NASA-TN-D-3797-1967 Large-scale wind tunnel tests of a subsonic transport with AFT engine nacelles and high tail《带有后体发动机短舱和高尾翼的亚音速运输机的大型风洞试验》.pdf

《NASA-TN-D-3797-1967 Large-scale wind tunnel tests of a subsonic transport with AFT engine nacelles and high tail《带有后体发动机短舱和高尾翼的亚音速运输机的大型风洞试验》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-3797-1967 Large-scale wind tunnel tests of a subsonic transport with AFT engine nacelles and high tail《带有后体发动机短舱和高尾翼的亚音速运输机的大型风洞试验》.pdf（49页珍藏版）》请在麦多课文档分享上搜索。

1、NASA TECHNICAL NOTE NASA TN D-37 97 - - h OI h M LOAN COPY: RETURN AFLJC (WLIL-2) KIRTLNdD AF-B, N MEX LARGE-SCALE WIND-TUNNEL TESTS OF A SUBSONIC TRANSPORT WITH AFT ENGINE NACELLES AND HIGH TAIL ,* by Kiyoslhi Aoyagi and WiZZiam H, ToZlhurst, Jr. Ames Research Center M offett Field, CuZzF . I NATIO

2、NAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. JANUARY 1967 1i i I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM IIlllll111111111101l1 lllllIll11 11111IllIll 0330504 NASA TN D-3797 LARGE-SCALE WIND-TUNNEL TESTS OF

3、A SUBSONIC TRANSPORT WITH AFT ENGINE NACELLES AND HIGH TAIL By Kiyoshi Aoyagi and William H. Tolhurst, Jr. Ames Research Center Moffett Field, Calif. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 221

4、51 - Price $2.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LARGE-SCALF: WIND-TUNNEL TESTS OF A SUBSONIC TRANSPORT WITH AFT ENGINE NACELLES AND HIGH TAIL By Kiyoshi Aoyagi and William H. Tolhurst, Jr. Ames Research Center The static longitudinal

5、 stability and control effectiveness at angles of attack above those for wing stall was investigated for a iarge-scale subsonic transport model with a 35 swept wing of aspect ratio 5.38. The model was tested with the nacelles in several locations and with wing leading- and trailing-edge high lift de

6、vices. Pitching moment and longitudinal control characteristics of the model and three-component longitudinal data are pre sented. Downwash angles and dynamic pressures in the horizontal tail plane location and nacelle inlet pressures are also presented. The static longitudinal stability and control

7、 effectiveness of the model was reduced substantially at angles of attack above that for wing stall. The nacelles did not decrease the longitudinal stability and control effectiveness of the model, compared to that without nacelles, for angles of attack up to 30. At larger angles the presence of nac

8、elles did reduce the stability of the model. Small changes in the nacelle locations or deflections of the trailing-edge flaps did not significantly improve the longitudinal stability or control effectiveness of the model. However, the use of leading-edge slats with or without trailing-edge flaps did

9、 improve both of these characteristics at angles of attack above the wing stalling angle. Sideslipping the model seemed to improve the pitching-moment characteristics. INTRODUCTION Flight tests of subsonic transports that have jet engines mounted at the rear of the fuselage and have the horizontal t

10、ail on top of the vertical tail indicate that airplanes having this general arrangement may inadvertently pitch up to angles of attack above that for wing stall. Studies (refs. 1 through 3) have shown that the high tail location usually increases the ten dency for the pitching-moment variation with

11、angle of attack to be unstable at and above the wing stalling angle. In order to study other factors affect ing the post -stall longitudinal stability and control of a configuration with a high tail and aft-mounted nacelles, NASA has undertaken a number of wind-tunnel investigations. References 4, 5

12、, and 6 present the effects of configuration variables on the longitudinal stability and control of a small scale model at Reynolds numbers of 0.810to 3.010. The present investigation was conducted to determine the post-stall static longitudinal stability and control characteristics of a large-scale

13、 research model at high Reynolds numbers and to explore methods of improving these characteristics. Results were obtained with the nacelles in several Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-locations and with wing leading- and trailing-edge

14、devices. The downwash and flow field at the horizontal tail plane and pressures at the nacelle inlet were measured at angles of attack above the stall angle. All of the data except those for variable Reynolds number were obtained at a Reynolds number of 6.5x106, based on a mean aerodynamic chord of

15、7.96 feet and a dynamic pressure of 20 pounds per square foot. NOTATION b wing span, ft C wing chord measured parallel to the plane of symmetry, ft c2 rolling-moment coefficient about stability axis, rolling moment qo,Sb dragCD drag coefficient, -lift CL lift coefficient, -QWS Cm pitching-moment coe

