氧化亚氮推进技术研究进展

随着环境保护的加强,人们越来越希望找到一种绿色推进剂来代替现有的肼类有毒推进剂.氧化亚氮作为一种绿色推进剂,无毒性,地面实验操作处理方便,不需要繁琐昂贵的防护;常温贮存性,贮箱几乎不需要主动热控制;饱和压力高,可采用自增压方式供应推进剂;绝热分解温度较高,可作为单组元和双组元发动机的推进剂.分析了氧化亚氮作为推进剂的性能及其主要应用领域,着重研究其在液体火箭发动机的应用.通过对氧化亚氮自增压供应系统,单组元推进的催化分解系统,克服催化床限制的氧化亚氮与燃料混合的NOFBXTM技术,以及氧化亚氮作为氧化剂的双组元推进系统的国内外研究进展进行综述,指出当前研究工作中存在的问题,以期为该方面的进一步研究提供一定的参考.     

Control aspects of efficient rocket propulsion systems

The means for enhancing the efficiency of rocket propulsion of spacecraft is considered in terms of the control of propellant consumption, trajectory planning, and in-flight operations management. Particular attention is given to techniques by which all the fuel and oxidizer are consumed completely by mixture adjustment as an example of a terminal control system providing significant propulsion efficiency.     

低功率水电弧加热发动机的初步研究

本文研究了以水为推进剂的电弧加热发动机。其中电弧加热发动机的流量为15mg/s、电流为8A,此时平均电压为80.5V、功率约为640W、平均推力为0.0975N,以此计算得到比冲为650s左右、效率为55%～60%。成功地验证了水做为电弧加热发动机的推进剂、产生推力的能力。     

A generalization of the rocket formula and its application to advanced space propulsion systems

A generalized rocket formula is derived from a first principles approach. The resulting expression of the thrust is applied to advanced space propulsion systems and a possible link between the asymptotic propellant velocity and the velocity at thruster exit is given. An estimation of the thrust modification due to spacecraft–plume interactions is also considered.     

推进系统并联贮箱均衡排放性能及其控制措施

推进系统在工作过程中,贮箱排放的不均衡性将引起飞行器质心偏移,产生干扰力矩。对空间推进系统并联工作的金属膜片贮箱均衡排放的影响因素进行了分析,结合均衡排放的控制措施对某型号推进系统工作过程中的排放不均衡量进行了预计,并与飞行试验结果进行了比对。飞行试验结果表明控制措施可行、有效。     

Mission analysis and systems design of a near-term and far-term pole-sitter mission

This paper provides a detailed mission analysis and systems design of a near-term and far-term pole-sitter mission. The pole-sitter concept was previously introduced as a solution to the poor temporal resolution of polar observations from highly inclined, low Earth orbits and the poor high-latitude coverage from geostationary orbit. It considers a spacecraft that is continuously above either the north or south pole and, as such, can provide real-time, continuous and hemispherical coverage of the polar regions. Being on a non-Keplerian orbit, a continuous thrust is required to maintain the pole-sitter position. For this, two different propulsion strategies are proposed, which result in a near-term pole-sitter mission using solar electric propulsion (SEP) and a far-term pole-sitter mission where the SEP thruster is hybridized with a solar sail. For both propulsion strategies, minimum propellant pole-sitter orbits are designed. In order to maximize the spacecraft mass at the start of the operations phase of the mission, the transfer from Earth to the pole-sitter orbit is designed and optimized assuming either a Soyuz or an Ariane 5 launch. The maximized mass upon injection into the pole-sitter orbit is subsequently used in a detailed mass budget analysis that will allow for a trade-off between mission lifetime and payload mass capacity. Also, candidate payloads for a range of applications are investigated. Finally, transfers between north and south pole-sitter orbits are considered to overcome the limitations in observations due to the tilt of the Earth's rotational axis that causes the poles to be alternately situated in darkness. It will be shown that in some cases these transfers allow for propellant savings, enabling a further extension of the pole-sitter mission.     

Development of hall thruster propulsion system for STSAT-3 application

Science and technology satellite-3 (STSAT-3) is being developed and is scheduled for launch in 2011. One of the primary objectives of its mission is to verify the performance of a hall thruster propulsion system (HPS) that uses xenon gas. According to its major functions, the HPS can be divided into several sub-modules. This paper presents the development and qualification of the hall effect thruster propulsion subsystem that includes a xenon feed system (XFS). The xenon feed system regulates the pressure down from the xenon propellant tank and supplies the xenon flows to the anode and cathode. The technology and xenon feed system developed for the STSAT-3 spacecraft will also be applicable to a variety of future electronic propulsion systems and micro-satellites. Details related to the overall development and performance results of the HPS are presented in this paper.     

