NASA-TM-X-1517-1968 Effect of boattail juncture shape on pressure drag coefficients of isolated afterbodies《船形尾部接缝形状对绝缘飞机后体压力阻力系数的影响》.pdf

《NASA-TM-X-1517-1968 Effect of boattail juncture shape on pressure drag coefficients of isolated afterbodies《船形尾部接缝形状对绝缘飞机后体压力阻力系数的影响》.pdf》由会员分享，可在线阅读，更多相关《NASA-TM-X-1517-1968 Effect of boattail juncture shape on pressure drag coefficients of isolated afterbodies《船形尾部接缝形状对绝缘飞机后体压力阻力系数的影响》.pdf（35页珍藏版）》请在麦多课文档分享上搜索。

1、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NASA TM X-1517 EFFECT OF BOATTAIL JUNCTURE SHAPE ON PRESSURE DRAG COEFFICIENTS OF ISOLATED AFTERBODIES By George D. Shrewsbury Lewis Research Center Cleveland, Ohio NATIONAL AERONAUT ICs AND SPACE ADM I

2、N ISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - CFSTI price $3.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EFFECT OF BOATTAIL JUNCTURE SHAPE ON PRESSURE DRAG COEFFICIENTS

3、 OF ISOLATED AFTER BODIES by George D. Shrewsbury Lewis Research Center SUMMARY A variety of afterbodies were tested on a sting-supported model of a closed-inlet nacelle. Jet effects were simulated with a cylinder positioned downstream of the after- body base. Axial-force coefficients were obtained

4、for a 7 conical boattail and various 15 boattailed afterbodies on which the boattail juncture with the cylindrical portion of the nacelle had been smoothed with different radii of curvature. Data were obtained over a Mach number range of 0.56 to 1.00 at angles of attack from 0 to 8. the occurrence o

5、f the transonic drag rise. With the 15 boattails, the sharp edge (R/DM = 0) configuration had a drag-rise Mach number near 0.6. Increasing the radius of curvature to R/DM = 1 delayed the drag-rise Mach number to approximately 0.8. For R/DM of 2.5 or greater, the drag-rise Mach number occurred slight

6、ly above Mach 0.9. The results indicate that increasing the boattail radius of curvature generally delays INTRODUCTION Supersonic airbreathing propulsion systems designed for Mach numbers up to 3.0 operate over a range of nozzle pressure ratios from approximately 2. 0 to 30.0. Efficient performance

7、of the propulsion system at all flight speeds requires variations in the noz- zle expansion ratio. If the configuration utilizes nacelle -mounted engines and divergent ejector nozzles, it may have a nearly cylindrical afterbody at the design Mach number. Because of the high nozzle pressure ratio at

8、the design Mach number, external flow ef- fects have little effect on nozzle performance. Off -design operation, however, requires a boattailed afterbody of the engine nacelle in order to provide this decrease in expansion ratio. The drag incurred by boattailing the nacelle afterbody can be a signif

9、icant portion of propulsion system net thrust, especially at subsonic cruise where the engine is throt- tled. Many supersonic aircraft missions may require that sizeable portions of the flight Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-be conduc

10、ted at subsonic Mach numbers. The subsonic cruise Mach number selection is influenced by the boattail transonic drag rise characteristics. If the transonic drag rise can be delayed, a higher subsonic cruise Mach number may be permissible. Conse- quently, the drag characteristics of the nacelle after

11、body become of significant impor- tance at subsonic and transonic Mach numbers, It has been demonstrated that circular arc afterbodies result in lower drag coeffi- cients than conical afterbodies for equal boattail angles and ratios of base diameter to maximum diameter (ref. 1). Since most supersoni

12、c aircraft nozzle system geometries are variable, the full circular arc afterbody, although desirable from a drag viewpoint, is mechanically dLfficult to transform into a smooth cylinder for design Mach number operation. Therefore, it became desirable to investigate intermediate transition radii of

13、curvature at subsonic and transonic Mach numbers. study the effect of varying the boattail transition radius of curvature on a 15 boattail with a ratio of base diameter to maximum diameter of 0.67. The jet was simulated with a solid cylinder which had a diameter equal to the afterbody base diameter.

