The impact of launch vehicle type and size on development cost

The technical development trend of future launch vehicle systems is towards fully reusable systems, in order to reduce space transportation cost. However, different types of launch vehicles are feasible, as there are

• —winged two-stage systems (WTS)

• —ballistic single-stage vehicles (BSS)

• —ballistic two-stage vehicles (BTS)

The performance of those systems is compared according to the present state of the art as well as the development cost, based on the “TRANSCOST-Model”. The development costs are shown versus launch mass (GLOW) and pay-load for the three types of reusable systems mentioned above.It is shown that performance optimization and cost minimization lead to different results. It is more economic to increase the vehicle size for achieving higher performance, instead of increasing technical complexity.Finally it is described that due to the essentially lower launch cost of reusable vehicles it will be feasible to recover the development cost by an amortization charge on the launch cost. This possibility, however, would allow commercial funding of future launch vehicle developments.     

1700℃高温、有氧及时变环境下隔热性能试验研究

针对高超声速飞行器面临极端高温热环境、飞行器外壳单侧面受热以及温度历程非线性时变的特点,自行设计并建立辐射式极端高温氧化环境下的单侧面试验加热装置,实现了1700℃高温有氧环境下对高超声速飞行器热防护材料的隔热性能试验测试。同时,对轻质陶瓷材料试验件和新型陶瓷、纳米材料复合结构在高达1700℃的高温氧化环境下的隔热性能进行试验测试,并对不同材料及其组合模式进行对比分析,优选高效能的隔热方案,发现陶瓷、纳米材料复合结构试验件比单层轻质陶瓷材料试验件的隔热效果提高了约50%。另外,生成了极端高温非线性时变热环境,并进行相应的隔热性能试验。通过建立极端高温、有氧、单侧面加热、非线性时变热环境试验系统及其实际应用研究,为高超声速飞行器的热防护设计提供重要的试验手段。     

Aeroshell design techniques for aerocapture entry vehicles

A major goal of NASA's In-Space Propulsion Program is to shorten trip times for scientific planetary missions. To meet this challenge arrival speeds will increase, requiring significant braking for orbit insertion, and thus increased deceleration propellant mass that may exceed launch lift capabilities. A technology called aerocapture has been developed to expand the mission potential of exploratory probes destined for planets with suitable atmospheres. Aerocapture inserts a probe into planetary orbit via a single pass through the atmosphere using the probe's aeroshell drag to reduce velocity. The benefit of an aerocapture maneuver is a large reduction in propellant mass that may result in smaller, less costly missions and reduced mission cruise times. The methodology used to design rigid aerocapture aeroshells will be presented with an emphasis on a new systems tool under development. Current methods for fast, efficient evaluations of structural systems for exploratory vehicles to planets and moons within our solar system have been under development within NASA having limited success. Many systems tools that have been attempted applied structural mass estimation techniques based on historical data and curve fitting techniques that are difficult and cumbersome to apply to new vehicle concepts and missions. The resulting vehicle aeroshell mass may be incorrectly estimated or have high margins included to account for uncertainty. This new tool will reduce the guesswork previously found in conceptual aeroshell mass estimations.     

Aeroshell design techniques for aerocapture entry vehicles

A major goal of NASA's In-Space Propulsion Program is to shorten trip times for scientific planetary missions. To meet this challenge arrival speeds will increase, requiring significant braking for orbit insertion, and thus increased deceleration propellant mass that may exceed launch lift capabilities. A technology called aerocapture has been developed to expand the mission potential of exploratory probes destined for planets with suitable atmospheres. Aerocapture inserts a probe into planetary orbit via a single pass through the atmosphere using the probe's aeroshell drag to reduce velocity. The benefit of an aerocapture maneuver is a large reduction in propellant mass that may result in smaller, less costly missions and reduced mission cruise times. The methodology used to design rigid aerocapture aeroshells will be presented with an emphasis on a new systems tool under development. Current methods for fast, efficient evaluations of structural systems for exploratory vehicles to planets and moons within our solar system have been under development within NASA having limited success. Many systems tools that have been attempted applied structural mass estimation techniques based on historical data and curve fitting techniques that are difficult and cumbersome to apply to new vehicle concepts and missions. The resulting vehicle aeroshell mass may be incorrectly estimated or have high margins included to account for uncertainty. This new tool will reduce the guesswork previously found in conceptual aeroshell mass estimations.     

