1、NASA M h cr) ? n z c TECHNICAL NOTE LOW-SPEED WIND-TUNNEL INVESTIGATION OF FLIGHT SPOILERS AS TRAILING- VORTEX-ALLEVIATION DEVICES ON AN EXTENDED-RANGE WIDE-BODY TRI-JET AIRPLANE MODEL Delwin R. Croom, Raymond D. Vogler, und John A. Thelander Langley Research Center Hdmpton, Vu. 23665 0-8373 -v NATI
2、ONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. DECEMBER 1976 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB. NM _ . 2. Government Accession No. 1 1. Report No. . . . NASA TN D-8373 4. Title and Subtitle LOW-SPEED WIND-
3、TUNNEL INVESTIGATION OF FLIGHT SPOILERS AS TRAILING-VORTEX-ALLEVIATION DEVICES ON AN EXTENDKD-RANGE WIDE-BODY TRI-JET AIRPLANE MODEL 7. Author(s) Delwin A. Croom, Raymond D. Vogler, and John A. Thelander NASA Langley Research Center Hampton, VA 23665 9. Performing Organization Name and Address 2. Sp
4、onsoring Agency Name and Address National Aeronautics and Space Administration Washington, DC 20546 5. Report Date December 1976 6. Performing Organization Code 8. Performing Organization Report No. L-I 1104 10. Work Unit No. 514-52-01-03 11. Contract or Grant No. 13. T pe of Report and Period Cover
5、ed itechnical Note 14. Sponsoring Agency Code 5. Supplementary Notes Delwin R. Croom and Raymond D. Vogler: Langley Research Center. John A. Thelander: Douglas Aircraft Company, McDonnell Douglas Corporation, - Long Beach, California. 6. Abstract An investigation was made in the Langley V/STOL tunne
6、l to determine, by the trailing wing sensor technique, the effectiveness of various segments of the exist- ing flight spoilers on an extended-range wide-body tri-jet transport airplane model when they were deflected as trailing-vortex-alleviation devices. On the transport model with the approach fla
7、p configuration, 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 25 to 45 percent at downstream distances behind the transport model of 9.2 and 18.4 transport wing spans. On the transport air
8、- plane model with the landing flap configuration, 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 35 to 60 percent at distances behind the transport model of from 3.7 to 18.4 transport wing
9、 spans, 18.4 spans being the downstream limit of distances used in this investigation. -_ 7. Keywords (Suggested by Author(s) Vortex alleviation Trailing-vortex hazard . 18. Distribution Statement Unclassified - Unlimited Subject Category 02 $4.25 20. Security Classif. (of this page) Unclassified 9.
10、 Security Clanif. (of this report) Unclassified _ *For sale by the National Technical information Service, Springfield. Virginia 22161 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LOW-SPEED WIND-TUNMEL INVESTIGATION OF FLIGHT SPOILERS AS TRAILING-
11、VORTEX-ALLEVIATION DEVICES ON AN EXTENDED-RANGE WIDE-BODY TRI-JET AIRPLANE MODEL Delwin R. Croom, Raymond D. Vogler, and John A. Thelander“ 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
12、 various segments of the existing flight spoilers on an extended-range wide-body tri-jet trans- port airplane model when they were deflected as trailing-vortex-alleviation devices. On the transport model with the approach flap configuration, the four combinations of flight-spoiler segments investiga
13、ted were effective in reducine; the induced rolling moment on the trailing wing model by as much as 25 to 45 percent at downstream distances behind the transport model of 9.2 and 18.4 transport wing spans, On the transport airplane model with the landing flap configuration, the fourcombinations of f
14、light-spoiler segments investigated were effective in reducing the induced roliing moment on the trailing wing model by as much as 35 to 60 percent at distances behind the transport model of from 3.7 to 18.4 transport wing spans, 18.4 spans being the dcwnstream limit of distances used in this invest
