Freestream

About: Freestream is a research topic. Over the lifetime, 3428 publications have been published within this topic receiving 56147 citations.

Direct simulation Monte Carlo calculation of a jet interaction experiment

Abstract: This study deals with the development of a methodology for numerically simulating the interaction of a reaction control system (RCS) jet with a low-density external flow. A European Space Agency experiment was chosen as a test case, since it provided experimental data that could validate some of the numerical results. The initial approach was to focus on several subproblems having direct relevance to the full interaction problem. This enabled different numerical methods to be investigated separately and validated for each part of the interaction problem. In this manner, the best methodology for solving the full interaction problem was developed. The subproblems considered in this study included typical RCS nozzle and plume flows, a flat plate at zero incidence, and the flow past the experimental test model without the control jet firing. Once these calculations were completed, a simulation was performed of the test model with the control jet operating at the experimental density. The results from this final simulation were compared with experimental measurements. Nomenclature = reference diameter, m = Knudsen number at nozzle throat = Knudsen number at nozzle lip = Mach number = number density, molecules/m3 = freestream number density, molecules/m3 = pressure, Pa = Reynolds number T^ef = reference temperature, K 7"waii = temperature of nozzle wall T^ = freestream temperature, K X, Y, Z = Cartesian body coordinates p , - density, kg/m3 Poo = freestream density, kg/m3 Kn\ip M

Shock Wave - Film Cooling Interactions in Transonic Flows

Abstract: Interactions between shock waves and film cooling are described as they affect magnitudes of local and spanwise-averaged adiabatic film cooling effectiveness distributions. A row of three cylindrical holes is employed. Spanwise spacing of holes is 4 diameters, and inclination angle is 30 degrees. Freestream Mach numbers of 0.8 and 1.10–1.12 are used, with coolant to freestream density ratios of 1.5–1.6. Shadowgraph images show different shock structures as the blowing ratio is changed, and as the condition employed for injection of film into the cooling holes is altered. Investigated are film plenum conditions, as well as perpendicular film injection cross-flow Mach numbers of 0.15, 0.3, and 0.6. Dramatic changes to local and spanwise-averaged adiabatic film effectiveness distributions are then observed as different shock wave structures develop in the immediate vicinity of the film-cooling holes. Variations are especially evident as the data obtained with a supersonic Mach number are compared to the data obtained with a freestream Mach number of 0.8. Local and spanwise-averaged effectiveness magnitudes are generally higher when shock waves are present when a film plenum condition (with zero cross-flow Mach number) is utilized. Effectiveness values measured with a supersonic approaching freestream and shock waves then decrease as the injection cross-flow Mach number increases. Such changes are due to altered flow separation regions in film holes, different injection velocity distributions at hole exits, and alterations of static pressures at film hole exits produced by different types of shock wave events.Copyright © 2001 by ASME

Numerical Study of Hypersonic Boundary-Layer Receptivity with Freestream Hotspot Perturbations

Abstract: This paper presents a numerical-simulation study of transient flow over a blunt compression cone under the effect of freestream hotspot perturbations. This study is motivated by concurrent wind-tunnel laser-spot experiments carried out at Purdue University. The flow conditions used in the simulation are based on the experimental conditions. The simulation is performed using a high-order shock-fitting finite-difference scheme. The simulation results show that the hotspot is able to excite second-mode instability, where the instability growth is found to be dominant in the boundary layer. The receptivity mechanism is investigated by comparing the simulated results with linear-stability theory. Fast acoustic waves generated by hotspot–shock interaction excite the boundary-layer disturbances. Also, the synchronization of mode F and mode S leads to the dominance of boundary-layer disturbances by the growing second mode.

Laminar boundary-layer separation over a circular cylinder in uniform shear flow

Abstract: By appealing to the classical boundary-layer theory, the present paper investigates the effect of the freestream shear on the separation of the laminar boundary layer around a circular cylinder. It is shown that on the side of the cylinder with faster freestream velocity the location of the separation point (point of vanishing wall shear) is virtually unaffected by the freestream shear, while on the other side of the cylinder a critical shear rate is observed. Below this critical value, separation occurs typically at the rear surface of the cylinder and is found to shift towards the downstream direction with increasing freestream shear. Above the critical shear rate, the boundary layer separates from the windward side of the cylinder. Further increase of the freestream shear then causes the separation point to move towards the upstream direction. The present findings may have important implication on the issue regarding to the orientation of the lift force exerting on the cylinder.