Thermal control materials covering a spacecraft for Mercury orbiter mission are required to have stability against the severe thermal environment. This paper describes a prediction method for the temperature dependence of the total hemispherical emittance and the incident angle dependence of the solar absorptance α of multilayer thermal control materials. The present method is performed by using data for the optical constants of thin polyimide films and vapor-deposited metal in the wavelength region from 0.25 to 100μm. This paper provides the values of in the temperature range from 173 to 773K and α in the incident angle region from 0 to 90° of three types of multilayer thermal control material, which is based upon a thin polyimide film coated with aluminum on the back surface. The predicted results are compared with the measured ones. These results agree well each other on the whole. According to these agreements, it has been demonstrated that this prediction method is an effective means to predict thermal radiation properties of multilayer thermal control materials.

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Effects of streamwise vortices on enhancement of supersonic mixing and combustion are examined in a scramjet combustor whose main flow is a Mach 2.45 vitiation air stream with a total temperature of 2200K. The fuel injector strut called “Alternating-Wedge Strut (AW-Strut)” is used to generate streamwise vortices and to inject hydrogen fuel into their core region. For comparison, a generic strut, “MO-Strut,” without streamwise vortex generation is also examined. Direct visualization of flame shows rapid formation of streamwise vortices and ignition/combustion within the vortices already upstream of the strut trailing edge. Wall pressure measurements along the combustor show that for the case of AW-Strut the wall pressure rise due to combustion is almost twice that of the MO-Strut case. Gas sampling at the combustor exit demonstrates that much more uniform fuel/air mixing is obtained for the case of AW-Strut. The present results verify the phenomenal performance of the supersonic streamwise vortices for enhancement of supersonic mixing and combustion.

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Heat flux on scramjet combustor wall is measured experimentally with thermocouples in a Mach 2 supersonic combustion wind tunnel, and numerical simulations are conducted to discuss the relation between combustion mode and the heat flux profiles. The combustion in the scramjet combustor is categorized into four modes; blow-off mode, weak combustion mode, intensive combustion mode and thermal choke mode. The heat flux profile in each mode reflects the corresponding characteristic flow pattern. If there is no combustion, the wall heat flux can be estimated by Reynolds analogy. When the fuel is injected into the main flow, the heat flux profile is influenced by the behavior of the free fuel plume in the absence of the shock wave system, whereas it is influenced by separation bubbles and/or large distortion of the main flow in the presence of the shock wave system.

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This paper reports results of an improved model of “ionospheric hole” caused by rocket exhaust injected into the F-region. We have investigated a practical “ionospheric hole” model by which a proper strategy of rocket launch can be examined from a view point of protection of the Earth’s environment. The analytical results of preceding models were published about 8 years ago. The model well included the trajectory of launch rocket and could evaluate the effect of rocket exhaust on the Earth’s environment for various trajectories. The analytical results, however, didn’t quantitatively agree with the observation data very well. The new model includes the effects of the exhaust velocity of effluents and the neutral wind that preceding model didn’t take into account, and the boundary conditions and parameters are improved. The analytical results by this improved model well agree with the observation data for Skylab–I and HEAO–C launching experiments not only qualitatively but also quantitatively.

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In the numerical simulation of hypervelocity impacts (HVIs) by means of a conventional SPH method, the loss of interactions among the particles causes numerical fractures. To prevent the numerical fractures, a 2-dimensional SPH method is improved using new particle generation and particle reducing techniques. In the new particle generation technique, new particles are generated in the case when the area with no particles or the distance between the adjacent particles exceeds a certain value. In the particle reducing technique, which is used to avoid the overabundance of particles, two particles approaching each other are combined to one when the distance between them attains a certain value. The perpendicular HVIs of a polyethylene projectile on an aluminum plate target are simulated numerically. Through the numerical results, it is shown that the improved SPH method can prevent effectively the numerical fractures without the overabundance of particles. The responses of oblique HVIs are also shown.

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It is known that when a rotor is flying over the ground at low speed, a ground vortex may appear in front of the rotor. The ground vortex, which can be formed by the interaction between the rotor downwash and the incoming flow in close proximity to the ground, will change the performance of the rotor. The flow environment becomes much more complex compared with that of O.G.E. flight. It is clearly necessary to investigate the basic physical phenomenon of the ground vortex and its effect on the aerodynamic characteristics of the rotor. Therefore, numerical study of the flowfields around a rotor in forward flight near the ground at low speed are carried out by solving the CFD code based on unsteady 3D compressible Euler equations with a moving grid system. In this paper, we present typical numerical results with respect to the formation of ground vortex, the induced velocity distribution on the rotor disc, and describe the effect of the rotor height and advanced ratio on the rotor thrust.

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