Igor Men'shov

Hybrid Explicit-Implicit, Unconditionally Stable Scheme for Unsteady Compressible Flows

A variable explicit-implicit scheme was proposed by Collins et al. [(CCG) scheme] for one-dimensional hyperbolic systems. By suitable blending of explicit and implicit schemes, this approach aims to ensure a maximum norm diminishing (MND) property at all Courant numbers. Although proven for linear equations, however, the CCG scheme, fails to maintain this property unconditionally for non-linear equations and requires a specific time-step restriction. The remedy for this limitation of the CCG scheme is shown, and a new general method for the design of unconditionally MND hybrid schemes for nonlinear hyperbolic equations is proposed. An extension to the compressible Navier-Stokes equations is also discussed. The method is tested and verified by calculation of the Burgers equation on a highly stiff nonuniform grid. It is also applied to simulate the reflection of a normal shock wave from the end wall of a two-dimensional channel, which is attended with unsteady, viscous interaction processes including shock bifurcation

Numerical simulations and experimental comparisons for high-speed nonequilibrium air flows

A computational fluid dynamics (CFD) technique is employed to study hypersonic high-enthalpy air flows around blunt bodies with the purpose of predicting convective heat transfer on the body surface for a range of flow velocities relevant to suborbital flight of re-entry vehicles such as the Space Shuttle Orbiter (USA), and the Buran (Russia). The method uses Parks two-temperature model for the description of thermochemical nonequilibrium processes in high-temperature air and solves the full Navier–Stokes equations for a model of multicomponent reacting gas mixture in the finite volume formulation. The calculations performed in this research are intended to simulate some experiments carried out in the high-energy shock tunnels of the DLR, Germany, and the CALSPAN, USA, where the heat flux distribution over a model surface was measured at several freestream conditions related to the range of velocities mentioned above. The main emphasis is on comparing numerical and experimental results in order to verify adequacy of the heat flux data predicted by the CFD technique for suborbital flight speeds of re-entry vehicles.

Detached-Eddy Simulation of Three Airfoils with Different Stall Onset Mechanisms

The detached-eddy-simulation (DES) method is applied to calculate prestall and poststall aerodynamic characteristics of three airfoils that have different flow separation patterns at the stall regime. The DES method combines strong points of two turbulent models that are based on time-averaged and space-filtered Navier-Stokes equations, respectively. Specifically, we employ the Spalart-Allmaras DES approach, which has a Reynolds-averaged Navier-Stokes formulation near wall and reduces to a subgrid turbulence model, if we go away from the wall. The comparison between experimental data and numerical results is presented, which shows that the stall angle and prestall and poststall aerodynamic characteristics can be predicted by the present method accurately for two airfoils: NACA63 1 -012 and NACA63 3 -018. However, a discrepancy between numerical and experimental data is observed for the third airfoil: NACA64 A -006. The reason for this discrepancy and also possible improvements of the numerical model are discussed.

Sound Emission from the Helical Mode of Supersonic Jet

A numerical simulation of sound emission from the helical mode of an under-expanded supersonic jet has been implemented. This simulation is still a challenging problem, since aeroacoustic noise from a supersonic jet is closely associated with the three-dimensional, complicated flow structure including various scale vortices, shock cell structure, and a turbulent mixing layer, which is very sensitive to the accuracy of computation. In the present study, an accurate numerical code is used; that is, the fourth order WENO scheme in spatial discretization to reduce numerical dissipation, and the third order Runge-Kutta method in time integration. The computed sound frequency of screech tone and the instability mode agree with the experimental data measured at our laboratory. The mechanism of sound generation along with the flow field is explained in this paper.

Numerical Investigation of SRB Ignition Overpressure

The numerical simulation of SRB ignition overpressure was performed to clarify the mechanism of ignition overpressure (IOP) under the same conditions as a solid rocket booster (SRB) of the Space Shuttle Columbia for the STS-1 mission. Two governing equations: 1) the Euler equations and 2) the linearized Euler equations (LEE) are employed to capture pressure fluctuations without attenuating their amplitudes as much as possible. It is assumed that gas is calorically perfect and that combustion gas, whose specific heat ratio is 1.18, fills the overall computational region. Computations are carried out for the axisymmetric case, where the ground is horizontal, and for the three-dimensional case, where the ground is eiter horizontal or oblique. The time history of pressure fluctuation shows reasonable agreement with the measured data on the SRB for the STS-1 mission. Moreover, the detailed mechanism of IOP was made clear, and the merit of using an oblique wall in a launch pad was assured to reduce the effect of overpressure waves on the body of spacecraft.

