A sensitivity analysis of a fully kinetic particle code was conducted to investigate the importance of uncertainties associated to physical parameters. A 500 W-class laboratory model magnetic-layer Hall thruster was used as the testbed. The sensitivities of the physical parameters, including thermal accommodation coefficient, anode/wall temperature, Bohm diffusion coefficient, electron injection current, cathode coupling voltage, and background pressure, were quantified one-by-one on a conservative possible range. The results suggest the wall erosion prediction is more sensitive to the physical parameters than the thrust or the discharge current. Among the physical parameters, sensitivity to the Bohm diffusion coefficient and parameters related to the neutral flow (i.e., thermal accommodation coefficient and anode/wall temperatures) were dominant. It was hence found that uncertainties in the physical parameters related to the neutral flow had comparable influence on the Bohm diffusion coefficient despite the low attention they attracted.
Considering the coupling interaction between oxidant system and fuel system in liquid-propellant rockets, the modeling and stability problem of Pogo vibration in liquid-propellant rockets with a two-propellant system are investigated. Firstly, the differential equations of the two-propellant equivalent system are derived through the physical characteristics and coupling mechanism of the basic elements in a propulsion system based on reasonable assumptions and simplification. Next, the dynamic equations of the Pogo analysis model are established by coupling a longitudinal structure mode and the two-propellant equivalent equations using the dimensionless methods. It is indicated that the simplified system contains 10 dimensionless parameters expressed by the combination of all physical factors. Furthermore, the critical parameter equation of this system is then explicitly formulated using Hurwitz criterion based on the characteristic equation of the Pogo analysis model. Moreover, the numerical solutions of the boundary surface of this system stability are extensively studied. The results show the effects of dimensionless and original parameters on Pogo stability. Finally, the correctness of the simplified Pogo stability analysis is confirmed using the simulation results of a certain type of Long March rocket.
A novel approach which leverages the model surface itself as the background pattern in the background-oriented Schlieren technique is applied to a hypersonic wind tunnel test. The ordinary background-oriented Schlieren approach assumes that the distance between the background pattern and the area of interest is constant. However, this is not generally true when the model surface is used as the background pattern. A new mathematical formulation of an algebraic reconstruction technique used in computerized tomography (CT) is therefore used to account for the varying distance. Another practical problem is the fact that only a very small displacement due to refraction can be expected due to the short distance between the point of refraction and the background plane. To solve this problem practically and to confirm the overall validity of the mathematical formulation, the approach has been applied to the visualization of the shock structures formed around a cylinder perpendicular to a sharp-nosed flat plate in a hypersonic flow of Mach 7. The CT reconstruction using the sand-blasted model surface as the background pattern and accounting for the varying distance shows clear shock structures even in regions close to the flat plate surface, demonstrating the effectiveness of the new concept.
A practical modeling method for conducting thermal analysis of a microsatellite with multilayer insulation (MLI) was proposed by introducing a new concept for the fitted conductivities of MLI blankets. A finite element (FE) model was first established, in which the MLI blankets are simplified as homogeneous materials with the thermal conductivities initially given. Subsequently, the conductivities were adjusted and determined using an optimization problem that minimized the root mean square (RMS) of temperature residuals between the test data and analysis results based on the current FE model. The FE model with the determined conductivities (i.e., fitted conductivities) was used for thermal analysis. By making comparisons between thermal balance test data and steady-state analysis results of a microsatellite, the rationality and validity of the proposed modeling method were evaluated. Based on the proposed method, the analysis model was further utilized for microsatellite on-orbit temperature prediction. The results revealed that the thermal control scheme with MLI meets the mission requirements.
A green monopropellant reaction system that substitutes discharge plasma for conventional catalysts has been studied. This reaction system was developed for use in the 1 N thrusters of spacecraft reaction control systems. However, the thrust generated by this system was found to be only tens of millinewtons, with a thrust-to-power ratio of about 0.2 mN/W. To improve the thrust and thrust-to-power ratio of the reaction system, we studied and evaluated the effects of electrode axial gap, discharge chamber diameter, discharge type, electrode diameter, and target combustion chamber pressure on thrust and thrust-to-power ratio. All five factors except the electrode diameter had a main effect and four types of interaction on the thrust were confirmed. However, only the target combustion chamber pressure had a main effect on the thrust-to-power ratio and two types of interactions were confirmed. A thrust of 322 mN with a thrust-to-power ratio of 0.95 mN/W was achieved through system optimization.
This paper describes the development and in-orbit demonstration of an electrical power system (EPS) for the 100 W/50 kg-class micro-satellite TSUBAME. Due to the high power consumption per surface area of the TSUBAME and the high power-to-mass ratio, we encountered several technical issues to be solved before the launch. TSUBAME's power demand varies drastically depending on the operation mode, from 2 W to 90 W. Therefore, we could not use the shunt regulator widely used for satellites. To suppress the exhaust heat from the dump power during low-power mode, we developed a series linear regulator that supplies less than 31 V to the main bus. Using this method, the operation point of the current-voltage curve for solar arrays was optimized automatically. However, this control law possesses a hysteresis which invokes a deadlock mode called battery lockup. In addition, the PWM current controller in the battery charger had 35 VPP spike noises, which resulted in system-wide failure. To resolve the fatal problem, we redeveloped the EPS, specifically the noise charger, taking into account physical mechanisms that emitted noise. After environmental and integration tests, the EPS placed on the satellite, which was launched and worked correctly on orbit.
The purpose of this study was to develop an advanced method to measure the properties of rocket combustion using the OH(2,0) band-excited Planer Laser-induced Fluorescence (OH-PLIF) method. This diagnostic method was applied to capture images of the high-pressure H2/O2 jet diffusion flames found in typical liquefied bi-propellant rocket combustion. In addition, axisymmetric numerical simulations of H2/O2 jet flames modeling the experimental conditions were conducted to evaluate the consistency of the OH-PLIF imaging results and to predict the OH chemiluminescence intensity and flame temperature at high pressure. Experimental results show that it is possible to detect the OH(2,1) band fluorescence effectively by eliminating the interference of OH(0,0)-band chemiluminescence under high-pressure conditions of up to 2.0 MPa. The OH fluorescence signal distributed near the injector face almost corresponded to the OH molar concentration distributions simulated by numerical simulations. Moreover, the simulated pressure dependence of the local OH* peak mole concentration reasonably corresponded to that of the local peak chemiluminescence intensity of the experimental chemiluminescence images.
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