figure 4.- Cmluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . . . . . (a) Plain wing, -0th. pi- 5 .- Low-sped aerodynamic data for the 0-4-0.6 wing at various Reynolds numbers. L Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I m Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
