NASA-TN-D-7670-1974 Longitudinal aerodynamic characteristics of an externally blown flap powered-lift model with several propulsive system simulators《带有若干推进系统模拟器的外部吹制襟翼动力提升模型的纵向空气动.pdf

《NASA-TN-D-7670-1974 Longitudinal aerodynamic characteristics of an externally blown flap powered-lift model with several propulsive system simulators《带有若干推进系统模拟器的外部吹制襟翼动力提升模型的纵向空气动.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-7670-1974 Longitudinal aerodynamic characteristics of an externally blown flap powered-lift model with several propulsive system simulators《带有若干推进系统模拟器的外部吹制襟翼动力提升模型的纵向空气动.pdf（143页珍藏版）》请在麦多课文档分享上搜索。

1、AND.33NASA TECHNICAL NOTE - NASA TN D-7670%oU. S. A.I(NASA-TN-D-7670): LONGITUDINAL AERODYNAMIC N74-34461)CHARACTERISTICS OF AN EXTERNALLY BLOWN FLAP POWERED LIFT MODEL WITH SEVERALPROPULSIVE SYSTEM SIMULATORS (NASA) Unclas143 p HC $4.75 CSCL 01C H1/01_ 52421 .LONGITUDINAL AERODYNAMIC CHARACTERISTIC

2、SOF AN EXTERNALLY BLOWN FLAPPOWERED-LIFT MODEL WITH SEVERALPROPULSIVE SYSTEM SIMULATORSby Danny R. HoadLangley DirectorateU.S. Army Air Mobility R however, the take-off and landing configurations required a high-lift tail.17. Key Words (Suggested by Author(s) 18. Distribution StatementExternally blo

3、wn flap Unclassified - UnlimitedPowered lift modelDaisy nozzle engineTriple-slotted flapEngine comparison STAR Category 0119. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price*Unclassified Unclassified 141 $4.75For sale by the National Technical Infor

4、mation Service, Springfield, Virginia 22151Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LONGITUDINAL AERODYNAMIC CHARACTERISTICSOF AN EXTERNALLY BLOWN FLAP POWERED-LIFT MODEL WITHSEVERAL PROPULSIVE SYSTEM SIMULATORSBy Danny R. HoadLangley Director

5、ate, U.S. Army Air Mobility R however, the take-off and landing configurations required a high-lift tail.INTRODUCTIONThe externally blown flap (EBF) concept considered for use on a jet-powered STOLtransport has been investigated on various engine and model configurations. (See refs. 1to 7.) Very few

6、 investigations to date have considered the effect of engine type, size,and fan-exit location on the aerodynamic characteristics of the configuration. This wind-tunnel investigation was conducted to determine these effects with five available enginesimulators. Three sets of engine simulators were de

7、signed to represent bypass ratio 3.2,bypass ratio 6.2, and bypass ratio 10.0 engines. The bypass ratio (BPR) as designated iscorrect; however, the physical size and locations are not necessarily optimum for thisProvided by IHSNot for ResaleNo reproduction or networking permitted without license from

8、 IHS-,-,-research model. The BPR of these engine simulators does not describe in any way thesize, horizontal position, or vertical position of the simulator. The data presented for theBPR 3.2 engine configuration were obtained from reference 8. One of these simulators wasdesigned to represent the ex

9、haust characteristics of a daisy-nozzle exhaust shape. Thisshape was representative of a design such as the ones in references 9, 10, and 11 whichreduce exhaust velocity near the flaps and therefore reduce jet-impingement noise. Thismultilobe nozzle, which will be referred to as the daisy-nozzle eng

10、ine, was sized so thatthe inlet and exit areas would match the BPR 6.2 engine. One parameter which was notheld constant was the relative chordwise position of the engine exhaust with respect to theflap system. In an attempt to determine whether there was an effect of the relativechordwise position o

11、f the engine exhaust, an extension was added to the BPR 6.2 fan cowlso that its exit would be positioned at the same relative chordwise location as the daisy-nozzle fan exit.The investigation was conducted in the Langley V/STOL tunnel. The longitudinaldata are presented at several thrust coefficient

12、s with flap deflections which represent acruise configuration, a take-off configuration, and a landing configuration. The thrust-removed longitudinal data and isolated engine exhaust pressure decay characteristics arealso presented.SYMBOLSThe longitudinal aerodynamic data in this report are referred

13、 to the stability axes.(See fig. 1.) The origin of the axes was located on the fuselage center line longitudinallyat 0.40 mean aerodynamic chord and vertically at the average center line of each engineconfiguration.The units for the physical quantities defined in this paper are given in both theInte

14、rnational System of Units (SI) and the U.S. Customary Units. Equivalent dimensionswere determined by using the conversion factors given in reference 12.c local wing chord, meters (ft)c mean aerodynamic chord, meters (ft)cs local chord, horizontal stabilizer, meters (ft)CD drag coefficient, Dragq0SCL

