1、RESEARCH MEMORANDUM THE EFFECT OF LIP SHAPE ON A NOSE-INLET INSTALLATION AT MACH NUMBERS FROM 0 Td 1.5 AND A METHOD FOR OPTIMIZING ENGINE-INLET COMBINATIONS By Emmet A. Mossman and Warren E. Anderson NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON May 7, 1954 Provided by IHSNot for ResaleNo r
2、eproduction or networking permitted without license from IHS-,-,-NACARMA54BO8 r AERONACS THEEFFECTOFIJR SHARE ONANOSE-INIETINS!PKLAITON ATMACHNUMHERSFROMO TO 1.5ANDAMETHODFOR OPTlXC!ZING ENGINE-INLET COMBINATIONS By Emmet A. Mossman and Warren E. Anderson An experimental investigation was made at su
3、bsonic, transonic, and supersonic speeds of the effect of lip shape on the drag, pressure recov- ery, and mass flow of a nose-inlet air-induction system. Four lips of varying degrees ofbluntness were testedonafuselage modelatMach numbers of 0 to 1.5 and at angles of attack of O“ to 120. In general,
4、blunting the lip increased the pressure recovery at all the speeds of this test. The improvement in pressure recovery due to rounding the lip was small at supersonfc and at high subsonic speeds, but resulted in marked improvement at the take-off condition. At supersonic speeds in the mass-flow-ratio
5、 range of normal operation (0.8 to maximum), going from a sharp lip to a slightly rounded lip had no significant effect on the drag. However, a more blunt lip, typical of a subsonic design, resulted in a considerable increase in drag. The rate of change of drag coefficient with mass-flow ratio was b
6、est predicted, in the supersonic speed range, by the theory of Fraenkel, An analysis was made by combinfng the pressure recovery and drag force into a single parameter (an effective drag coefficient), and by matching the inlet air flow with an assumed engine air flow. This analytical study showed li
7、ttle difference in the effective drag coeffi- cient for the sharp and slightly rounded lip shapes at supersonic speeds. It was indicated that these Snlets can operate efficiently over a wide ra.nge of mass-flow ratios at the supersonic speeds investigated, thus simplmng the engine-inlet matching on
8、this particular instsXation. From the standpoint of higher pressure recovery at t the leakage a-Lr flow through the sealwas calibratedand amounted to from 0.5 to 2.0 percent of the total air flow. Pertinent corrections were made. The pressures at the simulated compressor inlet were measured by a rak
9、e of 20 total pressure tubes and 2 static pressure tubes, and the pressures at the model exit were measured simultaneously by a rake of 20 total pressure tubes and 4 static pressure tubes (see fig. 1). Model base pressures were measured at 12 points. A three-component strain- gage balance inside the
10、 model was used to measure the forces, Tests were made for a range of mass-flow ratios from 0 to a maximum, angles of attack up to l2O, and Mach numbers of 0, 0.7, 0.8, 0.9, 1.23, 1.35, and 1.50. I$xcept for the statfc tests (M. P 0), all experiments were made with a constant tunnel stagnation press
11、ure of I.2 pounds per square inch absolute. The CorrespondingReynolds nuxiberper foot varied I-OBL 3.1310 to 3.8210% In the reduction of the data, the forces developed by the internal flow and the base forces were subtracted from the balance measured values. The internal-flow force is defined as the
12、 change in total momen- tum of the entering stream tube from the free stream to the exit of the model, and is thus consistent with.the usual definition of jet-engine thrust. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 RESULTS NACA RM A54BO8 A c
13、omparison of the pressure-recovery characteristics for the four lip shapes is given in figures 6, 7, and 8. The pressure recovery at O“ angle of attack for simulated take-off (Mo = 0), high-subsonic-speed (MO = 0.7, 0.8, 0.91, and supersonic-speed (Mo = 1.23, 1.35, 1.50) opera- tion is presented in
14、figures 6 and 7. The variation of pressure recov- ery for three of the lip shapes (lips 2, 3, and 4) tith angle of attack is shown in figure 8 for Mach numbers of 0.7 and 0.9. For the angle-of- attack range investigated at supersonic speeds (O“ to 50) there was no significant change in the pressure
15、recovery; at O“ are presented. consequently, only the data The results of the drag measurements are shown in figures 9 and 10. It will be noted that drag coefficients for supersonic speeds only are shown since the drag measurements at subsonic. speeds were inconsistent and were considered unreliable
16、. The variation of the drag coefficient with ml/nation (P c - Pot) dA 0 . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACARMA5 however, an accurate method for csl- culating the pressure drag is unavailable. lhalysfs To compare these individual li
17、p-shapes requires that their internal characteristics (pressure recovery) be related to an appropriate engine, and that their external characteristics (drag) be related to an assumed airplane. The J-57 jet engine has been chosen, and a wing area of 376 squake feet has been assumed for the airplane.
