基于气动辅助变轨的变气动外形飞行器概念研究--最优控制律问题

从一个天地往返飞行器的上升轨道和再入返回轨道的优化，以及适用不同飞行任务的变轨要求的气动外形问题，提出一项基于气动力辅助变轨的变气动外形飞行器的新概念研究。对于一个固定气动外形飞行器要同时满足上升轨道有效载荷最大和再入轨道热流峰值、过载峰值及机动性能约束下的成本最低往往是困难的。若同时满足不同飞行任务：飞往太空站的运输任务，空间拦截和交会机动巡航任务及星际探测任务，则更为困难，实际上是不可能的。文章研究基于气动力辅助变轨，在热流约束下，气动外形参数变化对最优控制的影响。其结论为：热流约束下的最优控制解，包括考虑推力协同变轨，除了在非约束弧的滚转角不直接受气动外形影响外，其余的控制律，升力系数和滚转角都是气动外形参数和攻角的函数。因而变气动外形可作为一项新技术，即通过气动外形参数变化和相应的变轨策略而获得性能和成本都最佳的用途很广的一种新型飞行器。     

Nonlinear feedback control of low-thrust orbital transfer in a central gravitational field

This paper investigates the problem of continuous-thrust orbital transfer using orbital elements feedback from a nonlinear control standpoint, utilizing concepts of controllability, feedback stabilizability and their interaction. Gauss's variational equations (GVEs) are used to model the state-space dynamics of motion under a central gravitational field. First, the notion of accessibility is reviewed. It is then shown that the GVEs are globally accessible. Based on the accessibility result, a nonlinear feedback controller is derived which asymptotically steers a spacecraft form an initial elliptic orbit to any given elliptic orbit. The performance of the new controller is illustrated by simulating an orbital transfer between two geosynchronous Earth orbits. It is shown that the low-thrust controller requires less fuel than an impulsive maneuver for the same transfer time. Closed-form, analytic expressions for the new orbital transfer controller are given. Finally, it is proven, based on a topological nonlinear stabilizability test, that there does not exist a continuous closed-loop controller that can transfer a spacecraft onto a parabolic escape trajectory.     

Solution of nonlinear optimal control problems by the interpolating scaling functions

This paper presents a numerical method for solving nonlinear optimal control problems including state and control inequality constraints. The method is based upon interpolating scaling functions. The differential and integral expressions which arise in the system dynamics, the performance index and the boundary conditions are converted into some algebraic equations which can be solved for the unknown coefficients. Illustrative examples are included to demonstrate the validity and applicability of the technique.     

The use of solutions to problems of spacecraft trajectory optimization in impulse formulation when solving the problems of optimal control of trajectories of a spacecraft with limited thrust engine: I

In this first part of our paper, it is suggested to use solutions to boundary value problems in the optimization problems (in impulse formulation) for spacecraft trajectories in order to obtain the initial approximation, when boundary value problems of the maximum principle are solved numerically by the shooting method. The technique suggested is applied to the problems of optimal control over motion of the center of mass of a spacecraft controlled by the thrust vector of jet engine with limited thrust in an arbitrary gravitational field in a vacuum. The method is based on a modified (in comparison to the classic scheme) shooting method computation together with the method of continuation along a parameter (maximum reactive acceleration, initial thrust-to-weight ratio, or any other parameter equivalent to them). This technique allows one to obtain the initial approximation with a high precision, and it is applicable to a wide range of optimal control problems solved using the maximum principle, if the impulse formulation makes sense for these problems.     

The use of solutions to problems of spacecraft trajectory optimization in impulse formulation when solving the problems of optimal control of trajectories of a spacecraft with limited thrust engine: II

As examples of application of the technique suggested in the first part of this work, the problems of optimizing the trajectories of spacecraft transfers between circular coplanar orbits are considered in this second part. During the transfer the spacecraft is controlled by the vector of thrust of a limited-thrust jet engine. The mass consumption is minimized for a limited time of transfer. Extreme trajectories with two and three powered sections (Homan-type and bi-elliptic transfer trajectories) are numerically determined. The solution of these well-studied problems allows one to compare the results of applying the suggested technique with the results of application of other previously used techniques.     

