An experimental investigation was conducted to determine the effect of bleed on a shock waveboundary layer interaction in an axisymmetric mixed-compression supersonic inlet. The inlet was designed for a free-stream Mach number of 2.50 with 60-percent supersonic internal area contraction. The experiment was conducted in the NASA Lewis Research Center 10-by 10-Foot Supersonic Wind Tunnel. The effects of bleed amount and bleed geometry on the bounduy layer after a shock wave -boundary layer interaction were studied. The effect of bleed on the transformed form factor Htr is such that the full realizable reduction in Htr is obtained by bleeding off a mass flow equal to about one-half of the incident boundary layer mass flow. More bleeding does not yield further reduction in Htr. Bleeding upstream o r downstream of the shock-induced pressure rise is preferable to bleeding across the shock-induced pressure rise since, for the latter, bleed flow from the high pressure side of the interaction can detrimentally reenter the boundary layer on the upstream low pressure side. Differences between upstream and downstream bleeding are detectable but slight and do not seem significant. Slanted holes yield lower values of Htr and of compressible displacement and momentum thicknesses than normal holes. Two bleed hole sizes were tested, and no difference in performance was detected.
2.7aForward Centerbody Bleed Region (Large S l a n t e d Holes)2.7b Forward Centerbody BleedRegion (Small S l a n t e d Holes) 2.8c Cowl Bleed Region ( S l a n t e d Holes)2.9 Schematic of T r a v e r s i n g T o t a l P r e s s u r e Probe
t a t i c P r e s s u r e D i s t r i b u t i o n 81 Run 5 Centerbody S t a t i c P r e s s u r e D i s t r i b u t i o n 82Run 6 Centerbody S t a t i c P r e s s u r e D i s t r i b u t i o n 8 3Run 10 Centerbody S t a t i c P r e s s u r e D i s t r i b u t i o n
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Schematic of Centerbody Bleed Regions 85
Comparison of Run 2 Centerbody S t a t i c P r e s s u r e D i s t r i b u t i o n w i t h t h e Method of C h a r a c t e r i s t i c S o l u t i o n
86Probe 1
Effects of nozzle scale, total temperature, and an afterburner on jet noise characteristics from a pre-cooled turbojet engine are investigated experimentally. In JAXA (Japan Aerospace Exploration Agency), a pre-cooled turbojet engine for an HST (Hypersonic transport) is under development. In the present study, 1.0%-and 2.4%-scaled models of the rectangular plug nozzle (Nozzles I and II) are manufactured, and the jet noise characteristics are investigated under a wide range of total temperatures. For Nozzle I, no air-heater is utilized and the total temperature is 290 K. For Nozzle II, a pebble heater and an afterburner (AB) are utilized upstream of the nozzle model, and the total temperature is varied from 520 K (pebble heater) to 1540 K (pebble heater + AB). The total pressure is set at 0.27 and 0.30 MPa(a) for both nozzle models. Jet noise is measured using a high-frequency microphone set at 135 deg from the engine inlet, and normalized jet noise spectra are obtained based on AU n j law and Helmholtz number. For cases without afterburner, the normalized spectra agrees well regardless of the nozzle scale and total temperature where the velocity index lies from n = 7.7 to 9.2, and the correlation factor between the two facilities is shown to be about 1 dB. For the case with afterburner, the normalized spectrum does not agree with other conditions where the velocity index n seems to be about 4.
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