Local Structural Stability of a Multidimensional Centered Rarefaction Wave for the Three-Dimensional Steady Supersonic Euler Flow around a Sharp Corner

TWENTY-FIVE years ago what was known on the dynamics of gases, if acoustics and the study of steady motion at speeds small compared with that of sound (when the flow resembles that of a liquid) be omitted, could have been expounded in a few dozen pages. Since then, and more rapidly since 1943, the subject has grown into one which no single book can describe in its entirety. The primary stimulus was the increase in aircraft speeds ; but a secondary stimulus, the belated realization of workers in pure science that the phenomena show fascinating differences from those of classical physics, has become equally important. Supersonic Flow and Shock Waves By R. Courant K. O. Friedrichs. (Pure and Applied Mathematics, a Series of Texts and Monographs, Vol 1.) Pp. xvi + 464. (New York and London: Interscience Publishers, Inc., 1948.) 42s.

Investigation on wall shear stress measurement in supersonic flows with shock waves using shear-sensitive liquid crystal coating

Supersonic axisymmetric expansion flow is a typical and fundamental issue in gas dynamics. It plays a vital role in the high-speed external and internal flow fields regarding the contour design and performance evaluation of supersonic/hypersonic vehicles and their propulsion systems. The supersonic two-dimensional (2D) planar expansion flow is dominated by the well-known Prandtl–Meyer (P–M) theory. However, no similar explicit relation exists for the supersonic axisymmetric expansion flow, and only the computational fluid dynamics results could be employed at present. Therefore, this work focuses on developing the analytical solution of supersonic axisymmetric flow around a sharp convex corner on the basis of the generic gasdynamic functions in a newly established coordinate system for addressing the aforementioned issue. Theoretical derivations and numerical results prove that the flow deflection angle and Mach number in supersonic axisymmetric flow around a sharp convex corner obey the identical law to the 2D planar situation, that is, the P–M theory, while the local axisymmetric expansion fan is not the simple wave flow despite the conical flow. Meanwhile, the method of characteristics is employed to further explicate the intrinsic connection and difference between the 2D and axisymmetric sharp convex corner flow. The equivalence of sharp corner and curved surface flows with the identical deflection angle is discussed, and three limitations of the proposed analytical solution are clarified.

In this paper, we study the transonic shock problem for the full compressible Euler system in a general two-dimensional de Laval nozzle as proposed in Courant and Friedrichs (Supersonic flow and shock waves, Interscience, New York, 1948): given the appropriately large exit pressure p e(x), if the upstream flow is still supersonic behind the throat of the nozzle, then at a certain place in the diverging part of the nozzle, a shock front intervenes and the gas is compressed and slowed down to subsonic speed so that the position and the strength of the shock front are automatically adjusted such that the end pressure at the exit becomes p e(x). We solve this problem completely for a general class of de Laval nozzles whose divergent parts are small and arbitrary perturbations of divergent angular domains for the full steady compressible Euler system. The problem can be reduced to solve a nonlinear free boundary value problem for a mixed hyperbolic–elliptic system. One of the key ingredients in the analysis is to solve a nonlinear free boundary value problem in a weighted Holder space with low regularities for a second order quasilinear elliptic equation with a free parameter (the position of the shock curve at one wall of the nozzle) and non-local terms involving the trace on the shock of the first order derivatives of the unknown function.

On the degeneration of supersonic flow due to interaction of centered compression and rarefaction waves: PMM, vol. 43, no. 1, 1979, pp. 45–50

Compressible flow past thin wings and slender bodies are studied. First, subsonic flow past thin wings are analyzed by means of potential flow theory. The kernel function method is introduced for arbitrary planforms undergoing simple harmonic oscillations. The spanwise polynomial approximation and the chordwise trigonometric function approximations which automatically satisfy the Kutta condition and inherit the leading edge singularity are used. The Doublet-Lattice method as a more general numerical approach is given for the analysis of the flow past additional surfaces like tail or store surfaces, which are not necessarily in plane with the wing surface or surfaces having spanwise deflections. Airfoil response to the arbitrary unsteady motion is also given for subsonic flows. A brief review of shock waves and Mach waves in a supersonic flow is given. Afterwards, unsteady supersonic potential flow is studied for a simple harmonically oscillating point source. First, unsteady flow past an airfoil is considered. Then, supersonic flow about thin wings are analyzed using the Mach box technique. Introduction to supersonic kernel function method is briefly provided. Arbitrary unsteady motion of an airfoil in a supersonic flow is presented. Slender body theory is introduced to analyze the cross flow past bodies of considerable fineness where the cross flow is shown to be approximately incompressible under certain conditions. In connection with the slender body approximation, the Munk’s airship theory is utilized to predict the stability derivatives of missile like bodies.

