NASA NACA-RM-A8A20-1948 An investigation of submerged air inlets on a 1 4-scale model of a typical fighter-type airplane《嵌入式进气道在典型战斗机的1 4比例模型上的研究》.pdf

《NASA NACA-RM-A8A20-1948 An investigation of submerged air inlets on a 1 4-scale model of a typical fighter-type airplane《嵌入式进气道在典型战斗机的1 4比例模型上的研究》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-A8A20-1948 An investigation of submerged air inlets on a 1 4-scale model of a typical fighter-type airplane《嵌入式进气道在典型战斗机的1 4比例模型上的研究》.pdf（48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-,-,-NACA RM No. 820 NADOmAL ADVISORY COMMITTEE FOR AERONAUTTCS RESEARCH MEMORANDUM AN INVESTIGA!LTON OF SUBMERGED AIEt lXGF,IS ON A l/hCm MOIEL OF A TYPICAL FIGHTEZ-TYPE AII(pLANE By Noel IC. Delany SUMMARY

2、 Wind-tunnel tests were made of submerged air inlets on the fuselage of a l/Ic-scale model of a typical fighter-type airplane. The results are presented for ramp plan forms with parallel and with diverging walls and show the effect of the duct-entrance location (forward of the wing and over the wing

3、), internal ducting efficiency, and deflectors. The air inlets having the ramps with djverging walls were satis- factory in both locations tested on the fuselage, providing high ram pressure recoveries at the simulated entrance to the compressor, high predicted critical Mach numbers, and low externa

4、l drags. The submerged air inlets with parallel ramp walls had lower ram pressure recoveries for the normal operating range. measured at the inlets were higher for the forward location of the inlets than for the aft location. For an assumed engine position, however, the aft location of the inlets wi

5、th the shorter, more efficient internal ducts gave the higher ram recoveries at the simulated compressor for the test conditions. The ram pressure recovery ratios INIBOIJUCTION The early development of NACA submerged air inlets was conducted with the submerged inlets installed in the flat wall of a

6、wind tunnel (references 1 and 2)0 The results of these tests indicated that it should be feasible to design an efficient air-induction system with twin submerged inlets installed on the sides of the fuselage. Placing the submerged inlets on the sides of the fuselage ahead of the jet engine results i

7、n a short, straight internal ducting system (references 3 and 4). not protrude outside of the basic fuselage cclntour they should-tend As the submerged inlets will Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM No, A8A20 to eliminate, by i

8、nertia separation, foreign material (shell cases, rocks, hail, etc.) from the air inducted into the motor. The results of reference 3 indicate that the relative location of the wing and the submerged inlets might be critical for inlet performance. The purpose of the tests presented in this report wa

9、s their characteristics. Two locations were tested, one forPaxd of the wing where the fuselage boundary layer was thin, and the other farther aft on the fuselage and over the point of maximum thickness of the wing, the effect of a tractor propeller on the ram recovery could be determined . . to inve

10、stigate the effect of the location of the duct inlets on The model was constructed so that, in later tests, The test results presented in this report were obtained in the Ames 7- by 10-foot wind tunnel No. 2 at the request of the Bureau of Aeronautics, Navy Department. A B D H M P 9 R r S v V SYMBOI

11、S The symbols used throughout this report are as follows: area, square feet depth of the ramp at the lip, inches drag, pounds total pressme, pounds per square foot Mach number static pressure, pounds per square foot dynamic pressure, pounds per square foot radius of duct, feet radius to a point, fee

12、t wing area, square feet stream velocity, feet per second local velocity, feet per second . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WA RM No. Am0 3 Y . di8,tance perpendicular to a surface, inches boundary-layer thickness, inches model angle

13、of attack with respect to the fuselage reference line, degrees P mass density of the air, slugs per cubic foot . The following subscripts have been used in conjunction with the above symbols : 0 1 2 3 cr av H-Po HxG f ree-stream duct entrance (1.5 in. behind lip leading edge) inlet to the compressor

14、 jet exhaust critical average The following ratios and coeff icienbhave been used: ram recovery ratio inlet velocity ratio v1 - TO CDintemal internal drag coefficient CDD external drag coefficient of inlet based on wing external drag coefficient of inlet based on inlet c?D Provided by IHSNot for Res

15、aleNo reproduction or networking permitted without license from IHS-,-,-NACA RM No. 820 4 h P qD the height of an area of unit width in which-the complete loss of free-stream ram pressure is equivalent to the integratedloss of the total pressure in unit width of the boundary layer (?I pressure caeff

