NASA NACA-RM-L51J04-1952 Low-speed longitudinal characteristics of a 45 degrees sweptback wing of aspect ratio 8 with high-lift and stall-control devices at Reynolds numbers from 1.pdf

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1、1 . c RESEARCH MEMORANDUM Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. bE NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEAIiCH NEXORANDUM LOW-SPEEI LONGITUDINAL CHARACTERISTICS OF A 45 SWEPTBACK WING OF ASPECT MTIO 8 WITH HIGH-LIFT AND STALL-CO

2、NTROL DEVICES AT REXX0L;DS NUMBERS FROM 1,500,000 TO 4,800,000 Ey George L. Pratt and E. Rousseau Shields SUMMARY The low-speed static longitudinal stability characteristics of a wing having 45 sweepback of the quarter-chord line, an aspect ratio to the air stream were investigated in the Langley 19

3、-foot pressure tunnel at Reynolds numbers from 1.5 X 10 to 4.8 x 10 . The effects of combinations of leading-edge and trailing-edge flaps, upper-surface flow-control fences, and a fuselage. on the longitudinal stability char- acteristics were determined. I of 8, a taper ratio of 0.45, and NACA 63101

4、2 airfoil sections parallel 6 6 The basic wing had a maximum lift coefficient of 1.01, exhibited a large degree of inetability throughout the lift range, and was unstable at maximum lift. With a combination 09 leading-edge and trailing-edge flaps and upper-surface fences, a maximum lirt coefficient

5、of 1.50 was obtained, the movement of the aerodynamic center was reduced to less than 6 percent of the mean aerodynamic chord throughout the lift range, and the pitching moment was stable at maximum lift. INTRODUCTIOET Previous investigations of sweptback wings (see, for example, refer- ences 1, 2,

6、and 3) have sham that as the aspect ratio and sweepback are stability throughout the lift range with the various devices used to control the stalling of sweptback wings. In order to extend these inves- tigations and to provide information in the low-speed range with which increased, it becomes incre

7、asingly difficult to provide longitudinal -b Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 -rc. to evaluate design configurations suitable for high-subsonic, long-range airplanes, sn investigation has been conducted in the Langley 19-foot pressur

8、e tunnel to determine the-.lar-speed longitudinal characteristics of a 45 sxeptback wing of aspect ratio 8. A wing of this sweep - aspect- ratio combination is well in the longitudfnally unstable region a8 set forth in reference 4, and on the basis of prqsent qagufacturing methods appears to be appr

9、oaching a- limit outsuration. 6 6 6 Results of measilrements of the pressLizontal tail on the lorqitudinal .stability are presented in references 5 and 6, respectively. The data are referred to a wind axis with the origin located at-the projection of the quarter-chord poin,t of the mean aerodynamic

10、chord on the plane of eymme-Standard EIACA symbols and coefficients are used. CL lift coefficient (Lin/qS) CD drag coefficient .(Drag/qS) bc, increment OCpitchingaoment coefficient resulting from the L/D 1ift-drag ratio . addition of the fuselage . U angle ohttack of whg chord plane.with.wind, degre

11、es 4. free-stream dynamic pressure, Gr.square *opt (g) *. Y L Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MACA RM L51J04 -.I f 3 c Reynolds number ( pVc /w) free-stream Mach number viscosity of air, slugs per foot-second. density of air, slugs pe

12、r cubic foot free-stream velocity, feet per second wing area, square feet c. mean aerodynamic local wing chord chord psrallel to parallel to plane plane of symmetry, feet of symmetry, feet wing span, feet - epanwise coordinate, feet local airfoil section maximum thickness, feet wing-fuselage inciden

13、ce, angle between wing chord plane and longitudinal axis of fuselage, degrees rate of change of pitching-moment. coefficient with lift coefficient MODEL The model tested in this investigation had 45O sweepback of the quarter-chord line, an aspect ratio of 8.02, and a taper ratio of 0.45 (see table I

14、). The wing was constructed of a steel core embedded in an alloy of bismuth and th tuthe plan form indicated in figure 1 and contoured to NACA 631A012 airfoil sections parallel to the plane of symmetry. The wing tips were 2.5 percent of the wing spn and were rounded to a parabolic curve plan form an

15、d cross section. The wing had no geometric twist or dihedral. Measurements were made of the torsional deflection due to aerodynamic loading at a Reynolds number of 4.0 x 10 6 (a free-stream dynamic pressure of approximately 120 pounds per square foot). The results indicated a nearly linear variation