太阳帆推进

太阳帆依靠反射太阳光光子产生推力,在飞行任务的整个过程推力连续作用于飞行器上而不需要任何推进剂。太阳帆可广泛应用于低成本大速度增量的太阳系飞行任务,具有其它推进系统无法替代的优点。简要介绍了太阳帆推进的机理及研究现状,给出了太阳帆飞行器的主要性能参数及需要解决的主要技术关键。     

太阳帆航天器姿态控制技术综述

太阳帆作为一种无工质消耗的新型航天器推进技术,逐渐受到国内外航天界的广泛关注。文章首先综述了太阳帆航天器控制技术的发展现状,重点分析了太阳帆航天器姿态控制技术的特点和难点;其次,按照执行机构的不同,分别介绍了近年来国内外航天领域提出的多种太阳帆航天器姿态控制方案,针对各个方案的优缺点加以分析;最后,对该领域未来的发展前景进行了展望。     

低轨航天器磁推进方法

基于地球空间磁场的磁作用效应,提出了一类低轨航天器无工质消耗的磁推进方法,并对航天器的高度保持进行了研究。该方法具有作用机理明晰、物理结构简单、控制策略灵活等特征。以磁场对磁体作用的磁矩理论为基础,建立了带磁航天器飞行的磁推力模型,提出了基于磁力线追踪策略的轨道高度保持方法。利用熟知的IGRF11地磁模型,通过数值仿真计算,验证了磁推进方法对于600～1000km圆轨道航天器进行高度保持的有效性。     

On the problem of optimal control of the thrust value of the electric propulsion rocket with solar energy source

Under consideration is the optimal control problem on a spacecraft motion in Newtonian central gravity field. With the use of the mathematical model of electrojet propulsion device (EPD) with solar energy source, proposed earlier in paper [1], the dependence of the EPD working substance choice on both the duration of the given dynamic maneuver and the propellant expenditures for its fulfillment is investigated. The efficiency evaluation is carrying out of optimal control of variable valued thrust as well as that for relay mode thrust and relay mode thrust with optimal fixed thrust value.     

Flight demonstration of new thruster and green propellant technology on the PRISMA satellite

The concept of a storable liquid monopropellant blend for space applications based on ammonium dinitramide (ADN) was invented in 1997, within a co-operation between the Swedish Space Corporation (SSC) and the Swedish Defense Research Agency (FOI). The objective was to develop a propellant which has higher performance and is safer than hydrazine. The work has been performed under contract from the Swedish National Space Board and ESA. The progress of the development has been presented in several papers since 2000.ECAPS, a subsidiary of the Swedish Space Corporation was established in 2000 with the aim to develop and market the novel “high performance green propellant” (HPGP) technology for space applications. The new technology is based on several innovations and patents w.r.t. propellant formulation and thruster design, including a high temperature resistant catalyst and thrust chamber.The first flight demonstration of the HPGP propulsion system will be performed on PRISMA. PRISMA is an international technology demonstration program with Swedish Space Corporation as the Prime Contractor.This paper describes the performance, characteristics, design and verification of the HPGP propulsion system for PRISMA. Compatibility issues related to using a new propellant with COTS components is also discussed. The PRISMA mission includes two satellites in LEO orbit were the focus is on rendezvous and formation flying. One of the satellites will act as a “target” and the main spacecraft performs rendezvous and formation flying maneuvers, where the ECAPS HPGP propulsion system will provide delta-V capability.The PRISMA CDR was held in January 2007. Integration of the flight propulsion system is about to be finalized.The flight opportunity on PRISMA represents a unique opportunity to demonstrate the HPGP propulsion system in space, and thus take a significant step towards its use in future space applications. The launch of PRISMA scheduled to 2009.     

