1、- LOW-SPEED CHARACTENSTICS OF A 45O SWEPTBACK WING OF ASPECT RATIO 8 FROM PRESSURE DISTmBUTIONS AND FORCE TESTS AT REYNOLDS NUMBERS FROM 1,500,000 TO 4 ,EcoO,OOO By Robert R. Graham NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON October 22, 1951 Provided by IHSNot for ResaleNo reproduction o
2、r networking permitted without license from IHS-,-,-UNCLASS t FIEO 1 NACA RM L5lHl3 -l?IemD- - .I NATIOHAL ADVISORY COMMITTEE FOR MROJYAUTICS LOW-SPEED CBARACTERISTICS OF A 45O ShTEPTBACK WING OF ASPECT RATIO 8 FROM PRESSURE DIGTRIBGTIONS AmD FOXE TESTS AT REYNOLDS NUMBERS FROM 1, WO, 000 TO 4,800,0
3、00 By Robert R. Graham SUMMARY Results are presented of 89 imeetigation in the Langley 19-foot pressure tunnel of tbe longitudinal characteristics of a wing ha- 45 sweepback of the quarter-chord line, an aspect ratio of 8, = taper ratio of 0.45, and WA 63L012 airfoil sections -parallel to the plane
4、of symmetry. The reeults were obtained from force measuretaents through a Fn long-range, high-speed airplanes has created a demand for irdormation on swept wings in the higher aspect-ratio range. Accordingly, the low-speed characteristics of a wing of aspect ratio 8 with the quarter-chord line mept
5、back 45O were investigated in the Laslgley .l9-foot pressure tunnel. The win; was untwisted and had a taper ratio of 0.45 and NACA 631012 airfoil sections parallel to the plane of symmetry. The investigation included the determination of the characteristics of the wing by force and pressure-distribu
6、tion measurements. The tests were made at a Reynold6 number of 4,000,000 and a Mach number of 0.19. The effects of wrying the Reynolds number from 1,500,000 to 4,800,000 and of leading-edge roughness were Investigated. The effects of one configuration of chordwise fences were also Fnvestigated. mplB
7、OIS The data are referred to the uind axe6 with the orfgin at projection on the plane of symmetry of the quarter chord of the mean aerodynamic Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L5m3 3 chord. The data have been reduced to nondime
8、nsiond coefficients which are defined as follows: cL Ilft coefficient (- or cz 5 a( The basic loading due to the spanwise angle variation ia preserrted in figure 5 and was obtained by multiplying the angles by the slopes of the section-lift curves obtained from the pressure measurements. The. basic
9、loading obtained from pressure measure- nients at zero lift is also presented in figure 5. The small differences between the two basic load- curves are probably due to slight hac- curacies in the conatruction of the model and experimental inaccuracies in measuring the air-stream angles and model pre
10、saure distribution. No satisfactory method is known for correctin; the individual pressure coefficients (table I and fig. 6) for the basic loading, but the force and moment coefficients integrated from the pressure- distribution data have been corrected at all angles of attack by sub- tracting the l
11、oading obtained at zero lift. The pitching-moment coef- ficients from the force tests have been corrected for the momrrt due to the basic loading on the BE#, wiq. The basic loading from the preaaure measurements was used far correcthg the data because possible model inaccuracies on the untwisted, un
12、cambered wing would be corrected for along with the angle variation. .I Mo correction was applied to take into account the spariwise fraria- tion of the jet-boundary-induced angle or the model twist-due to loading. Calculations of the spanwise variation of the induced angles and measure- ments of mo
13、del ,twist angles indicated them to be small (0.2 at CL = 1.0) and of the sam order of magnitude but opposite in direction. Presentation of Data The results of the pressure-distribution tests made on the plain wing at a Reynolds number of 4,000,000 are presented as pressure coef- ficients in table I
14、- and figure 6. The section-lift characteristics integrated from the pressure data of table I and ffgure 6 are presented in figure 7. Also presented in figure 7 are the section-lift character- istfcs integrated from pressure-distribution data for the plain wing at a Reynolds number of 1,0,000 and fo
15、r the wing with fences and with leading-edge roughness .at a Reynolds number of- 4,000,000. “ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 MCA RM L5lEl3 9 The results of force tests on the plain WFng through a Reynolds number range from 1,500,00
