NASA-TN-D-8360-1976 Low-speed wind-tunnel investigation of flight spoilers as trailing-vortex-alleviation devices on a medium-range wide-body tri-jet airplane model《飞行扰流板作为中程宽机身三喷气.pdf

《NASA-TN-D-8360-1976 Low-speed wind-tunnel investigation of flight spoilers as trailing-vortex-alleviation devices on a medium-range wide-body tri-jet airplane model《飞行扰流板作为中程宽机身三喷气.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-8360-1976 Low-speed wind-tunnel investigation of flight spoilers as trailing-vortex-alleviation devices on a medium-range wide-body tri-jet airplane model《飞行扰流板作为中程宽机身三喷气.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-8360 LOW-SPEED WIND-TUNNEL INVESTIGATION OF FLIGHT SPOILERS AS TRAILING-VORTEX-ALLEVIATION DEVICES ON A MEDIUM-RANGE WIDE-BODY TRI-JET AIRPLANE MODEL mad Geoffrey M. Williunzs Luizgtey Resed sch Center Ndiiptoi,Vu. 23665 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION W

2、ASHINGTON, 0. C. NOVEMBER 1976 ,d Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM IllllllIll1111111 11111 11111llllllllll1111Ill 1. Report No. 2. Government Accession No. NASA TN D-8160 I 4. Title and Subtitle LOW-SPEED WIND-TUN

3、NEL INVESTIGATION OF FLIGHT ISPOILERS AS TRAILING-VORTEX-ALLEVIATION DEVICES ON A MEDIUM-RANGE WIDE-BODY TRI-JET AIRPLANE MODEL 7. Author(s) Delwin R. Croom, Raymond D. Vogler, and Geoffrey M. Williams 9. Performing Organization Name and Address NASA Langley Research Center Hampton, VA 23665 12. Spo

4、nsoring Agency Name and Address National Aeronautics and Space Administration Washington, DC 20546 -3. Recipients Catalog No. 5. Report Date November 1976 6. Performing OrganizationCode 8. Performing Organization Report No. L-11103 10. Work Unit No. 514-52-01-03 11. Contract or Grant No. 13. Type of

5、 Repon and Period Covered Technical Note 14. Sponsoring Agency Code 7. Key Words (Suggested by Author(s) 18. Distribution Statement Vortex alleviation Unclassified - Unlimited Trailing-vortex hazard Subject Category 01 9. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No

6、. of Pages 22. Price Unclassified Unclassified I 46 I $3.75 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-,-.- LOW-SPEED WIND-TUNNEL INVESTIGATION OF FLIGHT SPOILERS AS TRAILING-VORTEX-ALLEVIATION DEVICES ON A MEDIUM-RANGE WIDE-BODY TRI-JET AIRPLAN

7、E MODEL Delwin R. Croom, Raymond D. Vogler, and Geoffrey M. Williams* Langley Research Center SUMMARY An investigation was made in the Langley V/STOL tunnel to determine, by the trailing wing sensor technique, the effectiveness of various segments .of the existing flight spoilers on a medium-range w

8、ide-body tri-jet transport air plane model when they were deflected as trailing-vortex-alleviation devices. The four combinations of flight-spoiler segments investigated were effective in reducing the induced rolling moment on the trailing wing model by as much as 15 to 60 percent at distances behin

9、d the transport model of from 3.9 to 19.6 trans port wing spans, 19.6 spans being the downstream limit of distances used in this investigation. Essentially all of the reduction in induced rolling moment on the trailing wing model was realized at a spoiler deflection of about 45. INTRODUCTION The str

10、ong vortex wakes generated by large transport aircraft are a poten tial hazard to smaller aircraft. The National Aeronautics and Space Administra tion is involved in a program of model tests, flight tests, and theoretical studies to determine the feasibility of reducing this hazard by aerodynamic me

11、ans. Results of recent investigations have indicated that the trailing vortex behind an unswept-wing model (ref. 1) or a swept-wing transport model (ref. 2) can be attenuated by a forward-mounted spoiler. It was also determined by model tests (ref. 3) and verified in full-scale flight tests (ref. 4)

12、 that there are several combinations of the existing flight-spoiler segments on the jumbo-jet transport aircraft that are effective as trailing-vortex-alleviation devices. The approach used in references 1, 2, and 3 to evaluate the effective ness of vortex-alleviation devices was to simulate an airp

13、lane flying in the trailing vortex of another larger airplane and to make direct measurments of 4 rolling moments induced on the trailing model by the vortex generated by the forward model. The technique used in the full-scale flight test (ref. 4) was to penetrate the trailing vortex wake behind a B