质量矩拦截弹的模糊滑模姿态控制系统设计

质量矩控制是一种全新的飞行控制方法，与广泛应用的空气动力控制相比，可以避免飞行器在超高马赫数飞行时舵面的气动加热问题，而且可以提高飞行器的机动性和敏捷性。以所建立的质量矩拦截弹数学模型为基础，通过对模型合理的简化，得到一个耦合的非线性动力学系统，考虑到系统的鲁棒性要求和三个滑块的协调控制问题，利用滑模控制设计了拦截弹大机动飞行时的姿态控制系统，同时采用模糊逻辑算法抑制系统的抖振现象。仿真结果表明系统性能指标满足设计要求，只需微调滑块位置，即可以实现拦截弹的姿态控制。     

Roadmap to a human Mars mission

We propose a new roadmap for the preparation of the first human mission to Mars. This proposal is based on the work of ISECG and several recent recommendations on human Mars mission architectures. A table is proposed to compare the possible benefits of different preparatory missions. Particular attention is paid to the possibility of qualifying important systems thanks to a heavy Mars sample return mission. It is shown that this mission is mandatory for the qualification of Mars aerocapture at scale-1, EDL systems at scale 1 and Mars ascent. Moreover, it is a good opportunity to test many other systems, such as the heavy launcher and the transportation systems for the trips beyond LEO. These tests were not mentioned in the last ISECG report. This strategy is facilitated in the case of the simplified Mars mission scenarios that have recently been presented because it is suggested that relatively small vehicles with small crew sizes are used in order to optimize the payload mass fraction of the landing vehicles and to avoid the LEO assembly. An important finding of the study is that a human mission to the surface of the Moon is not required for the qualification of the systems of a human mission to Mars. Since affordability is a key criterion, two important missions are proposed in the roadmap. The first is a heavy Mars sample return mission and the second is a manned mission to a high Earth orbit or eventually to the vicinity of the Moon. It is shown that both missions are complementary and sufficient to qualify all the critical systems of the Mars mission.     

基于模糊神经网络的质量矩拦截弹动态逆控制

质量矩控制是一种全新的飞行控制方法,与广泛应用的空气动力控制相比,可以避免飞行器在超高马赫数飞行时舵面的气动加热问题,而且,可以提高飞行器的机动性和敏捷性。以所建立的质量矩拦截弹数学模型为基础,通过对模型合理的简化,得到一个耦合的非线性动力学系统。考虑到系统的鲁棒性要求和三个滑块的协调控制问题,应用双时标分解的方法,提出将拦截弹动力学分离为快变状态动力学和慢变状态动力学。然后,设计了拦截弹模糊动态逆飞行控制系统,分别针对快变和慢变子系统,提出了一种基于模糊神经网络(FNN)的动态逆误差补偿方法,可以有效地抑制气动参数摄动而引起的控制系统性能的下降,提高了控制系统的鲁棒性。仿真结果表明系统性能指标满足设计要求,只需微调滑块位置,即可以实现拦截弹的飞行控制。     

无损检测技术在姿控发动机总装中的应用探析

阐述了目视检测、射线照相检测、泄漏检测等三种无损检测技术在姿控发动机上的应用情况。重点介绍了泄漏检测技术,针对发动机的结构特点、不同系统的性能要求及总装接头连接处各种密封形式在装配测试中分别采用了浸泡法、皂泡法及氦质谱简易包封积累法等不同的泄漏检测技术。经过对几种泄漏检测方法的对比分析,得出了针对不同的系统性能要求。需要采用不同的泄漏检测方法来满足设计要求的结论。     