16、0 to 4,800,000 are presented in figure 8. The results of spanwise integrations of the section characteristics are also presented in figure 8. A few of the chordwise pressure diagTams obtained, on the wing at a Reynolds number of 1,500,000 are presented in figure 9. The effects of fences and leading-
17、edge roughneas on the force character- istics of the wing at a Reynolds nuniber of 4,000, OOO are presented in figure 10. The pressure-distribution data for the wing at a Reynolds number of 4,000,000 have been integrated to give section pitching- moment and drag coefficients (fig. ll), span-loading
18、coefficients (fig. 12), spanwise, chordwise, and vertical centers of pressure (fig. 13), local chordwise centers of pressure (fig. 14), and wing bending- and twisting-moment coefficients (fig. 15) . Lift and PitchFng-Moment Characterlatics The section-lift curves for the phin wing at a Reynolds numb
19、er of 4,000,000 (fig. 7) show that the lift for the root sections increases nearly linearly with angle of attack up to the highest angle of the tests (31O). The lift curves for the tip sectiona show a decrease in lift-curve slope at low angles of attack (about 5O for the 0.96b/2 sta- tion) and a lev
20、euing off at around loo to 12. The combination of the linear variation at the root and the nonlinea variation at the tip causes a nonlinear pitch%-moment variation for the wing. (See fig. 8.) - The decreasing lift-curve slope at the tip sections cau8es a forward movement of the aerodynamfc center wh
21、ich begins at about 5 angle of attack and about 0.3 lift coefficient and continues to maximum lift. I The section pitching-moment characteristics (fig. 11) have a negligible effect on the WFng pitching-moment characteristics as demon- strated by the fact that at maximum lift the varfations of the se
22、ction pitching moments are stable while those of the wing pitchimg momenta are unstable. The decreasing llft-curve slope over the outboard sections is also reflected in the wing lift curve (fig. 8) where the slope starts decreasing at about 5O angle of attack. At about 20, the increasing lift at the
23、 root is just offset by decreasing lift over the outer por- tions of the wlng so that above an angle of attack of 2oo the lift coefficient is constant at about 1.01 up to the highest angle of the tests (31O) . I Effects of Reynolds number variation.- The effects of a reduction in Reynolds number fro
24、m 4,000,000 to 1,0,000 on the section lift Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 NACA RM L5lHJ-3 ch coefficients than were obtained at the same angles of attack at the higher Reynolds nrzlllber. As the angle of attack is increased furthe
25、r, the lift decreases again. Examination of the chordwiee pressure diagrams for the lower Reynolds number (fig. 9) reveals that the first loss of lift (a = l2.9*) was caused by trailing-edge seps,ratiC$I. The fncrease Fn lift following the Fnftial loss of lift (a = 14.9 ) occurred when separation wa
26、s complete over the Rzll chord of the 0.9Ob/2 and 0.96b/2 stations. At that -le of attack the loss of lift over the nose of the sections was more than offset by an increase in lift; over the rearward portions. At the 0.73/2 station the pressure diagram for 05 = 14.90 shows a widening out of the low
27、pressure area over the for- ward part of the section which more than offsets the loss of the peak at the leading edge. Although this diagram is similar to those obtained Fn the vortex type of flow (see reference 4), there was no evidence of that type of flow on this wing as shown by surveya with a s
28、ingle tuft of yarn on a long probe. The O.73/2 station waa obviously in a transi- tion region between separded and meparated flow where the separated- flow region ex%ended farther inboard at the leading edge than at the trailing edge. At 15. go angle of attack the stalled region moved inboard slight
29、ly so that the low pressure region was broader and the section lift coefficient was still higher. These lift variations over the tip sections caused the pitching-moment curve for the complete win “_ 1.26 1.31 “ “ :g 1: 2 1.28 1.21 1 .oo :,“E ;.a 1:9 122 1.23* 1.238 1.17 1.23 12 1.21 1:23 : : 3 :g 1.