14、oeing 747 aircraft with a Cessna T-37 aircraft and to evaluate the roll response and roll attitude of the Cessna T-37 airplane as an index to the severity of the trailing-vortex encounter. *Lockheed-California Company, Burbank, California. Provided by IHSNot for ResaleNo reproduction or networking p

15、ermitted without license from IHS-,-,-The purpose of the present investigation is to determine the trailing vortex-alleviation effectiveness of various segments of the existing flight spoiler of a medium-range wide-body tri-jet transport aircraft model. The direct-measurement technique described in

16、references 1, 2, and 3 was used with the trailing wing model from 3.9 to 19.6 transport wing spans behind the trans port aircraft model. (For the full-scale transport airplane, this would repre sent a range of downstream distance from 0.1 to 0.5 nautical mile.) SYMBOLS All data are referenced to the

17、 wind axes. The pitching-moment coeffi cients are referenced to the quarter-chord of the wing mean aerodynamic chord. b CD CL I ,TW m C -C it I 9 S X,Y ,Z x,y,z Ay ,Az U 2 wing span, m drag coefficient, Drag qsW lift coefficient, Lift qsW trailing wing rolling-moment coefficient, Trailing wing rolli

18、np moment TWTW pitching-moment coefficient, pitching moment 9SWFW wing chord, m wing mean aerodynamic chord, m horizontal-tail incidence, referred to fuselage reference line (posi tive direction trailing edge down), deg longitudinal distance in tunnel diffuser, m dynamic pressure, Pa wing area, m2 s

19、ystem of axes originating at left wing tip of transport aircraft model (see fig. 1) longitudinal, lateral, and vertical dimensions measured from trailing edge of left wing tip of transport aircraft model, m incremental dimensions along Y- and Z-axes, m angle of attack of fuselage reference line, deg

20、 (wing root incidence is 3O relative to fuselage reference line) Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 deflection, deg cb local streamline angle in tunnel diffuser relative to tunnel center line, deg Subscripts: flap transport aircraft mo

21、del flap max maximum t slat transport aircraft model slat spoiler transport aircraft model spoiler TW trailing wing model vane transport aircraft model vane W transport aircraft model MODEL AND APPARATUS A three-view sketch and principal geometric characteristics of the 0.05 scale model of a medium-

22、range wide-body tri-jet transport aircraft (Lockheed L-1011) are shown in figure 1. Sketches of the landing and approach flap con figurations are shown in figures 2 and 3, respectively. Figure 4 is a photo graph of the transport aircraft model sting mounted in the Langley V/STOL tun nel. Figure 5 is

23、 a sketch showing the location of the flight spoilers on the transport aircraft model. Photographs of the four combinations of flight-spoiler segments investigated are presented in figure 6. The test section of the Langley V/STOL tunnel has a height of 4.42 m, a width of 6.63 m, and a length of 14.2

24、4 m. The transport aircraft model was sting supported on a six-component strain-gage balance system which measured the forces and moments. The angle of attack was determined from an accelerom eter mounted in the fuselage, A photograph and dimensions of the unswept trailing wing model installed on a

25、traverse mechanism are presented in figure 7. The trailing model has a span and aspect ratio typical of small-size transport aircraft. The trailing model was mounted on a single-component strain-gage roll balance, which was attached to the traverse mechanism capable of moving both laterally and vert

26、i * tally. (See fig. 7.) The lateral and vertical positions of the trailing model were measured by outputs from digital encoders. This entire traverse mechanism could be mounted to the tunnel floor at various tunnel longitudinal positipns downstream of the transport aircraft model. 3 Provided by IHS

27、Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-TESTS AND CORRECTIONS Transport Aircraft Model All tests were made at a free-stream dynamic pressure (in the tunnel test section) of 430.9 Pa, which corresponds to a velocity of 27.4 m/sec. The Reyn olds number for th

28、ese tests was approximately 6.8 x IO5 based on the wing mean aerodynamic chord. No transition grit was applied to the transport aircraft model. The basic longitudinal aerodynamic characteristics were obtained through an angle-of-attack range of approximately -4O to 22. All tests were made with leadi

29、ng-edge devices and landing gear extended. 1Blockage corrections were applied to the data by the method of refer- ence 5. Jet-boundary corrections to the angle of attack and the drag were applied in accordance with reference 6. Trailing Wing Model JThe trailing wing model and its associated roll-bal

30、ance system were used iI as a sensor to measure the induced rolling moment caused by the vortex flow 1downstream of the transport aircraft model. No transition grit was applied to the trailing model. The trailing model was positioned at a given distance down stream of the transport aircraft model on

31、 the traverse mechanism which was posi tioned laterally and vertically so that the trailing vortex was near the center of the mechanism. The trailing vortex was probed with the trailing model. A large number of trailing wing rolling-moment data points (usually from 50 to 100) were obtained from the

