Effects of boundary-layers bleeding on unstart/restart characteristics of hypersonic inlets

Bao

Advances in Mechanical Engineering

2014

Supersonic film cooling is an efficient method to cool the engine with extremely high heat load. In order to study supersonic film cooling in a real advanced engine, a two-dimensional model of the hypersonic inlet in a scramjet engine with supersonic film cooling in the isolator is built and validated through experimental data. The simulation results show that the cooling effect under different coolant injection angles does not show clear differences; a small injection angle can ensure both the cooling effect and good aerodynamic performances (e.g., flow coefficient) of the hypersonic inlet. Under selected coolant injection angle and inlet Mach number, the cooling efficiency increases along with the injection Mach number of the coolant flow, only causing a little total pressure loss in the isolator. Along with the increase of the inlet Mach number of the hypersonic inlet, the cooling efficiency does not present a monotonic change because of the complex shock waves. However, the wall temperature shows a monotonic increase when the inlet Mach number increases. The mass flow rate of coolant flow should be increased to cool the engine more efficiently according to the mass flow rate of the main stream when the inlet Mach number increases.

Section: Physical Model and Computationsmentioning

confidence: 99%

Section: Physical Model and Computationsmentioning

confidence: 99%

Numerical Analysis of Supersonic Film Cooling in Supersonic Flow in Hypersonic Inlet with Isolator

Bao

Advances in Mechanical Engineering

2014

Topical Problems of Fluid Mechanics 2016

“…In this paper, we consider the flow in the channel with two expansion corners and with a central body. It should be noted that the problems of hysteresis of shock waves were studied in [4,5].…”

Section: Introductionmentioning

confidence: 99%

Bifurcation of Transonic Flow in a Channel With a Central Body

Рябинин¹,

Suleymanov²

2016

Abstract2D transonic flow in a channel of variable cross-section with a central body is studied numerically using a solvers based on the Euler and Reynolds-averaged Navier-Stokes equations. The flow velocity is supersonic at the inlet and outlet of the channel. Between the supersonic regions, there is a local subsonic region. Computations reveal a hysteresis in the shock position versus the inflow Mach number. In the certain range of inlet Mach numbers there are asymmetrical solutions of the equations.

Sci. China Phys. Mech. Astron.

“…These considerable benefits provided by bleed have led a number of investigators to explore its potentials in hypersonic devices both experimentally and numerically [9][10][11][12][13][14][15][16][17][18][19][20][21].…”

Section: Introductionmentioning

confidence: 99%

“…Häberle and Gülhan [11,12] deployed a passive boundary layer bleed at the throat of a fixed geometry hypersonic inlet to significantly diminish the lip shock-induced separation bubble on the ramp and reduce the risk of inlet unstart with 5% mass flow rate penalty. Falempin et al [13], Chang et al [14] and Pandian et al [15] also adopted perforation bleed as an efficient method to facilitate inlet starting and improve their overall characteristics. Herrmann et al [16] experimentally investigated the effects of different boundary layer bleed systems on the performance of a two-dimensional ramjet inlet and identified that the boundary layer bleed could significantly increase the inlet performance and stabilize the subcritical condition (i.e., later buzzing).…”

Section: Introductionmentioning

confidence: 99%

Aerothermal characteristics of bleed slot in hypersonic flows

Yue

et al. 2015

Two types of flow configurations with bleed in two-dimensional hypersonic flows are numerically examined to investigate their aerodynamic thermal loads and related flow structures at choked conditions. One is a turbulent boundary layer flow without shock impingement where the effects of the slot angle are discussed, and the other is shock wave boundary layer interactions where the effects of slot angle and slot location relative to shock impingement point are surveyed. A key separation is induced by bleed barrier shock on the upstream slot wall, resulting in a localized maximum heat flux at the reattachment point. For slanted slots, the dominating flow patterns are not much affected by the change in slot angle, but vary dramatically with slot location relative to the shock impingement point. Different flow structures are found in the case of normal slot, such as a flow pattern similar to typical Laval nozzle flow, the largest separation bubble which is almost independent of the shock position. Its larger detached distance results in 20% lower stagnation heat flux on the downstream slot corner, but with much wider area suffering from severe thermal loads. In spite of the complexity of the flow patterns, it is clearly revealed that the heat flux generally rises with the slot location moving downstream, and an increase in slot angle from 20° to 40° reduces 50% the heat flux peak at the reattachment point in the slot passage. The results further indicate that the bleed does not raise the heat flux around the slot for all cases except for the area around the downstream slot corner. Among all bleed configurations, the slot angle of 40° located slightly upstream of the incident shock is regarded as the best.