16、 in twist with increasing angle of attack to a maximum value of approximately 0.2 wash- out from the root to the tip at maximum lift (CL = 1.0). ,. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 - NACA RM L51J04 4 The dimensions and locations of t

17、he various high-lift and stall- control devices are shown in figure 2. The split-type trailing-edge flaps (fig. 2(a) ) were constructed of sheet steel with a chord equal to i 20 percent of the local wing chord in the undeflecw position and were deflected 50 from the lower surface of the wing paralle

18、l to the air stream (600 measured in a -plane perpendicular to the flap htnge line) Mounting brackets were constTuCted to simulate hinge-line locations of the trailing-edge flaps at 80 and 100 percent of the wing chord with spans of 35, 50, and 60 gercent f the wing span with the inboard end of the

19、flap located at .%he wing root-. The inboard 10 percent of the trailing- edge flaps was removed to permit installation of the fuselage. For - convenience in referring to the trailing-edge flaps, the flap pivoted about the 80-percent-chord line will be referred to as the split flap, and the flap pivo

20、ted about-the trailing edge will be referred to as the extended Split flap. . . The principal dimensions of the round-nose extensible-type leadlng- edge flaps and the span and spanwise locatipn are shown in figure 2(b). The flaps were constructed.of a wooden block having a sheet steel no6e rolled to

21、 approximately 8 3/8-inch diameter. When resolved parallel to the plane of symmetry, the leading edge flap dimensions presented in figure 2(b) resulted in a flap deflection of 30 with respect to the wing- chord plane and a constant chord of 2.75 inches. This chord is equal to . 16 percent of the loc

22、al wing chord at O.kb/2 and 27 percent at 0.975b/2. TIE upper-surface fences, were constructwi of 1/16-inh sheet steel. The 3 types of chordwise fences tested on the model are shown in fig- ure 2(c) . The hose fence“ extended aft 5 percent of the wing chord from the leading edge on the upper and low

23、er wing surfaces. The “chord fence “ extended along the upper surface from 0.05 to the trailing edge of the wing. The “cmpletfence“ is a combination of the first two fences. An additional segment of chord fence extending from 0.35 to the trailing edge was tested at 0.8gb/2. Uhless specifically state

24、d other- wise, the fences installed on the various configura“tions throughout the tests had a height (measured from the surface of the wing) equal to 0.6% at 0.575b/2 and 0.80b/2 and 0.7%- at 0:89b/2. The fknces will be referred to by type and spnwlse location. The fuselage was a body of revolution

25、having a fineness rstio of 10 with the nose and afterbody shapes as indicated intable I and shown in figure 1. Provisions were made to test the wing at wing-Puselage incidences of Oo and bo. Leading-edge roughness was obtained by applylng no. 60 carborundum granules to a thin coating of shellac on t

26、he 1ead.ing 0.08 chord of the wing measured along the upper and lower surfaces. For the flap-deflected combination, the roughness extended along the wing leading edge inbmrd of the leading-edge flaps. z Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,

27、-NACA RM L51J04 I_ 5 The wing mounted for testing on the two-support system of the Langley 19-foot pressure tunnel is sham in figure 3. h The tests were conducted in the Langley 19-foot pressure tunnel with the air in the tunnel compressed to approximately 33 pounds per square inch, absolute. Lift a

28、nd drag forces and pitching moments were measured through an angle-of attack range from -3.5O to 31, and unless stated otherwise, the tests were conducted at a Reynolds number of 4.0 x lo6. Scale-effect tests were made at Reynolds numbers from 1.7 x lo6 to 4.8 x lo6 for the plain wing and plain wing

29、 with fences and from 1.5 x 10 to 4.0 x 10 for one wing-flap combination. The Mch 6 6 numbers corresponding to the various Reynolds numbers are as follow: The lift, drag, and pitching-moment data have been corrected for support tare and .interference effects. As noted in reference 5, there was a spn

30、wise variation in the tunnel air-stream angle in the region occupied by the model. Inasmuch as only total wing-force coefficients are considered in this pper, an average air-stream misalinement correc- tion has been applied to the angle of attack and drag coefficients. The angle of attack and drag h

31、ave been corrected for jet-boundary effects and the pitching moment corrected for tunnel-induced distortion of the loading by the method of reference 7. These corrections are as follow and were all added to the data: mm = 0.0035 Provided by IHSNot for ResaleNo reproduction or networking permitted wi

32、thout license from IHS-,-,-6 The spanwise variation of the jet-boundary-induced angle was of the same magnitude and in,a direction opposite to the 0.2 twiat due to aero- dynamic loading. 0 RESULTS Fresentaticm of Results The longitudinal aerodynaniic charactmistics for the various con-. figurations

