4,372 research outputs found
On upstream influence in supersonic flows
The general problem of propagation of three-dimensional disturbances in viscous supersonic flows is considered in the framework of characteristic analysis. Unlike previous results for linear disturbances we deduce a condition determining nonlinear characteristic surfaces which is exact and therefore allows both qualitative and quantitative studies of the speed of propagation as a function of various physical phenomena. These include negative and adverse pressure gradients, and effects of wall cooling and suction–blowing, which are studied in this work as an illustration of the general theory
Pressure Bifurcation Phenomenon on Supersonic Blowing Trailing Edges
Turbine blades operating in transonic-supersonic regime develop a complex
shock wave system at the trailing edge, a phenomenon that leads to unfavorable
pressure perturbations downstream and can interact with other turbine stages.
Understanding the fluid behavior of the area adjacent to the trailing edge is
essential in order to determine the parameters that have influence on these
pressure fluctuations. Colder flow, bled from the high-pressure compressor, is
often purged at the trailing edge to cool the thin blade edges, affecting the
flow behavior and modulating the intensity and angle of the shock waves system.
However, this purge flow can sometimes generate non-symmetrical configurations
due to a pressure difference that is provoked by the injected flow. In this
work, a combination of RANS simulations and global stability analysis is
employed to explain the physical reasons of this flow bifurcation. Analyzing
the features that naturally appear in the flow and become dominant for some
value of the parameters involved in the problem, an anti-symmetrical global
mode, related to the sudden geometrical expansion of the trailing edge slot, is
identified as the main mechanism that forces the changes in the flow topology.Comment: Submitted to AIAA Journa
Stabilization of Hypersonic Boundary Layers by Porous Coatings
A second-mode stability analysis has been performed for a hypersonic boundary layer on a wall covered by a porous coating with equally spaced cylindrical blind microholes. Massive reduction of the second mode amplification is found to be due to the disturbance energy absorption by the porous layer. This stabilization effect was demonstrated by experiments recently conducted on a sharp cone in the T-5 high-enthalpy wind tunnel of the Graduate Aeronautical Laboratories of the California Institute of Technology. Their experimental confirmation of the theoretical predictions underscores the possibility that ultrasonically absorptive porous coatings may be exploited for passive laminar flow control on hypersonic vehicle surfaces
Investigation of nose bluntness and angle of attack effects on slender bodies in viscous hypersonic flows
Hypersonic flows over cones and straight biconic configurations are calculated for a wide range of free stream conditions in which the gas behind the shock is treated as perfect. Effect of angle of attack and nose bluntness on these slender cones in air is studied extensively. The numerical procedures are based on the solution of complete Navier-Stokes equations at the nose section and parabolized Navier-Stokes equations further downstream. The flow field variables and surface quantities show significant differences when the angle of attack and nose bluntness are varied. The complete flow field is thoroughly analyzed with respect to velocity, temperature, pressure, and entropy profiles. The post shock flow field is studied in detail from the contour plots of Mach number, density, pressure, and temperature. The effect of nose bluntness for slender cones persists as far as 200 nose radii downstream
A review of instability and noise propagation in supersonic flows
Originally analytical and numerical models were to be developed for noise production in supersonic jets, wakes and free shear layers. While the effort was concentrated initially on these aspects, other topics were also pursued, most were of interest to the Jet Noise Group of the Aeroacoustics Branch. An overview is given of subjects reviewed and the investigations that were carried out. A significant effort was devoted to numerically predicting the flow field of a turbulent supersonic wall jet. This information is necessary for computing the pressure in the far field. The wall jet was selected because it represents a generic flow that can be associated with plug nozzle in supersonic engines. It combines the characteristic of a boundary layer with that of a free shear flow. The spatially evolving flow obtained using Dash's code would form the input for the stability analysis program. This analysis would determine the large scale instability wave within the flow. The far field pressure can be computed from the shape of the evolving large scale structure by asymptotic methods. Flow characteristics obtained from a program that analyses the turbulent downstream supersonic flow in a nozzle are described and compared with experimental results
Effect of Mach number on the structure of turbulent spots
Direct numerical simulations have been performed to study the dynamics of isolated turbulent spots in compressible isothermal-wall boundary layers. Results of a bypass transition scenario at Mach 2, 4 and 6 are presented. At all Mach numbers the evolved spots have a leading-edge overhang, followed by a turbulent core and a calmed region at the rear interface. The spots have an upstream-pointing arrowhead shape when visualized by near-wall slices, but a downstream-pointing arrowhead in slices away front the wall. The lateral spreading of the spot decreases substantially with the Mach number, consistent with a growth mechanism based on the instability of lateral shear layers. Evidence for a supersonic (Mach) mode substructure is found in the Mach 6 case, where coherent spanwise structures are observed under the spot overhang region
The Numerical Simulation of Radiative Shocks I: The elimination of numerical shock instabilities using a localized oscillation filter
We address a numerical instability that arises in the directionally split
computation of hydrodynamic flows when shock fronts are parallel to a grid
plane. Transverse oscillations in pressure, density and temperature are
produced that are exacerbated by thermal instability when cooling is present,
forming post--shock `stripes'. These are orthogonal to the classic post--shock
'ringing' fluctuations. The resulting post--shock `striping' substantially
modifies the flow. We discuss three different methods to resolve this problem.
These include (1) a method based on artificial viscosity; (2) grid--jittering
and (3) a new localized oscillation filter that acts on specific grid cells in
the shock front. These methods are tested using a radiative wall shock problem
with an embedded shear layer. The artificial viscosity method is unsatisfactory
since, while it does reduce post--shock ringing, it does not eliminate the
stripes and the excessive shock broadening renders the calculation of cooling
inaccurate, resulting in an incorrect shock location. Grid--jittering
effectively counteracts striping. However, elsewhere on the grid, the shear
layer is unphysically diffused and this is highlighted in an extreme case. The
oscillation filter method removes stripes and permits other high velocity
gradient regions of the flow to evolve in a physically acceptable manner. It
also has the advantage of only acting on a small fraction of the cells in a two
or three dimensional simulation and does not significantly impair performance.Comment: 20 pages, 6 figures, revised version submitted to ApJ Supplement
Serie
On the linear stability of compressible plane Couette flow
The linear stability of compressible plane Couette flow is investigated. The correct and proper basic velocity and temperature distributions are perturbed by a small amplitude normal mode disturbance. The full small amplitude disturbance equations are solved numerically at finite Reynolds numbers, and the inviscid limit of these equations is then investigated in some detail. It is found that instability can occur, although the stability characteristics of the flow are quite different from unbounded flows. The effects of viscosity are also calculated, asymptotically, and shown to have a stabilizing role in all the cases investigated. Exceptional regimes to the problem occur when the wavespeed of the disturbances approaches the velocity of either of the walls, and these regimes are also analyzed in some detail. Finally, the effect of imposing radiation-type boundary conditions on the upper (moving) wall (in place of impermeability) is investigated, and shown to yield results common to both bounded and unbounded flows
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