Based on NACA airfoils,a new type of tandem buoyancy-lifting airships with both higher utility and economy efficiency was developed.With numerical simulations aerodynamic and configuration design for buoyancy-lifting airships was investigated and buoyancy-lifting characteristics were evaluated and analyzed synthetically.It shows that a tandem buoyancy-lifting airship has considerable volume efficiency.Its lift-drag ratio is increased 50% than single wing resulted from block of airflow by posterior flying-wing and speedup of airflow in overlay field.Its deadweight can be decreased and payload is increased as velocity is higher than 30 m·s^-1.At the same time,chord wise size of tandem airship is comparatively small so that surface tension of skin material can be reduced availably.It provides a wider scope for skin flexible material selection of high altitude aircrafts.

Magnetic field molding is a key technique in obtaining hish-performance permanent magnetic materials from strontium ferrite,Based on a mathematical model,we focus on effect of domain center and material geometrical center's relative position on magnetic properties of strontium ferrite.It shows that the external magnetic field atcts on magnetic domain center more obviously as the center shifts downward;as the sample is flat and the external magnetic field vertically acts on the sample absolutely.the effect is the greatest,and strontium ferrite in shape anisotropy can be obtained.

A scaled boundary finite-element method(SBFEM) is used for determining free surfaces in two-dimensional steady-state seepage problems.Equations of SBFEM for a seepage problem are derived with a transformation from Cartesian coordinates to scaled coordinates.Governing equations are discretized on boundary of computational domain by SBFEM,which reduces spatial dimension by one,and analytical procedure is applied in the reduced direction.As scaled centre is at intersection point of upstream and impermeable straight boundary of the dam bottom,only free surface and its downstream boundaries are discretized.A governing point method is used to control overflow point in the free surface.A 2-D steady seepage with a free surface in a homogeneous dam is analyzed and compared with experimental result.It shows that the method has satisfactory convergence,accuracy and less amount of data preparation.

Unsteady incompressible Navier-Stokes equations are solved with a pressure correction method.A series of numerical calculations are conducted for Low-Reynolds-Number aerodynamics of an inverse Zimmerman wing.It shows that a pair of strong leading-edge vortex exist beyond the wing.They have strong effect on the inverse Zimmerman wing aerodynamics and offer high nonlinear lift and large stall attack angle.Interaction between the leading-edge vortexes results in unsteady and asymmetry phenomenon of the wing.The inverse Zimmerman wing shows a better longitudinal stability and flow-field stability.

A pseudo-compressibility method developed by Rogers is used to solve the incompressible Navier-Stokes equation. Numerical flux is scattered through 3rd order Roe scheme and the 2nd order Harten-Yee TVD scheme separately. In order to accelerate convergence, several implicit methods(ADI-LU, LGS, LU-SGS) are accepted. Efficiency of different methods are shown as a classical flowfield around a cylinder in low Re(Re=200). Roe scheme is used to simulate two low speed unsteady problems:Low Re flowfield around a rotational cylinder(ω=1) and flowfield around a pitching NACA0015 airfoil with identical pitching rate.

A fully implicit Lower-Upper-Symmetric-Gauss-Seidel (LU-SGS) with pseudo time sub-iteration and the modified Jameson's central scheme are applied to solve thin layer Navier-Stokes(N-S) equations with laminar hypothesis and Baldwin-Lomax(B-L) model for the vortical flows around delta wings at high angles of attack (AOA).To validate the codes,a 65 degree swept delta wing is computed in transonic cases,where M_∞=0.85,α=10°,20° and Re=5×10⁶.Then asymmetrical vortex breakdown flows with M_∞=0.3,and Re=1.3×10⁶,at high AOA around an 80° swept delta wing are presented and discussed.The computational results are compared with experimental data available and good agreement is achieved.

A multi-layer hybrid grid method is constructed to simulate complex flow field around 2-D and 3-D configuration.The method combines Cartesian grids with structured grids and triangular meshes to provide great flexibility in discretizing a domain.It generates the body-fitted structured grids near the wall surface and the Cartesian grids for the far field.In addition,the triangular meshes are regarded as an adhesive to link each part of grids.The grid combination methodology is discussed,which can smooth hybrid grids and improve the quality of the grids.A uniform type of data is defined to manage and arrange grid data.Using the centered finite volume method,the inviscid and viscous flow allows computation by solving the Euler and Navier-Stokes equations.The computational results prove that grid generation and flow calculation are correct and feasible.

