NASA NACA-RM-L54C05-1954 Effect on the low-speed aerodynamic characteristics of a 49 degrees sweptback wing having an aspect ratio of 3 78 of blowing air over the trailing-edge fla.pdf

1、i “ 1 P I - CUUJ.DENTIAL COPY 6 RM L54C05 E- - RESEARCH MEMORANDUM EFFECT ON TEE LOW-SPEED AERODYNAMIC CHAFLACTERLSTICS OF A 49* SWEPTBACK WING HAVING AN ASPECT RATIO OF 3.78 OF BLOWING AIR OVER THE TRAILING-EDGE FLAP AND AILERON By Edward F. Whittle, Jr., and Stanley Lipson Langley Aeronautical Lab

2、oratory Langley Field, Va. x USSlRCATlON CANCELLED . FOR AERONAUTICS Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1P 8 * NACA F however, analysis of the effects due to blowing is necessazily IMted since only a few tests were conducted with blowing

3、 alone over the trailing-edge flap (zero suction at the midchord of the a). In view of the possibility, then, of incresskg wing lift by means of blowing over the trailing-edge flap, the method has been extended to the case of a highly swept, thin wing. Tests have been conducted In the Langley full-s

4、cale tunnel on a semispan .lo sweptback UFng having NACA 65A006 airfoil sections, an aspect ratio of 3.78, and a taper ratio of 0.59. Preliminary tests were conducted with a low-capacity blower and are presented in reference 3. Because of the very low preseure rise and quantity of flow of the blower

5、, no significant results were obtained. The investigation reported herein is a continuation of the full-scale- tunnel blowing tests but with a multistage, large-flow-capacity blower. The tests were made with and without a slat- and fences installed and with and without blow- over a 6railFng-edge fla

6、p or trailing-edge flap . * and aileron. Tests were ale0 made to determine the rolling effectiveness produced by blow- air over the aileron. In addition, some chordwise pressure distributions were obtained at the midspan of the trailing-edge flap in order to study the load change that occurred a a r

7、esult of the blowing method of bomdaq-layer control. i c The tests were m are .given in figure 1 and details of the slat, fences, and flap me given in figure 2. A photograph of the wing mounted on the reflection plane in the Langley full-scale tunnel is given as figure 3 and a description of the ref

8、lection plane is given in reference 4. The wing has 49. lo of sweepback at the leading edge, an aspect ratio of 3 48, a taper ratio of 0.39, and no geometric twist or dihedral. The airfoil sections parallel to the plane of symmetry are NACA 65A006 sections and the wing tip is half of a body of revol

9、ution based on the same airfoil section ordinates. The high-lift and stall-control devices .used (see figs. 1 and 2) are: a 0.266 inboard trailing-edge flap having a span of 0.469b/2; a 0.266 flap-type aileron, which only could be deflected down, located imediately outboard of the flap and having a

10、span of 0.234b/2; a 0.13 leading-edge slat having a span of 0.500b/2, measured inboard from the wing tip; and chordwise fences having a height of 0.06 and located at spanwise stations, measured outboard from the plane of symmetry, of 0.6b/2 or 0.6b/2 and 0.8b/2. - The nose and qper surface of the sl

11、at have the ordinates of the wing airfoil. The slat is not an integral paxt of the wing but is mounted directly onto the unmodified leading edge of the basic wing with the slat I brackets alined normal to the leading edge of the wing. The fences axe made of l/4-Fnch plywood and are mounted parallel

12、to the plane of symmetry. Just ahead of the trailing-edge flap and aileron is a slot (fig. 2) which opens into the upper portion of the gap between the airfoil and the flap and aileron. The slot is ueed for blowing a high-energy stream of air over the weer- surface of the flap and aileron. The wing

13、area affected by blow3ng over the flap is 76.4 sqwe feet and the wing mea affected by blowing over the aileron and flap is 108.0 square feet. At the midspan of the flap a thin strip of belt pressure tubing was glued to the surface of the flap perpendicular to the 0.50 I line (see fig. 1) at one spaa

14、rise station so that flap chordwise pressure distri- butions could be obtained for several of the configurations tested. Blower-ducting apparatus.- A modified compressor of a Jet engine, driven through a 2.6 to 1 ratio gearbox by two 200-horsepower electric motors in tandem, was used as the pumping

15、source for the boundary-layer- control air. The compressor was modified by removing three of the six stages in order to reduce the pressure rise and horsepower requirements for driving the compressor at high flow qwtities. The three remining Provided by IHSNot for ResaleNo reproduction or networking

