NASA NACA-RM-A8I29-1948 Ram-recovery characteristics of NACA submerged inlets at high subsonic speeds《在高亚音速下NACA嵌入式进气道的冲压恢复特性》.pdf

《NASA NACA-RM-A8I29-1948 Ram-recovery characteristics of NACA submerged inlets at high subsonic speeds《在高亚音速下NACA嵌入式进气道的冲压恢复特性》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-A8I29-1948 Ram-recovery characteristics of NACA submerged inlets at high subsonic speeds《在高亚音速下NACA嵌入式进气道的冲压恢复特性》.pdf（46页珍藏版）》请在麦多课文档分享上搜索。

1、1 -i * t f i RESEARCH MEMORANDUM RAM-RECOVERY CHARACTERISTICS OF NAGA SUBMERGED INLETS AT HIGH SUBSONIC SPEEDS By Charles F. Hall and Joseph L. Frank Ames Aeronautic81 Laboratory Moffett Field, Calif. “ “ “ “ - “-“ “ “ “ “ CummSD- NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON November 17, 1

2、948 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-“ - . By Charles F. Hall and Joseph L, Frank Results are presented of an experimental Investigation of the ratrecovery characteristics for the inlets in other configurations, the peeent report was p

3、repared. The symbole used in thia report and their definitions are aa f OllWS : d inlet depth, inches ,H average total pressure, pounds per eqwe foot H1-Po %-Po - r-ecovery ratio h the height of an area of unit width in which the oomplete loss of fre-tream ram pressure is equivalent to the integrate

4、d lose of the total pressure in unit width of M Mach nulziber L Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-3 - ml m, mass-flow ratio (the ratio of the mass flow through the inlet to the maes flow in the free stream through en area equal to the e

5、ntrance area) P static pressure, pounds per square foot Y increment of boundary-layer thickness, inches a, angle of attack uncorrected for tumel-wall effects (measured relative to the fuselage reference line), degrees 0 free stream 1 duct entrance A complete descriptidn of the model wa gfven in refe

6、rence 3. Briefly, the model (shown in figs. 1 and 2) #a8 patterned to represent a typical hi-peed ffghter drglane. Throughout the teste, a pair of identical inlets WBB used. They were disposed eymmetrically on each side of the fuselage and connected to a conmY3n plenum chamber in the aft part of the

7、 fuselage. The four longitudinal inlet locatione investigated (fig. 2) were at fuselage etatiom 34.25, 42.50, 50.75, and 59.00 corresponded, reepectively, to 16.7 percent of the root chord ahead of, and 8.3, 33.3,- and 58.3 percent of the root chord behind the wing-root leading edge. Dimensions of t

8、he ramp, lip, and boupbry-layer deflectors are shown in figure 3. To determine the effect of boundary-layer thiclmess, the boundary layer along khe fuseae surface was artificially increased from the natural thicknese to medium and thick by roughening the fuselage 5 inches from the nose by msam of sn

9、vrll, nails pmjecting from the surface. The boundary-layer thickness wa8 measured wikh three small rakes, each consisting of 10 total-pressure tubes, Pressure losses and flow rates at the intake were measured with a rake 2.1 inches behind the lip leading edge. The rake consisted of 30 tot eepare- ti

10、on; or to shock waves along the fuselage, in the wing-fueelage juncture, or on the rqs. In refere-3 it YBB indioated that aep mation oouurred at approximately fuselage station 50 at 0.30 Mach number and l2.5O angle of attaok and moved aft to fueelage etatfon 60 at lo angle of attaok aa Ma4h number i

11、mreaeed to 0.875. At low Maoh nmibers, the separation was oaused by poor flow in the wi- fuselage jumture at high angles of attack. At high Mach numbers the separation was due to the large inoreme in the boundary-layer thickness cawed by the shook wave at the wiefuselage junoturs. With the inlets in

12、 the two forward locations, the decrease In ram” reoovery ratio aa Mach nuniber hreaaed ie believed to be due prfmarily to the thickening boundary layer caused by a forward movement of the transition poht with Increasing Reynolds nmber, This effeot wae lndtcated in the motion discueelng the effeote

13、of boundary-layer thickness and also by the fact that the deoreaae of ram-reoovery ratio aa Mach number imreased ma fairly ate- throughout the Wh nMer range. Reference 3 shawed that critical epeede dong the ramg were barsly exoeeded at 0.875 Maoh number with the Inlet8 in the forward looatlon, thue

14、indicating that shock waves on the fuselage or the raq were not the caw8 of the decreaee of ram-recovery ratio. Reference 3 d.60 indicated that it was unlikely that critical speede would be reached on the Fangs of the Inlets at station 42.9 became the egeede in that region wlthout inlet8 were below

15、those in the region of station 34-25. With the inlet6 in the two aft locatio=, much of the pressure 108s can be attributed to the influence of- the bormdary lager. For example, when the boundary layer became thick and eeparated A-om the surface, pressure 108feB greater than fie-tream ram preseura we

