In this paper, we propose a control design method to suppress flexible appendage vibration while controlling the attitude motion of space structures simultaneously using two different actuators; one a “slow” actuator, and the other a “fast” actuator. Because the frequency of flexible appendage vibration is much higher than that of attitude motion, this system naturally exhibits a two-time-scale dynamic behavior. After describing this system as a singularly perturbed system, the control laws of each subsystem were determined to have the following properties: the slow actuator only controls the attitude motion of space structures, and the fast actuator only suppresses flexible appendage vibration. A composite state feedback control is obtained by combining the slow subsystem and fast subsystem control laws. Simulation results are presented to show the effectiveness of this control design method.

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Comprehensive numerical and experimental investigations of tip vortical characteristics were conducted for lateral tip jet flow over a fixed wing as a step to reduce blade vortex interaction noise. The tip vortex of a NACA0012 blade was measured and visualized for the fundamental study of tip vortical flow, and the results were compared with numerical data as a validation of numerical solvers. Three-dimensional compressible Euler/Navier-Stokes codes were used to calculate the effect of jet flow from the tip of an OLS (modified BHT 540) fixed blade at various freestream velocities and jet conditions. The results show that the jet flowing from the wing tip can diffuse the tip vortex enlarging the core size of tip vortex and weakening its strength. When applied to the blade vortex interaction phenomena, this enlarged and weak vortex can produce a lower pressure gradient on the blade surface, which means that the jet flow can effectively reduce blade vortex interaction noise.

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Aerospike nozzle flow fields are computationally analyzed by Navier-Stokes simulations. The secondary flow from the base surface is induced for the purpose of increasing total thrust. The computed result shows that base pressure efficiently increases when base bleeding is induced from the tip region of the base surface and injected radially inward towards the nozzle axis. The base bleeding induced at the tip of the base and injected horizontally with the axis produces maximum base thrust because the momentum thrust does not produce any divergence loss and the base pressure is kept high. Base bleeding induced against the existing recirculation flow results in lower base pressure distributions. Base bleeding that produces maximum thrust interacts with the exhaust flow and creates shock waves that emanate from the tip of the base. The shear layer is squeezed more towards the nozzle axis when base bleeding is induced and forces the stagnation point to move closer to the base surface.

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The MUSES-C mission is being prepared at the Institute of Space and Astronautical Science (ISAS). The MUSES-C spacecraft is equipped with four ion engines, and usually three ion engines are operated simultaneously. This paper describes the numerical studies of a plume exhausted from the ion engines of the MUSES-C. The direct simulation Monte Carlo (DSMC) method is employed for determining the flowfield of the neutral atoms, and the particle-in-cell (PIC) method is combined with the DSMC method to deal with the motion of the ions. The simulation results for the operation of a single ion engine are compared with experiments conducted at ISAS. It is revealed that the profiles of the ion beam are sensitive to the ion beam divergence angle at the engine exit and the electron temperature of the flowfield. Based on the simulation results of a single ion engine, a three-dimensional DSMC-PIC simulation for the simultaneous operation of three ion engines is carried out, and the interaction between plumes from the ion engines and the behavior of the CEX ions are investigated.

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A numerical tool is used to study the role of an ignition tube in the Hiroshima University ram accelerator (named HURAMAC), a unique rectangular shape ram acceleration tube utilized the optical visualization of flow fields. The finite difference method is used to obtain the solution for the reactive gas flow field. An adaptive multi-level grid system is employed for the purpose of effectively using fine mesh points. A detailed methane and oxygen reaction model is considered for the calculation of combustible mixtures. As shown in the simulation results, the process of shock induced ignition and flame propagation is achieved in the ignition tube. After the projectile is moved into the ram acceleration tube where the gas is diluted with a large amount of carbon dioxide, the combustion becomes weak. According to this simulation, it was found that the flame on the projectile surface becomes thinner and the front shock isn’t strong enough to pass over the shoulder which is the throat of the supersonic diffuser. The results also show that the flame is held at a stable state in the boundary layer of the projectile surface when burning in a low energetic mixture injected into the ram acceleration tube, successfully starting the ram accelerator at an early stage. The result is a sub-detonative operation mode.

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A comparative performance analysis based on design conditions has been made to give normalized algebraic expressions convenient to steady-state temperature and pressure predictions of a loop heat pipe (LHP). In analytical modeling of the LHP, a novel concept of pump efficiency is introduced to define the rate of heat loss or leakage through evaporator capillary wicks. Also introduced are general concepts of evaporator temperature effectiveness, condenser activeness, and subcooler temperature effectiveness. A simplified practical method is used to make a probable pressure loss estimate. The model was then mathematically evolved into a solution algorithm applicable to LHP off-design operation problems. In finding the solution, the radiation temperature is first calculated from specified heat load and sink conditions to give a possible condensation temperature, upon which other state variables are determined. Predicted thermohydraulic states are graphically shown in the figures to provide a better understanding of the operability of a commonly configured LHP. A few ground test results are also provided.

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In order to estimate the intrinsic accuracy of satellite reentry predictions, the residual lifetimes of 11 spacecraft and five rocket bodies, covering a broad range of inclinations and decaying from orbit in a period of high solar activity, were determined using three different atmospheric density models: JR-71, TD-88, and MSIS-86. For each object, the ballistic coefficient applicable to a specific phase of the flight was obtained by fitting an appropriate set of two-line orbital elements, while the reentry predictions were computed approximately one month, one week and one day before the final orbital decay. No clear correlation between the residual lifetime errors and satellite inclination or type (spacecraft or rocket body) emerged. JR-71 and MSIS-86 resulted in good agreement, with comparable reentry prediction errors (∼10%), semimajor axis residuals, and ballistic coefficient estimations. TD-88 exhibited a behaviour consistent with the other two models, but was typically characterised by larger reentry prediction errors (∼15–25%) and semimajor axis residuals. At low altitudes (<250 km), TD-88 systematically overestimated the average atmosphere density (by ∼25%) with respect to the other two models.

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The prototype of a solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft was completed and tested. The prototype has 10×10 φ0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 20×22 mm substrate. To realize such a dense array of micro-rockets, each ignition heater is powered from the backside of the thruster through an electrical feedthrough which passes along a propellant cylinder wall. Boron/potassium nitrate propellant (NAB) is used with/without lead rhodanide/potassium chlorate/nitrocellulose ignition aid (RK). Impulse thrust was measured by a pendulum method in air. Ignition required electric power of at least 3–4 W with RK and 4–6 W without RK. Measured impulse thrusts were from 2×10⁻⁵ Ns to 3×10⁻⁴ Ns after the calculation of compensation for air dumping.

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