Hayabusa2 arrived at the asteroid Ryugu in June 2018, and as of April 2019, the mission succeeded in conducting two rovers landing, one lander landing, one spacecraft touchdown/sample collection and one kinetic impact operation. This paper describes the initial nine months of the asteroid proximity operation activity of the Hayabusa2 mission, and gives an overview of the achievements thus far. Some important engineering and scientific activities conducted synchronously with spacecraft operations in order to complete all planned operations in time against unexpectedly harsh environment of Ryugu are also described.
Estimating the regolith properties of Phobos’ surface is of critical importance for the landing and performance of the Martian Moons eXploration (MMX) sample return mission. Regolith physical properties such as strength, regolith thickness, and the presence of regolith layers are related to morphologies of superposing impact craters. However, the accurate depths of Phobos’ sub-kilometer-diameter craters including irregularly shaped craters have not yet been fully characterized. Here, by using our high-resolution (20 m/pixel) digital elevation model of the nearside (or the sub-Mars side) of Phobos, we investigate the topographic profiles of the sub-kilometer craters. We confirm the presence of crater rims, and bowl-shaped, central-mound, and flat-floored crater geometries. The topography of one flat-floored crater is consistent with a boundary of regolith layers at a depth of ~160–180 meters. Morphometric measurements of 35 sub-kilometer craters show that their depth-to-diameter (d/D) ratios are in the range of 0.037 and 0.174 (mean value = 0.089, median value = 0.093). This suggests either a surface layer composed of rocky debris that effectively dissipates impact energy and causes a reduction in crater depth, or subsequent resurfacing events changed the original crater topography.
A newly constructed model calculates the starting position of the pseudo-shock, or shock train, in the divergent ducts. The model is based on the conservation laws of mass, energy and momentum of the inflow and outflow into/from the pseudo-shock region, and further uses a sonic condition in the core flow of the pseudo-shock. The model adopts an empirical growth rate of the low-velocity region, and the divergent angle of the nozzles becomes a design parameter. The lengths of the pseudo-shocks calculated show reasonable agreement with the experimental results. Pressure distribution with quick increase near the starting position is simulated. The supersonic core flow cross-section becomes smaller than that at the starting position of the pseudo-shock. The pressure at the sonic point rises to 0.7 of that at the duct exit. The length of the quick pressure-increase region becomes longer as the Mach number increases at the starting position.
The University of Tokyo has proposed a water resistojet thruster with a high certainty of liquid–vapor separation and low power consumption. In this propulsion system, liquid water is periodically vaporized in a pulsating manner to generate thrust. A vaporization chamber with a labyrinth-shaped flow path catches droplets using their surface tension to separate the liquid and vapor, and the droplets vaporize under normal temperature to reduce the input power by reusing the heat from the surrounding components. In this study, we designed and fabricated a flight model of the proposed propulsion system for 6U CubeSat and evaluated the performance of this propulsion system, including the control method. The results confirm the concept of the proposed liquid–vapor separation method and its low power consumption. Moreover, we revealed the relationships between the vaporizing duty cycle, input power, and thrust.
The aerodynamic, structure and propulsion interactions in a six-degree-of-freedom (6-DOF) flexible hypersonic vehicle bring great challenges to vehicle modeling and control design. This paper focus on the modeling of a 6-DOF flexible hypersonic flight vehicle. The flexible dynamic and three-dimensional aerodynamic loads are calculated by using or combining an assumed modes method, variation theory and shock/expansion theory. A control-oriented nonlinear model (COM) is obtained using curve-fitted approximation of the forces and moments on the vehicle. The COM is then linearized about a steady flight condition, and the rigid/flexible dynamics and coupling of the 6-DOF vehicle model with aerodynamic, structure and propulsion interactions are analyzed comprehensively.
The combustion characteristics of the cavity flameholder with a burned-gas injector at the cavity bottom wall in the scramjet model combustor was investigated experimentally. The flame structure in the cavity was investigated by direct imaging and OH-PLIF measurement. As the result, four combustion modes were identified: jet-plume mode, jet-wake mode, one-sided cavity mode, and two-sided cavity mode. In response to the experimental results, the effects of the airstream boundary layer thickness were additionally investigated numerically. Numerical results showed that an increase in the airstream boundary layer thickness adversely affected supported flameholding under conditions of a supersonic airstream with low total temperature. When the airstream boundary layer becomes thicker, interaction between airstream and burned-gas jet creates a wider boundary layer separation upstream of the cavity leading edge. If the width of the separation region is greater than the width of the jet-plume, additional air entrainment path from separation region decreases static temperature inside the cavity flameholder, which makes supported flameholding difficult. It was concluded that the scramjet combustor needs suppression of the interaction between the supersonic airstream and burned-gas jet or avoidance of the boundary layer separation to avoid supported flameholding failure in a supersonic airstream with a thick boundary layer.
The effects of uncertainty in flow conditions, namely angle of attack, Reynolds number, and freestream Mach number, on airfoil characteristics in the low-Reynolds-number regime are evaluated. The Ishii airfoil, a thin–cambered airfoil known to have high aerodynamic performance in this regime, is analyzed. The NACA0012 airfoil is also analyzed as a comparative study. The results for two sets of nominal flow conditions are compared to comprehensively characterize performance in the low-Reynolds-number regime. Statistical quantities of aerodynamic coefficients are computed by coupling the stochastic spectral projection method based on polynomial chaos expansion with two-dimensional flow simulations. The relative contribution of the uncertainty of each flow parameter to the variance of outputs is computed using Sobol's global sensitivity analysis. It is shown that, for the Ishii airfoil, the lift coefficient is highly sensitive to the uncertainty, while the lift-to-drag ratio has a high statistical mean and the pitching moment coefficient has low sensitivity. This indicates that, for thin–cambered airfoils, attention should be given to rapid degradation due to unexpected variation in the angle of attack.