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Oblique shock

About: Oblique shock is a research topic. Over the lifetime, 6551 publications have been published within this topic receiving 119823 citations.


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Journal ArticleDOI
TL;DR: In this paper, an investigation of the nature of cold gas has been made by an infra-red absorption technique, and the time of arrival of driver gas at the end plate is measured from the absorption records and calculations based on a modification of the shock bifurcation model agree substantially with these experimental results.
Abstract: Following the first interaction between the reflected shock from the end wall and the advancing contact surface in a shock tube, cold gas has been observed at the end plate at a time much earlier than that predicted on the basis of simple shock tube theory. Investigation of the nature of this cold gas has been made by an infra-red absorption technique. The driver gas hydrogen is "tagged" with a small quantity of infrared active gas, and shocks are driven with this gas mixture into an infrared inactive test gas. The time of arrival of driver gas at the end plate is measured from the absorption records and calculations based on a modification of the shock bifurcation model agree substantially with these experimental results. Use is also made of an infrared absorption/emission technique to determine the temperature of the infrared source.

34 citations

Journal ArticleDOI
TL;DR: In this paper, wind-tunnel tests are conducted to study the characteristics of the shock train in the curved variable-section diffuser for hypersonic inlets. But the results show that at an inlet Mach number of 2.59, the base frequency of the oscillations is about 19 Hz and the maximum fluctuating range of the instantaneous surface pressure is as high as 21.6% of the ideal pressure rise.
Abstract: Wind-tunnel tests are conducted to study the characteristics of the shock train in the curved variable-section diffuser for hypersonic inlets. The test model is equipped with a forebody, a contracting entrance, a dump mixing duct, and an aft plug. Tests are performed at nominal freestream Mach numbers of 4.0, 5.0, and 6.0, and the corresponding inlet Mach numbers of the diffuser are 2.05, 2.59, and 3.06, respectively. Results indicate that at an inlet Mach number of 2.05, the surface pressure distributions in the shock train are similar at different backpressure ratios and can be well predicted by the modified Waltrup formula, but the length of the shock train is increased by 32 % due to the curved duct, the incident shock waves, and the incident expansion waves at the inlet plane. At higher inlet Mach numbers, the similarity disappears and the distributions of the surface pressure are not easy to predict. At different measurement points in the shock train, the instantaneous surface pressures vary obviously and almost synchronously, suggesting the oscillatory motions of the shock train. At an inlet Mach number of 2.59, the base frequency of the oscillations is about 19 Hz and the maximum fluctuating range of the instantaneous surface pressure is as high as 21.6% of the ideal pressure rise of the shock train. The static pressure fluctuations at different points in the shock train correlate strongly but are almost unrelated with those of the exit survey point, which indicates that the oscillation of the shock train is not likely induced by the pressure fluctuations far downstream of the shock train.

34 citations

Journal ArticleDOI
TL;DR: In this paper, a hypersonic inlet with side compression was tested at a freestream Mach number of 6.0 to enrich the understandings of the inlet unstart process.
Abstract: A hypersonic inlet with side compression has been tested at a freestream Mach number of 6.0 to enrich the understandings of hypersonic inlet unstart. A flow plug is placed at the duct exit to simulate the combustion induced high pressure and to initiate the inlet unstart. High-speed schlieren imaging and time-resolved pressure measurements are used simultaneously to record the unsteady flow structures and surface pressures of the unstart process. The inlet operates in a big buzz mode with a base frequency of 30 Hz and exhibits a series of unsteady flow patterns similar to those of rectangular hypersonic inlets in a buzz cycle when the throttling ratio is 87.4%. During the upstream propagating process of the unstart shock system, the propagation velocity in the two ends of the duct is higher than that in the middle section of the duct, with a minimum value around the isolator. Once the separation bubble induced oblique shock is expelled over the cowl lip, a supersonic reverse flow with a Mach number of 1.5...

34 citations

Journal ArticleDOI
TL;DR: In this article, the authors studied the evolution of the ultrarelativistic shock wave in a plane-parallel atmosphere adjacent to a vacuum and the subsequent breakout phenomenon and derived the energy spectrum of the ejected matter as a result of the shock breakout.
Abstract: We study the evolution of the ultrarelativistic shock wave in a plane-parallel atmosphere adjacent to a vacuum and the subsequent breakout phenomenon. When the density distribution is a power law in distance from the surface, there is a self-similar motion of the fluid before and after the shock emergence. The time evolution of the Lorentz factor of the shock front is assumed to follow a power law when the time is measured from the moment at which the shock front reaches the surface. The power index is found to be determined by the condition for the flow to extend through a critical point. The energy spectrum of the ejected matter as a result of the shock breakout is derived, and its dependence on the strength of the explosion is also deduced. The results are compared with the self-similar solution for the same problem with nonrelativistic treatment.

34 citations

Journal ArticleDOI
TL;DR: In this paper, an experimental investigation was conducted to delineate the structure of the flowfield and temperature distributions in a shock wave/turbulent boundary-layer interaction with and without surface cooling.
Abstract: An experimental investigation was conducted to delineate the structure of the flowfield and temperature distributions in a shock wave/turbulent boundary-layer interaction with and without surface cooling. The Mach number upstream was about 3.5, and the wave angle was 23 deg. The wall to stagnation temperature ratio was 0.44 with cooling and 1.1 with heating. A detailed map of the interaction flowfields deduced from numerous boundary-layer traversing stations revealed the influence of wall cooling on the flowfield, wave structure, and size of the flow separation region. With surface cooling, the size of the separation region was much smaller, and the separation and reflected shock waves merged together near the edge of the velocity boundary layer, extending into the freestream as one wave.

34 citations


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Performance
Metrics
No. of papers in the topic in previous years
YearPapers
202369
2022142
2021106
202090
201992
2018102