共查询到20条相似文献,搜索用时 218 毫秒
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《Acta Astronautica》2001,48(5-12):651-660
The aim of this paper is to analyse an alternative scenario for Mars Sample Return Orbiter mission, where electric propulsion is used for Earth-Mars and Mars-Earth heliocentric cruises and for Mars orbit insertion / escape transfers, whereas chemical propulsion is used for final Mars rendezvous. The problem consists in minimizing the initial vehicle mass to obtain a specific final dry mass in reasonable time. The planetocentric phases correspond to continuous low-thrust trajectories, spiraling around Mars between a low orbit and the influence sphere altitude. The heliocentric phases consist of a succession of low-thrust and coasting arcs with specific departure and arrival conditions at the Earth. For these two types of transfer, efficient optimal control tools exist based on Pontryagin's maximum principle. Thanks to the coordination between planetocentric and heliocentric phases, the solution obtained with these two separate tools gives a good upper bound of the optimal solution in terms of propellant consumption and duration. This optimization procedure is described and finally applied to the proposed mission. The numerical results are presented and compared with the baseline chemical mission solution. The electric option could allow to decrease the spacecraft departure mass but may lead to rather long mission duration. 相似文献
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Mihai Popescu 《Acta Astronautica》1985,12(4):225-228
The transfer from the equilateral Lagrangian points of the Earth-Moon system is analysed. The final states of the velocity of the space vehicles and of the rotation velocity of the propulsion vector are assumed given. The trajectory which ensures the transfer in optimal time consists of three arcs. On this trajectory the rotation velocity of the direction of the propulsion has the extremal value or corresponds to the Lawden's tangent law. The use of the matching of the arcs together with transversality conditions and final conditions determines the constants of integration and the evolution time. The resulting parametric equations of the optimal trajectory are of integral form. 相似文献
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Low-thrust flights from high-elliptic orbits are of considerable interest, since they allow one to decrease (compared to high-thrust flights) the propulsion consumption and to reduce the flight duration. At the same time, in comparison with the spiral unwinding flights from low near-circular orbits, this scheme minimizes the harmful effect of the radiation belts. Based on the maximum principle, the problem of optimization is reduced to a two-point boundary value problem, which is solved numerically using the modified Newton method. A method is suggested to obtain the initial approximation for solving the boundary value problem. The method takes advantage of the idea of transition from an approximately optimal trajectory to the optimal one. Two problems, which have different low-thrust models, are considered: one with permanently acting low thrust and the other with the possibility of turning it on/off. In both cases no restrictions are imposed on the thrust direction. A comparison of these problems is made. We investigated (i) what gain in the final mass can be attained when passing from the first to the second problem, (ii) at the cost of what loss in flight duration this can be achieved, and (iii) what changes in the optimal program of control must be done in this case. 相似文献
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针对可重复使用运载火箭垂直回收轨迹优化问题,提出了一种带有最优终端时间估计策略的hp伪谱同伦凸优化在线轨迹规划算法。首先,考虑状态约束和过程约束的非凸性,采用无损凸化处理推力幅值约束;然后,结合同伦方法与不动点迭代思想将气动力与非凸质量约束转化为线性时变剖面,完成问题凸化;进一步基于hp flipped Radau伪谱法对问题进行离散化处理,将最优控制问题转化为参数优化问题,进而采用原-对偶内点法求解;最后,为进一步减少燃料消耗,提升经济效益,考虑最优终端时间难以在线确定的问题,结合解析推导与二次插值法,设计了最优终端时间快速估计策略。仿真结果表明所设计的轨迹优化算法最优终端时间估计速度快,收敛性能良好,具有较高的精度和计算效率,具备在线应用的潜力。 相似文献
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Hopes of sending probes to another star other than the Sun are currently limited by the maturity of advanced propulsion technologies. One of the few candidate propulsion systems for providing interstellar flight capabilities is nuclear fusion. In the past many fusion propulsion concepts have been proposed and some of them have even been explored in detail, Project Daedalus for example. However, as scientific progress in this field has advanced, new fusion concepts have emerged that merit evaluation as potential drivers for interstellar missions. Plasma jet driven Magneto-Inertial Fusion (PJMIF) is one of those concepts. PJMIF involves a salvo of converging plasma jets that form a uniform liner, which compresses a magnetized target to fusion conditions. It is an Inertial Confinement Fusion (ICF)–Magnetic Confinement Fusion (MCF) hybrid approach that has the potential for a multitude of benefits over both ICF and MCF, such as lower system mass and significantly lower cost. This paper concentrates on a thermodynamic assessment of basic performance parameters necessary for utilization of PJMIF as a candidate propulsion system for the Project Icarus mission. These parameters include: specific impulse, thrust, exhaust velocity, mass of the engine system, mass of the fuel required etc. This is a submission of the Project Icarus Study Group. 相似文献
