1、RESEARCH MEMORANDUM SOME OBSERVATIONS OF SHOCK-INDUCED TURBULENT SEPARATION ON SUPERSONIC DIFFUSKRS By T. J. Nussdorfer Lewis Flight Propulsion Lab Cleveland, Ohio Fay REFERENCE NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON May 4, 1954 Provided by IHSNot for ResaleNo reproduction or network
2、ing permitted without license from IHS-,-,- “ - NACA RM E5lL26 .“ “ . By T. J. Nussdorfer SUMMARY A survey of experimental data at supersonic speed indicates that shock-induced separation of a turbulent boundary layer will result for Mach numbers of approxrntely 1.33 or greater when a theoretical st
3、ream static-pressure-rise ratio of approximately 1.89 OCCUTS across a shock interacting with the boundary layer. The si the exact configuration is dependent upon any one of the deflect ions a, b, or c h the system. The action of the boundary layer is apparently the deter- mining factor in the orient
4、ation of the branched shock and concamitant 8 separation. “. . - “. - -. . . -. . - Exgerwtd reports on linear expamion nozzles (ref. 4 and unavail- a Ln “ - able reports) indicated that when a boundary layer was present the branched shock occurred for Mach nunibers greater than about 1.35 to 1.4, d
5、ependbg upon the nozzle expansion angle. For Mach numbers less than these, a normal shock without separation was observed. Therefore, it appears that the existence of boundary-layer separation is dependent upon the stream static-pressure-rise ratio. The work reported in relerence 4 is for turbulent
6、boundary layers. Fram the results of reference 5, a marked difference in the type of separation and point of separation should be expected between turbulent and laminar boundary mers. Inasmuch as turbulent mixing is much more effective than molecular mixing in transferrhg momentum within a boundary
7、layer, separation would be expected for a lamlnar boundary layer for smaller value6 of pressure rise than that required for a tur- bulent boundary layer. Extension of Gruschwitz calculations to cover separation in transonic flow with shocks is included in reference 6. A more complete discusaion of s
8、eparation is given in reference 7. In the absence of a theoretical explanation of shock-induced sqar- ation of a turbulent boundary layer, an engineering criterion obtained from a survey of experimental data has been deduced This report; pre- sents the tentative criterion, which relates separation o
9、r nonseparation of the boundary layer to the theoretical static-pressure-rise ratio across an imposed shock. The significance of the criterion is discussed with regard to stzpersonic diffusers for ram-jet and turbojet engine application. The criterion presented in this report was developed at the NA
10、CA Lewis laboratory in 1951, but publication -was withheld at that time be- cause of pwallel studies presented in reference 8. The fnformatl6n * contained in reference 8 has since been superseded by reference 9. The recent work of reference 10, which includes different criteria for predicting shock-
11、induced boundary-layer separation from those of reference 9, supports the conclusions presented herein. Release of this paper in substantially the original form is, therefore, considered appropriate. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NA
12、CA RM E5lL26 II c 3 DISCUSSICffS Ln this report separation was distinguished by the presence of a branched shock. Separation was most eas- recognized frm a schlieren or interferometer photopph, but velocity and total-pressure profiles and static pressures in the region of the boundary layer were als
13、o use- ful. Most of the data presented (ref s. ll to 14) were obtained frcan studies on supersonic diffuser inlets. Investigation of these inlets over a range of stream Mach nunibers provides a convenient method of studying the interaction of shocks of varyhg strength upon the boundary layer. The fi
14、rst inlets studied were of the two-dhemional ramp type where the angle X which the ranrp makes with the free stream adequately desoribes the inlet for this study. For a given free-stream Mach nuuiber, a theoretical static-pressure-rise ratio across the normal shock may be obtained for any given ramp
15、 . Rep. 9961-12, Aug., SeTt., and Oct. 1950, Aero. Lab., Univ. Southern Cal., Nov. 7, 1950. Navy Contract Hoa(s) 9961.) 16. Dailey, C. L. : Diffuser Instability in Subcritical Operation. Univ. Southern Cal., Sept. 26, 1950. 17. Moeckel, W. E.: Flow Separation Ahead of Blunt Bodies at Supersonic Spee
16、ds. NACA lIIN 2418, 1951. 18. Moeckel, W. E. : Flow Sepaxation Ahead of a Blunt Axially Symmetric Body at Mach numbers 1.76 to 2.10. MclCA RM E5lI25, 1951. - 19. Moeckel, W. E.; and Evans, P. J., Jr. : Prelimhary Investigation of Use of Conical Flow Separation for Efficient Supersonic Diffusion. - N
17、ACA RM E51J08, 1951. 20. Nitzberg, Gerald E., and Crandall, Stewart: A Study of Flow Changes Associated with Airfoil Section Drag Rise At Supercritical Speeds. HACA TN 1813, 1949. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-* (b) Curved shock.
18、NACA RM E5lL26 * (c) Branched shock. Q527 Figwe 1. - Types of shock interacting with boundary layer in superBonic flow (refs. 1 and 2). Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM E5lL26 - 9 35 30 26 M d 20 .% x 0) - 2 k? P; 10 5 0 1.0 1.
19、4 1.8 2.2 2.6 3.0 Stream Mach number Figure 2. - Relation of ranq angle, Mach number, and theoretical static- pressure-rise ratio across normal shock on two-dimensional inlets. Ratio of specific heats, 1.4. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS
20、-,-,-10 - NACA RM E5m6 Stream Mach number Figure 3. - RelatioCOf.Acoiie half aiig1e;Mach number, and theoretical static- pressure-rise ratto across normal shock on cone surface of three-dimensional conical inlets. Ratio of specific heats, 1.4. . - ., Provided by IHSNot for ResaleNo reproduction or n
21、etworking permitted without license from IHS-,-,-NACA RM E5Z26 11 (a) Curved shock; Wch number, 1.YY; point A af figure 2. (b) Branohed ehock; Mach number, 1.57; p0-t B of figure 2. (a) Branched shock3 Msch nmber, 1.83; point C of flgure 2. Figure 4. - Shock ptteny on 6 ramp. Two-dimensional inlet.
22、I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- NACA RM E51 L26 1.0 . 1.2 1.4 - 1.6 1.8 - 2.0 Mach rlumber ahead of shock interact- with boundary layer Figure 5. - Cmelatlon of thaoretloal static presaure rise ratio and Mach number with shock-indu
23、cedseparatlon; ratio of specific heats, 1.4. . “ ,. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA 3iM E5SL26 - 4 a 24 4 12 20 28 36 44 52 Cone half angle, Bc, deg Figure 6. - Theoretical pressure recovery for various two- and three-dimensional
24、 inlets. Ratio of specific heats, 1.4. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 - NhCA RM E5lL26 NACA-LUIR - 5-4-54 - 360 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-t Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