32、lateral traverses at several vertical locations to ensure good definition of the vortex wake, In addition, certain test condi tions were repeated at selected intervals during the test period and the data were found to be repeatable. Trailing wing rolling-moment measurements were made at downstream s

33、cale distances from about 3.9 to 19.6 transport wing spans behind the transport air craft model, All trailing wing rolling-moment data at distances downstream greater than about 3.9 spans were obtained with the trailing model positioned in the diffuser section of the V/STOL tunnel, These data were r

34、educed to coef ficient form based on the dynamic pressure at the trailing wing location. For these tests, the dynamic pressures at the 3.92, 9.81,and 19.61 span locations were 430.9, 287.0, and 88.38 Pa, respectively. The trailing wing location rela tive to the wing tip of the transport aircraft mod

35、el has been corrected to account for the progressively larger tunnel cross-sectional area in the dif-t fuser section. The corrections to the trailing wing location in the diffuser were made by assuming that the local streamline angles in the tunnel diffuser section are equal to the ratio of the dist

36、ance from the tunnel center line to # the local tunnel half-width or tunnel half-height multiplied by the diffuser half-angle. Corrections to the trailing model locations are as follows: Ay correction or Az correction = I tan $ where By correction and Az correction are, respectively, the corrections

37、 to the measured lateral and verti cal locations of the trailing model relative to the tip of the transport air craft model, I is the longitudinal distance in the tunnel diffuser, and $ 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-is the local s

38、treamline angle in the tunnel diffuser relative to the tunnel cen ter line. RESULTS AND DISCUSSION Transport Aircraft Model The longitudinal aerodynamic characteristics of the transport aircraft model in the landing flap configuration with spoilers retracted (see fig. 2) are presented in figure 8. T

39、hese data were obtained over a range of horizontal.-tail incidence sufficient to trim the model through the range of lift coeffi cient. These data indicate that the transport aircraft model was statically stable up to the stall. The static margin, dCm/dCL, for the model was about -0.24. The longitud

40、inal aerodynamic characteristics of the transport aircraft model with flight-spoiler segments 1 and 2, 2 and 3, 3 and 4, and 1 and 4 deflected symmetrically through a spoiler deflection range of from 0 to 60 are presented in figures 9, 10, 11, and 12, respectively, for the landing flap configuration

41、 and in figures 13, 14, 15, and 16, respectively, for the approach flap configuration. These data were obtained with it = 0. For both of these configurations, there was essentially a linear increase in drag with spoiler deflection. For the landing flap configuration, about 50 percent of the lift los

42、s at a given angle of attack occurred at a spoiler deflection of about 15. For the approach flap configuration, about 50 percent of the lift loss at a given angle of attack occurred at a spoiler deflection of about 30. The vari ation of pitching-moment coefficient with angle of attack was generally

43、more linear when the spoilers were deflected than when they were retracted. The longitudinal aerodynamic characteristics of the transport aircraft model with the four combinations of flight spoilers on each wing deflected sym metrically 45 are presented in figures 17 and 18 for the landing flap conf

44、igu ration and the approach flap configuration, respectively. These data indicate that for the landing flap Configuration (fig. 17) a nominal lift coefficient of 1.2 can be maintained with an increase in angle of attack of no more than 3 for any of the spoiler configurations tested. It can also be s

45、een in figure 17 that the maximum increase in drag coefficient at CL = 1.2, for any of the spoiler configurations, was about 0.05 and that the maximum lift coefficient was reduced by about 0.14. For the approach flap configuration (fig. 18) a nom inal lift coefficient of 1.2 can be maintained with n

46、o more than a 2 increase in angle of attack for any of the spoiler combinations investigated. The drag penalty due to spoilers was no more than 0.04 and the reduction in maximum lift coefficient was no more than 0.08 for any of the spoiler configurations investi 1 gated. It can also be seen in figur

47、es 17 and 18 that, for both .flap configura tions, the variation of pitching-moment coefficient with angle of attack was generally more linear when the spoilers were deflected than when they were retracted. 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from I

48、HS-,-,-Trailing Wing Model The maximum rolling-moment coefficient measured by the trailing model and the position of this model relative to the left wing tip of the transport air craft model are presented as a function of flight-spoiler deflection for the various combinations of flight spoilers inve

49、stigated in figures 19 to 22. The data presented in figure 19 are for the approach flap configuration with the trailing model positioned 9.8 transport wing spans behind the transport air craft model. The data presented in figures 20, 21, and 22 are for the landing flap configuration with the trailing model positioned 3.9, 9.8, and 19.6 wing spans, respec