16、 ic ient internal ducting efficiency DESCRIPTION OF MODEL This investigation of twin NACA submerged air inlets was conducted with a l/kcale model of a typical high-speed, turbo- propeller driven, fighter-type airplane. In this series of tests the propeller was not used, The pertinent model dimension

17、s and a three-view drawing of the airplane are presented in Appendix A and figure 1, respectively. A photograph of 6he model mounted in the wind tunnel is shown in figure 2. The submerged air inlets investimted were designed from the results of reference 2 which indicated that an entrance aspect rat

18、io of 4 and a ramp having an angle of 7O with respect to the fuselage surface and curved diverging walls should produce optimum chmacteristics. The ramps were submerged in the fuselage so that the ordinates of the rmp below the basic fuselage contour (fig. 3) were equal to those for a 7 ramp below a

19、 plane surface. The ramp plan forms tested are given in figure 3 and correspond to those of reference 2. The lips of the duct inlets tested (fig. 4) were the sam.e as the untilted lip of reference 2 but with the mean camber line tilted in 3O. on the center line of the ramps and lips of the air intak

20、es. Flush static-pressure orifices were installed Two inlet positions, on the sides of the fuselage, were tested. For both positions the horizontal center plane of the inlets was in the horizontal fuselage reference plane (figs. 1 and 3) which was 7.1 percent of the root chord of the wing above the

21、wing upper surface at the point of maximum thickness of the wing at the root. For the forward position of the inlets, the leading edge of the lip was 19.3 percent of the root chord of the wing ahead of the wing leading edge, and for the aft position of the inlets the leading edges of the lips were a

22、bove the point of maximum thickness of the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM No. 820 5 wing-root sec*ion (35.6 percent chord), One location of the jet motor was assumed for the airplane. This location allowed a short internal du

23、cting system for the aft location of the inlets and a longer internal ducting system for the forward location. These two internal ducts are shown assembled for preliminary bench tests in figure 5. The long internal duct . consisted of the short internal duct with a 14.2Finch, constant- mea section a

24、dded to extend it forward. comparison of the duct entrances, the forward inlets were also tested with the short internal ducting system, The area ratio between the simulated face of the turbo-jet compressor and the submerged inlets was 1,336 for both the short and long internal ducts. To provide a m

25、ore complete Deflectors (reference 2) were investigated on only the inlets with divergent ramp walls. deflectors installed on the model are shown in figures 6 and 7, respectively. and aft locatiomof the inlets while various modifications were investigated for the forward location of the inlets. Coor

26、dinates and photographs of the The normal deflectors were tested at both the forward I JBST METHOD3 AM) REIUCTION OF DATA The quantity of air flow through the submerged air inlets of the model was varied and controlled by a centrifugal pump located outside of the wind tunnel, The pump was connected

27、to the duct system by a pipe attached to the rear of the model. The length of the pipe (fig. 2) attached to the model and passing through the wind-tunriel floor was flexible to allow the angle of attack of the model to be changed. A standard sharp-edged ASME orifice meter was used to measure the qua

28、ntity of air drawn through the submerged air inlets., In determining the inlet velocity ratio from the measured quantity of flow, the free-fltream air density was used. This intro- duced a maximum error of 2.0 percent in the inlet velocity ratio. Ram pressure recovery, at the duct inlets and at the

29、simulated entrance to the compressor, was measured by rakes of pressme tubes. There were 36 total-pressure and 5 static-pressure tubes in each inlet and 40 total-pressure and 4 static-pressure tubes at the simulated entrance to the compressor. In computing the mean ram recovery ratio at the inlets H

30、1-po/Ho-po the reading of each tube was weighted (reference 2) in accordance with the variation of the mass flow across the duct inlets. As. the vmiations in the velocity weresmall at the simulated entrance to the compressor, an Provided by IHSNot for ResaleNo reproduction or networking permitted wi

31、thout license from IHS-,-,-6 NACA RM No. A8A20 arithmetical mean of the tube readihgs we used to determine the ram recovery ratio H-p/Ho-po at this position, The external drag of the submerged inlets was determined for only the forward location of the inlets. determine the drag: and (2) measurements

32、 of the momentum of the air just behind the inlet location. The force-test drag was measured with the flexible pipe (fig. 2) at the aft of the fuselage removed (fig. 8) while air was allowed to bleed through the internal ducting system. The inlet velocity ratio V1/Vo was changed by varying the outle