16、 permitted without license from IHS-,-,-NACA RM LwO5 7 stages of the modified compressor produced a pressure rise of 1.2 at the maximum compressor speed tested. A calibrated entrance bell, Fnstalled at the compressor inlet, was used to determine the mass flow of air. A shielded thermocouple and a sh

17、ielded total-pressure tube were wed to obtain the temperature and pressure of the boundazy-layer-control air at the wing root. !These temperature and pressure measurements were used in conjunction with the hewn flow weight Fn order to determine the flow quantity of the boundary-layer-control air. 3

18、The blower is connected to the blowing slot ahead of the flap and aileron by a duct inside the wing which extends through the reflection plane at the wing root. A mercury seal was used beneath the reflection plane between the uing duct and the stationary blower duct in order to prevent transmission

19、of forces from the stationary duct to the wind- tunnel scale system. The blowing-slot gap could be varied by manually adjusting a spanwise series of throttling plates. As a result of springing of the wing upper surface at the blowing slot, the blowing-slot gap, with the blower operating at 9,600 rpm

20、, was about 0.004 when the flap was deflected and about 0.003 when the flap asd aileron were deflec- ted. A rake of shielded total-pressure tubes was employed to check the resulting velocity distribution along the blowing slot. The velocity of the air exiting from and perpendicd-m to the blouFng slo

21、t ahead of the flap (aileron blowing slot sealed) varied from 415 ft/sec at the out- board end of the flap to 450 f%/sec at the inboard end of the flap to air exiting from and perpendicular to the blowing slot ahead of the aileron and flap Varied from 388 ft/sec at the outboasd end of the aileron to

22、 448 ft/sec at the inboard end of the flap to give an Integrated average velocity of 404 ft/sec. The largest vmiation occurred over about the inbomd 30 percent of the flap span, with the highest velocity at the very inboard end of the flap. - * give an integrated average velocity of 423 ft/sec. The

23、velocity of the Tests.- An index of the test conditions and configurations tested is given in table I. Data were obtained through an angle-of-attack range from approximately -ko to 31. Force measurements were made to determine the lift, drag, pitching moment, and spanwise center-of-pressure varia- t

24、ion of the basic wing and the wing with various combinations of the high-lift and stall-control devices without and with blowing a hfgh- energy stream of air over the flap or flap and aileron. The rolling- moment characteristics of the aileron were determined with the trailing- edge flap neutral and

25、 deflected, and with end without blowing. With blarlng, the flow coefficient CQ was varied by varying either blower rotational speed or tunnel velocity. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 I NACA 24 L5kCOj Chordwise pressure distributio

26、ns were obtained on the trailing-edge flap at the midspan station for several test-conditions. Flow studies, using woolen tufts attached to the upper surface of the wing, were made for several of the wlng configurations. The tests were made at Reynolds numbers of 2.9 x 106, 4.4 x 106, and 6.1 x 106

27、corresponding to Mach numbers of 0.03, 0.07, and 0.10, respectively. . V Corrections.- The data have been corrected for airstream misaline- ment, blocking effects, and jet-boundary effects. The jet-boundary corrections follow the method outlined in reference 5 for semispan wings. The rolling-moment

28、correction for the effects of the reflection plane, as discwsed in reference 4, was obtained from unpublished results based on the methods of references 6 and 7. h-esentation of drag data.- In comparing the drag characteristics of a wing egploylng boundary-layer contro1,by blowing with the drag ry o

29、f the more significant results. Figure 22 illustrates the variation of Cp obtained with CQ for the subject wFng, I. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2P 1 A wide range of values of the mmentllm coefficient CP could be obtained during

30、this investigation only by using Mge values of CQ and blowing-slot gap, since the compressor and power available for blowing limited the available pressure rise and thus restricted this investiga- tion to testing at mderate values of blowing-slot exiting velocities. Even though the maximwn values of

31、 CQ may be unrealistically high, it fs felt that the effects obtained me indicative of those that would be obtained at similar values of Cp produced by combining high blowing- slot exiting velocities with low flow rates typical of bleed systems having high pressure and small mass flow that are curre

32、ntly adapted from turbo jet -engine installations. Since “Q be seen that for any given combination of Cp and CQ the ratio V Vo is fixed. For this fixed ratio of V Vo, then, there is some value of V, which, if exceeded, will require a supersonic Vj. The most critical CQ - Cp combination tested herein

33、 was the case of CQ = 0.042 and Cp = 0.68 for which condition the limiting value of Vo would be a Mach nuniber of 0.12. Therefore, for -ding or take-off speeds above a Mach number of 0.12, a supersontc blowing jet would be required to obtain this af0r.e- mentioned CQ - Cp combination. It is of inter