18、The performance com- parison of-the lip shapes is made by combining the drag force and the pressure recovery into a single parameter. In this analysis, the loss in pressure recovery Ptc i.e., l- is converted to a thrust loss pto . . through use of a vsLue of 1.21 for the factor mbisen upto , this va
19、lue being appropriate for the J-57 jet engine for the assumed flight conditions. This thrust loss is then combined with the drag to give an effective drag coefficient based on wing area. The inlets must also be compared at their actual operating points. At the operating (or %atched“) condition, the
20、air supplied by the inlet must be equal to the ak required by the engine. The operating mass- flow ratios and the corresponding pressure recovery and drag coefficients were obtained for each inlet at several assumed inlet areas. Themethod used is outlined in Appendix B, and typical curves for one li
21、p shape are given in figure 13. It should be mentioned that as the inlet area is reduced, the body pressure drag is increased slightly and the pre-entry drag is decreased. These effects on the drag are not included; but for the range of inlet areas of the analysis (+a, = 520 to 640 sq In,), a study
22、of the body pressure forces by the method of reference 9 and con- sideration of the pre-entry drag show these force changes to be neglf- glble, and also of opposite sign. The results of the analysis at super- sonic speeds are given for each lip in terms of an effective drag coef- ffcient based on wi
23、ng area (figs. 14 and 15). From figure 14 it can be seen that a subsonic lip shape (lip 4) results in a considerable effective drag penalty at supersonic Mach num- bers. Rounding of the lip and some internal contraction (lip 3) is beneficial. Because there is only a small variation in CD tith inlet
24、sea for most of the lip shapes, it can be said that the performance of the inlets IS not sensitive to changes in inlet area from 540 to 640 square inches. In figure 13 it is shown that an inlet area of 5, the cP,R. I., Merlet, C. F., and Putland, L. W.: Flfght Determina- tion of Drag of Normal-Shock
25、 Nose Inlets with Various Cowling Profiles at Mach Numbers from 0.9 to 1.5. NACA RM L5p25a, 1953. 8. Frick, Charles W., and Olson Robert N.: Flow Studies in the Asym- metric Adjustable Nozzle of the Ames 6- by B-Foot Supersonic Wind Tunnel. NACA RM AgE24, 1949. 9. Brajnikoff, George B., and Rogers,
26、Arthur W.: Characteristics of Four Nose Inlets as Measured at Mach Numbers Between 1.4 and 2.0. NACA RM A51Cl2, 195l. 10. Sibulkin, Merlin: Theoretical. and Expertiental Investigation of Additive Dr8g. NACA RM E5lB13, 1951. ll. Fkaenkel, L. E.: The External Drag of Some Pit a = O“ Provided by IHSNot
27、 for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 54x18 .96 t? P a .84 .76 .72 / ! sr (b) I+, = 0.80 Figure I.- Continued. 27 UP 2 ,3 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 NACARMAsBO8 .96 .84 .76 .
28、72 -2 4 .,6 .8 Mass- ffow mtlo, m,/m, (4 M, = 0-w Figure 7.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-e .96 .92 f Q St- .88 2 h 3 .84 h $ E! .80 .76 .72 I 0 .2 4 .6 .8 /.u Mass-f 10 w r otio, m/m, (d) M, = 1.23 Figure 7.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