33、tested are presented in figwes-4 to 31. Table I1 presents a smary of the maximum lift and pitching-moment characteristics. Basic Wing Plain wing.- The lift curves shar 8 decreasing slope and the pitching- moment curves shar a positivk increase in dCm/dCr, with an increase in angle of attack beginnin

34、g at a low angle of.attack (fig. 4) . At the low Reynolds number and above a lift coefficient of 0.7, there was a rapid increase in lift-curve slope. which became much less pronounced and occurred at a higher lift coefficient as the Reynolds number was increased. This increase in lif%-cume slope was

35、 accompanied by a stable break in the pitching-moment. .curve which also became less severe at“the higher Reynolds numbers. In the region near maximum lift-the lift curves tended to level off, snd the pitching moments were highly unstable. In general, an increase in Reynolds number in the range inve

36、stigated caused the lift curve to be more nearly linear and reduced the variation of dCm/dCL throughout the lift range. “ “he pressure-distribution surveys presented .in reference 5 indicate- that the decreased lift-c.um slope and positive increase fn dCm/dCL with increasing angle of attack result f

37、rom a loss in lifedue td trail%=- edge separation which began at .low angles of attack .over .the tip, sections of the wing; The increased lift and stable m however, Considerable undesirable changes in stability remained through- out-the lift range. At maximum lift, the pitching-moment curves broke

38、in a stable direction. A comparison of the lift characteristics of the plain wing (fig. 4(a) and the wing with leading-edge flaps deflected (fig. 13) shows that the leading-edge flaps reaulted in a higher lift-cur-ve slope through the moderate and upper lift- coefficient range and produced an increm

39、ent of maximum lift coefficient-of approximately 0.2. A change in leading- edge flap span from O.3Tb/2 to 0.375b/2 resulted in only small changes ir-maximum lift. As can be seen from the curves of figure 13, there was t i Provided by IHSNot for ResaleNo reproduction or networking permitted without l

40、icense from IHS-,-,-Y an initial break in the lift curve at a lift coefficient of approximately 1.1 and a small increase in lift with further increase in angle of attack. The change in lift-curve slope at a lift coefficient of approxi- mtely 1.1 is associated with the unstable break in pitching mome

41、nt obtained for the shorter spans of leading-edge flap at the same lift coefficient and results from 8 loss in lift over the wing inboard of the inboard end of the leading-edge flap, as indicated by wool tuft studies and pressure distribution measurements (data not published). The longer spans of fl

42、ap move the initial stall inboard and reduce the loss in lift behind the center of moments, thereby reducing the instability. The effectiveness of the leading-edge flaps in providing stability appears to result frum their ability to maintain lift over the outboard portion of the wing. By the selecti

43、on of the proper flap span, the stalled and unstalled areas may be balanced to provide the desired stability. Combinations of leading-edge and trailingee flaps.- When the leading-edge and trailing-edge flaps were tested in combination, the model exhibited varying degrees of instability which were de

44、pendent on the span of both the leadingedge and trailingddge flaps (figs. 14 to 16). In general, the longer spans of leadingedge flaps and the moment characteristics near maximum lift. The chordwise location of the trailing-edge flaps had little effect on the longitudinal stability characteristics w

45、ith the leading-edge flaps installed. . shorter spans of trailing-edge flaps provided the most favorable pitching- An examination of figures 14 to 16 indicates that, for many combina- tions (particularly the configurations having the longer s-ns of leading- edge flaps), the initial leveling off or b

46、reak in the lift curve is followed by a small increase in lift at higher angles of attack. For purposes of comparison, the maximum value of lift coefficient obtained will be used in discussing the maximum lift characteristics of the wing with flaps deflected, although it is realized that this may no

47、t be a usable value of lift coefficient from the standpoint of longitudinal stability. In most cases, mmirnum lift occurs after the pitching- moment curves have broken in a stable or unstable direction. The maximum values of lift coefficient obtained are presented in figure 17 for the various combin

48、ations of flaps. With the leading-edge flaps deflected, the 0.6b/2 split flap produced only an increment of maximum lift coefficient of approximately 0.10 to 0.15. Several of the shorter spsns of split flap actually produced a decrease in maximum lift over that obtained with the leading-edge flaps a

49、lone. The extended the 0.6b/2 trailing-edge flap resulted in m increment of miximum lift coefficient of approximately 0.25 with the leading-edge flaps installed. . split flaps improved the maximum lift characteristics appreciably, and L Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,