With a sub-iterative technique for fluid dynamic equations and rigid-body dynamic equations,a coupling numerical method is proposed to explore the unsteady movement characteristics of aero-crafts and the effect of coupling time accuracy.It is shown that the motion finally becomes a self-maintained stable limit cycle oscillation.The amplitude with an increasing incidence angle shows a jump,which is quite close to the experimental result.The coupled numerical result with the second-order time accuracy depends hardly on the physical time step for this multi-system coupling.On the contrary,the uncoupled numerical result with one step time lag depends obviously on the physical time step selected.A longer time step results in unreal numerical results.

Firstly, from the calculations of the heating rate on a blunt body in hypersonic flow and separation of shock/boundary layer interac tion flow, the capability of upwind TVD scheme for computing viscous flows is st udied. It is found that the computed heating rate and separated region are lower than the experimental data. The problem is the entropy correction formula. Then a new entropy correction formula is proposed, and the agreeable results compared with experiment are obtained. Secondly, using modified scheme, the study of control of separated flow over a smooth surface by body wall temperature is completed. The results show that the change of wall temperature can control the locati on of separation point and separation region effectively.

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.

TVD scheme is one of the most advanced schemes for investigating supersonic and hypersonic complex flows in which shock wave is one of important properties up to date. In present paper, axisymmetric viscous flows about three indented nosetips and hypersonic flow about a blunt-cone body (θ=10°) at angles of attack were simulated with a two-order windward TVD scheme. High quatitive shocks and surface pressure distributions agreed with experiments and results from space-marching method were obtained. It is also demonstrated the unique advantages of present method for computing complex flows over re-entry vehicles.