30、02 “ 1.w .a4 “ “ 1.06 1.21 -95 “ “ “ “ 1.15 “ 1.25 “- “ 1 .I1 1.16 1.228 “ 1 :* ; t? 1 :2i 1.7 1.23 . 1.17 1.11. 1.07 a = 0.6 7 f- 0.61 -72 3; 1:2$ 1.2% 1.01 12 1.21 1.13 1.00 .9f 1 1.12 1.22 1.26 1.2 1.25 :z 1.28 1.9 , 1-19. 1.11 I“ I .60 80 .9f “ “ - e3 1.01 1.10 “ “ “ 1.19 “ -91 1 .os “- .93 1.10
31、 “ 1.17 “ ;:OO i “ - .0?90 - .0350 - .o - .0579 - . 1.10 1.15 1.11 1.13 1.09 I:% 0 -73 “ “ “ “ “ “ a = 3.70 e:2 1.41 1.29 1-51 iif 1.41 1. 1.36 1.29 1.21 1.12 3 -9 -62 “ “- “ -91 1.02 1.10 “ :3 :2 1.13 1.00 .* -59 1 1 1 1.10 :z 1-73 1.68 1.56 1.61 :E 1-37 1.20 I .ll LO3 -95 -64 .el .94 “ “- “- “ q 1
32、.1 1.10 1.15 :3 44 .61 -77 .ss “ “ 1.02 “ ;g 1.11 1.06 1.02 :3 “- -63 .z 1.23 1.16 1.25 1.15 1.05 “ -40 .61 -75 “ “ 1.00 .e8 1.06 1 .I1 1.13 1.08 1 -11 1.04 -99 a-= 4.7 “ I 1.28 .99a :a2 I “ 1 .79 I “ -1 “ . 97p I 1.06 1.41 : aa :ai :g 3; 1.44 1.39 1.32 1.13 1.Z l:;z “ -58 “ “ “ .e5 1% 1.11 1.13 1 .
33、ll 1.03 1.M - 99 -95 ;:J% :% :3 :x; ;:$ 3: 1.61 1.21 1 .1? 1.05 .97 .56 .68 .81 -95 1.z “ “ “ “ :4 “ a = 6.8 I 2.06 1.23 1.13 l:;2 -58 .60 “ “ -71 “ 05 “ 1.02 1.a .97 : : z ;:Do3 .:za 1-93 i.8 1. 1.40 “ 1.20 1.11 1.03 -95 .60 .74 -86 1 .ll 1.01 “ “- “_ :zj : : 3 1.03 -95 1.00 33 1:g :z: ;:.% : 3 1.9
34、0 1-94 3.199 , 1.38 1.09 1.01 .93 -57 -70 -83 “ “ I“ 1:8 1.11 1.12 1.10 1 .06 1.0P -98 .94 T i I “ -77 .60 .72 “ “ :3 1. -97 -95 “ .60 -74 .e “ “ “ .bo “ a65 -77 “ 13 1.10 1-09 1.05 1.08 1.03 :Bil bout 0.00350 frm rlng surface. w Provided by IHSNot for ResaleNo reproduction or networking permitted w
35、ithout license from IHS-,-,-WA RM L5lEl3 PAW I.- VALUES 02 EXPERIY%!WT B = ,000,000 ortrim loaatf an Preaaure coefilaient. 9 1.26 I 3.00 3% x: ;:a :z; 1-74 1.31 1.18 1.09 1.03 1.00 .66 -56 .a .79 -91 1.02 -98 1.03 1.02 1.00 -98 .98 “- “ “ “ “ “ I i “ 1.10 1.21 !:% .68 .59 -68 “- “- 1.09 1.01 -98 -97
36、 “ -2 5 -45 “_ -I- .60 .66 -55 -63 -7-7 *ss “ “ “ “- ;:g “ 1.30 1.02 -98 - .O -2 1.02 .99 - I .89. a = t:; 2% :-P 1:2;5 : 3 4 -53 4.25 3.82 3.u 1.18 1.20 1.18 1.21 .a5 - 58 .62 1.02 -93 “ “- “- -78 y7 13 :z 1.08 O:g i.Y 1% 1.10 1-33 1.62 1.62 1.21 1.07 . 11 0 2 .i- 1.41 1-78 I.* 1.30 t.27 1:g 1.31 1
37、.30 _“_ 1.11 “ -57 -53 “- :% ;:3 .93 1.01 1.06 “ -63 :;.? -95 -99 -97 1.01 “ “ - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-20 HACA RM L5lHl3 I I - .04 .04 -.02* -.wgo -.020z “ “ -44 - -0579 .74 -.O 2.78 “ “ “ 1.01= “ “ 47 “ “ .79 .93a .go. “ “
38、“ 3.30 “ “ “ 1-45. “ “ 1.00 :be 2.66 2.03 1.51 1.36 1.31 1.31 1.31 1.30 1.30 1- 33 :3 1.26 .74 .53 * 56 “ - “- “ 1. 1.06 1.09 1.08 1.11 1.16 . aThese pressurea manured with atatlo-4 survu). tubs ibout 0.00jgc from mlng surface. =Ev- Provided by IHSNot for ResaleNo reproduction or networking permitte