15、igation. INTRODUCTION The strong vortex wakes generated by large transport airplanes are a potential hazard to smaller aircraft. The National Aeronautics and Space Administration is involved in a program of aodel tests, flight tests, and theoretical studies to determine the feasibility of reducing t
16、his hazard by aerodynamic means. Results of recent investigations have indicated that the trailing vor- tex 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 deter- mined by model tests (ref. 3) and verified in
17、full-scale flight tests (ref. 4) that there are several combinations of the existing flight-spoiler segments on the jumbo-jet airplane that are effective as trailing-vortex- alleviation devices. The approach used in references 1, 2, and 3 to evaluate the effectiveness of vortex-alleviation devices w
18、as to sinulate an airplane flying in the trailing vortex of another larger airplane and to make direct measurements of rolling moments induced on the trailing model by the vortex . . *Douglas Aircraft Company, McDonneil Douglas Corporation, Long Beach, California. Provided by IHSNot for ResaleNo rep
19、roduction or networking permitted without license from IHS-,-,-generated by the forward model. The technique used in the full-scale flight tests was to penetrate the trailing vortex wake behind a Boeing 747 airplane with a Cessna T-37 airplane and to evaluate the roll attitude and roll rate of the C
20、essna T-37 airplane as an index to the severity of the trailing- vortex encounter. The purpose of the present investigation was to determine the trailing- vortex-alleviation effectiveness of various segments of the existing flight spoiler on an extended-range wide-body tri-jet transport airplane mod
21、el. The direct-measurement technique described in references 1, 2, and 3 was used with the trailing wing model from 3.7 to 18.4 transport wing spans behind the transport model. (For the full-scale transport airplane, this would repre- sent a range of downstream distance from 0.1 to 0.5 nautical mile
22、.) SYMBOLS All data are referenced to the wind axes. The pitching-moment coeffi- cients are referenced to the quarter-chord of the wing mean aerodynamic chord. b cD cL 1 ,TW Cm C - C it I 9 S wing span, m drag coefficient, 9% lift coefficient, - Lift qsW trailing wing rolling-moment coefficient, Tra
23、iling wing rolljng mome?nt qSb pitching-moment coefficient, moment qswcw wing chord, m wing mean aerodynamic chord, m horizontal-tail incidence, referred to fuselage reference line (positive direction trailing edge down), deg longitudinal distance in tunnel diffuser, m dynamic pressure, Pa winp area
24、, m 2 X,Y,Z system of axes originating at left wing tip of transport airplane model (see fig. 1) 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-x ,y ,z ay ,Az a 6 6 longitudinal, lateral, and vertical dimensions measured from trail- ing edge of le
25、ft wing tip of transport airplane model, m incremental dimensions along Y- and 2-axes, m angle of attack of fuselage reference line, deg (wing root inci- dence is 3 relative to fuselage reference line) deflection, deg local streamline anple in tunnel diffuser relative to tunnel center line, deg Subs
26、cripts : flap transport airplane model flap max maximum slat transport airplane model slat spoiler transport airplane model spoiler Tw trailing wing model W transport airplane model MODEL AND APPARATUS A three-view sketch and principal geometric characteristics of the 0.047- scale model of an extend
27、ed-range wide-body commercial tri-jet transport air- plane (McDonnell Douglas DC-10-30) are shown in figure 1. Sketches of the landing and approach flap confieurations are shown in figures 2 and 3, respec- tively. Figure 4 is a photograph of the transport model mounted in the Langley V/STOL tunnel.