Four turbulence models:algebraic Baldwin-Lomax model with Degani-Schiff modification, two versions of Johnson-King model (J-K90A and J-K92) and two-equation k-g model, are described and evaluated for missile supersonic flow, NASA TN D-712 standard model and the civil airplane wing-body configuration (two-block C-O mesh) transonic flows. The 3-D compressible Reynolds averaged Navier-Stokes (RANS) equations are integrated numerically by central-difference with artificial viscosity scheme, finite volume formulation and explicit multi-step Runge-Kutta algorithm. The turbulence equations of k-g model are explicitly solved in the same way as the RANS. Results show that all models can perform well for attached and mildly separated flows. For large angle-of-attack separated flow over the missile, k-g and J-K92 models match the experimental results much better than those with B-L and J-K90A models. According to the B-L model with some modification, the shock location logs behind those with the other three models for the civil airplane geometry. Model k-g, because of its advantages such as without using normal-to-wall distance, simple source terms and straightforward boundary conditions, performs best in the four models for complex configuration with multi-block mesh system or multi-wall interference. B-L and J-K models, despite of cheapness and robustness, both base on empirical Prandtl mixing-length hypothesis and need to calculate the distance normal to the wall, undoubtedly limiting the application scope of these models for multi-block mesh system and complex geometry.

A method for the adaptive refinement of a Quad/Cartesian grid for the solutions of the Euler and N-S equations is presented.An adaptive Quad/Cartesian hybrid grid method,which retains the advantages of the Cartesian grid approach and, at the same time, touches the viscous flow problems,is provided.The body fitted Quad grids are generated around the body and the Cartesian grids are used in the remains of the flow field.The flow solver of the Euler and N-S equations is performed by a conventional algorithm,including the finite volume method and the Runge-Kutta time-stepping scheme.The grid adaptation and the flow solver achieve reasonable efficiency and accuracy with minimum computer resources.Based on the above approach,the numerical analysis of flows around the single-element and the multi element airfoils is given.The numerical results presented show the flexibility of this approach and the accuracy attainable by the solution-based refinement.

Grid generation and flow analysis methods for multi-element airfoil based on multi-block point-to-point patched grid technology are developed. Some strategies are developed to improve the efficiency, accuracy and applicability of the method.Several flow cases about multi-element airfoils are computed with the present method in order to validate it. Numerical results are in agreement with experiment data. Flow field visualization shows that the present method has good resolution in flow details.

A new unstructured Cartesian cutted grid generation method is introduced. The grid generation of some airfoils and wing is performed with the tree data structure. Euler equations are solved using the cell-centred finite volume method. The results of flow field computation are in agreement with the experiments. These facts show that the grid generation and the numerical calculation approach is correct and reasonable. Based on the above work, a flow field numerical simulation of wing and store separation is finished. This process provides a new way and tool for the more accurate solver of the unsteady flow problem.

Because of the complexities in both configurations and flow phenomena,there are great difficulties in the numerical simulation of flow about wing with lift-enhancing devices and control surfaces.By using cell-centered finite volume method,it deals with Renolds-averaged Navier-Stokes equations on multi-block structural grids generated by elliptic method.Two kinds of patching schemes,point-to-point and patch-on-common-surface,are employed to establish the information exchange among blocks.In the later one,the overlap cell area weighted interpolation is used to guarantee the conservation.Ramshaw algorithm is also used to pre-determine the overlap areas.Flow fields about multi-element airfoils and delta wing with deflected flap are computed with the present method.Numerical results are all in good agreement with experiment data.Flow field visualization shows the present method has good resolution in flow details.

A Fast grid generation procedure named parabolic equation method for complex aircraft is developed. Grid is obtained by marching from body surface to outer boundary. Successful control of grid orthogonality on body is achieved by means of the introduction of algebraic grid prediction. A numerical algorithm based on a finite volume explicit scheme is employed for calculating transonic flow around a realistic aircraft configuration. Comparisons of the computational results with experimental data show very good agreement in terms of trends.

The explicit and implicit methods are employed in the numerical simulation of unsteady transonic NACA10012 airfoil flow,and comparison analyses have been made in the respects of computational results,efficiency,and sensitivity to grid quality.

The second order moment conservation and Bottadvection schemes that are responsible for being positive definite and high accurate are introduced into the MM4(meso scale model) for moisture advection.When separately used with the MM4 with observed windfields available a range of scenarios have been subjected to theoretical test and numerical experiments done with more than one rainstorm event.Nurnerical results show that the improvement of calculation accuracy of model moisture advection is the key link of raising the prediction of weather calamities like heavy rain.

