Development of navigation guidance and control technology for Indian launch vehicles

The orbital injection accuracy of any payload depends on the calibre of the inertial guidance system used on board the launch vehicle. This paper outlines the mission definition and the rationale for the selection of a proper guidance system to meet final mission objectives. The functions and the architecture of the navigation, guidance and control are discussed. The developmental aspects of the sophisticated inertial sensors, inertial systems, associated complex electronics, on-board computers, control actuators and systems are reported. The complexity of the on-board control and guidance software and the test and evaluation procedures used for their validation are included. The general scheme of the inertial guidance systems and the critical role played by them in the realization of Indian satellite launch vehicles SLV-3, ASLV and PSLV are presented in brief.     

Computer simulation methods for launch vehicle mission and control problems

The flight control system of a launch vehicle is the result of the right tradeoff between different objectives, such as the interaction between the control, guidance and performance aspects of a mission with specified end conditions and the analysis of the mission trajectories and vehicle systems under a variety of normal and failure modes. Hence an evaluation of the design and performance of such a system is not feasible through purely analytical means even with simplified models. This, together with the necessity for step-by-step refinement of the models used for the vehicle and its environment, calls for the computer simulation approach. The various considerations involved in developing and selecting the simulation model and implementing it on a computer are discussed. To illustrate the approach, a hybrid simulation evaluation of the performance of the first stage control system of a satellite launch vehicle and that of the controlled vehicle under different operational modes is presented.     

Modeling and reduction of shocks on electronic components within a projectile

Electronic components within a projectile are subjected to severe loads over an extremely short duration during the launch process. Failure of these components during launch can result in negative effects on the mission of the projectile. While experimental data can be helpful in understanding failure of electronic components within a projectile, collecting such data are usually difficult. There are also limitations on the reliability of sensors under these circumstances. Finite element modeling (FEM) can offer a means to better understand the behavior of these components. It can also be used to develop better shock mitigation features into the projectile design. This research has two objectives. The first objective is to develop an FEM that one describes the interaction of a typical projectile with the gun barrel during launch. The projectile includes a payload of a one-pound mass representing a typical electronic package supported by a plate. The second objective of this work is to investigate the use of composite plates to support electronic payload as a means to reduce the transmitted shocks during the projectile launch event. The proposed plate has carbon fibers embedded in an epoxy matrix. A parametric study of the effects of varying the thickness of the supporting plate and the fiber volume fraction on the accelerations and stresses is included. Results of the study are used to reach general recommendations regarding reducing failure of electronic components within a projectile.     

基于数据融合和小波分析的结构损伤诊断

基于数据融合和小波分析理论，提出一种新的结构损伤诊断方法。采用改进的一致性算法融合多传感器的测量数据，克服了一致性算法中两传感器在测量精度不同时置信距离不同的缺点，对支持矩阵进行模糊化处理，避免了人为定义阈值而产生的主观误差。利用小波分析的降噪和多尺度分辨能力对多传感器的数据进行分析处理，从而对结构损伤作出诊断识别。通过数值算例，验证了该方法可以充分利用所有传感器的有效信息，能够在部分传感器性能降低（如受到噪声影响），甚至是完全失效的情况下，对结构损伤作出正确诊断。     

Cryogenic propulsion for the Titan Orbiter Polar Surveyor (TOPS) mission

Liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic propellants can dramatically enhance NASA’s ability to explore the solar system due to their superior specific impulse (I_sp) capability. Although these cryogenic propellants can be challenging to manage and store, they allow significant mass advantages over traditional hypergolic propulsion systems and are therefore enabling for many planetary science missions. New cryogenic storage techniques such as subcooling and the use of advanced insulation and low thermal conductivity support structures will allow for the long term storage and use of cryogenic propellants for solar system exploration and hence allow NASA to deliver more payloads to targets of interest, launch on smaller and less expensive launch vehicles, or both. These new cryogenic storage technologies were implemented in a design study for the Titan Orbiter Polar Surveyor (TOPS) mission, with LH2 and LO2 as propellants, and the resulting spacecraft design was able to achieve a 43% launch mass reduction over a TOPS mission, that utilized a traditional hypergolic propulsion system with mono-methyl hydrazine (MMH) and nitrogen tetroxide (NTO) propellants. This paper describes the cryogenic propellant storage design for the TOPS mission and demonstrates how these cryogenic propellants are stored passively for a decade-long Titan mission that requires the cryogenics propellants to be stored for 8.5 years.     