28、Figure 5 is a sketch showing the location of the flight spoilers on the transport model. Photographs of the four combinations of flight-spoiler segments investigated are presented in figure 6. Spoiler segments 1 and 2 (fig. 6(a) and 3 and 4 (fig. 6(c) were deflected as units by internal electrical m
29、otors. No provision was made in the model to deflect the segments separately; therefore, spoiler segments 2 and 3 (fig. 6(b) and 1 and 4 (fig. 6(d) were simulated with wedges. 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 14.24 m. The transp
30、ort model was strut sup- ported on a six-component strain-gage balance system which measured the forces and moments. The angle of attack was determined from an accelerometer mounted in the fuselage. A photograph and dimensions of the unswept trailing wing model installed on a traverse mechanism are
31、presented in figure 7. The trailing model has a span and aspect ratio typical of small-size transport airplanes. It was 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-mounted on a single-component strain-gage roll balance, which was attached to th
32、e traverse mechanism capable of moving the model both laterally and verti- cally. (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
33、positions downstream of the transport airplane model. TESTS AND CORRECTIONS Transport Airplane 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 Reynolds number for these tests was approximately 6.4
34、 x IO5 based on the wing mean aerodynamic chord. No transition grit was applied to the transport airplane model. The basic longitudinal aerodynamic characteristics were obtained through an angle-of-attack range of approximately -4 to 22. tests were made with the leading-ed.qe devices extended. The l
35、anding gear was retracted for the approach flap configuration and was extended for the landing flap configuration. All Blockage 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
36、6. No corrections were applied to the data for any possible strut interference effects. Trailing Wing Model The trailing wing model and its associated roll-balance system were used as a sensor to measure the induced rolling noment caused by the vortex flow downstream of the transport airplane model.
37、 No transition grit was applied to the trailing model. The trailing model was positioned at a given distance downstream of the transport model on the traverse mechanism which was posi- tioned laterally and vertically so that the trailing vortex was near the center of the mechanism. The trailing vort
38、ex 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 lateral traverses at several vertical locations to insure good definition of the vortex wake. In addition, certain test conditions were repeated at select
39、ed intervals during the test period and the data were found to be repeatable. Trailing wing rolling-moment measurements were made at downstream scale distances from about 3.7 to 18.4 transport wing spans behind the transport airplane model. All trailing wing rolling-moment data at distances down- st
40、ream greater than about 3.7 spans were obtained with the trailing model positioned in the diffuser section of the V/STOL tunnel. These data were reduced to coefficient form based on the dynamic pressure at the trailing wing location. For these tests, the dynamic pressures at the 3.69, 9.19, and 18.3
41、9 span locations were 430.9, 253.1, and 85.5 Pa, respectively. The trail- ing wing location relative to the wing tip of the transport model has been 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-corrected to account for the progressively larger t
42、unnel cross-sectional area in the diffuser 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 distance from the tun- nel center line to the local tunnel half-width
43、or half-height multiplied by the diffuser half-angle. Corrections to the trailing model locations are as follows: Ay correction or Az correction 1 tan I$ where Ay correc- tion and Az correction are, respectively, the corrections to the measured lateral and vertical locations of the trailing model re
44、lative to the left wing tip of the transport airplane model, I is the longitudinal distance in the tunnel diffuser, and 4 is the local streamline angle in the tunnel diffuser relative to the tunnel center line. RESULTS AND DISCUSSION Transport Airplane Model The longitudinal aerodynamic characterist
45、ics of the transport airplane model with the double-slotted flaps and leading-edge slats in the landing configuration (see fig. 2) and in the approach configuration (see fig. 3) are presented in figures 8 and 9, respectively. These data were obtained with the horizontal tail off and over a range of
46、horizontal-tail incidence suffi- cient to trim the model throughout the range of lift coefficient. These data indicate that the transport model wit-h either flap configwation was stati- cally stable up to the stall. with either flap configuration was about -0.24. The static margin, aCm/aCL, for the
47、model The longitudinal aerodynamic characteristics of the transport model with flight-spoiler segments 1 and 2, 2 and 3, 3 and 4, and 1 and 4 deflected sym- metrically through a spoiler deflection range of from 0 to 60 are presented in figures IO, 11, 12, and 13, respectively, for the landing flap c
48、onfigura- tion and in figures 14, 15, 16, and 17, respectively, for the approach flap configuration. For any one of these configurations, there is essentially a linear increase in drag with spoiler deflection. For the landing flap con- figuration, about 50 percent of the lift loss at a given angle o
49、f attack occurred at a spoiler deflection of only 15; whereas 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. For both flap configurations, the variation of pitching-moment coefficient with angle of attack was more linear when the spoilers were deflected than when they were retracted.