16、fficient about 0.44C,pitching moment q SF co Cn yawing-moment coefficient about stability axis, yawing moment sooSb side-force coefficient, side force gs00 it horizontal-tail incidence angle, deg pT total pressure, in. Hg q dynamic pressure, lb/sq ft qt dynamic pressure at the horizontal-tail plane,

17、 lb/sq ft vwcR Reynolds number, 2, S wing area, sq ft VW free-stream air velocity, ft/sec Y spanwise distance perpendicular to the plane of symmetry, ft 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a angle of attack of wing, deg 6f trailing-edge

18、 flap deflections measured normal to the hinge line, deg 6s slat deflection measured perpendicular to the leading edge, deg downwash angle at the tail location with respect to free stream, deg rl wing semispan station, -Y b/2 A,/ 4 sweep angle of quarter chord line, deg V free-stream kinematic visco

19、sity, ft2/sec Subscripts t tail U uncorrected 03 free stream MODEL AND APPAW1TUS In figure 1 the model is shown mounted in the Ames 40-by 80-foot wind-tunnel. Pertinent dimensions of the basic model configuration are given in figure 2( a). Wing 0The wing had a quarter chord sweep of 35 , an aspect r

20、atio of 5.38, a taper ratio of 0.23, and a dihedral of 3. The airfoil section was an NACA 65-412section from the tip to 0.37 of the wing semispan. Inboard of 0.37 semispan, a chord extension added at the trailing edge changed the trailing-edge sweep from 23 to Oo (see fig. 2(b). High Lift Devices Co

21、nventional leading-edge slats and trailing-edge flaps were provided for the wing as shown in figures 2(c) and (d), respectively. The slats extended either the full span or half the span of the wing with a 6s of 20. Single slotted flaps extended from 0.11 to 0.53 of the wing semispan and were deflect

22、ed 40. 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Fuselage The fuselage had a constant 4-foot diameter except at the nose and tail. Both of these sections had elliptical outlines with circular cross sections that decreased from 4 feet to a sma

23、ller diameter. Nacelles Nacelle details and locations are shown in figure 2(e). The nacelles could be located at the rear of the fuselage at three longitudinal and two spanwise positions. The longitudinal positions were varied by moving the nacelle and pylon together fore and aft on rails; the extre

24、me positions were physically limited because of structural interference. The spanwise positions were varied by moving the nacelle support strut normal to the model center- line. Tail The geometry of the horizontal and vertical tails is described in figure 2(a). Pitch control was provided by an all-m

25、ovable tail that was variable from -20 to +20 and by a 25-percent-chord elevator that was also variable -20 to +20. The vertical tail was fixed. Instrumentation Forces and moments were measured on the wind-tunnel six-component balance. Dynamic pressure and flow direction were measured at the pivot a

26、xis of horizontal-tail plane by directional pitot-static probes at four spanwise stations, Nacelle inlet flow distortions were measured with four total pressure rakes located close to the inlet lip and spaced 90 apart around the circum ference. Each rake contained four pressure probes located at the

27、 center of equal areas. TESTING AND PROCEDUFE Forces and moments were measured for the model through an angle-of-attack range from Oo to bo0. Pitching-momentdata were computed about a moment center located at O.44C. This center was chosen to represent the static margin of a typical high-tailed, rear

28、-engined transport with an aft center of gravity location. All tests except those to show Reynolds number effect were conducted at a Reynolds number of 6.5x106, based on a mean aerodynamic chord of 7.96 feet and a dynamic pressure of 20 pounds per square foot. Provided by IHSNot for ResaleNo reprodu

29、ction or networking permitted without license from IHS-,-,-Tests were conducted with the basic configuration at several tail inci dences and elevator settings. Similarly, tests were conducted with several nacelle positions (see fig. 2 e and with the nacelles removed. Tail inci dence ranged from -10

30、to +216,)ind the elevators were set at Oo and 20. Sideslip angle ranged from 0 to -9O, and Reynolds number was varied from 3.810to 8.Cn106, based on a mean aerodynamic chord of 7.96 feet and dynamic pressure range of 5 to 30 pounds per square foot. Maximum angle of attack at a Reynolds number of 8.O

31、X1O6 was 20 because of a model load limitation. COFEUXTIONS All data were corrected for strut tares and wind-tunnel-wall effects. Drag and pitching-moment tares due to the support struts were based on data obtained with the struts alone. Corrections for wind-tunnel-wall effect were as follows: XD =