A body-fitted orthogonal grid around complex configurations including wing,body,pylon and nacelle has been suc-cessfully generated by applying the multi-block grid technique and the method of solving elliptic and parabolic partial differential e-quations.The transonic flow field and aerodynamic characteristics of the advanced twin engine transport aircraft are also simulated numerically by using finite volume methods to solve the Euler equations with high efficiency,and the viscous effect of wing bound-ary layer is considered.Results of the calculation show that pylon and nacelle have significant effect on the supercritical wing aerody-namic characteristics.the numerical results are in good agreement with the experimental wing surface pressure distribution data.

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.

Deconvolution methods based on the pulse compression principle are analyzed,and a more generalized l_p-norm criterion for wide use is proposed,which in cludes as special cases pulse deconvolution,minimum-entropy deconvolution l_t-norm deconvolution and D-norm deconvolution.According to the p-norm properties of the finite-dimensional space and the geometric properties of the higher-dimensional space,an analysis is made on the effect of the parameters in the generalized l_p-norm method,and a criterion function is designed to compare pulse compression of different methods.It is also theoretically proved why for nonminimum-phase signals the l₁-norm approach yields better pulse compression than pulse deconvolution,and why the D-norm approach has more "simplicity" than general minimum-entropy methods.New deconvolution methods can be readily constructed,and analyzed,by use of this generalized l_p-norm criterion.The results provide a better understanding of the role these deconvolution methods play in enhancing the resolution and the theoretical basis for further study of the optimal pulse compression criterion and the algorithms to find its optimal solution in deconvolution.

Some aerodynamic characteristics of an oscillating trail-edge flap were investigated by numerical simulation, such as the unsteady effect of flow field and flow field comparison between different frequences or amplitudes. The Euler equations were solved by LU-ADI algorithm, the grid is C type grid. The numerical results showed that: (1) with the oscillation of trail-edge flap, the shock strength and its location are nonlinear to the motion of the trail-edge flap, (2) the higher the reduced frequence, the less the influence to variable pressure jump, (3) with the increase of the amplitude, the variations of shock strength and its location become more seriously.

This paper utilizes a series of analytical transformations to generate two-dimensional curvefitted, orthogonal meshes. The generated meshes can be regulated and controlled according to the variation of the flow parameters in the flow field so as to obtain the optimal grid distribution. As an example, taking unsteady transonic flows over oscillating NACA 0012 airfoil, the effects of transformation accuracy on computation accuracy of flow field, and how to control the grid generation corresponding with the variation of the flow parameters are analyzed in detail.The proceeding techniques have been applied to solve Euler equations with succes and the numerical results are satisfactory.

In this paper, LU-ADI scheme and Baldwin-Lomax turbulent model were applied to simulate turbulent transonic flow around twodimensional airfoils. By a series of numerical experiments, it is proven that the method used in this paper is successful and the numerical results are satisfactory.

In this paper, the Computational methods for determing the D-set defined in [1] are Considered, particularly, for the,D-set presented in [2] with more practical value, we give a efficient algorithm and discuss its mathematical principle and implement procedure. Otherwise, Some simple properties aboutδ((P_i^s_i)(i=1,2,…,K)are described and a numerical example .to show the correctness of the procedure is given.

In this paper, the method to determine all parameters a for MCS(x_nΞax_n-1(modM),n ≥ 1) with any Period m is given on the basis of the theorems in[1] We deduce the parameters a from seeking m orders unit root for modul M' (M'=M/d, (x₀,M)=d)and propose the simplified algorithms to do so for the composite m. A numerical example shows the correctness of the method.

The author presended the sufficient and necessary Conditions for MCS With maximum-period δ(M) for given modul M in[1]. In this paper, the result in[1] is extended to any period m(≤δ(M)). At last, an example verifies the correctness of the conclusions given in this paper.

In this paper a mathematical model for the extraction-injection of geothermal reservoir in multiple wells is developed, It silplidies a complicated tridimentionsl heat conduction problem to a bi-dimentional one, The latter is a problem of coupling equations, one of wbich is a one-dimentional heat conduction equation in vertical, and another is a one-dimentional convection-conduction eqyation in horizontal, It deduces several different schemes for coupling solutions as the differential equations are discrtized, and discusses the procecdure of computation, We have prepard the software for drwing the groundwater has been exploited with double wells, three wells and five wells.