Studies On Hypersonic Multi-field Coupled Computation And Its Acceleration Using GPU

Hypersonic flow with shock-boundary layer interaction, high temperature flow and othercomplex phenomena is significantly different from low speed flows, which bring lots of difficultiesand new problems for vehicle design, especially in aerothermoelastic behavior. Solving theseproblems must take into count the effects of high temperature, boundary layer, structure thermalconduction and thermal radiation, which is in excess of tridiantional aeroelastic problems. Thesemulti-field coupled and high arithmetic intensity problems are depend on the cooperation ofcomputational fluid dynamics (CFD), computational structure dynamics (CSD) and computationalthermal dynamics (CTD), and produce a big challenge for analysis and computing capacity. Thecontributions to the state-of-the-art made in this dissertation are summarized below:A details description of the finite speed non-equilibrium chemical reaction Navier-Stokesequations is present and computational methods of gas transport properties and chemical reactionsource term are provied. The formulas of explicit and implicit time discretisation method are derived,the full system Jacobi matrix which contains convective/viscous/source terms is also provided. Acomputer program, which using Roe and AUSM+-up schemes and Menter's SST two equationsturbulence model and implicit LUSGS method, is developed for solving hypersonic CFD problems,and also verified by some test cases.Based on the features of the GPU architecture, an implicit data-parallel scheme has beendeveloped for solving CFD problems. The presented method is applicable to structured andunstructured mesh and uses upwind scheme to achieve more accurate results. The method has beenimplemented on NVIDIA GTX280GPU by employing CUDA technology and compared with IntelCore2Quad3.0GHz CPU. The results indicate that the implicit scheme proposed in this paper is6times faster than the explicit scheme with same hardware and the computation is speed-up to28.7x byusing GPU and implicit scheme, which will be more efficient for larger scale problems. At last, theresults provide good agreement with the existing experimental data.Based on analysis of geometric relationship and interpolation features, an improvement forpresent transfinite interpolation (TFI) method with a rotation correction is proposed to solve theorthogonal problem with large mesh deformation. The computational results of typical two and threedimensional viscous grids indicate that good orthogonal and smoothing properties can be achieved byrotation correction for large mesh deformation. In addition, the computational efficiency is slightlydecreased than the traditional TFI method, but improved by1or2orders of magnitude compared tothe spring analogy method.The thin-plate spline (TPS) or infinte-plate spline (IPS) methods are not suitable for complexstructure, and have no physical meanings for temperature interpolating. A new interpolating methodusing high order isoparameter finite element shape function is presented for solving these problems,and validated by some test cases. A local conservative remapping method is presented for thermal fluxand aerodynamic loads interpolation, based on analysis of the element features of finite element andfinite volume. It's availability is also identified by some test cases.A multi-field coupled computing platform using multi-zone iteration method is developed for solving multi-disciplinary problems. Shared memory method is employed for faster data exchange forgeneral finite element analysis (FEA)/computational fluid dynamics (CFD) software. The problems ofconjugate heat transfer for a cooled converging-diverging nozzle and a cylindrical leading edge inhypersonic flow are studied. Effects of mesh density, nonlinear material properties and radiationeffects are considered during the computation,and the results indicate good agreement with theexisting experimental data. The relationship between stagnation temperature, cooling power and thethickness of nose thermal protection structures (TPS), which resemble X-34hypersonic vehicle, underhypersonic cruise condition are emphatically investigated. The results indicate that the thicknessvariations have much less influences on stagnation temperature, while the cooling power dropssharply as the thickness increases. Furthermore, the nonlinear material emission properties havesignificant influences on the analysis results.A coupled CFD/CSD method was used to solve supersonic and hypersonic panel flutter problemin time domain using Euler and Navier-Stokes equations. Flutter dynamic pressure was calculatedunder different boundary layer thickness and Mach number; the results show that boundary layerthickness has a large stabilizing influence on the flutter of flat panels. The effect on flutter dynamicpressure is maximum near Ma=1.2and decrease rapidly with increasing Mach number, which agreedwell with experimental data. The dynamic pressure is20%higher than the Euler results at Ma=8, theinfluence of turbulent boundary layer thickness can not be neglected at hypersonic flow.Aeroelastic behavior of a typical double-wedge airfoil in hypersonic flow was investigated. Theeigenvalue method using3rdpiston theory, Euler equations, Navier-Stokes equations with adiabaticand isothermal wall boundary were employed to determinate the flutter boundary under differenceflight altitude. The results indicated that Euler equations is very closed to3rdpiston theory, butsignificant difference from Navier-Stokes, and result with difference wall temperature conditaionusing Navier-Stokes is agree well with each other. This conclusion results from two reasons:shock-boundary layer interaction changes the characteristics of flow field, and the thickness ofboundary layer modifies the geometric shape indirectly. Therefore, the viscous effects play a key rolein this aeroelastic problem and can not be neglected.Analysis of aerothermoelastic behavior for the nonsymmetric wing structure of typicalairbreathing hypersonic vehicle was accomplished using multi-field coupled method. It can be dividedinto static aerothermoelastic trim and transient aerothermoelastic response. A new iterative methodusing transient coupling method instead of steady method is present to avoid some numericaldifficulties of convergence. The free-stream condition is10km standard atmosphere. Ideal gas andnon-equilibrium chemical reaction gas model are adopted, piston theory and Euler equations are alsoemploy for comparsion. The results indicated that flutter speed calculated with ideal gas is6%higherthan non-equilibrium chemical reaction Navier-Stokes equations. A very large error exists betweenEuler/3rdpiston and Navier-Stokes. The main reason is the modification of structure dynamicproperties produced by the thermal-stress and variation of materials properties under high temperature.The non-equilibrium chemical reaction Navier-Stokes equations must be used for thisaerothermoelastic problem.