32、0.0092 CL AC, = 0.0144 CL (tail on tests only) RFSULTS Table I is an index to the configurations.andvariables tested and the figures in which the results are presented. Figures 3 through 10 show the variation of pitching-moment coefficient with angle of attack; figures 11 through 18 show three-compo

33、nent longitudinal force and moment data. Fig ures 1-9 through 23 present lateral characteristics and show downwash and dynamic pressure data at the tail location and total pressure distortion at the nacelle inlet. DISCUSSION Longitudinal Characteristics Basic configuration characteristics.- Figure 3

34、 shows that for less than loo tail incidence the model was statically stable to an angle of attack of about 16O. Tuft studies and the reduction in lift coefficient (fig. 11) indicated that the wing stalled at about 160. From 160 to 20 angle of attack the model was unstable with the tail off or on. T

35、he progression of air-flow separation from the tip of the wing inboard and the increase of downwash angle with angle of attack (fig. 21) caused the instability. From an angle of attack of about 20 to 24O the model was statically stable for tail incidences above 15 because the large downwash angle un

36、stalled the tail. 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-At angles of attack above 24 static margin was largely a function of the tail moment contribution since the tail-off configuration was neutrally stable (fig. 3(a) . This contribution

37、 decreased substantially above 24 angle of attack due to a reduction of dynamic pressure at the tail (shown in fig. 21). Tail effectiveness and elevator control power were reduced substantially at angles of attack above that for wing stall (see fig. 3). For example, at an angle of attack of 30, the

38、effectiveness was approximately half that at an angle of attack of 16O (when the tail was unstalled) because of the large reduction in dynamic pressure at the tail. For tail incidences of loo or more, the tail stalled at angles of attack below that for wing stall. Consequently, control effectiveness

39、 was less than that at lower tail inci dences. At angles of attack above that for wing stall the large downwash angle unstalled the tail; thus, longitudinal control power increased at angles of attack above 24O for tail incidences of 150 and 20. This could be used as a possible means of recovery fro

40、m extremely high angles of attack. The results presented are for the center of gravity at 0.44Fwhich gave a static margin of 27.5 percent at angles of attack below that for wing stall. An aircraft with a less stable static margin could be unstable at all angles of attack above that for wing stall. E

41、ffect of nacelle and location.- The effect of nacelles on the variation of pitching moment and longitudinal control characteristics of the model with angle of attack (fig. 4) was small up to an angle of attack of approximately 30. At larger angles the nacelles reduced the stability; limited downwash

42、 data above 30 angle of attack (fig. 21) indicate that the nacelles increased downwash angles at the tail. The variation of nacelle spanwise and longitu dinal positions (figs. 4 and 5), within the range tested, had little effect on the pitching-moment characteristics of the model. Results obtained f

43、or a larger range of longitudinal positions, 0.936 forward and 0.607 aft (ref. 6), show an effect of nacelle location. Effect of high-lift devices.- The effect of high-lift devices on the pitching-moment and longitudinal control characteristics was explored. Con figurations tested included leading-e

44、dge slats alone and trailing-edge flaps with and without slats. The data in figures 6 and 7 are for the half-span leading-edge slats located on either the inboard or the outboard portion of the wing. In comparing these data, it should be noted that they are for dif ferent nacelle positions. Neither

45、slat location eliminated the instability at wing stall, but the outboard slats did increase the angle of attack for wing stall (see fig. 15). The inboard slats increased the longitudinal control power at angles of attack above the wing stalling angle, but the outboard slats did not. The effect of tr

46、ailing-edge flaps, both with and without slats, on the pitching -moment variation with angle of attack is shown in figure 8. With-the flaps deflected and the slats off, this variation was similar to that of the basic configuration. When the flaps were combined with the full-span leading-edge slats,

47、the pitching-moment variation was neutrally stable from an angle of attack of 13O to 20. At larger angles of attack the pitching-moment variation with the half-span and the full-span slats was similar. 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-

48、,-Pitching and Rolling Moments With Sideslip Angle The effect of sideslip angle on the variation of pitching and rolling moments with angle of attack is shown in figures 9 and 19. Sideslip angleincreased rolling moments and reduced pitching-momentvariation in the angle of-attack range above that for

49、 wing stall. When longitudinal control is limited, both these characteristics could be advantageous in recovery from deep stall, but further investigation of lateral and directional controls at angles of attack above the wing stalling angle are required. Total Pressure Distortions at the Nacelle Inlet Total pressure distortio