基于精密压力传感器气密性检测装置设计

针对某型装备润滑系统、燃油系统、助推器等气密性检测要求,分析了常用气密性检测方法的特点,介绍了一种基于精密压力传感器的直压气密性检测技术方案,采用DPS8000精密压力传感器作为系统压力检测单元,内嵌了无油干式压力泵作为压力源,采用了STM32F103VE微控制器,提出了基于最小二乘法的气压泄漏率计算模型和温度补偿模型,扩展了设备温度适应范围,提高了检测准确度,实现了气密性自动检测,检测压力和保压时间数字设定,具有操作方便、显示直观、灵敏度高、便携性好的特点。     

冷氦增压系统的研制

冷氦增压在我国是首次研制，并成功地应用于长征三号甲型号。冷氦增压是长征三号甲研制中的关键技术之一。介绍了冷氦增压系统方案，冷氦增压方案的优点，箭上冷氦增压系统几个部件的主要设计技术要求，冷氦增压系统地面设备及主要技术要点、箭上冷氦增压系统安全措施，冷氦增压系统的试验及主要结果，冷氦增压计算，长征三号甲火箭冷氦增压试验结果以及国外冷氦增压技术等。     

中国的载人运载火箭

高可靠、高安全和高质量，确保航天员绝对安全，是对载人运载火箭的最主要要求。简要介绍了CZ-2F载人运载火箭的总体方案、系统组成以及设计的基本原则。     

某多管火箭炮车炮刚柔耦合极限路况行驶动力学研究

采用刚柔耦合的方法,对某新型车载防空多管火箭炮进行了通过垂直凸台、水平壕沟、陡坡等极限路况的行驶动力学仿真研究.建立了越野车与防空多管火箭炮的刚柔耦合动力学模型,并考虑轮胎的作用,计算了越野车前后轮通过垂直凸台的极限高度和水平壕沟的极限宽度,建立各种极限路况的计算模型,对车载防空多管火箭炮进行了行驶动力学仿真,得到全系...     

Solar sailing: mission applications and engineering challenges

Solar sailing is emerging as a promising form of advanced spacecraft propulsion, which can enable exciting new space-science mission concepts. By exploiting the momentum transported by solar photons, solar sails can perform high-energy orbit-transfer manoeuvres without the need for reaction mass. Missions such as planetary-sample return, multiple small-body rendezvous and fast missions to the outer Solar System can therefore be enabled with the use of only a modest launch vehicle. In addition, new families of highly non-Keplerian orbits have been identified that are unique to solar sails, and can enable new ways of performing space-science missions. While the opportunities presented by solar sailing are appealing, engineering challenges are still to be solved before the technology finally comes to fruition.     

整星隔振防摇装置的研究

在研究整星隔振过程中发现,整星的摇晃振动模态和纵向振动模态密切耦合,根据整星隔振系统基频下降的设计限制,仅采用一般的隔振器,其纵向隔振性能不易满足设计指标.针对某卫星的隔振设计要求,提出由七连杆防摇机构和隔振器组成的整星隔振防摇方案,通过有限元模型的模态分析和缩比卫星模型的性能测试,验证此方案的可行性,即在不提高整星的纵向振动模态频率下,提高摇晃振动的模态频率.     

Phased mission modelling of systems with maintenance-free operating periods using simulated Petri nets

A common scenario in engineering is that of a system which operates throughout several sequential and distinct periods of time, during which the modes and consequences of failure differ from one another. This type of operation is known as a phased mission, and for the mission to be a success the system must successfully operate throughout all of the phases. Examples include a rocket launch and an aeroplane flight. Component or sub-system failures may occur at any time during the mission, yet not affect the system performance until the phase in which their condition is critical. This may mean that the transition from one phase to the next is a critical event that leads to phase and mission failure, with the root cause being a component failure in a previous phase. A series of phased missions with no maintenance may be considered as a maintenance-free operating period (MFOP). This paper describes the use of a Petri net (PN) to model the reliability of the MFOP and phased missions scenario. The model uses Monte-Carlo simulation to obtain its results, and due to the modelling power of PNs, can consider complexities such as component failure rate interdependencies and mission abandonment. The model operates three different types of PN which interact to provide the overall system reliability modelling. The model is demonstrated and validated by considering two simple examples that can be solved analytically.     