14、 Four-inch diameter models were tested with six radii of curvature ranging from 0 (sharp corner) to 4.84 DM (tangent ogive). A 7 conical boattail with a L/DM the same as the 15 conical boattail was also investigated. Data were obtained over a Mach number range of 0.56 to 1.00. The models were tested

15、 at angles of attack ranging from 0 to 8. The test sec- 6 6 tion Reynolds number ranged from 3.610 per foot to 4.610 per foot. An investigation was conducted in the Lewis 8- by 6-foot Supersonic Wind Tunnel to SYMBOLS A a cP D L M P q R V X 2 area axial-force coefficient, axial for ce/qo AM pressure

16、 coefficient (p - po)/qo diameter model length from afterbody base Mach number static pressure dynamic pressure boattail juncture radius of curvature velocity axial distance aft of model afterbody interface Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS

17、-,-,-Y radial distance from model surface Q! model angle of attack, deg P boattail trailing-edge angle, deg 6 boundary layer thickness Subscripts : a axial b afterbody base e nozzle exit conditions L local M maximum 0 f ree-stream conditions S sting P boattail surface APPARATUSAND PROCEDURE The comp

18、lete afterbody model configurations, as installed in the Lewis 8- by 6-foot Supersonic Wind Tunnel, is shown in figure 1. The basic model was a sting-supported 4-inch-diameter (10.16 cm) cylindrical section with a 10 half -angle conical forebody. The length of this cylindrical section was varied to

19、evaluate the effect of boundary layer thickness ahead of the afterbody region. Figure 2 is a sketch of the model installation showing the location of the short and long models in the perforated test section. The lo- cation of the forebody remained fixed, and the position of the afterbody moved aft w

20、hen the model length was changed from short to long model configurations. The model length from the forebody shoulder to the model-afterbody interface varied from 5.91 to 10.91 model diameters. Mach number to minimize tunnel wall interference effects, a constant value of 3.1 percent was selected for

21、 this study. Other unpublished data from the 8 by 6 tunnel indicate that this is an acceptable compromise with 4-inch (IO. 16 cm) diameter models. Model block- age was 0.18 percent at a 0 angle of attack. The cylindrical portion of both model lengths was pressure instrumented at 2-inch (5.08 cm) int

22、ervals along the top and side. A boundary layer rake was installed on both model lengths to survey the local flow field ahead of the afterbody region and to measure 0 Although reference 2 indicates the desirability of variable tunnel wall porosity with 3 Provided by IHSNot for ResaleNo reproduction

23、or networking permitted without license from IHS-,-,-boundary layer thickness. The boundary layer survey plane was located 1 inch (2.54 cm) forward of the model-afterbody interface, The total pressures from the rake were used with static pressures located at 90 and 180 from the rake to compute value

24、s of V/Vo using the Rayleigh-pitot equation. Details of the boundary layer rake are shown in fig- ure 3. The afterbody geometries investigated are shown in figure 4. The afterbody geom- etries included a cylindrical afterbody, a 7 conical boattail, and 15 boattails with radii of curvature of 0 (shar

25、p edge), 0.5, 1.0, 2.5, 3.5, and 4.84 model diameters. Details of the afterbody geometries are shown in figure 5. The cylindrical afterbody was a heavily instrumented extension of the cylindrical portion of the model. It was investi- gated to determine the static-pressure environment of the afterbod

26、y region as influenced by terminal shock waves from the conical forebody, wall reflected expansion and shock waves, and wall-generated disturbances. The 7 conical boattail was investigated to evaluate boattail angle effects. The boattail L/DM of the 7 afterbody was the same as the 15 conical boattai

27、l. All 15 boattails had a ratio of base diameter to maximum di- ameter of 0.67. On boattails with radii of curvature, the curvature was tangent to the cylindrical portion of the afterbody. Since the ratio of base diameter to model diameter was held constant, increasing the radius of curvature increa

28、sed the length of the boattail. The 4.84 R/DM boattail is a tangent ogive configuration with the entire boattail being a curved surface. Since the boattail axial-force coefficients were determined from boattail pressure measurements, extensive pressure instrumentation was located on the afterbodies.

29、 In- strumentation details for the 15 conical boattail are shown in figure 6. Instrumentation of all afterbody configurations was similar. The axial projection of the boattail was di- vided into 10 equal annular areas. Pressure taps were located around 180 of the cen- troid line of each annular area

30、 at 30 intervals. It was assumed that the local flow field would be symmetrical about a vertical plane through the model axis, so pressures were located only on one side of the boattail. Extra pressure taps were locatednear the corner of the boattails to help define boattail pressure distribution. T

31、hese pressures were not used for drag determination. coefficient can be computed. This average pressure coefficient is then used to compute the axial-force coefficient. The boattail axial-force coefficient computed in this manner does not include the afterbody base drag or afterbody skin friction dr