N2O微推力器性能及喷管热结构分析

针对采用氧化亚氮推进剂的单组元微推力器开展了比冲性能影响因素的分析,分析结果显示微推力器比冲与氧化亚氮分解效率及喷管扩张比有着密切关系。利用有限元分析法对高空及地面试验两种工况下氧化亚氮单组元微推力器喷管的结构温度场开展了数值仿真计算,并在结构温度场仿真计算的基础上进一步对地面试验用喷管的结构应力场进行了分析。初步试验表明,所设计的微喷管在地面工况下工作良好。     

Single crystal mo solar thermal thruster for microsatellites

One potentially attractive propulsion concept offering significant payload gains for orbit transfer from LEO to higher orbits, station keeping and attitude control of spacecraft is thermal propulsion using light gas (typically hydrogen) as propellant and various kinds of heat energy. Solar Thermal Propulsion (STP) is a typical thermal propulsion with high Isp (500 – 1,000 s) in an appropriate thrust magnitude range and provides possibly much less space pollution than conventional chemical propulsion.

This paper presents the test results of a 30 mm dia. (medium-sized) windowless type of single crystal Mo thruster for orbit transfer of 50 kg class microsatellites. The cavity dia. is 20 mm, double the size of the previous model, and can apply to a primary solar reflector of up to 3.5 m dia., which is the maximum size containable in the H-II rocket fairing without segmentation. The performed mission analyses indicate that this size of STP is suitable to orbit transfer of 50 kg class microsatellites, such as LEO to GEO, or only multiple apogee kicks from GTO to GEO or deep space missions.     

Optimal Control of a Programmed Turn of a Spacecraft

A problem of optimal turn of a spacecraft is considered. The time of turn is minimized, as well as the functional having a meaning of the propellant consumption. An analytical solution to the problem stated is derived. It is demonstrated that the solution optimal in this sense belongs to a class of two-impulse controls, under which a spacecraft executes the turn along the trajectory of its free motion. The solution obtained in this paper differs from earlier available solutions considerably. The estimations of the propellant consumption for a realization of the programmed turn are made.     

叠氮复合固体推进剂技术研究

分析了航天器和导弹武器系统对固体推进剂提出的新的要求和实现这些要求的技术途径。介绍了国外研制低特征信号叠氮固体推进剂的主要原材料叠氮粘合剂、含能增塑剂、高能高密度氧化剂的发展概况和关键技术。分析认为，由以上材料组成的叠氮复合固体推进剂具有含能量高、密度大、发动机排气羽烟对微波、激光和可见光的透过率高等特征，因此这是一种很有前途的新型推进剂。     

NanoSail-D: A solar sail demonstration mission

In the early to mid-2000s, NASA made substantial progress in the development of solar sail propulsion systems. Solar sail propulsion uses the solar radiation pressure exerted by the momentum transfer of reflected photons to generate a net force on a spacecraft. To date, solar sail propulsion systems were designed for large robotic spacecraft. Recently, however, NASA has been investigating the application of solar sails for small satellite propulsion. The NanoSail-D is a subscale solar sail system designed for possible small spacecraft applications. The NanoSail-D mission flew on board the ill-fated Falcon Rocket launched August 2, 2008, and due to the failure of that rocket, never achieved orbit. The NanoSail-D flight spare is ready for flight and a suitable launch arrangement is being actively pursued. This paper will present an introduction solar sail propulsion systems and an overview of the NanoSail-D spacecraft.     

国外深空探测推进技术发展及启示

推进系统为探测器的飞行提供所需的控制力和控制力矩,一定程度上决定了探测器的规模、最远飞行距离,乃至任务的成败,是探测器的重要分系统之一。为满足不同的深空探测任务需求,国外发展了多种形式的推进技术,但总体上仍以化学推进为主,部分采用了电推进系统,并在发展高性能低温推进技术等。对国外典型探测器推进系统进行了叙述,分析了其技术特点和发展趋势,并分别针对无人探测和载人探测应用探讨了对我国开展深空探测推进技术研究的启示。     

肼-70胶体推进剂的特性及制备方法

胶体推进剂是火箭推进和其它燃气发生器的新型待选推进剂。常规液体推进剂在储存过程中泄漏危险性较高,将液体推进剂制成介于固体与液体之间的胶体,可大大增加其粘度与表面张力,降低泄漏危险性。文中对肼-70胶体推进剂的国内外研究现状作了概述,着重对其制备工艺进行了具体介绍,包括胶凝剂的选择、分散工艺等方面,并通过实验对比发现对于肼-70体系而言,添加相同含量的B类胶凝剂成胶效果优于A类胶凝剂。     

Design of optimal earth pole-sitter transfers using low-thrust propulsion

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.