A 3d LU-SGS computational code is developed to solve the hypersonic flows, which is modeled by the compressible Reynold averaged three-dimension Navier-Stokes equations and with a use of a finite volume formulation, the convective parts are discretized with the Harten TVD scheme, turbulent effects are modeled utilizing Menter SST tow equation turbulence model.

In order to simulate the flow field around parachute during terminal descent, a finite volume method and Spalart-Allmaras turbulent model are used to solve N-S equations.

Transonic flow over an oscillating airfoil has been computed by solving the Euler equations. A finite volume scheme is used to spatially discretize the integral form of the Euler equations for a moving domain, and a symmetric TVD scheme is presented to construct dissipative terms.

(2) Issues of the nonlinear flutter of the 3-D wing which stiffness is the hard spring and freeplay nonlinear are studied by using the mothod of finite element and the theories of the qusi-steady aerodynamics and the Hopf bifurcation.

This paper used the finite element method to compute the temperature field, and combined finite element with analytic solution to develop a calculation of thermal stresses.

Through static, modal and crack stress intensity factor analysis, the vibration stress of blade under rotating state before fracture can b e calculated out finally.

In all these cases we actually show that Γ=π1(M) has a finite index subgroup which is mapped onto a nonabelian free group.

Using these monomial bases we prove that the image of the transfer for a general linear group over a finite field is a principal ideal in the ring of invariants.

We prove the following result: LetG be a finite irreducible linear group.

The first one is a conjecture of Ian Hughes which states that iff1, ..., fn are primary invariants of a finite linear groupG, then the least common multiple of the degrees of thefi is a multiple of the exponent ofG.

More precisely, each orbit of the above action intersects one componentX ofQ in a finite number of points and the action of PGL4 restricted on each componentX is equivalent to the action of a finite groupGX onX which can be explicitely computed.

A general program using finite-difference method is written to calculate the flow field of supersonic axisymmetric nose inlet. This program is suitable to the direct problem of aerodynamic design of inlets e. g. pitot-type, single-cone, bicone, triple-cone and isentropic cone etc. Since "seperation singularity" difference method and implicit-explicit difference scheme are adopted in the calculation of inviscid flow field in the inlet, results of calculation are obtained with second order accuracy at boundary...

A general program using finite-difference method is written to calculate the flow field of supersonic axisymmetric nose inlet. This program is suitable to the direct problem of aerodynamic design of inlets e. g. pitot-type, single-cone, bicone, triple-cone and isentropic cone etc. Since "seperation singularity" difference method and implicit-explicit difference scheme are adopted in the calculation of inviscid flow field in the inlet, results of calculation are obtained with second order accuracy at boundary points and internal points, as well as near singular points. This program can offer required internal and external flow characteristics of inlets.Numerical results for five examples are in satisfactory agreement with corresponding results obtained from the method of characteristics and experimental data.

A general formulation for oscillatory subsonic potential flows around three-dimensional bodies of various configuration and its application to the calculation of dynamic stability derivatives of the aircraft are presented. By applying the Green function method, we obtained an integro-differential equation relating the perturbation velocity potential to its normal derivatives on the surface of the body. In order to solve this equation, the surface of the body and its wave are divided into small quadrilateral...

A general formulation for oscillatory subsonic potential flows around three-dimensional bodies of various configuration and its application to the calculation of dynamic stability derivatives of the aircraft are presented. By applying the Green function method, we obtained an integro-differential equation relating the perturbation velocity potential to its normal derivatives on the surface of the body. In order to solve this equation, the surface of the body and its wave are divided into small quadrilateral elements. The unknown φ and its derivatives are assumed to be constant within each element. Thus the integro-differential equation reduces to a set of differential-delay equations in time. This set of equations can be used as the basis of a general method for the fully unsteady flow calculation. For oscillatory subsonic potential flow, this set of equations further reduces to a set of linear algebraic equations which is solved numerically to yield the values of φ; at the centroid of each element. The pressure coefficient is evaluated by the finite difference method. The lift and the moment coefficients are determined by numerical integration of the pressure coefficient. The dynamic stability derivatives are obtained from the imaginary parts of the lift and the moment coefficients.The formulations in this paper are embedded into a general computer program. Several typical numerical results have been obtained by means of this program. Figure 2 shows the distribution of lift coefficient CL along the middle section for a rectangular wing oscillating in pitch with λ =2, τ =0.001, M∞ = 0, K = 2 .The result is identical to the original calculation by Merino. Figure 3 shows the distribution of pressure coefficient Cp for a harmonically oscillating spheroid witha/b= 8, M∞=0.5, K=2 . The result is in good agreement with the analytical solution of wave equation.Figures 5 , 6, 7 show the distributions of lift coefficient CL at various stations of an aircraft (wing-body-tail combination) oscillating in pitch with M∞ = 0.6, K -0.005, 0.01. Vable 2 shows the dynamic stability derivatives CLa, Cma of the aircraft. The. results are in good agreement with the experimental data.

In order to calculate the lifts of rectangular wings with different winglets at subsonic speeds, we have adopted the finite element method which divides spanwise lattice and determines spanwise locations of control points by means of a constant roll-angle method. The induced drags are also calculated by using combined flow field method.As the results of calculating various configuration of a winglet, we have found out some rules affecting the lift and drag characteristics of wings with winglets and picked out...

In order to calculate the lifts of rectangular wings with different winglets at subsonic speeds, we have adopted the finite element method which divides spanwise lattice and determines spanwise locations of control points by means of a constant roll-angle method. The induced drags are also calculated by using combined flow field method.As the results of calculating various configuration of a winglet, we have found out some rules affecting the lift and drag characteristics of wings with winglets and picked out a favourable configuration from them. The aerodynamic mechanism of winglets is also discussed.