33、t area A3 of the duct for the force-test drag measurements. Two methods were used to (1) force measurements on the complete model, The drag attributed to the submerged inlets was Wen as the difference in the drag, measured by the wind-tunnel balances, with the duct entrances installed and removed le

34、ss the internal drag. The internal drag was calculated from the loss of momentum per unit time of the air flowing through the internal ducting. The internal drag coefficient was computed.with the following equation: -“( 2) 1 - p)( “( 1 + L) 1+2n Vo A3 Dinternal - s The value of the constant n was fo

35、und to be 0.44 X.(Ax/her with the inlets in the forward than in the aft posit i on. 4. The ram recoveries were higher at the simulated entrance to the compressor, with some inlet velocity ratios, for the aft location of the inlets with the short internal ducting than for the forward location with th

36、e longer internal ducting. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 WCA RM No. 820 5. The drag of the fuselage with submerged duct inlets operat- ing with inlet velocity ratios greater than drag of the basic fuselage; however, with inlet ve

37、locity ratios below 0.70 there was an appreciable increase in the drag attributable to the inlets. 0.70 was less than the 6. The external drag of the deflectors more than offset the improved ram recovery they provided on this model. 7. For the high-speed flight condition, the predicted critical Mach

38、 number of the inletswas hi#er for the forward location than the aft location. Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics, Moffett Field, Calif. 1. Wick, Charles W, Davis, Wallace F, kandall, hwos M, and Mossman, Met A.: Submerged-Duct Entrances. An Experimental Invest

39、igation of NACA NACA ACR No. 5120, 1945. 2. Mossman, Emmet A+, and Randall, Lauros M.: An Experimental Investigation of the Design Variables for NACA Submerged Duct Entrances. NACA No. A7130, 1947. 3. Gault, Donald Ee : An Experimental Investigation of NACA Submerged Air Inlets on a 1/wcale bdel of

40、a Fighter Airplane. NACA RM No. 7106, 1947. 4. Mossman, Emmet A., and Chiit, Donald E.: Development of NACA Submerged Inlets and a Comparison with Wing Leadingddge Inlets for a l/bcale Model of a Fighter Airplane. NACA RM No. A7A31, 194% 5. von E MOIIEL OF A TYPICAL FIGH!I!IE+-!IXPE AIRPLANE Model W

41、ing area . 14.519 sq ft Aspect ratio 4.98 Wingspan . 8.50ft Wing section 631-110 Root chord . 2.30 ft Tip chord 1.15 ft Wing incidence 0 Submerged Inlets tmpa;ngle,. 7 Aspect ratio of inlet 4 Total crost3+3ectional area of both inlets measured l$ inches behind lip leading ed.ges . . 0.0718 sq ft Dep

42、th of the ramps at the lip leading edges . . . 1.720 in. Distance of duct-lip leading edges from wing leading edge Forward location 19.3 percent root chord ahead Aft location . 35.6 percent root chord behind Distance of inlet center lines above the wing at the fuselage juncture . e . ., 7.1 percent

43、root chord Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-16 NACA RM No, A8A20 Length of the internal duct from lip leading edge to the simulated entrance of the compressor Short internal ducting 15,25 in. Long internal ducting .29.5 ina kea ratio (

44、2) 1.336 APPEM)M B Determination of Internal Drag For the determination of the external drag of the twin submerged duct inlets the drag of the internal ducting had to be determined. The internal drag was computed from the inlet velocity ratio V1/Vo, wing area S, duct inlet area AI, and the duct exit

45、 area bo The internal drag was taken as the free-stream ran drag minus the momentum of the air per unit time exiting from the tail pipe (reference 2) The first term of equation (1) is readily evaluated, For the second term, surveys were made at the exit across one diameter to determine the variation

46、 of the velocity v3 across the outlet, The velocity distribution was assumed equal on all diameters. The experimental velocity prof ilea were plotted and matched by a mathematical curve where n was found to be equal to 0.44 Ax/ velocity across the kxit is Substituting the value of v3 from equation (

47、B2) and integrating equation (B3) v3max V3 = (l+n) From the continuity equation for an incompressible fluid VlAl = V3&. Substituting this in equation (B4) Subatituting the value of v3max from equation (B5) in equatfon 02 1 n Substituting the value of v3 and integrating from equation (B6) in equation

48、 (Bl) The internal drag may now be computed from the inlet velocity ratio V1/Vo, wing area S, duct inlet area Al, and the exit area A3. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-J FIGURE I- A THUEE- VIEW DffAWfNG OF THE AIRPLAN Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted