NASA NACA-TN-1071-1946 Wind-tunnel investigation of boundary-layer control by suction on the NACA 653-418 a = 1 0 airfoil section with a 0 29-airfoil-chord double slotted flap《带有0 .pdf

1、FOR AERONA.UTICS TECHNICAL NOTE PO l 10-f 1 ?IXD-TUISNEL IBV3STIGATIO OF EXfBCARY-IkYER CONTROL BYSUCTIOM tingley Xemorfal Uemnautical Laboratory Lmsley Fielci, Va. . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NATIaXAL ADVISORY COMIIITTEE FOR

2、AERONAUTICS TECHXICAL NOTE NO. 1071 WIND-TUKREL INVESTIGATION OF BOUITDARY-LAYER CONTROL BY SUCTION 0X TRE NACA 653418, a = 1.0 AIRFCIL SECTION WITH A C+AIRFOIL-CHCRD DOUBLE SLOTTED FLAP By John H. Quinn, Jr. Tests have been made to find the maximum lift of the NACA 653-41t3, a = 1.0 airfoil section

3、 equipped -Jvith a 0.2?-airfoil-chord double slotted flap and a boundary- layer suction slot located at 0.45 airfoil chord. The tests were mzde at Reynolds numbers of 1.9, 3.%ithout boundary-layer control, roughness decreased the maximum lift c.Iefficient from 3.11 to 2.84. At a flap deflection of 6

4、50, Reynolds number had little effect on the maximum lift attainable with boundary-layer control above a flo;v coefficient of Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA TN No. 1071 approximately 0.012 at least at Reynolds numbers between

5、1,y X lo6 and 6.0 x 106. Throughout the range ,of flow rate for which data were obtained, maximum lift coeffi- cient increased with increasing flow coefficient. In no case did the section angle of attack for maximum lift of any of the configurations tested with boundary-layer con- trol exceed by mor

6、e than 2“ or 3O the section angle of attack for maximum lift at a Reynolds number of 6.0 x 104 for the airfoil with flap retracted and no boundary-layer control. 1NTRODTJCTION k recent investigation (reference 1) was conducted on the NACA 653- 018 airfo:l section with boundary-layer control by sucti

7、on to n that in general greater maximum lift coefficients may be obtained with hLgh lift devices on relatively thick highly cazibered airfoil sections than on thin low-cabered sections, and that lai?!inar separation often limits the maxfmum lift attainable with the thin low-cambered sections. It see

8、med . likely that further development of boundary-layer control for high lift would result from tests of a cambtjrsd aping. Tests v;ere made, therefore, in the Lanbley two- dimensional low-turbulence tunnel and the Langley two- dimensional low-turbulence pressure tunnel of the NACA 65.3ic18, a =l.O

9、airfoil section with a single boundary- layer suction slot located at 0.15 airfoll ckorti that Is, profile-drag coefficient equivalent to power required to discharge at free-stream total-pressure ai.r removed from boundary layer - Bb) total drag coefficient ( Ch + cd ob ) local velocity outside boun

10、dary layer local velocity inside boundary layer perpendicular distance above airfoil surface Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 6 6 ib 8 H =0 6f X R boundary-layer total thickness boundary-layer displacement t,hickness boundary-layer m

11、omentum tticknsss boundary-layer shape parameter ( 6*/e ) section angle of attack deflection of flap chordwise distance measured from leading edge Reynolds number MODEL AND TESTS The airfoil used in this investigation was of 3-foot cnor14 and was built to the ordinates of the KkCh 655-415, a 72 1.0

12、airfoil section. The model aas constructed of laminated mahogany with laminations running in the chcrd- wise direction. Ordinates for this airfoil section are presented in table I. The model was equipped with a 0.23 double slotted flap and a suction slot located at 0.45c. A schematic drawing of the

13、model showins the suction slot, wing duct, and double slotted flap is presented as figure 1. Ordinates for the flap and vane are presented in tables II and III, respectively. The tests were made in the Langley two-dimensional low-turbulence tunnel (designated LTT) and in the Langley two-dimensional

14、low-turbulence pressure tunnel (designated TDT) . The LTT was used for the development of the best flap configuration and for the detailed boundary-layer surveys and pressure measurements; the T3T was used for tests of the most promising configurations at the higher Reynolds numbers. Roth the LTT an

15、d TDT have test sections 3 feet wide and 7.$ feet high and were designed to test models completely spanning the jet in two-dimensicnal flow. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN No. 1071 5 Lifts were measured by an arrangement desi

16、gned to inte- grate the pressures along the floor and ceiling of the tunnel test section. External drag was measured by the wake-survey method. Air was sucked off the upper surface of the model through the suction slot and into the wing duct. Zrom the wing duct it passed through the tunnel wall and