34、est to note that, for the subject wing, the required blowing-slot pressure coefficient Cp could be accurately estimated by the method of reference 8 which hdicates that Cp for a blow3n.g arrangement of the type tested may be considered as being approximately equal to VJ Vo . The values of Cp compute

35、d by this simple relationship are 16, 34, and 66 as compared to measured values of 15, 31, and 70 for the case of the flap deflected 53O and correspondlng values of CQ of 0.02, 0 .O3, and 0.04. JI J/ (1) Lift Chaxacteristics A sumnary of the miation of %* and with flow coef- ficient CQ and momentum

36、coefficient Ccr is presented in figures X) and 21, respectively. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 I NACA RM LwO5 With the slat and fences installed, the maximum-lift gah6 obtained for the 0.47b/2 flap deflected 53O and the 0.7Ob/2 f

37、lap (that- is, flap plus aileron) deflected 53O are as follows: c 0.47 0 -39 0*12 0 0 - 94 63 .042 .68 .46 .16 o 0 .68 t I I I I I t I I .68 r The increase in CL obtained by applying suction on a flap may be attributed to the increased circulation around the wing associated with alleviation of separ

38、ation on the flap. When applying suction to a flap, then, the maxirrfum increase Fn CL is limited to that associated with obtaining the theoretical flap effectiveness. By blowing a high-energy stream of air over a flap, however, the maximum increase in CL can be greater than the increase.associated

39、with obtaining the theoretical flap effectiveness. This additional increase in CL is probably associated with (I) for a Oo, a lift component due to the thruet of the ejected air and (2) an increase in circulation around the wing due to a flow condition simulating a physical extension of the flap cho

40、rd and Fesulting from the momentum of the ejected ab. It was of Fnterest to determine whether the theoretical lift incre- ment for EL 0.4p/2 flap deflected 53O (44 in the streamwise direction) was realized by blowing air over the upper surface of the trailing-edge flap. It was calculated, by means o

41、f reference 9, that the theoretical lift increment was 0.70, and this lift increment was obtained during the tests for a = 00 at C = 0.25 (CQ = 0.025). The results presented in figure 21 show that, at a given value of Cp, increasing the flap deflection from 45O to 53O produced a small increase in bu

42、t actually slightly reduced Eh in the Cp - range tested. From a study of cer*ain pitching-moment data, as discussed in the section on “Pitching-Moment Characteristics,“ it ie believed that, for the higher values of CQ and flap angle the blowlng air was not properly impinging on the upper surface of

43、the flap. It may weU be, then, that for a flap deflection of 530 a more efficient slot arrange- ment would not only have resulted in a higher . LO - By deflecting the aileon in combination with the 0.47b/2 flap, a continuous flap spen of 0.7Ob/2 could be obtained. A comparison of the relative effect

44、iveness of the two different flap arrangercents with blaring can be made either on the basis of equal CQ or equal air quantity Q. When compared on the basis of equal CQ, and are larger for the larger flap span than for the smaller flap span for the range of CQ tested (fig. X). It should be noted tha

45、t, on the basis of equal compressor air quantities, vales of CQ of approximate- 0.03 (c = 0.35) and 0.04 (Cp = 0.61) for the flapdeflected configuration correspond to values of CQ of about 0.02 Cp = 0.21) and 0.03 (Cp = 0.46), respec- tively, for the configuration with the flap plus aileron deflecte

46、d. Deflecting the aileron, then, reduced the quantity being eJected over the flap and probably reduced the local circulation on the flapped portion of the wing. However, the increased lift on the outboard part of the wing containing the aileron was such that at low and mderate angles of attack the o

47、verall wing lift was greater than that obtained by blowkg the total air flow over the flap alone. A comparison of f Gures ll and 12 and fig- ures 20 and 21 indicates that, for a given air flow, Cko was Increased about x) percent to 30 percent by deflecting the aileron. Beyond an angle of attack of a

48、bout “0, the wing lift-curve slope for the O.7Ob/2 flap configuration was reduced as compared to the case of blowing over the 0.47b/2 flap. This difference appears to be a result of a more rapid reduction in flap load at the moderate angles of attack for the 0.7Ob/2 flap configuration due to its low

49、er CQ (see section on “Pitching-Moment Characteristics“). The rougher flow obtained at the higher angles of attack over the deflected flap and aileron for the O.7Ob/2 flap arrange- ment, as compared to the 0.47b/2 flap configuration, is evident in fig- ure 13. The net result, then, was that deflecting the aileron to 53O, without increasing Q, and employing it as a high-lift device produced only a small increase in Ck. %FO ( Pro