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Jonathan F.C. Herman Aline K. Zimmer Johannes P.J. Reijneveld Kathryn L. Dunlop Yu Takahashi Simon Tardivel Daniel J. Scheeres 《Acta Astronautica》2014
This paper presents a mission analysis comparison of human missions to asteroids using two distinct architectures. The objective is to determine if either architecture can reduce launch mass with respect to the other, while not sacrificing other performance metrics such as mission duration. One architecture relies on chemical propulsion, the traditional workhorse of space exploration. The second combines chemical and electric propulsion into a hybrid architecture that attempts to utilize the strengths of each, namely the short flight times of chemical propulsion and the propellant efficiency of electric propulsion. The architectures are thoroughly detailed, and accessibility of the known asteroid population is determined for both. The most accessible asteroids are discussed in detail. Aspects such as mission abort scenarios and vehicle reusability are also discussed. Ultimately, it is determined that launch mass can be greatly reduced with the hybrid architecture, without a notable increase in mission duration. This demonstrates that significant performance improvements can be introduced to the next step of human space exploration with realistic electric propulsion system capabilities. This leads to immediate cost savings for human exploration and simultaneously opens a path of technology development that leads to technologies enabling access to even further destinations in the future. 相似文献
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Sellers Jerry Jon Paul Malcolm Sweeting Martin 《Journal of Reducing Space Mission Cost》1998,1(1):53-72
The paper summarizes research into cost-effective propulsion system options for small satellites. Research into the primary cost drivers for propulsion systems is discussed and a process for resolving them is advanced. From this analysis, a new paradigm for understanding the total cost of propulsion systems is defined that encompasses nine dimensions – mass, volume, time, power, system price, integration, logistics, safety and technical risk. This paradigm is used to characterize all near-term propulsion technology options. From this effort, hybrid rockets emerges as a promising but underdeveloped technology with great potential for cost-effective application. A dedicated research program was completed to characterize this potential. This research demonstrated that hybrid rockets offer a safe, reliable upper stage option that is a versatile, cost-effective alternative to solid rocket motors. Finally, an innovative technique was derived to parametrically combine the diverse cost dimensions into a useful, quantifiable figure of merit for mission and research planning. Overall, it is shown that the most cost-effective solution is found by weighing all options along the nine dimensions of the cost paradigm within the context of a specific mission. 相似文献
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The problem of optimal control over many-revolution spacecraft orbit transfers between circular coplanar orbits of satellites is considered. The spacecraft flight is controlled by a thrust vector of a jet engine with restricted thrust (JERT). The mass expenditure is minimized at a limited time of flight. The optimal control problem is solved based on the maximum principle. The boundary value problem of the maximum principle is solved numerically using the shooting method. A modified computation scheme of the shooting method is suggested (multi-point shooting), as well as a method (correlated with the scheme) of choosing the initial approximation with the use of a solution to the optimization problem in the impulse formulation. The scheme and method allow one to construct many-revolution spacecraft orbit transfers. 相似文献
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This paper deals with the design, development and experimentation of a new test stand for the accurate and precise characterization of small cold-gas nozzles having thrust of the order of 0.1 N and specific impulse of the order of 10 s. As part of the presented research, a new cold-gas supersonic nozzle was designed and developed based on the quasi one-dimensional theory. The test stand is based on the ballistic-pendulum principle: in particular, it consists of a suspended gondola hosting the propulsion system and the sample nozzle. The propulsion system consists of an air tank, pressure regulator, solenoid valve, battery and digital timer to command the valve. The gondola is equipped with a fin, immersed in water, to provide torsional and lateral oscillation damping. A laser sensor measures the displacement of the gondola. The developed test stand was calibrated by using a mathematical model based on the inelastic collision theory. The obtained accuracy was of ~1%. Sample experimental results are reported regarding the comparison of the new supersonic nozzle with a commercially available subsonic nozzle. The obtained measurements of thrust, mass flow rate and specific impulse are precise to a level of ~3%. The broad goal of the presented research was to contribute to an upgraded design of a spacecraft simulator used for laboratory validation of guidance, navigation and control algorithms for autonomous docking manoeuvres. 相似文献