Near-optimal geostationary transfer maneuvers with cooperative en-route inspection using hybrid optimal control

This research investigates the performance of bi-level hybrid optimal control algorithms in the solution of minimum delta-velocity geostationary transfer maneuvers with cooperative en-route inspection. The maneuvers, introduced here for the first time, are designed to populate a geostationary constellation of space situational awareness satellites while providing additional characterization of objects in lower-altitude orbit regimes. The maneuvering satellite, called the chaser, performs a transfer from low Earth orbit to geostationary orbit, during which it performs an inspection of one of several orbiting targets in conjunction with a ground site for the duration of the target?s line-of-site contact with that site. A three-target scenario is used to test the performance of multiple bi-level hybrid optimal control algorithms. A bi-level hybrid algorithm is then utilized to solve fifteen-, and thirty-target scenarios and shown to have increasing benefit to complete enumeration as the number of targets is increased. Results indicate that the en-route inspection can be accomplished for a small increase in the delta-velocity required for a simple transfer to geostationary orbit given the same initial conditions.     

MDO approach for early design of aerobraking orbital transfer vehicles

This paper presents a new multidisciplinary design optimization (MDO) methodology for preliminary design of an aeroassisted orbital transfer vehicle (AOTV) performing a two-way transfer between a low-Earth “parking” orbit and a high-energy orbit. This work has been performed in the frame of Onera's CENTOR [N. Bérend, C. Jolly, F. Jouhaud, D. Lazaro, Y. Mauriot, C. Monjaret, J.M. Moschetta, M. Parlier, J.L. Pastre, Y. Servouze, J.L. Vérant, Project CENTOR: Preparing the design of future orbital transfer vehicles; IAC-07-D.2.3.07, in: 58th International Astronautical Congress, 24–28/09/2007, Hyderabad, India] project whose objective is to prepare tools and methodology for studying and designing future space transportation systems for new kinds of missions such as on-orbit servicing (OOS), payload ferrying, or in-situ observation of space-debris. Using simplified models and an appropriate low-dimension formulation for the optimization problem the method makes possible to obtain rapidly and easily a global view of the trade-off between the payload mass and the total mass. It also makes possible to discuss the feasibility of the vehicle with regard to different multidisciplinary constraints and technology hypotheses for the heat shield. This approach is illustrated by eight different AOTV design studies, considering two different missions (LEO–MEO and LEO–GEO), two different propulsion technologies (LOX-LH2 and LOX-CH4) and two different thermal protection system (TPS) characteristics. In each case, we discuss the feasibility and characteristics of the lightest vehicle carrying a prescribed 100 kg payload, and, conversely, a heavy vehicle with a prescribed 18 ton total mass, carrying the heaviest possible payload.     

MDO approach for early design of aerobraking orbital transfer vehicles

This paper presents a new multidisciplinary design optimization (MDO) methodology for preliminary design of an aeroassisted orbital transfer vehicle (AOTV) performing a two-way transfer between a low-Earth “parking” orbit and a high-energy orbit. This work has been performed in the frame of Onera's CENTOR [N. Bérend, C. Jolly, F. Jouhaud, D. Lazaro, Y. Mauriot, C. Monjaret, J.M. Moschetta, M. Parlier, J.L. Pastre, Y. Servouze, J.L. Vérant, Project CENTOR: Preparing the design of future orbital transfer vehicles; IAC-07-D.2.3.07, in: 58th International Astronautical Congress, 24–28/09/2007, Hyderabad, India] project whose objective is to prepare tools and methodology for studying and designing future space transportation systems for new kinds of missions such as on-orbit servicing (OOS), payload ferrying, or in-situ observation of space-debris. Using simplified models and an appropriate low-dimension formulation for the optimization problem the method makes possible to obtain rapidly and easily a global view of the trade-off between the payload mass and the total mass. It also makes possible to discuss the feasibility of the vehicle with regard to different multidisciplinary constraints and technology hypotheses for the heat shield. This approach is illustrated by eight different AOTV design studies, considering two different missions (LEO–MEO and LEO–GEO), two different propulsion technologies (LOX-LH2 and LOX-CH4) and two different thermal protection system (TPS) characteristics. In each case, we discuss the feasibility and characteristics of the lightest vehicle carrying a prescribed 100 kg payload, and, conversely, a heavy vehicle with a prescribed 18 ton total mass, carrying the heaviest possible payload.     