16、re obtained at subcritical speeds with the inlets in the aft loca- tion. (See fig. 18 for results at Eo angle of attack and a Mach number of 0.a for which ooditlone reference 3 indicated sub- critical speeds and a tuck, poseibly separated, boundary layer on the fuselage surface without inlets.) For

17、oonditione having a similar boundary-layer growth at supercritical speede, it le believed that laxge loeeee aleo would be c8med primarib by the thick boundarg layer. (See fige. 6 and 8 for results at the hlgheet wee of attack at Mach nMere of 0.70 and 0.80.) When the Provided by IHSNot for ResaleNo

18、reproduction or networking permitted without license from IHS-,-,-9 . boundary layer on the fuselage $id not thicken, 88 indicated by the boundary-layer data obtdned withait inlets (fig. 61, some of the losses might be attributed to boundary-layer and shockaave Inter- action on the ramp. For example

19、, in figure 8 the results snow that the increase in losses with angle of attaok at high Mach-nuuibers was larger at 0.a than 0.80 mass-flow ratio. This characteristic wa8 probably due to the interaction of the shock wave and the thicker boundary layer on the ramg caused by the mre adverse pressure g

20、r- dient at 0.60 mass-flow ratio because the shock waves on the ramp were probably weaker at 0.60 mas-f low ratio. Reference 3 showed increase in static pressure from the point of minimum pressure to e the inlets wa larger asd the meximum airspeeds were lower at 0.60 than 0.80 mas-flow ratio. The ef

21、fect of the boundary layer in the presence of shock waves would be less severe with a thinner boundary lager at the beginning of the ramp. This effect, together with .the faL.ecovery ratio for the inletrs in the forward 100- tions (16.7 peroent of the root chord ahead and 8.3 peroent of the root cho

22、rd aft of the wing-root leading edge) raried only slightly aa Wh nuniber increased. ITor the two aft locations (33.3 and 58.3 peroent of the root chord af% of the wlngeoot leading edge) large deareases in rausrecoverg ratfo ocourred at high Wh Ilumbera and angles of attaak above 2 . 2. The highest r

23、wecovery ratios were obtained with the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA FU4 no. 829 11 inlets in the forward location. 3. kcreasing the boundary-layer thiclmess decreased the r- recovery ratio. 4. In general, the ram-recovery rati

24、o decreased with inoreasing angle of attack. b 5. With 110 deflectors on the ramp the r-ecovery ratio inoreased greatu aa mass-floW ratio increased to approzlmately o .a, reached a maximMl betweep 0.60 and 0.80 mass-flaw ratio, and slowly decreased for greater flow rates. 6. “he boundary-layer defle

25、ctors increased the meximum r8m- recovery ratio and the mass-flow ratio at which it occurred. They redwed the r-ecovery ratio between approximately 0.40 and 0.70 mass-flow ratio aad also reduced the change in rmcovery ratio with angle of attack for inlets in the two forwmd locations. Ames Aeronautic

26、al Laboratory, National Advisory Connittee for Aeronautios, Ibffett Field, Calif. KEKEREKES 1. Frick, Charles W., Davis, Wallace F., Randall, Lauros M., and bbssman, Emmet A.: An Erperimental Investigation of mAcA Submerged-Duct Entrances. WA ACR No. 5I20, 1945. 2. nafurd baundary layer. Provided by

27、 IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-./ .l 0 0 I .2 .I 0 0 20 40 60 b 0 20 40 60 .2 I 0 0 PO 40 60 ,2 .2 J ./ 0 0 0 204060 0 2040a Fuselage Natural boundury layer sfafiim, inches from nose Medium Thk4 .5 .6 .7 :8 .9 I. .9 B -7 .3 .4 .5 .6 .7 .8 .9 Ma

28、ch number, M C.“ “.“ Figwe X - Effect of buundury-/uyer thickness on ram-recovety rafio for 0.70 mass-fhw ratio. fnlef of fuselage station 34.25; no deflecfors UR ramp. .-.$AU Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EACA RM k. A8129 23 -lo 9

29、.8 .3 .4 . c Ang/e of offack, a nafurd bwndary hyer. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 10 9 .8 . “ . . . .- 9 /n/ef posihn .8 mse. sf0 -3425 “4250 “50.75 “5900 .7 .6 -3 0 .3 .4 .5 .6 .7 .8 .9 Angle of uffuck, cu, deg Much number, a t

30、bl mt/t?e , 0.80 -?7 Fiwm 8. - Conchded. I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-(b) Fusehge sfot/on 4 2.50 MUSS-OW rotio, - 4 /c/fuse/oge stuf/on 50.75 (dl Fusdage station 5900 m, v Figure 9. - Effecf qf boundor -/ 8r deflectors on ram-rec

31、overy rofm for 0. ?6 %cb number. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-medwm boundary /ayer. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.?. P a% -2“ O0 /= 2“ 3“ 4“ 6O Provided by IHSNot

32、 for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.r 0 Moss-f/ow GI -2“ 0 0“ A /“ v 2“ 3“ 6“ a 40 Figure /2 - Conduded Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-