For hybrid rocket engines, there are some unique relationships between the parameters, such as the fuel regression rate, oxidizer mass flow rate, equivalence ratio and so on, which determine the burning properties. These relationships are strongly related to the engine performance. Our final goal of this study is to establish the optimum design method which considers the engine performance of hybrid rocket engines for practical hybrid rockets varying with time. In this paper, we focused on the time variation of ballistic parameters of the swirling-oxidizer-flow-type hybrid rocket engine. We establish the new simulation method, the Two-Step combustion model, in which at the fuel grain head the fuel regression rate is independent of the oxygen mass flux and in the rear region depends on the flux according to the ordinary hybrid combustion formula. The model succeeded to predict the time variation of the burning properties and ballistics parameters, which coincided with the experimental burning values and time behavior. These results show the possibility that there are two different combustion mechanisms in the burning process of the swirling-oxidizer-flow-type hybrid rocket engine.
This paper presents a new guidance and control system based on an adaptive backstepping method for a space transportation system. In recent years, many studies of flight control systems using feedback linearization combined with time-scale separation have been carried out. Since this type of control system does not depend on the design points along a predetermined trajectory, the designed system can be applied to an extensive flight region. However, in this method, control performance tends to deteriorate with changes in the control gains and parameters because it is difficult to guarantee the stability of the system. Additionally, since it is not easy to obtain a priori knowledge about disturbances, aerodynamic characteristics, an estimation mechanism must be added to the system. For this problem, we propose an adaptive flight control system by combining with feedback linearization, backstepping method, and disturbance observers. A disturbance observer is effective at estimating the effect of extraneous signals. In the proposed system, by appropriate redesign of the disturbance observer, it becomes possible to guarantee the stability of the entire system including the estimation mechanism. Numerical simulations are performed to verify the effectiveness and robustness of the proposed system when applied to an automatic landing problem.
Use of liquid propellants, such as water or ethanol, as an alternative to the most common propellant, solid Teflon, for a pulsed plasma thruster (PPT) is reported to improve the performance of PPT by excluding the drawbacks of uncontrollable propellant feed rate and timing, while maintaining the advantages of digital control features. In this study, we propose to use dimethyl ether (DME) as for the liquid propellant PPT, because DME has a low freezing point (131K at standard pressure) and an expedient vapor pressure (6 atm at standard temperature), both of which eliminate the need for a heater and pressurant when using a water or ethanol propellant. A series of experiments using a prototype DME-PPT with parallel-plate electrodes demonstrated reliable operation with a tendency for the electromagnetic impulse bit Ibit to increase in proportion to the stored capacitor energy Ec up to Ibit = 53μNs at Ec = 13J.
In this paper, the optimal guidance strategy for simultaneous attack by multiple missiles is proposed; the strategy consists of following steps: (1) designation of optimal final impact time which is shared among all missiles participating in the simultaneous attack; (2) individual guidance of missiles by optimal control input which enable missiles to impact the target at the designated time. The optimization for this strategy is conducted on a non-linear model without linear approximation. Hence, although proposed strategy must be based on numerical optimization, optimal guidance of our strategy can be achieved by finite number of calculation. This is because all cases of optimal trajectories can be consolidated in finite set of normalized parameters by considering the necessary optimality conditions. Proposed strategy is effective for not only the missile guidance but also the cooperative control of multi vehicle systems.