NASA NACA-RM-A9F16-1949 A comparison of two submerged inlets at subsonic and transonic speeds《在亚音速和跨音速时两个嵌入式进气道的对比》.pdf

1、RESEARCH MEMORANDUM A COMPARISON OF TWO SUBMERGED INLETS AT SUBSONIC AND TRANSONIC SPEEDS By Emmet A. Mossman Ames Aeronautical Laboratory Moffett Field, Calif. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON Sedember 15. 1949 UNCLASSIFIED Provided by IHSNot for ResaleNo reproduction or netwo

2、rking permitted without license from IHS-,-,-UNCLASSIFIED NATIONALADVISoRy COMMImFaRmONAUTICS A COMPARISONOFTWO SUBMERGED IXWZTSAT SUBSONIC AND !llEumoNIc SPEEDS By Emmt A. Mossman Operation of two submerged-type inlets has been simulated in a 2.1- by 7. consequently, for the simulated NACA sublmjrg

3、ed inlet the losses due to turbulent mixing, as explained in reference 4, are not included. However, the nreasurements in this center plane should qualitatively indicate the inlet characteristics at high subsonic and transonic Mach numbers. Static-pressure distributions down the center line of the r

4、smp leading to the entrances were measured with flush orifices connected to a multiple-tube manometer. Measurenrenta for computing wind-tunnel Mach _ number distributions were also obtained from flush static orifices distributed over a steel plate mounted on one side of the test section. Visual flow

5、 studies were made with a schlieren apparatus and with a shadowgraph apparatus utilizing a Libessart spark. For this report, the free-stream Mach number is defined as the Mach number naeasured on the center of the tunnel floor one-uarter inch forward of ramp station 0. This location on the winddunne

6、l floor was selected so that the inlet would have the least effect on the free- stream MBch number measurement. A direct-ad- nomographic Mach Illster, explained in reference 5, was used to indicate the wind-tunnel speed in terms of free+tream Mach nuniber. Both inlets were tested from 0.20 Mach numb

7、er to the msximum that could be obtained with this wfnd tunnel. The Mach number limit was 0.94 with the psraUel-xalled inlet, and 0.96 with the divergent-walled inlet. The maximumMach number attainable with the parallel-wslled inlet installed in the wind tuzmel was determined by power limitatfons of

8、 the wind-tuunel mtor-compressor unit; whereas with the divergent-walled inlet the limiting factor appeared to be the errtablishment of sonic velocity across thewind tunnel back of ramp station 0. The range of mass-flow ratios varied with Mach number and inlet configuration. The following table indi

9、cates the mass-flow ratios that were obtainable during these tests: Mach nuciber Range of mass-flow ratio, mx/rao Parallel walls Divergent walls 0.20 0 to 1.2 0 to 1.2 :Z 0 to 0.8 1.2 0 to 1.2 1.0 .80 0 to 0.8 0 to 0.8 .90 0 to 0.2 0.4 to 0.8 0.6 0.4 to 0.8 mm- 0.4 to 0.8 . Provided by IHSNot for Re

10、saleNo reproduction or networking permitted without license from IHS-,-,-NACA HM AgF16 t 5 RESULTS The Mach nmiber distribution in the wind-tunnel test section, calc+ la-ted from static pressures measured on one test-section wall, is shown in figure 4 for free-stream bout the inlet was constrained b

11、y the wind-tunnel walls. Also, the ratio of the inlet area of the duct to the cross-sectional area of the wind tunnel was relatively large (1 to 15). Consequently, the Mach nuniber distribution in the test section was affected by mass-flow ratio (fig. k(a). However, the data presented should be usef

12、ul qualitatively. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA RM A9F16 Shadowgraphs with either inlet installed indicate that at Mach - numbers up to 0.93 the oblique shock disturbance, origiuating at the beginning of the test-section expa

13、nsion, was weak; consequently, the shock is believed to have had a negligible effect on the conditions downstream. (See fig. IO(g).) Conp?srism of Pressure Recovery It should be remmibered that the pressure recoveries presented in this report were obtained. only in a line normal to the duct width an

14、d passing through the duct center line. The transverse. variation of pressure recovery has not been determined. . A significant effect of Mach-nuniber on the pressure recovery of the parallel-+ralled iulet was evident at Mach numbers between 0.75 aud O. conse- quently, its growth was less rapid than

15、 for the two-dimensional flow which existed with the parallelralled inlet reference 6). However, for both inlets, the decrease in pressure recovery at the center section with a decrease of mass-flow ratio was due to a thickening of the ramp boundary layer. This thickening was, in turn, a consequence

16、 of increased adverse pressure gradients along the ramp. The pressure recoveries given in figures 5(a) and 5(b) are an indication of the relative boundary-layer thicknesses of the two types of Wets. Measurements of the velocity profile just behind the beginning of the parallelalled ramp showed that

17、the boundary layer was turbulent. It should be noted in figure y(a) that the curves showing rBTlt- recovery ratio for the parallel-walled iulet are extrapolated. In the Mach number range between 0.79 aud 0.94 the air flow in the duct was unstable and it was not possible to obtain consistent data in

18、this ranQe. However, the pressure recovery did decrease markedly, and it was impossible to obtain mass-flow ratios greater than 0.20 with the test equipment. For a Mach number of Oi94 the air flow became steady at a mass-flow ratio of 0.60. Provided by IHSNot for ResaleNo reproduction or networking

19、permitted without license from IHS-,-,-NACA RM AgF16 7 . Shadowgraph studies of the air flow with the parallel-walled inlet did not show the presence of strong shock waves for Mach numbers just before snd after the sharp decline in pressure recovery (M. about 0.75 snd 0.82, respectively). There was

20、evidence, however, of a shock dis-, turbsnce which extended only a short distance above the ramp surface at the beginning of the ramp and coincided with a thickening of the bound- arylayer along the rw surface (figs. 10(b) and 10(d).) Figure i(a) shows a decrease in pressure recovery for the psralle

21、lralled inlet as the free-etreamMach number was increased from 0.76 to 0.80. At greater Mach numbers, visual observations of the multiple manometer registering the rressure recovery indicated the unstable nature of the air flow in the duct system. Visual schlieren studies showed boundary-r separatio

22、n along the ramp for Mach numbers just greater than those at which the sharp decrease of pressure recovery occurred. A return to a more stable type of boundary-layer air flow is indicated by the shadowgraph for a free-o Divergent - wa leading edge. L Provided by IHSNot for ResaleNo reproduction or n

23、etworking permitted without license from IHS-,-,-20 - NACA RM A9F16 .8 .6 A2 Qw8 $ .6 3 k .4 -1 m 0 / 2 3 4 5 6 7 . . -/ 0 / 2 3 4 5 6 7 L2 .8 .6 .4 -1 0 / 2 3 4 5 6 7 Rump stution, in. v (ii PufuUe/-wu4ed Wet. . l a- . . Figure z p 0.8. (b) s = 0.80; e = 0.6. (cl M$ 0.82; y$ = 0.2. (a) hfo= 0.82; :

24、 = 0.6. v A-14019 FQpre lO.- Inlet. Shadowgraphs of the air flow along the ramp of the pamlleled 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-,-,-(63)

25、q = 0.93; 2 (-teaa. (h) I z = 0.6. - A-14018 Figure lo.- Concluded. 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-,-,-(a) iq, = 0.900; 2 = 0.6. (c) Id,-, = 0.92; 2 = 0.8. (b) do = 0.90; f$ a 0.8. A-14020 Figure Ill.- inlet. Shadowgraphs of the air flaw dongthe ramp of the divergend Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-