宽频段一体化硬件处理平台设计

为解决传统通讯卫星面临的问题,针对微纳卫星发展趋势和方向,将具备核心功能的载荷微小型化,提出了一种宽频段多功能一体化硬件平台设计方案,搭建的硬件平台具有体积小、重量轻、功耗低、集成度高、可重构的特点。为解决传统硬件平台功能固化,不能满足功能多、快速响应的问题,提出了一种基于软件无线电设计思想,构建宽频带可重构通用信号处理平台,最大限度使功能软件化,通过加载存储芯片中不同程序实现不同的算法功能,通过一体化设计有效实现高功能密度,大幅度地降低成本,缩短研制周期,通过部组件产品模块化、接口标准化,制定即插即用标准规范,从而构建面向不同应用场合和需求的硬件平台,为传统硬件平台提供了一种新思路,是未来硬件平台的发展趋势。     

一种可兼容多型号运载火箭的可重构式活动发射平台（英文）

A launch pad is an important part of the launch vehicle system. It bears the weight of launch vehicles and payloads, and is burnt due to the high-temperature and high-pressure gas flow during launch. This study analyzes the development of launch pads in the world, reviews a reconfigurable mobile launch pad for multiple launchers, then proposes common points and development trends of future launch pads in the world.     

几何结构可调的亚燃冲压发动机性能研究

液体亚燃冲压发动机结构简单、推重比高,是高动态临近空间飞行器的最佳动力装置,临近空间飞行器的飞行速度范围宽、距离远,亟需采用几何结构可调技术来提高冲压发动机的性能。本研究对具有固定、连续可调进气道和尾喷管的冲压发动机性能进行了计算和比较。结果表明,采用连续可调喷管的冲压发动机的性能大大优于固定几何结构的冲压发动机,进气道可调带来的冲压发动机性能增加远小于喷管连续可调带来的发动机性能增加。     

电磁脉冲弹对雷达“前门”损伤研究

分析了电磁脉冲弹从"前门"损伤雷达的机理,给出了雷达"前门"电磁损伤判断公式,并提出了一些雷达系统"前门"电磁加固措施.可作为电磁脉冲弹威胁下雷达系统生存性研究和战场抢修研究的理论参考.     

低轨卫星可重构通信系统设计

提出一种适合于在轨重构的低轨卫星通信系统软、硬件架构方案,从工程实施角 度探讨包括天线、RF前端和基带处理单元在内的可重构硬件平台、可重构策略与软件控制流 的可行性,分析了可重构卫星通信系统实现的关键技术。利用软件无线电技术,分析了基带 和射频前端分离,可裁减、复用式平台实现的基本路径,以及基带、射频天线单元重构的实 现方法。通过各常规通信模式射频前端和数字部分的灵活组合,同时支持多频点、多模式的 通信制式,达到节省硬件成本和提高系统灵活性的目的。最后研制了一个可重构卫星通信地 面测试原理样机,模拟了卫星有效载荷和地面通信站在IS-95体制下两种CDMA加密码字间的 动态重构,有效验证了本文提出的可重构策略的正确性与合理性。
     

基于亲和传播算法的车辆自组织网络分簇组网方法

随着5G通信和自动驾驶汽车的发展，车辆自组织网络VANET（Vehicular Ad Hoc Network）作为一种新型的移动自组织网络，因其在改善道路安全、为驾驶员和乘客提供便利方面的潜力，而引起学术界和工业界的广泛关注。它可以通过提供交通流量、事故通知、危险警告、可能存在的定位偏差、天气等信息来增强道路安全，从而提高交通效率。由于车辆的高速移动会经常引起网络拓扑中断，因此一个设计良好的路由协议至关重要。提出了一种适用于车辆自组织网络环境的基于分簇的路由协议，设计并实现了一种基于亲和传播算法的分簇组网模型V-APC（VANET-Affinity Propagation Clustering）。通过重新定义亲和传播算法的相似函数，设计了簇头的选择过程、簇的形成过程还有簇的维护过程。结果表明：采用上述方法形成的簇，在簇内的通信性能和簇稳定性方面具有显著优势，使得该协议在路由延迟和数据包转发成功率方面表现优异。     