17、was ducted through a Venturi to the inlet of a blower. The volume rate of flow Q was obtained from measurements of the total and static pressures in the throat of the Venturi. For the no-flow condition, the slot was faired over with plastelfne. The loss in total pressure incurred fn sucking the air

18、through the slot plus the total-pressure deficiency of the boundary layer was obtained by measuring the pressure inside the wing duct. For some tests the local dynamic pressure outside the boundary layer just ahead of the slot was determined by p1acing.a static pressure tube at O. the vane -.vas rem

19、oved at these deflections to simplify the tests. . An arbitrary flap path ;Nas chosen to retract the flvp into the wSn 2 . the 650 and The flap moved slightly fortTar between Oo deflections, pivoted about a Doint near the nose of the vane between deflections of 6b=) and ver control were planned. to

20、find not only the effect of boundary-layer control on the lift and drag , characterfstics of the airfoil but also the relation between changes in the lift and drag characteristics and .l changes in the nature of the flow in the boundary layer. - The discussion is therefore divided into three parts.

21、The first two narts deal with the e, “ect of flo:v rate on the lift and drag characteristics of the wing with various fla=, deflections and at different Reynolds numbers bnC the-third Dart,with the effect of boundary-layer control on the variations of the boundary-layer displacement thickness and sh

22、ape ,Faram;eter and the pressure losses in the suction slot. Lift Characteristics Variation of lift coefficient with angle of attack.- The lift characteristics of the ZLCA b52-41L airfoil sectlon w.lth boundary-layer control at 6arious flap deflections and Reynolds numbers are Fresented in figure 5.

23、 The vredo.mlnter than the angle o of 6.0 x 10 8 attack for maximum lift at a Reynolds nu!?bor (fig. 5(b) for the plain wing. Consistent Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN No. 1071 ? increases in maximum lift coefficient were fou

24、nd with increasing rate of flow and wFth increasing flas deflec- tion up to flap deflections of 45*. Ait a Reynolds number of 1.9 x 106, little change in maximum lift was found with increasing flap deflection above a deflection of 45O. Kost of the lift data presented in figura 5 shon that the lift-c

25、urve slope and angle of zero lift for the wing with boundary-layer control-differ somewhat fram the values found for the no-control condition. In general the lift- curve slope tends to increase and the angle of zero lift tends to become more negative with fncreasing flow coef- ficient. The lift-curv

26、e slope probably increases because the boundary layer becomes thinner over a large Tart of the M.ng as the flow rate increases. The thinner boundary laver had an effect similar to that of increased camber an3 brought about the downward shift 2n the angle of zero lift. mfect of roughness.- Lift data

27、are presented in figure 6 for the airfoil with leading-edge roughness at a flap def16ction of 650 and with different flow rates. The roughness consisted of Carborundum grains having an average diameter of O.Oll-inch applied to both surfaces of the airfoil as far back as 0.078. in figure 6, As may be

28、 seen increaslng the flow rate above a value of 0.016 brought about only a small change in maximum lift. Co,pmison of these curves with those for the smooth wing presented in figure 5(t) shows that roughness decreased the maximum lift coefficient for the no-flow condition from 3.11 to 2.64, and from

29、 3.88 to 3.16 at a flow coefficient of 0.024. Turbulent separation uro5ably occurred upstream of the slot at angles of attack greater than that at which the lift coefficient of 3.16 iras obtained. The angle at which maximurfi lift occurred, approximately 6”, was very low compared with tha angle of a

30、ttack for maximum lfft of 17o for the smooth v:ing at the ssme flow rate, flap deflection, and Reynolds number. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Vdriations of czmLax with flap deflection.- The variations of maximeasure of the amount of

31、 the boundary layer ahead of the slot that is being removed at various flow coefficients, the ratio f 6.020 the value of Q/W*b was equal to 0.4. In reference 1 it was found that the suc.tion slots were operatrng at their maximum effectiveness irrhen Q/lw-b Wnic pressure are presented in figure 12 fo

32、r several flap Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I 10 . Nl(HO Horsepower = - Hb) 550 where Q is in cubic feet Ter second and FI, and Hb are in pounds per square foot. Boundary-layer shape parameter and displacement thickness.- The resul

33、ts of boundary-layer surveys at a FiGdefTection of 6!j“ and an angle of attack of 16O are presented in figure 13. The variation of the shape parameter H is presented in figure 13(a) and ,“;a;,o,“ the boundary-layer disclacement thickness 6$ sented in figure 13(b). AS far back as 0.25 little change i