15、 lift coefficient, Lift2Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Cm pitching-moment coefficient, Pitching momentC thrust coefficient, Thrustq OSit horizontal-tail incidence angle (positive direction, trailing edge down), degreesqj effective je

16、t-exhaust dynamic pressure, newtons/meter2 (lb/ft2)qm 0free-stream dynamic pressure, newtons/meter2 (lb/ft2)r radius, meters (ft)S wing area, meters2 (ft2)t/c airfoil thickness ratioT static thrust, newtons (lb)x longitudinal distance from leading edge of wing (positive when measured aft ofleading e

17、dge of wing), meters (ft)X,Z body reference axes (see fig. 1)a angle of attack, degrees5e elevator deflection (positive when deflected down), degrees6f wing trailing-edge flap deflection (positive when deflected down), degreesbj jet-exhaust deflection (measured from body reference axis X,positive wh

18、en deflected down), tan- 1 ormal forcedpositive when deflected down), Axial force , degrees6sh horizontal stabilizer leading-edge slat deflection (positive when deflecteddown), degrees6sw wing leading-edge slat deflection (positive when deflected down), degrees77 static thrust recovery efficiency, (

19、Normal frce)2T+ (Axial force)23Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MODEL AND APPARATUSA three-view drawing of the model with dimensional characteristics is presented infigure 2. This model was identical to the model in reference 8. (See r

20、ef. 8 for all airfoilordinates.) Photographs of the model showing the daisy-nozzle engine simulator installedand the triple-slotted flap system are presented in figure 3.The offtboard leading-edge slat element had an airfoil section that was a 25-percent-chord St. Cyr 178 modified to t = 0.0065c at

21、the trailing edge. This slat element extendedfrom the outboard engine pylon to the wing tip. Two inboard slat elements had an airfoilsection that was a 15-percent-chord St. Cyr 178 modified to t = 0.0065c at the trailingedge. The innermost slat element extended from the fuselage to the inboard engin

22、e pylon,and the third slat element extended between the inboard and outboard engine pylons. Inthe deployed position, the slat gap was 0.015c.The wing flaps were triple-slotted, full-span flaps with 15-, 20-, and 22.5-percentlocal wing chord for the first, second, and third elements, respectively. Th

23、e first andsecond flap elements had a St. Cyr 178 airfoil section modified slightly to provide a finitetrailing-edge thickness and to fit the upper surface contours of the wing in the retractedposition. The third flap element had a NACA 4412 airfoil section modified to t = 0.0045cat the trailing edg

24、e. All flap-slat gaps were 0.015c. The geometric characteristics ofthe wing leading-edge slats and flaps are presented in figure 4.The geometric characteristics of the horizontal tail are shown in figure 5. Thehorizontal tail was pivoted about 0.555 root chord with an incidence range of 150 in50 inc

25、rements. It had a 15-percent local-chord leading-edge slat set at -400. The35-percent local-chord elevator had three deflections relative to the tail chord: 00-250, and -500.Four air ejector engine simulators were used to represent each fan-jet propulsionsystem. Each engine simulator was a two-part

26、ejector with individual air-supply linesand control valves designed to provide the efflux of the fan- and gas-generator stages. Atypical ejector assembly is presented in figure 6. Each engine simulator was fitted withfive separate cowl assemblies intended to represent the five different engine confi

27、gura-tions. (See fig. 7.)The bypass ratio (BPR) is defined as the ratio of total fan-exit mass flow to totalgas-generator-exit mass flow. The five separate cowl assemblies fitted to these ejectorsimulators were designed to represent a BPR 6.2 daisy-nozzle engine, a BPR 6.2 engine,a modified BPR 6.2

28、engine, a BPR 10.0 engine, and a BPR 3.2 engine. The daisy-nozzle4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-engine simulator was designed to represent an engine whose exit and inlet area areidentical to those of the BPR 6.2 engine. This type of

29、 nozzle exit is designed to increasethe velocity decay characteristics of the engine and thereby reduce the velocity at theflaps, which, in turn, reduces flap-impingement noise. Although the inlet and exit areasof the BPR 6.2 and daisy-nozzle engines were identical, the longitudinal positions of the

30、irfan exits relative to the flap system were not. The BPR 6.2 engine simulator was modi-fied during the investigation. An extension was added which relocated the fan exit at thesame relative location as that of the daisy-nozzle engine simulator. This engine simulatorwill be referred to as the modifi

31、ed BPR 6.2 engine simulator. (See fig. 7.)The model was mounted in the Langley V/STOL tunnel on a sting-supported six-component strain-gage balance for measurements of the total forces and moments.TEST AND CORRECTIONSThis investigation was conducted in the Langley V/STOL tunnel. The free-streamdynam