航行体水下垂直发射空泡脱落条件研究

航行体水下垂直发射出水时空泡的脱落能够影响溃灭的压力以至结构的设计载荷,因此研究发射条件对空泡脱落和空泡状态的影响具有重要的意义。该文首先从回转体空泡脱落的特征出发,确认回射流是导致空泡脱落的控制因素,进而通过对空泡脱落的影响分析,提出回射流运动时间和航行体运动时间的比值大小能够作为空泡脱落和空泡状态的判据;接着对不同深度发射工况流场与空泡演化过程开展了系统数值模拟并与典型试验结果对比分析,验证并给出了该判据其在典型发射工况下的量化表达式;进而讨论了空化数、弗劳德数、深度、发射速度、泡内压力等条件对空泡断裂脱落的影响,给出了出水空泡溃灭时产生较强随机性高压力脉冲的发射区间。     

Multidisciplinary design and optimization of an air launched satellite launch vehicle using a hybrid heuristic search algorithm

A multidisciplinary design and optimization strategy for a multistage air launched satellite launch vehicle comprising of a solid propulsion system to low earth orbit with the implementation of a hybrid heuristic search algorithm is proposed in this article. The proposed approach integrated the search properties of a genetic algorithm and simulated annealing, thus achieving an optimal solution while satisfying the design objectives and performance constraints. The genetic algorithm identified the feasible region of solutions and simulated annealing exploited the identified feasible region in search of optimality. The proposed methodology coupled with design space reduction allows the designer to explore promising regions of optimality. Modules for mass properties, propulsion characteristics, aerodynamics, and flight dynamics are integrated to produce a high-fidelity model of the vehicle. The objective of this article is to develop a design strategy that more efficiently and effectively facilitates multidisciplinary design analysis and optimization for an air launched satellite launch vehicle.     

一种整星隔振器的研制

在不改变卫星和运载火箭结构前提下,针对某型号的配重卫星模型的隔振要求和约束条件,研制出一种新型的整星隔振器,安装在适配器和星箭界面之间,阐述隔振器的设计路线,建立合理的有限元模型,进行隔振器的结构参数灵敏度分析和频率响应计算,最后对整星隔振系统的进行动态性能测试。有限元计算结果和试验数据表明,所研制的整星隔振器能够有效地隔离卫星在宽频带上横向和纵向振动,降低发射时运载火箭传递给卫星的环境载荷。     

美苏氢温度介质下材料物理力学性能研究

概述氢温度介质环境下材料物理力学行为及其研究的生,介绍了美国和前苏联在研制大型运载火箭等航天器过程中,对液氢及氢介质条件下使用的结构材料的物理力学性能研究的情况,强调了开展氢介质力学性能研究的必要性和重要性。对我国大型运载火箭材料低温性能的需求作了分析和预测。     

装甲车辆状态信息远程监控系统设计与实现

从装甲车辆实车实验需求出发,分析了装甲车辆的传感器设备,以车辆状态采集器、GPS(全球定位系统)接收机为信息采集端,利用数据电台轮询方式时分多址传输的机制,实现了对装甲车辆实车实验时车辆状态、车辆位置等信息的显示和远程监控,并能在车辆状态超出设定阈值后向实验车辆发送报警信息.     

Vacuum Radiance-Temperature Standard Facility for Infrared Remote Sensing at NIM

As infrared remote sensors are very important parts of Earth observation satellites, they must be calibrated based on the radiance temperature of a blackbody in a vacuum chamber prior to launch. The uncertainty of such temperature is thus an essential component of the sensors’ uncertainty. This paper describes the vacuum radiance-temperature standard facility (VRTSF) at the National Institute of Metrology of China, which will serve to calibrate infrared remote sensors on Chinese meteorological satellites. The VRTSF can be used to calibrate vacuum blackbody radiance temperature, including those used to calibrate infrared remote sensors. The components of the VRTSF are described in this paper, including the VMTBB, the LNBB, the FTIR spectrometer, the reduced-background optical system, the vacuum chamber used to calibrate customers’ blackbody, the vacuum-pumping system and the liquid-nitrogen-support system. The experimental methods and results are expounded. The uncertainty of the radiance temperature of VMTBB is 0.026 °C at 30 °C over 10 μm.     

A profile of aerodynamic research in VSSC with application to satellite launch vehicles

Some of the aerodynamic and fluid dynamic problems associated with the satellite launch vehicles of the Indian Space programme are discussed in this paper. Taking into account the aerodynamic flight profile of a launch vehicle and also the configurations of interest, the topics considered are vehicle lift-off fluid dynamics, strap-on aerodynamics, strap-on separation problem, stage separation aerodynamics, boat-tail aerodynamics, hypersonic flow problems and some special fluid dynamic problems of interest like nozzle flows and shock wave-boundary layer interaction. The topics are mostly related to external fluid dynamics though certain aspects of internal ballistics are also touched upon. More emphasis is placed on the results and their applications rather than on the methodology used. Also, the main features of two special wind tunnel facilities developed at the Vikram Sarabhai Space Centre (VSSC) — the hypersonic wind tunnel and the heat transfer facility — are discussed briefly. It is hoped that this paper will give an idea of the profile of aerodynamic research being conducted inVSSC with application to satellite launch vehicles in view.