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Recent studies have shown the feasibility of an Earth pole-sitter mission using low-thrust propulsion. This mission concept involves a spacecraft following the Earth's polar axis to have a continuous, hemispherical view of one of the Earth's poles. Such a view will enhance future Earth observation and telecommunications for high latitude and polar regions. To assess the accessibility of the pole-sitter orbit, this paper investigates optimum Earth pole-sitter transfers employing low-thrust propulsion. A launch from low Earth orbit (LEO) by a Soyuz Fregat upper stage is assumed after which solar electric propulsion is used to transfer the spacecraft to the pole-sitter orbit. The objective is to minimize the mass in LEO for a given spacecraft mass to be inserted into the pole-sitter orbit. The results are compared with a ballistic transfer that exploits manifold-like trajectories that wind onto the pole-sitter orbit. It is shown that, with respect to the ballistic case, low-thrust propulsion can achieve significant mass savings in excess of 200 kg for a pole-sitter spacecraft of 1000 kg upon insertion. To finally obtain a full low-thrust transfer from LEO up to the pole-sitter orbit, the Fregat launch is replaced by a low-thrust, minimum time spiral, which provides further mass savings, but at the cost of an increased time of flight. 相似文献
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The possibility of using the statistics of recurrence time for extreme events is studied in this paper having in mind the
problems of control and prediction of failures in spacecraft operation. The information about failures onboard satellites
of various types presented by the US National Geophysical Data Center was analyzed. It was found that the probability density
of recurrence intervals followed a power law of the Pareto type with an index equal to 2.3. The obtained result is consistent
both with the theory of normal catastrophes and with the principle of self-organization of criticality for metastable active
heterogeneous environment. A practical consequence of the obtained result consists in the fact that predictions of these extreme
events should not rely on traditional models with the second-order Pearson statistics. To make predictions, the models are
necessary that take into account the power law distribution of recurrence intervals for failures on satellites. The failures
should be considered in these models as extreme events connected with manifestation of the space environment factors. 相似文献
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Power-limited systems with variable Isp, which have been studied theoretically since the beginning of astronautics, are getting closer to practical applications thanks to recent technological advances in the field of magnetosplasma rockets, such as Ad-Astra’s VASIMR concept. This type of propulsion system is considered for high-speed interplanetary transfers, such as Mars missions, with demanding payload fractions that would be compatible with manned missions. This paper explores the problem of the optimization of a power-limited propulsion system through simple performance models, and investigates the trade-off between the technological requirements, the transfer time and the payload fraction1. Following previous works existing in literature, we model the technological characteristics of the vehicle through a small number of parameters, the most important of which being the specific weight (or mass-to-power ratio) of the power generation system. Also, we use in our models the classical “trajectory characteristic” parameter (defined as the integral over time of the squared thrust acceleration) which represents – under certain hypotheses – the propulsion requirements for an orbital or interplanetary transfer with a given time and a given thrust strategy. In this paper, we first give a review of existing methods in literature, then we present the equations of a new class of optimal design which maximizes the payload fraction, for a given transfer time and given technological characteristics. This class of optimal design is described through very simple equations that make possible to study more straightforwardly than existing calculations the links between the main mission requirements (transfer time and payload fraction) and the main technological requirements (specific weight of the power generation and structure mass ratio of the whole vehicle, excluding the power generation system). One important result obtained from these equations is a simple expression which estimates the theoretical upper limit of the power source’s specific weight as a function of transfer time and the payload mass ratio. In the last part of this paper, we apply this simple performance model to discuss the feasibility of a fast Earth-to-Mars transfer using a power-limited system. 相似文献
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以过驱动航天器的推力器控制分配误差最小、推力器负载均衡等为设计目标,构建航天器推力分配混合优化模型,并将其转化为线性规划模型进行求解,提出了一种考虑负载均衡的航天器推力器动态分配算法。该算法在确保分配误差最小前提下,能够降低各推力器的最大分配推力之差,有效均衡各推力器总工作时长和开关次数,进而延长推进系统的整体工作寿命。进一步定义了表征负载均衡性能的推力平衡度和干扰敏感度性能指标,并在此基础上给出了一种分配算法负载均能能力的定量化评价方法。在仿真验证中,采用平衡度和敏感度对算法性能进行定量评估,结果表明该方法在保证控制性能和控制分配误差的前提下,能够有效均衡各推力器最大推力,提高了系统的平衡度和对扰动力矩的鲁棒性。 相似文献