When a spacecraft reenters the Earth's atmosphere, it becomes overheated by the strong shock waves generated ahead of the vehicle. To protect the vehicle from severe aerodynamic heating, one can apply a thermal protection system using a technique known as the ablation method. However, the properties and behavior of ablation remain poorly understood, necessitating further research. The present study conducts ablation experiments on silicon carbide using two kinds of narrow band-pass filters with different measuring wavelengths. During the ablation, the radiation intensity distribution was calculated from the spontaneous emission images of the test piece acquired through each filter. Based on the relationship between the radiation intensity distribution ratio and temperature, the estimated surface temperature of the test piece after 10 s of heating ranged from 1,800 K at the rear to 3,000 K at the tip.

This study considers the non-linear optimal control of spacecraft formation flying in eccentric orbits. A control law based on the Tschauner-Hempel equations is usually used to reconfigure a spacecraft formation. However, the solutions of the Tschauner-Hempel equation are no longer valid when the formation size is large. This study focuses on the effects of the relatively long-distance between two spacecraft. Accordingly, a state equation is obtained by taking the differences of the orbital elements between the two spacecraft as state variables, with long-distance effects as perturbations. The quasi-optimal control input is then analytically derived by applying Hamiltonian perturbation theory, including the long-distance effects. The input can be easily calculated when the initial and target states are given. The usefulness of the control input is shown by numerical simulations from the viewpoint of the accuracy of the relative position and velocity.

To simulate the motion of a rotor while accelerating during slow descent, a simulation model that corrects errors in data obtained experimentally is developed in this paper. The effect caused by acceleration is expressed as errors in induced velocity and external thrust force, and defined by two correction factors. Best-fit correction factors that provide good approximations for experimental flight data of a bamboo dragonfly are searched and the effectiveness of the method proposed is shown.

This paper proposes a new strategy for improving deployment repeatability in deployable structures using kinematic couplings. In this strategy, the order of the pressing and the pressing load applied to the coupling are controlled to meet the necessary conditions. The basic theory for latch completion conditions is derived based on our previous study, which revealed the mating completion conditions of kinematic couplings. An experimental model that consists of two circular plates and three parts of two-degrees of restraint (2-DOR) (balls with grooves) with pressing load mechanisms using tension springs is developed. The shape repeatability in the experimental model is evaluated by measuring the relative displacements of two plates in each installation while changing the pressing order and load applied. The effects pressing order and load applied on shape repeatability are evaluated through these experiments, and the effectiveness of the strategy for pressing load adjustment is demonstrated.

A 1,000-hours preliminary life test of a 6-kW-class xenon Hall thruster for Engineering Test Satellite-9 (ETS-9) was conducted by using a bread board model thruster. Thrust degraded by approximately 2% before and after the test, which is mainly attributed to erosion and contamination of discharge channel made of boron nitride. The thruster was disassembled at an accumulated operational time of 652 hours to find 2.3-mm erosion at the channel's downstream surfaces. From 652 to 1,012 hours, low erosion rates about 10 μm/khr were found for the channel and the cathode; exception was found in the case of the insert, whose erosion rate was relatively high as 60 μm/khr. Thruster's estimated lifetime extrapolated from the test is more than 10,000 hours, which is satisfactory for all-electric propulsion satellites.