In this paper, we study the expansion problem which arises in a two‐dimensional (2D) isentropic pseudo‐steady supersonic flow expanding into vacuum around a sharp corner. The problem catches interaction of an incomplete centered simple wave with a backward planar rarefaction wave as well as interaction of a simple wave with rigid wall boundary of the 2D self‐similar Euler equations. Using the methods of characteristic decompositions and invariant regions, we mainly deal with the interaction of two symmetric non‐planar simple waves of the 2D self‐similar Euler equations to obtain the global existence of the solution for the interaction of the simple wave with rigid wall boundary. It follows the global existence of the solution up to the interface of supersonic flow with vacuum of the expansion problem.

Lattice Boltzmann simulation for unsteady shock wave/boundary layer interaction in a shock tube

: A Navier-Stokes code to solve the aft end flow field of missile type configurations is presented. The consistently split linearized block implicit method of McDonald and Briley is employed in modified form to handle L-shaped domains with sharp reentrant corners. Appropriate boundary conditions are applied for the supersonic flow in particular at the outer boundary so that waves generated within the flow field are allowed to pass out of the computational domain without reflecting back into it. An adaptive grid option has been incorporated into the code and has been exercised by following the shear layer in a model backstep problem. Results are presented for the supersonic turbulent flow over a nozzle boattail configuration with and without jet exhaust and the results are compared with experiment. Calculations of the 2-D turbulent supersonic flow over a right angle back step with shear layer reattachment on a 20 deg ramp are also shown, and compared with experiments. The computation shows the qualitative physical behavior of the flows and there is generally good agreement with the experimental velocity profiles through most of the free shear layer and the ramp reattachment zone. (Author)

This study presents flow simulation over the reentry capsule at supersonic and hypersonic speeds. Numerical algorithms solve for the CFD method, which is produced using help ANSYS Fluent 19.2. The using GPU core to get a solution faster. The main purpose – flow simulation and numerical analysis reentry capsule; understand the behavior of supersonic and hypersonic flow and its effect on the reentry capsule; compare temperature results for the range Mach numbers equals 2–6. This study showed results on velocity counters, on temperature counters and vector of velocity for range Mach numbers equals 2–6. This study demonstrates the importance of understanding the effects of shock waves and illustrates how the shock wave changes as the Mach number increases. For every solves, the mesh had adapted for pressure gradient and velocity gradient to get the exact solution. As a result of the obtained solution, it is found that a curved shock wave appears in front of the reentry capsule. The central part of which is a forward shock. An angular expansion process is observed, which is a modified picture of the Prandtl- Mayer flow that occurs in a supersonic flow near the sharp edge of the expanding region. It is revealed that with an increase in the Mach number, the shock wave approaches the bottom of the reentry capsule, and there is also a slope of the shock to the flow direction, with an increase in the Mach number. The relevance and significance of this problem for the design of new and modernization of old reentry capsules.

Subject of study.We consider the Riemann problem for parameters at collision of two plane flows at a certain angle. The problem is solved in the exact statement. Most cases of interference, both stationary and non-stationary gas-dynamic discontinuities, followed by supersonic flows can be reduced to the problem of random interaction of two supersonic flows. Depending on the ratio of the parameters in the flows, outgoing discontinuities turn out to be shock waves, or rarefactionwaves. In some cases, there is no solution at all. It is important to know how to find the domain of existence for the relevant decisions, as the type of shock-wave structures in these domains is known in advance. The Riemann problem is used in numerical methods such as the method of Godunov. As a rule, approximate solution is used, known as the Osher solution, but for a number of problems with a high precision required, solution of this problem needs to be in the exact statement. Main results.Domains of existence for solutions with different types of shock-wave structure have been considered. Boundaries of existence for solutions with two outgoing shock waves are analytically defined, as well as with the outgoing shock wave and rarefaction wave. We identify the area of Mach numbers and angles at which the flows interact and there is no solution. Specific flows with two outgoing rarefaction waves are not considered. Practical significance. The results supplement interference theory of stationary gas-dynamic discontinuities and can be used to develop new methods of numerical calculation with extraction of discontinuities.

A calculating method of shock wave oscillating frequency due to turbulent shear layer fluctuations in supersonic flow

Generalised correlations of blow-off and flame quenching for sub-sonic and choked jet flames

The limitations of high-power electrical lasers due to heating of the gas and the instability of the glow discharge can be alleviated by the flow of the lasing medium. We describe here the results of two experiments which combine glow discharges with gas flow. In the first experiment we study the effects of supersonic flow and shock waves on glow discharges. The results of some numerical calculations of the variation of the gas properties along the discharge tube are then described. In the second experiment the influence of a highly swirled and turbulent flow on electrical discharges is studied. It is observed that both flows have a favorable effect on the discharge stability.

Pulse volume discharge behind shock wave in channel flow with obstacle