39、d without license from IHS-,-,-NACA RM L5lEl3 21 a = 19.0 r .e2 2.05 1-92 2.00 1.96 1.65 1.61 38 :a :a2 :g 1.d I .2 .a2 -13 -25 -38 .68 “ “ “ :a: ;g .91 3 -78 “ I “ -52 -62 :% :2 1.00 -94 I .a 1.21 a = 21.0 I 0 4-56 “ 1 .oo 1.71 ?.lo 3 :zz 1.45 1-95 1.3 1.12 .a2 .10 .20 “ “ “ .001 .w25 .01 5 .0050 .
40、0250 :E :t5 2; :i5 .95 .I5 -25 “ 1.53, “ 1.02a 1.02 .54 -74 “- .52 “I -70 -54 -55 “ -.dm “ .63a “ “ .7bL Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-22 EJACA RM L5UU3 . TAB= I.- VALUES OB EXPERIMENTAL PRESSURE COEFFICIENT - Conoluded Uncorrected
41、for baslc loading due to apamlas varlatfon of tunnel stream angle; R = 4,OOO,OOO1 I 6.68 “ “ “ 4 -89 ;:s? ?:a 2 :;f: 2.96 “ 2.26 1 - 55a 1.68 -7 5 36 “ “ “ “ .i?l .20 .2Ba .46a .61a .9* “ “ “ “ a = 3a.oo 2.24 2 -23 2.23 2.19 2.17 2.17 2.15 2.13 2.21 2.12 2.1 2 -16 2.10 2.06 2.08 2.06 2.1% 1.37 .67 .
42、44 -40 .60 “ “ “ - -“ :g ;:3 “ i .l+4 1.68 : 2 ;:$ 1.64 1.62 1.63 1.63 1.63 1.63 kg 1.61 1.60 -91 -9 -46 1.60 “ “ “ 2; 2; -90 1 .oj * 97 1.22 1.29 aLmese pressures measured with statio-pressure survey tube about 0.00350 frm rlng mrfaco. Provided by IHSNot for ResaleNo reproduction or networking perm
43、itted without license from IHS-,-,-MACA RM L5l.Hl-3 . O-OJPl Typicd chordwise orifice foeations Figure 1.- Geometric characteristics of model. Aspect ratio 8.02; taper ratio 0.45; atrfoil section 631AO12; wing area 14.021 sq ft. (Dimensions n inches eXcept as noted. 1 Provided by IHSNot for ResaleNo
44、 reproduction or networking permitted without license from IHS-,-,-Figure 2. - kdel in tunnel for force tests. (Front dew. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a 1 a Figure 3.- Model in tunnel for pressure-distribution teats. (Rear view.)
45、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-26 NACA M L5lH13 . Figure 4. - Detail6 of chordwise fences. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-UCA RM 5x13 . 27 cc Z& C Determfned from: Sp
46、an loading at zero lift - -.- -Air a tream surveys .04 0 -. 04 0 .2 .4 .6 - 2Y b .8 1.0 v. Figure 5.- Basic loading aid angle-of-attack distribution BCTOSB left wing of model. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 NACA RM 5113 8 6 4 2 S
47、0 8 4 S P 0 6 4 S 0 0 .20 .40 .60 .80 i.00 0 .PO .40 .60 .810 1.00 we x/. (a) a = -0.4. Figure 6.- Chordwise pressure diagrams for plain rring. R = 4,000,000. (Flagged syibols denote pressures measured utth survey tube. ) Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RACA RM L5lHl3 6 4 S &? 0 8 6 S 4 2 8 6 S 4 a 0 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
