1、. C-I I co$y 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . NACA RM A5lP27 - ?XATIONALADVTSCRY CO- FOR AERONAl3rICS RESEARCH MEMomuM CHARACTERISTICSTHROzRHOUTTHE SUBSONIC SPEEDRANGE OF APLAIEWINGAXDCFACAM was cambered for a design lift coeffic
2、ient of 0.4 and twisted to relieve the loading at the tip which accompanies sweepback. The airfoil sections normal to the quarter-hord line were the NACA 64010 for the plane wing and the NACA 64A810 for the cambered and twisted wFng. The cambered and twisted wing had 8.70 of washout between the root
3、 and the tip. The tests were made at Mach numbers from 0.25 to 0.94. At each Mach number the -3 Reynolds number was varied over as -wide a range as possible within the limitations of wind-tunnel power and wind-tunnel pressure. At Mach nun+ bers above 0.70, the maximumReynolds number was 2,OOC,OOC; a
4、t a Mach number of 0.25; the maximumReynolds number was lO,OCKl,OOO. The effects of a fuselage, of boundary-layer fences, and of surface roughness were also investigated. At a Reynolds number of lO,OOO,C!CO and a Mach number of 0.25, the combined camber and twist were effective in delaying extensive
5、 separation on the wing to a higher lift coeffictent. At the lower Reynolds numbers, the effectiveness of camber and twist in delaying extensive separation was seriously reduced. The aerodynamic characteristics of both wings were seriously influenced by dynamic-scale effects. At Mach nuln models had
6、 a fineness ratio of 12.5. The equation defining the coordinates of the body is given in figure 1. The plane wing was mounted with its root chord co which is a closed-throat variable-density wind tunnel with a low turb* lence level closely approximating that of free air. As shown in figure 3, the mo
7、dels were mounted with the wing plane perpendicular to the floor which served as a reflection plane. The gap between the body and the tunnel floor was maintained between l/32 and 1/16 inch. No attempt was made to remove.the tunnel-floor boundary layer which, at the location of the model, had a dis-
8、alone ana alone ana alcme and alone bW boas bow e- 3.0072 0.0061. -i 0.0065 A074 .0063 0.0072 .oo6g .oo79 067 .0077 .0073 -o086 :z: .oq8 .oogo c- o.oo76 .0081 .ooa8 .0092 I - -L-L PlaneWing 1 Effects of Mach nmiber.- In figures 4 through 7, lift, drag, and pitchlng-mmm t aata for the plane wUg are p
9、resented for Mach numbers from 0.a to 0.94 and a RepUs number of 2,000,000, for M;Bch nuder8 from 0.25 to 0.70 and a Reynolds nuniber of 3,000,000, and for Mach nunitmrs from 0.15 to 0.60 and a Reynoa number of ,OOO,OOO. The data in figure 4 jndicate that, as Mach number increased, there was an Fncr
10、eaee in tbs lift-curve slope for lift coef- ficients less than 0.4 throughout the Mach number range at Reynolds nunibera less than 4,ooO,ooO and a decrease in the maximum lift for Mach numbers up to 0.90. In figure 6, the variation of drag coefficient with Mach nuuiber at a Reynolds n and second, th
11、at some high Reynolds number probably exists at which little impro-vement in the wing characteristics will result from the addition of fences. At a Reynolds nuwhile on the cambered and twisted wing, separation originated in th8 turbulent boundary layer at the trailing edge. The marked difference in
12、the 8ff8c- Wmuxs of the fences on the two-s suggests that even at a Reynolds nmmr of 2,33,WO and at Mach nu Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-HACARMA5lD27 29 .04 .02 n 11 iii f i (c) c,=o.2 04 02 01 ” ” i I I- s - I I! I! I! c-c; I I I ” ” ” (bj Gp0.i .04 .02 Figure 6.- The measured vurioth of drug coefficienf wifh MU& number of the phe wing compared with fhe variafh ca/cu/ated fmm secfion dofa. R, 2*ouu,ooo. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