CFD Analysis of Aerodynamic Interference between a Delta Wing and a Hemisphere-cylinder

A computational study of aerodynamic interferences around a conceptual Two-Stage-To -Orbit (TSTO) model in hypersonic flow is presented. The TSTO model proposed in our laboratory consists of a delta wing as a booster and a hemisphere-cylinder as an orbiter. The objective of this study is to analyze the flow field in detail using CFD. Firstly, we made comparison between the computational results with experimental data and good agreement is confirmed. Using the computational results, the flow field and the mechanism of high heat flux on the nose of the orbiter are analyzed. Results show that the heat flux increment is dominated by the distance between the bow shock and the hemisphere-cylinder.

Panel Flutter Analysis with a Fluid-Structure Coupled Scheme

A numerical scheme to treat the problem of fluid-structure coupling has been developed in the present study. A high order scheme based on the finite volume method is used for fluid calculation, along with the finite element method for structure calculation in order to make more detailed analysis of flutter phenomena. Time evolution using the Runge-Kutta method has been applied to the problem of panel flutter that is a coupled problem including fluid and structure. In the low supersonic region, the flutter boundaries were computed, compared with other computational and some experimental results. In the case of M=1.2, flutter shows a rapid increase in amplitu de into a limit cycle oscillation with the first mode due to the nonlinear characteristics of aerodynamics force.

Aeroacoustic Simulation around Automobile Rear-view Mirror

A method for performing aeroacoustic simulation, where the acoustic field is split from the flow field, is applied to the flow around an automobile rear-view mirror placed on a large solid plane. The flow can be obtained by performing large eddy simulation for the incompressible flow equations on colocated grids, while the acoustic equations are defined by the difference between the compressible and the incompressible flow equations, and solved by the finite volume method with the fourth order WENO scheme. The non-reflecting boundary conditions with a perfectly matched layer are applied to the outer boundary. Calculated results show clear sound emissions, where acoustic streamlines as propagation path are depicted based on acoustic intensity vectors. In this paper, two kinds of mirror shape, i.e., the eases with and without a strong vortex at the tip of a rear-view mirror are compared. Calculated results show some differences between them in the acoustic field. The sound pressure related with the pressure frequency of the generated vortex is increased in all directions, though the propagation path dose not show a large difference except near part of the body surface.

Unsteady-Nonequilibrium Shock Wave / Boundary Layer Interaction with Explicit/Implicit Hybrid Approach

The flow fields of a shock tube and a nozzle starting process for both perfect gas and flow with thermochemical nonequilibrium has been simulated. This flow is produced in high enthalpy impulse facilities such as free piston shock tunnel. The governing equations are the axisymmetric, compressible Navier-Stokes equations. For the case of thermochemical nonequilibrium, Park’s two-temperature model, where air consists of 5 species, is used for defining the thermodynamic properties of air as a driven gas. The numerical scheme employed here is the hybrid scheme of explicit and implicit methods, which was developed at our laboratory, along with AUSM + to evaluate inviscid fluxes. In the present simulation, the Mach number of an incident shock wave is set at Ms = 2:6;5:3. The results clearly show the complicated shock wave/ boundary layer interaction in the part of shock tube. These suggest that the retention of gas accounts for the growth of bifurcated shock structure. The discharge of vortex from the bifurcation surpresses the magnification of size of the bifurcation.

Numerical Study of Unsteady Shock Waves in Hypersonic Nozzle Flows

The flow field with thermal and chemical nonequilibrium in a hypersonic nozzle has been simulated. This flow is produced in high enthalpy impulse facilities such as free piston shock tunnel. The objectives of this study is make clear the complicatedunsteady process of the nozzle starting. The governing equations are the axisymmetric, compressible Navier-Stokes equations. In this study, Park’s two-temperature model, where air consists of 5 species, is employed for defining the thermodynamic properties of air used as a driven gas. The numerical scheme is a hybrid scheme combining explicit and implicit methods, which were developed in our laboratory, implemented with the AUSM+ to evaluate inviscid fluxes. In the present simulation, the Mach number of an incident shock wave is set at M, = 10.0. It corresponds to a specific enthalpy, h,, of 12MJ/kg. The results clearly show the flow structures around shock waves: viscous interaction with the wall of a shock tube and the initial stage of nozzle starting process. They also suggest that the phenomenon of a nozzle melting might be associated with a flow separation a t the nozzle inlet.