一种姿轨控发动机地面试验控制系统设计

控制系统在试车过程中发出指令信号,使发动机和试车台各工艺系统按照预定的程序完成规定试车任务。为了适应某型号姿轨控发动机试验对控制精度要求的提高,需要对现有控制系统进行改造,基于现场可编程门阵列和高速固态继电器设计了一种新的控制系统,其控制模块用于操作人员对试验系统上的各路电磁阀进行手动与自动控制,其复记模块可以实时记录系统状态以便于指挥人员了解系统工作状况,并能够进行数据分析判断。该系统建成后可以满足该型号姿轨控发动机试验的要求,控制精度为0.1 ms。     

空空导弹的一种优化复合控制方案研究

通过研究空空导弹推力矢量和气动复合控制与一般气动控制方式，讨论了新一代近距格斗型空空导弹的一种控制模式，并参照某型导弹的数学模型及有关气动参数对两种控制方式进行了数字仿真。结果表明，在新一代格斗型空空导弹上采用复合控制方式可明显地改善导弹的总体性能。     

H2 optimal control solution for a combined energy and attitude control system

This addendum is an extension of a series of research work on the combined energy and attitude control system (CEACS) for small satellites. Previous works appeared in Acta Astronautica showed that the CEACS is able to simultaneously perform the attitude control and energy storage task. This addendum focuses on the CEACS attitude control enhancement by employing the H₂ optimal control method. Governing equations describing the CEACS H₂ attitude control architecture are established. Numerical treatments are performed in order to validate the attitude control option.     

Application of SDRE technique to orbital and attitude control of spacecraft formation flying

This paper proposes the application of a nonlinear control technique for coupled orbital and attitude relative motion of formation flying. Recently, mission concepts based on the formations of spacecraft that require an increased performance level for in-space maneuvers and operations, have been proposed. In order to guarantee the required performance level, those missions will be characterized by very low inter-satellite distance and demanding relative pointing requirements. Therefore, an autonomous control with high accuracy will be required, both for the control of relative distance and relative attitude. The control system proposed in this work is based on the solution of the State-Dependent Riccati Equation (SDRE), which is one of the more promising nonlinear techniques for regulating nonlinear systems in all the major branches of engineering. The coupling of the relative orbital and attitude motion is obtained considering the same set of thrusters for the control of both orbital and attitude relative dynamics. In addition, the SDRE algorithm is implemented with a timing update strategy both for the controller and the proposed nonlinear filter. The proposed control system approach has been applied to the design of a nonlinear controller for an up-to-date formation mission, which is ESA Proba-3. Numerical simulations considering a tracking signal for both orbital and attitude relative maneuver during an operative orbit of the mission are presented.     

Adaptive environmental control for optimal results during plant microgravity experiments

The SVET Space Greenhouse (SG)--the first and the only automated plant growth facility onboard the MIR Space Station in the period 1990-2000 was developed on a Russian-Bulgarian Project in the 80s. The aim was to study plant growth under microgravity in order to include plants as a link of future Biological Life Support Systems for the long-term manned space missions. An American developed Gas Exchange Measurement System (GEMS) was added to the existing SVET SG equipment in 1995 to monitor more environmental and physiological parameters. A lot of long-duration plant flight experiments were carried out in the SVET+GEMS. They led to significant results in the Fundamental Gravitational Biology field--second-generation wheat seeds were produced in the conditions of microgravity. The new International Space Station (ISS) will provide a perfect opportunity for conducting full life cycle plant experiments in microgravity, including measurement of more vital plant parameters, during the next 15-20 years. Nowadays plant growth facilities for scientific research based on the SVET SG functional principles are developed for the ISS by different countries (Russia, USA, Italy, Japan, etc.). A new Concept for an advanced SVET-3 Space Greenhouse for the ISS, based on the Bulgarian experience and "know-how" is described. The absolute and differential plant chamber air parameters and some plant physiological parameters are measured and processed in real time. Using the transpiration and photosynthesis measurement data the Control Unit evaluates the plant status and performs adaptive environmental control in order to provide the most favorable conditions for plant growth at every stage of plant development in experiments. A conceptual block-diagram of the SVET-3 SG is presented.     

稀疏拟谱最优控制法求解Goddard火箭问题

提出了一种新的基于直接转化法的求解基于常微分方程(ODE)和微分代数方程(DAE)的最优控制问题的数值方法.该方法通过Legendre-Gauss拟谱法同时离散化状态变量和控制变量,把最优控制问题转化为一个非线性规划问题,并利用改进的多相处理方法避免优化无控段,同时基于稀疏矩阵探索其一阶导数信息.数值结果表明,与传统的直接转换法相比,该方法是一种通用高效的精度较高的ODE/DAE最优控制直接数值求解法.最后,从工程观点出发,应用该方法成功求解了终端自由有路径约束的奇异最优控制问题Goddard火箭问题.     