32、ic pressure for the entire investigation was 814 N/m2 (17 lb/ft2). The Reynoldsnumber (based on wing Z and free-stream velocity) was approximately 0.697 x 106.The data presented in this report are not corrected for wind-tunnel wall effects. Sincethe corrections calculated by the method of reference

33、13 were found to be small and thisreport is a comparison of the engine-model configurations of the test, it is felt that thedata are valid.Calibrations were made to determine the thrust, inlet mass-flow rate, and primarymass-flow rate of the fan- and gas-generator stage of each engine simulator asse

34、mblyseparately as a function of their respective plenum pressures. These data were runat zero airspeed and reflect the static thrust only. The values of thrust coefficient arebased on this static-thrust calibration and are presented as the conventional thrust coef-ficient, that is, static thrust non

35、dimensionalized by the product of free-stream dynamicpressure and wing area (Cg =The static inlet and primary mass flow rates were used to set the desired pressurein the plenum of each separate stage of each engine to provide the correct ratio of totalfan exit mass flow rate to total gas generator e

36、xit mass flow rate (BPR).Isolated engine-exhaust-wake surveys were conducted. Dynamic-pressure measure-ments were made with a pressure rake positioned so that the probes were alined with theflow and parallel to the engine geometric center line. The probes were alined along aradial line from the geom

37、etric center line. Four radial positions were chosen for the5Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-daisy nozzle and two radial positions were chosen for the other four engine simulators.These profiles were repeated at various downstream loc

38、ations to obtain dynamic-pressure decay characteristics.Jet-deflection angles 6j and static-thrust recovery efficiency 77 were determinedfrom measurements of the normal and axial forces made in the static thrust conditionwith flaps deflected and leading-edge slat deployed.Each of the five engine con

39、figurations was tested in the cruise configuration, thetake-off configuration, and the landing configuration at an angle-of-attack range from -40to 240. Each configuration was tested at a thrust coefficient range from 0 to 4 and atvarious horizontal-tail incidences.The cruise configuration was defin

40、ed as the model with 00 flaps, no wing leading-edge slats, no horizontal-tail leading-edge slats, and elevator set at 00. The take-offconfiguration was defined as the model with the elements of the three-element flap sys-tem set at 00, 200, and 400; wing leading-edge slat deployed at 500; horizontal

41、-tail leading-o 0edge slat deployed at -400; and elevator set at -25. The landing configuration wasdefined as the model with the flap system elements set at 150, 350, and 550; wing leading-edge slats deployed at 500; horizontal-tail leading-edge slats deployed at -400; and ele-vator set at -250. The

42、 wing cross sections for these configurations are presented infigure 4.PRESENTATION OF RESULTSResults of the present investigation are presented in the following figures:FigureFlap static turning effectiveness forthrust of 1219 newtons (274 lb) 8Effect of engine type on the variation of flap static

43、turningeffectiveness parameters with thrust . . . . . . . . . . . . . . . . . . . . . 9Effect of wind-tunnel wall corrections 10Effect of thrust coefficient on longitudinal aerodynamic characteristics for -Daisy nozzle:Cruise configuration 11Take-off configuration 12Landing configuration 136Provided

44、 by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FigureBypass ratio 6.2:Cruise configuration . 14Take-off configuration 15Landing configuration 16Modified bypass ratio 6.2:Cruise configuration . 17Take-off configuration 18Landing configuration 19Bypass ratio

45、10.0:Cruise configuration . 20Take-off configuration 21Landing configuration 22Bypass ratio 3.2:Cruise configuration . 23Take-off configuration 24Landing configuration 25Effect of thrust coefficient and tail incidence on pitching-momentcharacteristics for -Daisy nozzle . . . . . . . . . . . . . . 26

46、Bypass ratio 6.2 . . 27Modified bypass ratio 6.2 . 28Bypass ratio 10.0 . . . . 29Bypass ratio 3.2 . . . . 30Effect of engine type on longitudinal aerodynamic characteristics for -Cruise configuration, tail off 31Cruise configuration, it 00 . . . 32Take-off configuration . . . . . . . . . . . . . . .

47、 . . . . . . . . . . . . . . . 33Landing configuration . 34Effect of engine type on variation of lift coefficient with thrust coefficient . . . . . 35Effect of thrust coefficient on thrust-removed lift coefficient anddrag coefficient for -Daisy nozzle . . . . . . . . . . . . . . . . . . . . . . . .

48、. . . . . . . . . . . 36Bypass ratio 6.2 . . 37Modified bypass ratio 6.2 . 38Bypass ratio 10.0 . . 39Bypassratio 3.2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40Effect of engine type on thrust-removed lift coefficient and drag coefficient . . . . 41Isolated engine effective d

49、ynamic-pressure decay . . . . . . . . . . . . . . . . . 427Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-DISCUSSION OF RESULTSThe reader should be reminded that each engine simulator used in this investigationis not necessarily indicative of the geometric size of the respective full-scale engine whoseby