34、n the shape ,parameter was found to occur between flow coefficients of 0.010 and 0.017. kt 0.2oc 1; had attained a value of 1.66. From this point up to the suction slot the value of H decreased, the amount of the decrease depending upon the flow rate. In refer- ence 2 It was pointed out that se arat

35、ion was itnmjnent I-or values of 8 H greater than 1. . Because at 0.2Oc H had attained a value close to 1.8, it is possible that at a slightly higher angle of attack than that for which data are presented separation would occur close to 0.20. As the flow coefficient was increased, the slot might hav

36、e an appreciable effect in the neighborhood of 0.20 and serve to delay separation to a slightly higher angle of attack. Tuft studies showed that, as the flow coef- ficient was increased, a tendency for separation to occur near the trailing edge was eliminated and smooth flow was observed over the en

37、tire wing. As the angle of attack was increased in this condition, no fluctuation of the tufts was apparent until the flow apgaared to separate from the leading edge. Increasing the flow coefficient. still further brought about no change in the nature of the stall but did increase the maximum lift c

38、oefficient and extend the straight part of the lift curve to a slightly higher angle of attack. Further straightening of the lift curve, even after turbulent separation at the rear had been eliminated by the boundary-layer control, is ascribed to the reduction of boundary-layer thickness toward the

39、rear. . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. KRCA TN No . 1071 11 The boundary-layer displacement thickness (fig. 13(b) was affected by the suctFon slot in much the same manner as the shape parameter, because t-he slot exerted an influ

40、- ence on the displacement thickness as far forward as approxfmately 0.2Oc, and directly behind the slot the dtsplacement thickness was extremely small. The variations with flow coefficient of Wring conclusions: 1. B maximum section lift coefficient of 4.16 was obtained at a flap deflection of 650 f

41、ora Reynolds number of 3.4 x 106 with boundary-la ir er control. The flow coef- ficient for trrber of 1.9 X 106, an3 a flap deflection of 65”, r.oughneso applied to the leading edge of the ling reduced thz maxi- mum lift coefficient from 5.813 to 3.16. Rithout boundarg- layer control, the maximum li

42、ft coefficient was reduced from 3.11 to 2.Euinn, John II., Jr.: Tests of the NACA 65 5 -018 hirfoiL Section with Boundary-Layer Control by uction, NACA CR No. :1 4.8;e 25 8 6$j 1?i3: -1:1 2 1.1 8 l 33 9.02 ?I 1. AP DEFI;EC!XIONS NATIONAL ADVISORY COMMllTEE FOR AERONAUTICS Provided by IHSNot for Resa

43、leNo reproduction or networking permitted without license from IHS-,-,-NATHWL AOWSOIIY CoJlNlnEI Fm AEmNurrIcs ,pi R = 1.9 x 106, tests, LPT 402, 406. Fle 5.- Lift oharaoterlstlos of the NACA 65+8 airfoil motion with a 0.290 double alotted flap and how-layar oontrol. . . Provided by IHSNot for Resal

44、eNo reproduction or networking permitted without license from IHS-,-,-1 . _ . r I HACA TN No. 1071 Fig. 5b . (b) Q = o”; taut, TDT 892. Fiwe 5.- OO?ithuad. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.g. 5c NACA TN No. 1071 Fj (0) Of = 100; R = 1

45、.9 x 106; teat, Lrn 402, 406. Flgur0 5.- ctontlnu6d. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c NACA TN No. 1071 Fig. 5d . (d) 8f p 20; 2 = 1.9 X 10 6; toot. LTplpJ2. 406. Figurr 5.- CantlmnQ. Provided by IHSNot for ResaleNo reproduction or ne

46、tworking permitted without license from IHS-,-,-Fig. 5e NACA TN No. 1071 f (0) Of = 30; R = 1.9 x 106; t08ts, LTT 402, 406, Flgurs 5.- Contlnuad. c . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN No. 1071 Fig. 5f . . . .= .- f,;j i-.1:- i I

47、 -t i-l- i 1 I- t .t t t t t- t t- ;: r ._ i i : : . i-l (f) a* = 40: R = 1.9 X 106; teats, I,Tl l - .,:, . I :I +-t-t-t (9) bF = 45; R = 1.9 X 106; tests, LIT 402, 40s. igirs 5.- aontinwd. l , . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA T

48、N No. 1071 Fig. 5h . . (h) Q = 4505 test, DT 892. a Figure 5.- OontLnusd. / Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Pig. 51 NACA TN-No. 1071 tion angle of attaok, (i) ef = 500; R = 1.9 x 10 6 ; tests, LIT 402, 406. Flgu-e 5.- Ormtlnwh. c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-