Linear parabolized stability equations(LPSE) considering model curvature are solved with finite difference method.Stationary cross-flow instabilities in boundary layers of an infinite swept-wing are analyzed.LPSE and experimental results are compared.It is shown that at early development of stationary cross-flow disturbances,LPSE simulates flow structure and disturbance profiles well and predicts the N-factors accurately.When disturbances are amplified enough,high order terms can not be omitted and linearization assumption of LPSE is no longer suitable.Moreover,effects of model curvature and boundary layer non-parallelism are investigated.It shows that curvature and non-parallel terms have significant effects on stability analysis of stationary cross-flow instabilities on a swept-wing,and the effects are independent of Reynolds numbers.In the model investigated,inclusion of curvature has a stabilizing effect and non-parallelism shows destabilizing effects on disturbances.

Plasma actuator is modeled as body force vectors resulting from external electric field and net charge distribution in plasma. Leading edge separation flows around NACA0015 airfoil under control of plasma actuator is simulated by Navier-Stokes equations with body force source which are solved by detached eddy simulation (DES) and delayed-detached eddy simulation (DDES). With actuator on, the separation flow reattached through an asymptotic progress in DDES simulation. An unsteady response occurs in DES due to grid induced separation.

An LU-SGS(Lower-Upper Symmetric Gauss-Seidel) scheme and improvement for three-dimensional unstructured grid Euler computation are presented.Grid reordering for efficient implementation of the improved LU-SGS and the original LU-SGS is proposed.In order to test the feasibility and efficiency of the improved LU-SGS algorithm,two numerical examples of transonic inviscid flow around an ONERA M6 wing and an LANN supercritical wing are presented.It shows that the numerical results agree well with the experimental data.Moreover,the efficiency of the improved LU-SGS scheme is two times higher than that of the original LU-SGS scheme and seven times higher than that of the explicit Four-Stage-Runge-Kutta scheme.

The building mechanism and artifical dissipation models of hybrid flux difference schemes are analyzed. The explicit algorithm turbulent model (EASM) is combined with the Reynolds average N-S equations to simulate turbulent flows. Resolution and robustness near shock discontinuities are investigated with low speed laminar/turbulent flows over a flat plate, transonic convergence-divergence diffuser flows and hypersonic inviscid flows over blunt-nose body. Compared with experimental data it shows that the hybrid flux difference schemes obtain similar resolution as a Roe-FDS scheme, and are free of non-physical numerical solutions as FVS schemes.

Based on a distributed parallel computing system,a new approach to the transonic flutter of a flexible wing is presented by coupling the computational fluid dynamics (CFD) method with the computational structural dynamics(CSD)method. by using a transfinite interpolation(TFI) scheme to generate a multi-block C-H moving grid,the aerodynamic analysis is carried out by solving the three-dimensional unsteady Navier-Stokes equations. By coupling with the flutter motion equations,the response of the time history of generalized coordinates is found.The critical velocity is got on the base of the response. by considerally decreasing the computing time,a multidisciplinary analysis code runs on a distributed parallel computer system by using the PVM library.Computational results are presented for the transonic flutter of wings with deformed shapes as found in flutter simulations.

An analytical non-linear model including surface-state effect is proposed for 4H-SiC power MESFET's with which the impact of suface damage at the ungate recess region caused by the dry-etching process on the output steady-state characterization can be illustrated clearly.The model has the advantage in very simple calculations over the 2D numerical simulations, therefore suitable for process analysis of SiC power MESFET's.

An analysis method of the airfoil flutter system with nonlinear structure at transonic speeds is presented. The unsteady aerodynamic forces are calculated by Navier-Stokes equations. The time histories of structure responses of airfoil are solved by coupling the Navier-Stokes equations and flutter equations. The systems with gap stiffness, cubic type stiffness and cubic structure damping are investigated. The calculated results show that the influence of the structure nonlinearity on flutter speed is sensitive. Due to both the structure and aerodynamic nonlinearity in the flutter system, the oscillation characterstics become very complex.

The 2-D time-dependent compressible Navier-Stokes equations are adopted as the governing equations. The unsteady lift coefficients of airfoil are calculated by Navier-Stokes equations with the implicit LU-NND algorithm, Baldwin-Lomax turbulent model and C-type grids. The fluctuating values of lift coefficient are increased with the enlargement of incidence and Mach number. The buffet onset boundary is defined as the beginning point where the fluctuating values of aerodynamic coefficients are increased suddenly. The fluctuating values of aerodynamic coefficients expresses the buffet strength. The buffet characteristics of supercritical airfoil DFVLR-R2 and airfoil NACA0012 are investigated. The calculating results show that:in the design Mach number range of supercritical airfoil, there are larger buffet onset incidence and less buffet loads. When Mach number is larger than the design Mach number, the buffet onset incidence of supercritical airfoil decreases quickly and there are rather large buffet loads. In the convention airfoil NACA 0012, there are much less buffet onset incidence, larger buffet loads and separated range.

A dual-time scheme is used to solve the unsteady compressible Navier-Stokes equations to obtain the needed airload.The model structural equations of motions are computed simultaneously using the Runge-Kutta method to calculate the aeroelastic static deformations and responses of wings in the time-domain.The O-H type grid around the deforming wings is generated using the improved transfinite interpolation method so that grid generation in each time-step is not too time-consuming.The numerical results show the transonic flutter speed drop and nonlinear characteristics.

It focuses on an approach to generating 2D unstructured grids for the solution of Euler equations.Unlike the generation of unstructured grids before with background grids,the background information needed in Avancing Front Method is obtained directly with the distance to the airfoil surface.In solving the Euler equations,the cell centered finite volume method is used for space discretization,and four step Runge Kutta method is adopted for time marching.Two kinds of boundary condition are presented and three kinds of boundary conditions are applied and analyzed.The calculated results are given and discussed.

A technique on generation of unstructured grids for three dimensional complex geomtries is presented.Complex configurations are described by several typical kinds of surface patches.A 3D surface patch is first mapped to a 2D segment,then triangulated using a 2D grid generator based on the advancing front method,and finally mapped back to the original 3D shape.Tetrahedral grids are generated by an advancing front method and smoothed by a nodal relaxation technique and a Deluanay transformation.A mesh refinement technique is also given to obtain more accurate solutions.

A numerical method is presented for calculating the unsteady transonic viscous flows around oscillating aerofoil and wing based on the unsteady Navier-stokes equations.It adopts an implicit NND algorithm and Baldwin-Lomax turbulence model in solving procedure.The computational grid is C and C-H types in a body-fitted coordinate system generated by an algebraical method. The transonic viscous flows around the NACA 0012 aerofoil and M6 wing with pitching oscillation have been calculated.Numerical results are in good agrement with experimental data.

The aeroelastic behaviour of wings with different types of separated vortex at high incidences is investigated.The unsteady aerodynamic equations and the equations of motion are solved simultaneously in the time domain.The separated vortices shed-ding from the leading-edge and/or side-edge of wings are taken into account by an improved Nonlinear Vortex-Lattice Method.The equations of motion are integrated by Runge-Kutta Method.Wings with different planforms at different basic incidences are consid-ered.It is found that (1) the separated vortex shedding from the leading-or side-edges results in the critical flutter speed of the wing lower,and the higher the basic incidence,the lower the critical flutter speed.(2) the smaller the sweep of wing,the smaller the in-fluence of separated vortex on the aeroelastic characteristics.(3) at certain flight speeds and basic incidences,stable limit cycles ap-pear,and the amplitudes of limit cycles even keep constant when the initial distu rbances of response vary.