Aero-optic effect is investigated in supersonic/hypersonic flow. Flat plate, ramp and cave model in supersonic/hypersonic flow condition are studied with large eddy simulation (LES). Turbulent density pulse is predicted to verify coefficient of density pulse model-mixing length (ML) model. It indicates that LES method simulates turbulent density pulse well in flat plate boundary layer flow, oblique shock wave/boundary layer interaction and strong separation flow. Density pulse data obtained by LES is used to verify ML coefficient and set a quantitative result. The modified ML was used to predict aero-optical effect on missile successfully.

Heat transfer distribution is analyzed on a rough wall in high speed turbulent compressible flow via computational fluid dynamics(CFD) method and analytical correlations, focusing on heat transfer with different roughness number and roughness element shape. It shows that, in all cases, heat flux augmentation predicted with CFD increases with reduction of roughness element density and levels off after roughness shape density is small, which differs with data of three analytical correlations. Predicted heat fluxes are same if same roughness element density and equivalent height are imposed on analytical correlations distinguishing from tendency of CFD results, which changes with roughness element shapes.

Three-dimensional Navier-Stokes equations in high temperature real gas model and perfect gas model are solved for hypersonic entry process. Good agreement is achieved between numerical results, reference values and flight data of Viking, which validates physical-chemical models and numerical methods. Aerodynamic prediction of Mars Pathfinder proves that results of real gas model are close to LAURA. Perfect gas model with effective specific heat ratio is capable of computing lift and drag coefficients. At 2° angle of attack, MPF exhibits static instability along trajectory, which perfect gas model failed to catch on. It is considered that main reasons of static instability are sonic line shifting around shoulder in windward, subsonic area varying and subsonic bubble occurring in leeward. They result in different pressure change in shock layer and in transportation process from expansion zone to upstream.

Test particle Monte Carlo (TPMC) method is presented. Return flux on four geometric surfaces due to ambient scatter of outgassing molecules is simulated. Return flux ratio (RFR) obtained for flow past a sphere is in good agreement with DSMC results. RFR on outgassing and freestream conditions for flows past three geometric bodies, including a circle flat plate, a convex and concave hemisphere, is investigated. RFR for flows past a circle flat plate and a concave hemisphere is much greater than that for flows past a convex hemisphere. Outgassing molecules collide on outgassing surfaces directly forming direct flux contamination for flows past concave surfaces, which is much greater than RFR. Thus, using convex outgassing surfaces in spacecraft design can decrease return flux contamination effectively.

Numerical analysis is made to study thermodynamic models for surface thermal environment of Mars entering vehicles. It shows that flow in stagnation region stays almost in thermal equilibrium. As flight altitude rises,flow changes to thermal nonequilibrium. Heat transfer rate keeps the same in different thermodynamics models,if non-catalytic condition is set on the wall. In simulations with full-catalytic wall conditions,heat transfer rate in thermal non-equilibrium model become higher,due to difference of characteristic temperature in chemical reactions. The stronger the effect of thermal non-equilibrium is,the greater the characteristic temperature change and the heat transfer rate will be.

A three-dimensional computational fluid dynamics algorithm is developed to study chemical nonequilibrium flowfield in a standard model, ELECTRE, for which flight data exits. A finite rate chemistry model based on Dunn and Kang's model is implemented. An extensive investigation on catalysis modeling relevance simulations with non-catalytic wall conditions as well as with full catalytic boundary conditions is made. The method developed is used to compute detailed flow features of hypersonic flow around forebody of a sphere-cone vehicle under different attack angles and altitudes. It is shown that reacting gas numerical results are consistent with flight data. It shows that real gas effects are prominent in thin shock wave layers closing to the wall surface. It makes outstanding distance of shock wave short. Heat flux under full catalytic wall boundary condition is higher than that under non-catalytic condition. The greater the attack angle,the more obvious this variance and the less electron number density on the wall. The higher the flight altitude,the lower dissociation of oxygen and nitrogen,and the lower heat flux on stagnation point.

Hypersonic flows around a blunt body are studied with numerical simulation and theoretical analysis.Correlations between plasma prediction and gas components are obtained.It shows that it is Mach number which affects the peak number density of electrons.Numerical simulation agrees with theoretical analysis well.Differences between two gas models get greater as Mach number increases.The differences follow same trend at back taper with stagnation region.A 11-species chemical model should be applied to increase accuracy when reentry capsule flight at height of 60 km and Mach number is over 23.

Three-dimensional Navier-Stokes equations in real gas models are solved with a parallel code to analyze aerodynamic characteristics of Mars Science Laboratory in hypersonic entry in Martian atmosphere. Good agreement between numerical results and flight data of Viking validates physical-chemical models and numerical methods. It shows that impacted by real gas effect,shock layer thickness is reduced; drag coefficient rises,lift coefficient is almost unchanged. Difference of trim angle between real gas and perfect gas is about 2.2°; As keeping altitude,greater Mach number results in greater drag and pitching moment coefficient. Difference of trim angle varies from 1.6° to 2.6°. Increasing Mach number enhances real gas effect. As keeping Mach number,increasing altitude weakens chemical reactions behind the shock,but it has weak influence on aerodynamic coefficient.