皮卫星星箭分离机构设计分析方法及关键技术研究
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摘要
皮卫星是基于MEMS微机电系统技术的一种全新概念的微小卫星,它具有研发周期短、制造发射成本低、功能针对性强、载荷接口灵活等常规卫星无法比拟的优点。而星箭分离机构作为皮卫星项目中的重要组成部分,直接决定着皮卫星发射任务的成败。论文结合“十一五”预研项目“皮卫星技术*****”,采用理论分析、计算机模拟和实验研究相结合的方法,系统、深入地开展了皮卫星星箭分离机构基础理论和关键技术的研究,设计并研制了皮卫星星箭分离机构,2010年9月22日,“皮星一号A”卫星的成功发射验证了星箭分离机构适用性、有效性和可靠性。
论文首先阐述了课题的研究背景与意义,详细论述了皮卫星星箭分离机构国内外研究现状,明确了星箭分离机构设计分析的关键技术。在此基础上,提出了论文的主要研究内容。
针对皮卫星对星箭连接分离的技术指标要求和功能需要,设计并研制了皮卫星星箭分离机构。为实现星箭无干涉分离,论文分别对动力系统、防干涉系统等关键部件进行了理论计算和数值分析。
星箭分离机构主要由薄壁件组成,在振动环境中容易引起刚柔耦合振动,对薄弱部件产生破坏作用。为保证星箭分离机构在复杂振动环境中的连接可靠性,必须分析其振动特性。皮卫星和星箭分离机构整体模型很复杂,求解难度大,根据边界连接关系将整体系统分解为皮卫星、舱门、框架、后板四个子结构,分别对各子结构进行了理论建模分析,然后通过坐标转换和方程联立,实现皮卫星及其星箭分离机构整体求解。通过对皮卫星和星箭分离机构组合体的模态和静力学分析,得到了组合体整体力学性能。
发射过程中,除了复杂力学环境外,还需考虑热真空环境影响,卫星朝阳面和背阴面产生的截面温差将导致结构的热变形,影响到皮卫星的正常分离。为了解皮卫星在发射过程中的热变形情况,以皮卫星及其星箭分离机构组合体为研究对象,根据卫星发射轨道参数,计算皮卫星及其星箭分离机构在卫星发射过程中所受到的内外热流,并以其所受到的内外热流为输入条件,采用有限元软件对其进行热-结构模拟,分析了卫星发射过程中皮卫星外壳与导轨面的温度场和热变形,以达到合理设计皮卫星与导轨面间配合间隙的目的。
皮卫星能源有限、姿控能力差等特点决定了其对星箭分离初始姿态要求较高,为合理设计星箭分离机构,确保星箭成功分离,在对皮卫星分离过程中所受外力和外力矩进行分析的基础上,建立了皮卫星星箭分离动力学仿真模型,并通过地面分离试验对模型进行了修订,并对卫星入轨初始姿态进行了预测。
最后,通过对星箭分离机构的系列化试验研究,验证了理论分析与数值模拟的正确性和可行性。在试验中,对皮卫星及其星箭分离机构组合体进行了力学环境试验和分离试验研究,并将试验结果与理论分析、仿真计算结果进行比较,达到对星箭分离机构动态设计模型进行校核的目的。
Pico-satellite is a new concept of micro-satellite based on MEMS technology. Compared with conventional satellite, it owns the advantages of shorter development cycle, lower manufacturing and launching cost, stronger functional targeted and load interface flexibility. As an important part of the Pico-satellite project, the separation mechanism directly determines the success or failure of the pico-satellite mission. Supported by the'Eleventh Five-Year'pre-research project 'pico-satellite technology*****', a pico-satellite separation mechanism is developed. The research work of this dissertation is carried out through the methods of theoretical analysis, computer simulation and experimental verification. On September22,2010,'ZDPS-1A' satellite was successfully launched, and download pictures of the earth taken by built-in camera. Downlink data show that the pico-satellite work properly.
Paper first discusses the background and significance of the research. The current research situations of the pico-satellite separation mechanism, key technologies of design and analysis theory are expatiated. And then, the research contents of this paper are proposed.
According to the technical requirements and functional needs, the proposed separation mechanism is designed and developed. In order to fulfill the task of no-interference separation, key components such as guidance system, power system, and anti-interference system are theoretically calculated and analyzed with simulation softwares.
The pico-satellite separation mechanism is mainly composed of thin-walled parts, which are subject to rigid-flexible coupling vibration. Such vibration can cause damages to weak parts. In order to ensure the reliability of the connection between the satellite and the rocket, it is necessary to analyze the vibration characteristics. As the model is too complex to solve as a whole, it is decomposed into four sub-system, namely the pico-satellite, the hatch door, the frame, and the back plate. Each sub-system is theoretically modeled and analyzed respectively. Then coordinates transformation and equations are employed to realize an integrated solution.
In addition to complex mechanical environment, thermal vacuum environment is also needed to pay attention to in the process of launch. Temperature difference between sunshine side and shady side will cause thermal deformation, which will affect normal separation of the pico-satellite. Based on the consideration of geometry relations of the satellite, the earth, and the sun in the process of launch, the internal and external heat flow absorbed by the 'ZDPS-1A' satellite and its separation mechanism is analyzed. Temperature field and thermal deformation are got by finite element analysis. The clearance fit value between the satellite and guide rail is rationally designed based on above analyses.
Pico-satellite energy is limited, and pico-satellite attitude control ability is poor, so pico-satellite attitude control system make high claims on orbit injection initial attitude. In order to design the separation mechanism rationally, ensure successful separation between the satellite and the rocket, the pico-satellite separation dynamic model is established based on the analysis of external forces and external moments acting on the pico-satellite. The virtual prototyping model is modified by comparing the results of ground tests and simulation results. After that, the initial attitude was predicted.