火箭入轨的大偏航非线性鲁棒自适应控制方法

火箭入轨通常是沿标准轨道面的飞行控制,常规发射任务只需侧向小偏航角校正,但当今一些特殊的入轨任务要求火箭制导控制能侧向大偏航角飞行,以克服较大初始侧向偏差对末级火箭入轨的影响。文中提出了一种末级火箭的侧向大偏航非线性自适应组合制导控制方法,结合土星-5火箭IMG方法和航天飞机LTG方法各自的优点,进行了大偏航角的非线性耦合补偿修正,并对动力飞行过程的迭代算法进行了鲁棒稳定性改造。基于姿态喷嘴开关控制的六自由度数值仿真表明,提出的控制策略和算法简单可靠、稳定性好、精度高,在火箭入轨控制和空间飞行器变轨控制中具有参考和应用价值。     

Profiles of quick access to the orbital station for modern spacecraft

The problems of decreasing the duration of the autonomous flight of a spacecraft (SC) before the docking with an orbital station (OS) are considered in this paper. Modern SCs should be docked with the International Space Station (ISS) with an arbitrary initial phase angle; for this reason, the rendezvous of the Russian Soyuz-TMA spacecraft with the ISS is performed for 2 days. The paper presents to consideration some new flight profiles with essentially smaller duration. The results of modeling the developed rendezvous profiles are presented and solutions to emergency situations are considered.     

Quasioptimal control with feedback for multiorbit low-thrust transfer between noncoplanar elliptic and circular orbits

The problem of synthesizing stable feedback control is considered based on solving the problem of time minimization for a multiorbit transfer between noncoplanar elliptic and circular orbits in a Newtonian gravitational field. The problem is solved using asymptotic properties and symmetries of optimal control in the unperturbed problem. Stability of the obtained control against external perturbations, deviations of initial conditions, and errors in thrust vector realizations is demonstrated. The obtained quasioptimal control with feedback can be used as an onboard algorithm of spacecraft control and when performing design and ballistic analysis.     

相对振型转换法求解分段线性边界条件连续体振动响应的优点

为了研究航天领域关注的非线性连接结构对整体结构振动响应的影响,文章利用变分原理,重新推导了基于振型转换的求解具有任意分段线性及可以转换为分段线性边界条件的连续体振动响应的方法——相对振型转换法,并与文献[10]中的混合方法的推导过程进行了对比,印证了相对振型转换法的正确性及其优点。针对一个施加均布载荷的端点单侧带阻挡的悬臂梁算例,分别应用相对振型转换法及力积分法进行了求解,通过悬臂梁端点的分岔响应研究了系统的非线性响应,讨论了两种方法得到的结果。     

Families of periodic solutions to Hamiltonian systems: Nonsymmetrical periodic solutions for a planar restricted three-body problem

A monodromy matrix calculated at a single arbitrary point of the periodic solution to a Hamiltonian system allows one to obtain both the direction of continuation for the family of solutions of the first (in Poincarés sense) kind and the multiplicity and direction of branching for periodic solutions of the second kind. In case of resonances 1 : 1 and 1 : 2 one needs to take into account the structure of elementary divisors of the monodromy matrix. Using the planar circular restricted three-body problem as an example, the infiniteness of the process of branching for a nonintegrable system and its finiteness for an integrable system are demonstrated. It is proved that periodic solutions of both first and second kinds which are obtained by continuation of symmetric periodic solutions of a restricted problem are also symmetric. The only exception is the case of resonance 1 : 1 and two second-order cells of the monodromy matrix in the Jordanian form. In this case, all periodic solutions of the second kind turned out to be nonsymmetrical. Examples of the families of nonsymmetrical periodic solutions are given.__________Translated from Kosmicheskie Issledovaniya, Vol. 43, No. 2, 2005, pp. 88–110.Original Russian Text Copyright © 2005 by Kreisman.     

Variation of the orbital elements for parabolic trajectories due to a small impulse using Gauss equations

Firstly we derive Gauss’ perturbation equation for parabolic motion using Murray–Dermott and Kovalevsky procedures. Secondly, we easily deduce the variations of the orbital elements for the parabolic trajectories due to a small impulse at any point along the path and at the vertex of the parabola.