A fully reusable two-stage-to-orbit launch vehicle with ethanol-fueled rocket-based combined cycle (RBCC) engines has been studied in Japan as a promising option for future space transportation system. In this paper, a conceptual design study of such a vehicle is conducted using multidisciplinary design optimization (MDO) techniques in order to clarify a technology goal for related technology development activities. An MDO framework composed of coupled analysis disciplines (vehicle geometry, mass property, aerodynamics, propulsion, and trajectory) is constructed. In particular, consideration is given to the development of a simplified numerical model for evaluating the airframe-propulsion integration that can be incorporated into MDO studies, in contrast to costly CFD simulations. Vehicle design and ascent trajectory are then simultaneously optimized with the aim of minimizing the gross mass of the mated vehicle (booster and orbiter). The gross mass of the obtained optimal design is 581 t for the assumed mission of transporting an 800 kg payload into a low Earth orbit. A detailed inspection of the solution reveals that an external nozzle of the engines enhances not only the propulsion performance, but also longitudinal static stability of the vehicle during hypersonic flight.

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In medium-Earth-orbit, geostationary, and deep space missions, the effects of solar radiation pressure (SRP) on the motion of a spacecraft are dominant relative to other perturbations. Recently, several spacecraft have actively used SRP torque for attitude control, and the importance of calculating the SRP torque accurately is increasing. The present paper introduces a newly developed SRP analysis tool called the FEM radiation analysis tool (FRAT). This tool uses an element-based ray tracing strategy, and each element is assigned different optical parameters derived from its material properties. This tool calculates the SRP applied to each mesh element taking shadows into consideration. In the present paper, the accuracy of this calculation is evaluated using flight data of the HAYABUSA 2 asteroid exploration spacecraft. Then, two application examples in which FRAT has played an important role are presented. The first example involves HAYABUSA 2, the attitude of which has been stabilized by actively using the SRP. In the second example, the shape of the sail of IKAROS, a solar sail spacecraft, was estimated using FRAT in order to reproduce the flight data.

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An extensible shearable elastica is a rigorous mathematical model of the Timoshenko beam, of which the cross-sections remain planar, but not necessarily normal to the beam axis after deformation. First, the principle of virtual work for the extensible shearable elastica expressed in terms of the extensional and shear strains of the axis and the rotation of the cross-section in Engesser's approach is derived from the virtual work in three-dimensional solid mechanics. Then, utilizing linear constitutive equations between generalized stresses and strains, we derive the principle of stationary potential energy, also expressed in terms of the extensional and shear strains and rotation. Finally, from the criterion of Trefftz on the second variation of the potential energy, we obtain the buckling equations for the extensible and shearable elastica, which show the effect of the axial and shear stiffness on the buckling load for a cantilever elastica subjected to compressive end load.

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The author and K. Kondo derived an expression for the potential energy of extensible shearable elastica based on Engesser’s approach in a previous paper. In this paper, a simple method to calculate the post-buckling deformation of extensible shearable elastica is proposed. The potential energy is directly minimized by the “Solver” optimization tool of MS-Excel. The compression buckling loads calculated by the method are consistent with the buckling load equation in the previous paper. The relationship between applied load and deflection is calculated for beams with representative combinations of axial stiffness and shear stiffness. Snap-through behavior is observed in the load-displacement curves of beams with low axial stiffness.

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Supersonic drag reduction performance using repetitive pulse energy depositions over blunt bodies was experimentally studied under two Mach numbers. The normalized drag reduction and energy deposition efficiency of Mach-1.92 over a 10-mm-dia. blunt-cylinder model were 8% and 1.2 at most, respectively. On the other hand, these values at Mach-3.20 over the same model were 22% and 6.2, respectively. The shock-wave deformation period using single-pulse energy deposition at Mach-3.20 was 64 μs. This duration was shorter than that of 80 μs at Mach-1.92, but the deformation magnitude on the model center axis of 40% at Mach-3.20 was larger than that of 15% at Mach-1.92. These experimental characteristics were consistent as solutions of the Riemann problem. Moreover, a drag reduction performance was much improved with a larger model diameter of 20 mm. Therefore, it has been experimentally demonstrated that the drag reduction performance due to energy deposition improves much at a high Mach number and with large model dimensions.

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This study investigated the anti-spoofing capability of a Commercial-Off-The-Shelf (COTS) Global Positioning System (GPS) receiver and proposes a spoofing detection method with minimum user complexity. A spoofing test environment with a GPS simulator was developed to test a conventional GPS receiver with an in-line radio frequency (RF) signal. The tests on a geodetic-grade receiver revealed a vulnerability in the COTS GPS receiver against spoofing attacks. Receiver Autonomous Integrity Monitoring (RAIM) was not capable of handling a spoofing attack sufficiently, but could manage a low-level attack. Thus, a spoofing detection method based on the simulation results is proposed as a feasibility check.

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Three different types of high power Hall thrusters—anode layer type, magnetic layer type with high specific impulse, and magnetic layer type with dual mode operation (high thrust mode and high specific impulse mode)—have been developed, and the thrust performance of each thruster has been evaluated. The thrust of the anode layer type thruster is in the range of 19–219 mN, with power in the range of 325–4500 W. The thrust of the high specific impulse magnetic layer type thruster was 102 mN, with specific impulse of 3300 s. The thrust of the bimodal operation magnetic layer thruster was 385 mN with specific impulse of 1200 s, and 300 mN with specific impulse of 2330 s. The performance of these thrusters demonstrates that the Japanese electric propulsion community has the capability to develop a thruster for commercial use.

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