32、ag, but pertains only to those pressure forces acting on the boattail surface. The afterbodies were tested in the presence of a cylindrical section extending from the afterbody base. The purpose of the cylinders was to approximate the local flow field that would exist if a jet were present with a va

33、lue of pe/pL = 1.0. The simulator diam- eters were equal to the afterbody base diameters. Details of the jet simulators for the By instrumenting the boattail in this manner, an area weighted average of pressure 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license fr

34、om IHS-,-,-7 and 15 boattails are shown in figure 7. The afterbodies were also tested with only the 1.62-inch (4.115 cm) support sting present to evaluate the effect of the jet simulator. RESULTS AND DISCUSSION Jet Simulator and Sting Effects The experimental and calculated effects of different stin

35、g diameters on the drag of the 4.84 R/DM (ogive) boattail are shown in figure 8. The boattail axial-force coefficients for the ogive boattail were corrected for sting effects by using the method developed in reference 3. The ogive boattail was used because the corrections made using this method are

36、valid only for circular arc boattails, parabolic boattils, or conical boattails with angles less than 8. In addition, the corrections are valid only for Mach numbers below the transonic drag rise. By utilizing this method, the experimental data taken with the 1.62-inch (4.115 cm) sting were used to

37、calculate the no-sting axial-force coefficients. The no-sting results were then adjusted for the jet simulator effect (by the same method), and a comparison with the experimental data was reasonably good as indicated by figure 8. for the entire Mach number range. Data are shown with and without the

38、jet simulator. The presence of the jet simulator results in a decrease in boattail axial-force coefficient for all Mach numbers investigated. The effect of jet simulator on afterbody pressure distribution is shown for a 15 conical boattail in figure 10. The presence of the jet simulator creates a st

39、ronger re- compression region at the boattail trailing edge than exists with only the support sting present. At subsonic speeds the stronger recompression region causes higher pressures to propagate forward on the boattail surface. This effect was not as apparent at Mach 1.0. figure 11. The axial-fo

40、rce coefficients for the cold flow model are for a slightly dif- ferent configuration in that Db/DM = 0.65. The cold jet data shown were interpolated values for pe/po = 1.0. Since the local pressure at the base plane is nearly equal to free-stream static, a pe/po of 1.0 should result in a nearly cyl

41、indrical jet. In general, the data are in good agreement except at Mach 1.0, which indicates that the effects of afterbody shape on drag-rise Mach number are valid utilizing this solid jet simulation technique. Boattail axial-force coefficients for the 15 conical boattail are shown in figure 9 A com

42、parison of jet simulator data and cold jet data from reference 4 is shown in Pressure Environments of Short and Long Model Configurations The cylindrical afterbody was tested with both short and long model configurations to determine the presence of terminal shocks, reflected expansion and shock wav

43、es, and 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-wall-generated disturbances. The effect of afterbody location on cylindrical afterbody model pressure distributions is shown in figure 12. For the subsonic Mach numbers, the afterbodies of the

44、 short and long models are in nearly identical pressure environments. At Mach values of 0.67, 0.74, and 0.80, the local static pressures on both model lengths were slightly below free-stream pressure, This resulted in an increase in Mach number of approximately 0.01, At transonic Mach numbers, over-

45、expansion at the shoulder of the forebody is severe enough that recompression along the cylindrical portion of the model is partially accomplished by a normal shock wave called a terminal shock. At Mach 1.0, the terminal shock wave? is located approximately 4.0 model diameters forward of the after -

46、 body base of the short model. The terminal shock wave occuring on the long model is located farther upstream from the afterbody region and is not seen in figure 12. The in- fluence of terminal shock on afterbody pressures is small for both long and short models. With the cylindrical afterbody confi

47、guration, an inadvertent 0.003-inch (0.0076 cm) for- ward facing step existed at the model-afterbody interface. Disturbances from this step apparently created small pressure increases at Mach 1.00 approximately 1.5 model diam- eters forward of the afterbody base. Since this disturbance existed on bo

48、th the long and short models, it could not have been a result of tunnel wall effects. This disturbance probably had little effect on the pressures in the afterbody region however, since the ratio of step height to model diameter is relatively small. Similar surface irregularities did not exist on th

49、e other afterbody configurations. Boundary Layer Boundary layer profiles for both short and long model configuration are shown in figure 13. The experimental data are compared with a theoretical seventh-power bound- ary layer profile. Both short and long model configurations have well developed turbu- lent profiles at all Mach numbers investigated. Boundary layer thi