Finally, experiments are used to verify the correctness of the theoretical analysis and feasibility. The pico-satellite and its separation mechanism is used to carry out mechanical environment tests and separation tests. The dynamic design model of the separation mechanism is verified through comparisons between test results and results of theoretical analysis as well as numerical simulation.
论文首先阐述了课题的研究背景与意义,详细论述了皮卫星星箭分离机构国内外研究现状,明确了星箭分离机构设计分析的关键技术。在此基础上,提出了论文的主要研究内容。
针对皮卫星对星箭连接分离的技术指标要求和功能需要,设计并研制了皮卫星星箭分离机构。为实现星箭无干涉分离,论文分别对动力系统、防干涉系统等关键部件进行了理论计算和数值分析。
星箭分离机构主要由薄壁件组成,在振动环境中容易引起刚柔耦合振动,对薄弱部件产生破坏作用。为保证星箭分离机构在复杂振动环境中的连接可靠性,必须分析其振动特性。皮卫星和星箭分离机构整体模型很复杂,求解难度大,根据边界连接关系将整体系统分解为皮卫星、舱门、框架、后板四个子结构,分别对各子结构进行了理论建模分析,然后通过坐标转换和方程联立,实现皮卫星及其星箭分离机构整体求解。通过对皮卫星和星箭分离机构组合体的模态和静力学分析,得到了组合体整体力学性能。
发射过程中,除了复杂力学环境外,还需考虑热真空环境影响,卫星朝阳面和背阴面产生的截面温差将导致结构的热变形,影响到皮卫星的正常分离。为了解皮卫星在发射过程中的热变形情况,以皮卫星及其星箭分离机构组合体为研究对象,根据卫星发射轨道参数,计算皮卫星及其星箭分离机构在卫星发射过程中所受到的内外热流,并以其所受到的内外热流为输入条件,采用有限元软件对其进行热-结构模拟,分析了卫星发射过程中皮卫星外壳与导轨面的温度场和热变形,以达到合理设计皮卫星与导轨面间配合间隙的目的。
皮卫星能源有限、姿控能力差等特点决定了其对星箭分离初始姿态要求较高,为合理设计星箭分离机构,确保星箭成功分离,在对皮卫星分离过程中所受外力和外力矩进行分析的基础上,建立了皮卫星星箭分离动力学仿真模型,并通过地面分离试验对模型进行了修订,并对卫星入轨初始姿态进行了预测。
最后,通过对星箭分离机构的系列化试验研究,验证了理论分析与数值模拟的正确性和可行性。在试验中,对皮卫星及其星箭分离机构组合体进行了力学环境试验和分离试验研究,并将试验结果与理论分析、仿真计算结果进行比较,达到对星箭分离机构动态设计模型进行校核的目的。
Pico-satellite is a new concept of micro-satellite based on MEMS technology. Compared with conventional satellite, it owns the advantages of shorter development cycle, lower manufacturing and launching cost, stronger functional targeted and load interface flexibility. As an important part of the Pico-satellite project, the separation mechanism directly determines the success or failure of the pico-satellite mission. Supported by the'Eleventh Five-Year'pre-research project 'pico-satellite technology*****', a pico-satellite separation mechanism is developed. The research work of this dissertation is carried out through the methods of theoretical analysis, computer simulation and experimental verification. On September22,2010,'ZDPS-1A' satellite was successfully launched, and download pictures of the earth taken by built-in camera. Downlink data show that the pico-satellite work properly.
Paper first discusses the background and significance of the research. The current research situations of the pico-satellite separation mechanism, key technologies of design and analysis theory are expatiated. And then, the research contents of this paper are proposed.
According to the technical requirements and functional needs, the proposed separation mechanism is designed and developed. In order to fulfill the task of no-interference separation, key components such as guidance system, power system, and anti-interference system are theoretically calculated and analyzed with simulation softwares.
The pico-satellite separation mechanism is mainly composed of thin-walled parts, which are subject to rigid-flexible coupling vibration. Such vibration can cause damages to weak parts. In order to ensure the reliability of the connection between the satellite and the rocket, it is necessary to analyze the vibration characteristics. As the model is too complex to solve as a whole, it is decomposed into four sub-system, namely the pico-satellite, the hatch door, the frame, and the back plate. Each sub-system is theoretically modeled and analyzed respectively. Then coordinates transformation and equations are employed to realize an integrated solution.
In addition to complex mechanical environment, thermal vacuum environment is also needed to pay attention to in the process of launch. Temperature difference between sunshine side and shady side will cause thermal deformation, which will affect normal separation of the pico-satellite. Based on the consideration of geometry relations of the satellite, the earth, and the sun in the process of launch, the internal and external heat flow absorbed by the 'ZDPS-1A' satellite and its separation mechanism is analyzed. Temperature field and thermal deformation are got by finite element analysis. The clearance fit value between the satellite and guide rail is rationally designed based on above analyses.
Pico-satellite energy is limited, and pico-satellite attitude control ability is poor, so pico-satellite attitude control system make high claims on orbit injection initial attitude. In order to design the separation mechanism rationally, ensure successful separation between the satellite and the rocket, the pico-satellite separation dynamic model is established based on the analysis of external forces and external moments acting on the pico-satellite. The virtual prototyping model is modified by comparing the results of ground tests and simulation results. After that, the initial attitude was predicted.
Finally, experiments are used to verify the correctness of the theoretical analysis and feasibility. The pico-satellite and its separation mechanism is used to carry out mechanical environment tests and separation tests. The dynamic design model of the separation mechanism is verified through comparisons between test results and results of theoretical analysis as well as numerical simulation.
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