Aero-optic effect is investigated in supersonic/hypersonic flow. Flat plate, ramp and cave model in supersonic/hypersonic flow condition are studied with large eddy simulation (LES). Turbulent density pulse is predicted to verify coefficient of density pulse model-mixing length (ML) model. It indicates that LES method simulates turbulent density pulse well in flat plate boundary layer flow, oblique shock wave/boundary layer interaction and strong separation flow. Density pulse data obtained by LES is used to verify ML coefficient and set a quantitative result. The modified ML was used to predict aero-optical effect on missile successfully.

A method is established to calculate aero-optical effects.It is valid for nonuniform flow fields.A three-dimensional spatially developing mixing layer is simulated by large eddy simulation(LES).Favre-filtered Navier-Stocks equation is discretized with a high order compact difference scheme.Transition from laminar to turbulence is simulated successfully.And it is compared with that of direct numerical simulation.Aero-optical effects caused by mixing layer is calculated to study relation between flow field and imaging distortion.It shows that imaging distortion is the severest at transition region.

Shock/turbulent boundary layer interaction(STBLI) in a M_∞=3.0 supersonic compression ramp flow are studied using a spatial large eddy simulation method. At upstream, flat plate transition and developed turbulent boundary layer are simulated in which Shock/turbulent boundary layer interaction shows at the ramp. Analysis on mechanism of turbulent boundary layer/shock interaction indicates that linear unstable disturbance leads transition rapidly. In the process of STBLI, average separation range of the ramp is smaller than that in the laminar flow. However, extent of the separation shows different characteristics in different span-wise positions due to high velocity strip and low velocity strip.

A uniform method for unsteady flows at arbitrary Mach number is developed.The preconditioned dual-time compressible Navier-Stokes equations for unsteady flows are obtained by incorporating a pseudo-time derivative term and preconditioning.The inviscid flux vectors and the viscous flux vectors are differentiated by a Roe-scheme of three-order and a general central difference scheme,respectively.A pseudo-time subiteration with an implicit ADI-LU scheme is implemented based on state-vector flux linearization.A two-order precision for physical-time step is achieved.Taking low speed flows as examples,the flow over a circular cylinder and the dynamic stall of the NACA0015 airfoil with identical pitching rate are simulated.Computed results agree with published results.It shows that the method simulates unsteady flows at arbitrary Mach number well.

A hybird scheme is constructed by Steger-Warming Splitting and its modified algorithm.The enlargment factor of pressure difference is used to prompt ability of flow field resolution.Meanwhile,a special matrix splitting method is adopted and makes matrices very easy to compute.Three examples of two-dimensional fluid flows containing shocks and/or boundary layer demonstrate the efficiency and accuracy of this new scheme.

Based on a static non-overlapped domain decomposition method,the problem is studied about the parallel implementation of a Harten-Yee implicit TVD difference algorithm on a MIMD multi-computers parallel system.The convergence and stability of the parallel processing techniques used here are manifested well through numerical simulation of the steady "super-supersonic" jet flow fields,and the parallel parameters obtained are promising.

Three-dimension, compressible, internal flow solution obtained using a modifiedBeam-Warming implicit factorization scheme is presented. Steady-state solution is obtained by solving numerically the full Navier-Stokes equations from given initial conditions until the time-dependent terms become negligible. The configuration considered is a rectangular cross-section, S-shaped centreline diffuser duct with an exit/inlet area ratio of 2.25. TheMach number at the duct entrance is 0.9 with a Raynolds number of 5.82×10⁵. Two regions of separated flow exist within the diffuser. The detail flow field especially the developement process of secondary velocity patterns in transverse plane are pressented. Comparing with the results ofRef.[1]. they have a good agreement. The secondary velocity patterns are more reasonable.

The interaction between supersonic external flow and an exhaust jet forms a complex flowfield region near the missile afterbody. In this paper, the research about this interaction flows was completed with finite-difference method to solve N-S equations. The flowfield structures observed in experiment and it's relation with jet pressure ratio p_j/p_∞ were obtained. For freestream, M_∞=1.94, R_e∞=2.2×10⁵, for jet flow, M_j=3.0, three p_j/p_∞(1.03, 0.527 and 0.15) were selected. The numerical algorithm used is a modified Beam-Warming scheme. The nozzle base pressure coeficient and shock reflection length were compared with experimental data, good agreements were reached.

In this paper, we present a modified implicit ENO scheme and combine it with CSCM-like implicit supra-characteristic boundary treatment. Numencal experiment with this scheme for 2-D channel fupersonic flow aboat 4% circular arc has damenstrated the sharp shock-capturing capability of the scheme. The results are also compared with those of TVD scheme. The futher application to the viscid flow in simplified inlet is made. They show a good agreelment with the results of SIP/Beam-warming schemes except nearby the reflection point of shock. But our results seem to be more reasonable.