Economic benefits of the OTV program

A new upper stage for the Shuttle called Orbiter Transfer Vehicle (OTV) is planned by the National Aeronautics and Space Administration (NASA) for a broad range of missions including transfer of very large spacecraft, unmanned and manned servicing at Geosynchronous orbit (GEO). Leading OTV configurations use 13 to 34 tonnes of cryogenic propellants in vehicles based on the existing Centaur or new designs. These OTVs can deliver to Geosynchronous orbit more than double the payload possible with the solid propellant Intertial Upper Stage (IUS), which is currently being developed. This high performance reduces the number of shuttle launches required to deliver a given total mass of payloads. After delivery of current size spacecraft, OTV could be returned to the Orbiter for reuse, saving the cost of building a new stage. OTV performance and flexibility will create the opportunity for the next generation of spacecraft such as Geostationary Platform. In these three ways, the high-performance OTV will provide economic benefits to Space Transportation Systems.     

用于航天器贮箱的复合材料层合板低速冲击损伤及其渗漏规律研究

得益于优异的力学性能和减质优势,贮箱复合材料化已成为新一代航天器的重要特征之一,而复合材料贮箱的冲击后渗漏问题须要重点关注。针对一种用于航天器贮箱的含表层机织布复合材料层合板,通过依次开展低速冲击试验、C扫描损伤检测和氦质谱渗漏检测,获取不同能量冲击后层合板的内部分层损伤和渗漏率,并对比分析了机织布分别置于冲击侧和背侧时层合板的渗漏规律。结果表明,将机织布层置于冲击背侧时,层合板的冲击渗漏门槛值显著提高,且发生渗漏时出现目视可见损伤。     

高超声速飞行器基于路径跟踪的制导方法

针对具有高位置精度要求的高超声速飞行器制导问题,提出一种基于路径跟踪和自适应动态面控制（ADSC）方法的轨迹跟踪制导的方法。根据高超声速飞行器动力学建立数学模型,同时为了方便跟踪控制对数学模型进行了部分线性化处理。针对高超声速滑翔飞行器三维轨迹跟踪存在的欠驱动问题,引入路径跟踪的思路将其转化为二维的跟踪控制问题;针对高超声速飞行器建模不确定性强的特点,使用具有较强鲁棒性的ADSC方法实现鲁棒的路径跟踪。能够证明所提出的方法可以以任意小的误差实现参考轨迹跟踪。仿真结果表明,该制导方法具有较高的跟踪精度并且具有较强的克服模型不确定性的能力。     

考虑高空粘性干扰效应的乘波体气动性能工程预测方法研究

粘性干扰效应是飞行器在高空、高马赫数飞行状态下所面临的诸多重要物理效应之一,对飞行器在这一区段飞行时的气动性能有着极其重要的影响.本文基于粘性干扰理论,结合参考温度方法提出了一种能够考虑粘性干扰效应的高超声速乘波体气动性能的工程预测方法,克服了传统工程预测方法不能计及粘性干扰效应的不足.文中对该方法的合理性进行了理论分析,并在飞行高度30～70km,飞行马赫数15～20范围内,通过本文提出的方法与传统工程方法以及计算流体力学(CFD)方法计算结果的比较,验证了本文所提出的方法的有效性.     

一种高速可变形飞行器智能变形决策方法

针对一类高速可变形飞行器(HMFV)的变形决策问题，提出一种基于深度确定性策略算法(DDPG)下考虑综合性能指标最优的智能变形决策方法。首先，以一类后掠角可连续变化的高速飞行器为研究对象，给出变形飞行器动力学模型，分析模型特性及变形量与关键气动参数之间的定性关系。其次，基于关键气动数据特征分析，考虑包含气动性能、控制误差在内的综合性能指标，设计一种基于DDPG算法的智能变形决策方案。再者，针对带有标称控制器的HMFV进行变形决策训练，实时获得滑翔过程中不同飞行状态下的最优构型。最后，仿真结果表明所设计的智能变形决策算法收敛效果好，且具备较好的泛化性能。相比于固定外形，可通过变形使得在不同状态下的升阻比保持最优，且与考虑单一决策指标相比，考虑综合指标最优的变形决策可进一步缩小姿态动态跟踪误差。