This paper considers deployment control of a tethered satellite system (TSS) for rendezvousing a subsatellite with a target object orbiting at a lower altitude. Contrary to past studies that considered a linear quadratic regulator (LQR), in the present study, an LQR-based controller is designed using a pole assignment shift technique, in which the acceleration of the tether deployment or retrieval is treated as the control input and no propellant is used. For simplicity, the TSS is represented as a dumbbell model. The tether tension is calculated based on the tether state variables (tether length, tether length rate, tether angle, tether angle rate, and acceleration of the tether deployment treated as the control input) in order to confirm whether or not tether slack occurs. The experimental results show that the proposed control law contributes to reducing the difference between the desired and experimental tether librational motions.

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Employing airplanes for Mars exploration will lead to novel scientific discoveries. Airplanes can fly over several hundreds to thousands of kilometers and obtain high-resolution data, which is impossible for rovers and satellites to achieve. Such airplanes would require deployment mechanisms to be compact and produce enough lift force to fly in the thin Martian atmosphere. However, existing conceptual design methods do not consider deployment mechanisms. This paper provides the conceptual design method of an airplane intended for Mars exploration that includes a basic estimation of the specifications related to the aerial deployment mechanisms. This conceptual design method deals with two types of deployment mechanisms: a folding type and an inflatable type. The specifications of the airplanes with different deployment mechanisms are compared. The design results quantitatively show that an inflatable type wing is lighter for short wingspan design and a folding type wing is lighter for long wingspan design. The mass of the inflatable wing increases in long wingspan conditions since the required inflation pressure increases owing to the small spar radius and large bending moment. Finally, feasible specifications of the Mars airplane with deployable wings are proposed as a result of the constructed method.

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The effects of radiation heat transfer in rocket engine combustion chambers are studied analytically. The fuel is hydrogen, methane or ethanol, and the oxidizer is oxygen. Radiative heat fluxes are estimated using empirical equations, and convective heat flux is estimated using flux on a flat plate, with modification of circumferential length in convergence to, or divergence from, the throat. The calculated total heat flux including radiation and convection showed reasonable agreement with the flux measured experimentally. The ratio of radiative heat flux to total flux is increased up to 30% in the cylindrical section, whereas it is less than 10% at the throat. The effect of radiation on the total amount of heat transferred to the chamber is remarkable when increasing the length of the cylindrical section and diameter, respectively. It is also made clear that the conventional estimation method of heat flux based on the pipe flow model can estimate larger heat flux in small chambers, in spite of ignoring the effects of radiation.

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The generation of shock waves is inevitable during supersonic cruising, which results in the generation of wave drag as well as sonic boom on the ground. Some innovative concepts, such as the supersonic biplane concept and supersonic twin-body fuselage concept, have been proposed recently to reduce the supersonic wave drag dramatically. In this study, these two concepts are adopted, and then the aerodynamic and sonic boom performance of innovative supersonic transport (SST) wing-body configurations are discussed using numerical approaches. This study is performed to obtain design knowledge for the innovative SST using an optimization method. In this research, the number of design variables is limited to only three in order to obtain fundamental design knowledge of the innovative SST configuration. The three design variables are utilized to deform the wing section shape. The wing section shape of a Busemann-type-biplane/twin-body model is optimized under the conditions of a design Mach number of 1.7 and angle of attack of 2 degrees. The optimized results show the tradeoff relationship between lift-drag ratio and maximum overpressure of sonic boom distribution on the ground. To obtain detailed knowledge of the design space, analysis of variance and visualizations of response surfaces of objective functions are also performed.

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In order to solve the supersonic inlet control problems during supersonic cruising state, an integrated model and inexact one-dimensional search algorithm are devised to design the inlet control law. An online inlet ramp angle mapping model is also proposed using recursive reduced least squares support vector regression and BP neural network methods. Finally, a proportional and integral (PI) controller is designed to control the inlet ramp. The simulation results indicated that an effective inlet ramp controller would greatly improve the engine installed performance and protect the engine from surge and other dangerous operations.

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The laminar to turbulent transition in a pipe flow is studied from the viewpoint of momentum conservation. The transition is presumed to happen downstream of the laminar flow development. The inflow conditions to the transition region are presumed not to change at the moment of transition. The exit pressure is presumed not to change either. In the present model, the turbulent transition is a function of the ratio of the pipe length to its diameter, and becomes larger as the pipe length ratio increases. The natural transition Reynolds number calculated shows reasonable agreement with previous experimental results. The critical transition Reynolds number calculated is 1,752, and is close to the critical number of 1,760 measured previously. The distribution of difference in pressure drop between the laminar and turbulent flows corresponds to the reported location of the puffs and slugs in the Reynolds number under forced transition.

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In this study, we present a potential method to predict the highest surface temperature reached during the heating by Raman analysis for CFRP ablator, which is commonly used in aerospace fields as a material of thermal protection system. The specimens are heated using various treatments such as static heating and dynamic heating by arc jet in inert and active environments followed by Raman analysis. The results obtained in this study indicate that the relationship between the highest surface temperature and the Raman parameters is independent of the heat treatment. The estimation method suggested for the highest surface temperature is based on Raman spectroscopic analysis.

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