Michiko Furudate

Unified Calculation of Hypersonic Flowfield for a Reentry Vehicle

A unified computational fluid dynamics (CFD) method is developed that integrates numerical methods for solving specific problems related to aerodynamic heating phenomena and ablative heatshield. With the use of this unified CFD code, trajectory-based analysis on the aerodynamic heating environment for the MUSES-C superorbital reentry capsule is conducted. Converged solutions can be obtained by loosely coupling the CFD code and the charring materials ablation code within a few iterations

Behavior of Two-Temperature Model in Intermediate Hypersonic Regime

Shock standoff distance for a sphere is calculated to examine the behavior of the existing two-temperature model in the intermediate hypersonic flow regime. Calculations are carried out for three binary scaling parameter values, corresponding to nearly frozen, nonequilibrium, and nearly equilibrium flows, respectively. The obtained shock standoff distances are compared with the experimental data obtained in a ballistic range. The two temperature model reproduces the shock standoff distances in the intermediate hypersonic flows fairly well but tends to lose its accuracy where vibrational excitation occurs but chemical reactions are nearly frozen

Coupled Rotational-Vibrational Relaxation of Molecular Hydrogen at High Temperatures

The internal energy relaxation processes of molecular hydrogen in the temperature range of 1000-50,000 K are investigated by integrating the master equation considering all of the rotational and the vibrational energy levels simultaneously. The state-to-state and the state-specific dissociation rate coefficients used in the calculation are determined by a quasi-classical-trajectory analysis. The obtained rates are validated against the existing experimental data. Recombination rate coefficients for bulk gas are also calculated from the detailed rates and compared with the existing studies. From the calculated evolutions of energies in the rotational and the vibrational modes, the relaxation times of these two modes for the Landau-Teller equation are derived. The obtained relaxation times agree fairly well with the existing experimental data for the temperatures up to 1000 K for the rotational relaxation and up to 3000 K for the vibrational relaxation. It was revealed that these two energy modes are strongly coupled at high temperatures. The effective collision numbers required for equilibration of the two modes are found to vary from 6 to 600. The approximate expressions describing energy relaxation of the coupled ro-vibrational mode are derived for temperatures from 5000 to 50,000 K.

Experimental Investigation of Electrodynamic Heatshield Effect in a Reentry Plasma

We conducted an experimental verification of the electromagnetic effect on the high enthalpy flow which mimics the reentry condition. For this purpose, the expansion tube was made use of, which enables us to generate a hypersonic flow around a model with magnetic field around it. For measurement, the sequential images of the flow around a model were recorded. To investigate the experimental result, we carried out, first, the expansion tube simulation based on the quasi-one-dimensional model. This enable us to get insight into the test flow condition which was produced by the expansion tube. Based on the expansion tube simulation, the test flow time is determined and the experimental flow images during the test flow time was investigated. Our experimental result shows that the shock layer width increases with the magnetic field strength as expected. The numerical MHD flow simulation around the magnetized model shows a good agreement with the experimental results, at least qualitatively.

Calculation of Shock Shapes over Sharp Cone in Intermediate Hypersonic Airflow

Shock shapes over a sharp cone in the intermediate hypersonic e ow regime are calculated to examine the validity of the existing two-temperature thermochemical model. Two different apex angles are considered in the calculations: one with a half-angle of 30 deg and the other with 45 deg. The calculations for these geometries are carried out for several different static pressure values at the e ight velocity of about 3.0 km/s. The calculated shock layer thickness is compared with the corresponding experimental data obtained in a ballistic range. The resultsshowthatthetwo-temperaturemodelwellreproducestheexperimentaldataforthee owconditionsinwhich chemicalreactionsaswellasvibrational excitationsareabsent.However,thecalculated shock layerthicknesstends to be thinner than the experimental data for the e ow conditions in which vibrational excitation begins to occur. It is implied that vibrational relaxation has a close connection with the thinner shock layer. The study cone rms our previous results that the shock layer thickness over a sphere in the same velocity range can be underestimated in the calculation using the existing two-temperature model.

RANS Simulation Using High-Order Spectral Volume Method on Unstructured Tetrahedral Grids

An aerodynamics simulation code for the Reynolds averaged Navier-Stokes (RANS) equations is developed using the spectral volume method for unstructured tetrahedral meshes. The turbulent viscosity is modeled by the Spalart-Allmaras (SA) one-equation model. The developed scheme is validated for turbulent flow over a flat plate and assessed for transonic flowfield over a wing compared with available CFD results. Then, the developed flow solver is applied to obtain complicated flows over high-lift devices. By comparing obtained results with the corresponding experimental data and the available numerical results, the capability to predict complicated flowfields has been favorably shown. And the applicability to large scale industrial problems has been also indicated.

Numerical Computation of Radiative Heating Environment for Huygens Probe Entry Flight

A trajectory-based analysis for obtaining aerodynamic heating environment for the Huygens probe is carried out using the thermochemical nonequilibrium computational fluid dynamics code. Radiative heat transfer is accounted for in computational fluid dynamics calculations where a modern multiband radiation model is employed. In this study, we first compare the radiative heat flux obtained by the ray-tracing approach in three-dimensional space with that given by the tangent-slab approximation, to determine how the difference in obtaining radiative heat flux can alter the overall aerodynamic heating environment. We then explore the radiative cooling effect on surface heat flux through radiation coupled computational fluid dynamics calculations. It is shown that the radiative heat flux value at the stagnation point obtained by the ray-tracing approach becomes about 17-19% smaller than that given by the tangent-slab approximation due to body curvature at all the chosen trajectory points. Furthermore, it is also shown that the radiative cooling effect can reduce the surface radiative heat flux by almost the same amount at the stagnation point when computational fluid dynamics calculation coupled with radiation is conducted. It is therefore confirmed that, even for the relatively lower radiative heating rate such as for the Huygens entry flight, we need to employ the ray-tracing approach instead of the tangent-slab approximation, and also need to account for radiative cooling effect through radiation coupled computational fluid dynamics calculation, to evaluate the surface heating condition accurately.

Nonequilibrium Calculation of High-Temperature Radiating H2-He Flowfield

Chemical kinetics of high-temperature hydrogen-helium gas mixture behind a shock wave is numerically investigated by integrating rate equations for species concentration in time. State-to-state transition rates are used to determine quasi-steady-state rate coefficients for atomic hydrogen ionization. The electron concentrations in front of the shock wave are deduced by solving the radiative heat transfer equation of molecular hydrogen. The computed incubation time of avalanche ionization is compared with the experimental data and those appeared in past studies. It is found that the precursor photoionization and the associative ionization of the molecular hydrogen are important to determine the ionization time behind the shock wave. The present chemical kinetic models are found to reproduce the shock tube experimental data of the ionization time reasonably well.

Calculation of Intermediate Hypersonic Flow Using Multi-Temperature Model

Our previous study revealed two-temperature model reasonably well that the existing reproduced the shock stand-off distance for a sphere in the intermediate hypersonic range, except for the cases where chemical reactions were almost frozen but vibrationally highly excited. In order to explain the cause of distinction, a four-temperature description of the intermediate hypersonic flowfield has been attempted. The model assigns a translational temperature and three different vibrational temperatures, one for each of three molecular species, 02, Nr, and NO. Calculated shock stand-off distances for several conditions, for which the discrepancy was seen in the two-temperature calculation, are compared with the corresponding experimental data measured with a ballistic range and the calculated results using a two-temperature model. The agreement, however, is only slightly improved by the use of the present four-temperature model. A possible explanation of the modest improvement is indicated.

Three-Dimensional Aerodynamics Study for Mars Aeroshell in Nonequilibrium Flow

Three-dimensional calculations are carried out to clarify the effect of high temperature real-gas phenomena on the aerodynamic characteristics of Mars aeroshell geometry. The calculated aerodynamic coefficients are compared with the experimental date obtained in the hypersonic wind tunnel test. Under the frozen flow assumption, the calculated aerodynamic coefficients agree well with the experimental data. The results in the nonequilibrium calculation show that the real gas effects alter the position of recompression zone in the shock layer, consequently the aerodynamic coefficients of the aeroshell.