工程知识管理系统模型研究

针对工程知识管理系统模型多样化的现状,从需求、功能和应用三个方面论述了对工程知识管理系统模型的理解和认识以及需求、功能及应用模型的含义和组成。提出工程知识管理系统模型由知识需求、系统功能和知识应用构成。在工程知识管理系统模型的基础上,成功设计开发出工程知识管理系统实例,并在航天企业中形成了一定规模的应用。     

Structure design and mechanical measurement of inflatable antenna

This paper deals with the initial shape analysis, cutting-pattern analysis, surface accuracy measurement and modal testing of high-precision inflatable antennas reflectors that are intended for spaceflight applications. The initial shape analysis method, formulated on the basis of membrane theory and elastic mechanics, determines the required as-manufactured shape of the reflective surface of the antenna reflector. On the other hand, the cutting-pattern analysis method, with its formulation based on spring-mass representations, numerically calculates the required cutting pattern of the planar membrane gores that are to be assembled to form the 3-dimensional reflective surface. To validate the effectiveness of the proposed analysis methods, a 3.2-m antenna reflector model was designed, manufactured, and assembled for ground demonstration and testing. The reflective surface accuracy of this demonstration reflector model was measured by a photogrammetric measuring system. Shape adjustments of the reflective surface were performed by systematically adjusting the tension in the cables that were used to mount the reflector to its support structure. It was found that the reflective surface accuracy of the reflector model, as defined by its RMS error from a best-fit parabolic shape, was less than 1 mm. In addition, dynamic and RF tests were also performed on the demonstration reflector model. The test results indicated that the first-mode frequency of the reflector model agreed well with the corresponding analytical prediction, and its radiation pattern was also well focused.     

End-to-end validation process for the INTA-Nanosat-1B Attitude Control System

This paper describes the end-to-end validation process for the Attitude Control Subsystem (ACS) of the satellite INTA-NanoSat-1B (NS-1B). This satellite was launched on July 2009 and it has been fully operative since then. The development of its ACS modules required an exhaustive integration and a system-level validation program. Some of the tests were centred on the validation of the drivers of sensors and actuators and were carried out over the flying model of the satellite. Others, more complex, constituted end-to-end tests where the concurrency of modules, the real-time control requirements and even the well-formedness of the telemetry data were verified. This work presents an incremental and highly automatised way for performing the ACS validation program based on two development suites and an end-to-end validation environment. The validation environment combines a Flat Satellite (FlatSat) configuration and a real-time emulator working in closed-loop. The FlatSat is built using the NS-1B Qualification Model (QM) hardware and it can run a complete version of the on-board software with the ACS modules fully integrated. The real-time emulator, running on an industrial PC, samples the actuation signals and emulates the sensors signals to close the control loop with the FlatSat. This validation environment constitutes a low-cost alternative to the classical three axes tilt table, with the advantage of being easily configured for working under specific orbit conditions, in accordance with any of the selected tests. The approach has been successfully applied to the NS-1B in order to verify different ACS modes under multiple orbit scenarios, providing an exhaustive coverage and reducing the risk of eventual errors during the satellite's lifetime. The strategy was applied also during the validation of the maintenance and reconfiguration procedures required once the satellite was launched. This paper describes in detail the complete ACS validation process that was performed and it shows the most relevant errors detected and fixed during testing. Finally it summarises some of the most significant conclusions.     

Optimal reconfiguration maneuvers for spacecraft imaging arrays in multi-body regimes

Upcoming National Aeronautics and Space Administration (NASA) mission concepts include satellite arrays to facilitate imaging and identification of distant planets. These mission scenarios are diverse, including designs such as the terrestrial planet finder-occulter (TPF-O), where a monolithic telescope is aided by a single occulter spacecraft, and the micro-arcsecond X-ray imaging mission (MAXIM), where as many as 16 spacecraft move together to form a space interferometer. Each design, however, requires precise reconfiguration and star tracking in potentially complex dynamic regimes. This paper explores control methods for satellite imaging array reconfiguration in multi-body systems. Specifically, optimal nonlinear control and geometric control methods are derived and compared to the more traditional linear quadratic regulators, as well as input state feedback linearization. These control strategies are implemented and evaluated for the TPF-O mission concept.     

A new real-time guidance strategy for aerodynamic ascent flight

Reusable launch vehicles are conceived to constitute the future space transportation system. If these vehicles use air-breathing propulsion and lift taking-off horizontally, the optimal steering for these vehicles exhibits completely different behavior from that in conventional rockets flight. In this paper, the new guidance strategy is proposed. This method derives from the optimality condition as for steering and an analysis concludes that the steering function takes the form comprised of Linear and Logarithmic terms, which include only four parameters. The parameter optimization of this method shows the acquired terminal horizontal velocity is almost same with that obtained by the direct numerical optimization. This supports the parameterized Liner Logarithmic steering law. And here is shown that there exists a simple linear relation between the terminal states and the parameters to be corrected. The relation easily makes the parameters determined to satisfy the terminal boundary conditions in real-time. The paper presents the guidance results for the practical application cases. The results show the guidance is well performed and satisfies the terminal boundary conditions specified. The strategy built and presented here does guarantee the robust solution in real-time excluding any optimization process, and it is found quite practical.     

Response of a pentagonal PZT element as a component of a 4π-real-time detector

The detector characteristics of a pentagonal element were studied by colliding it with hypervelocity micro-particles. A charge-sensitive amplifier was developed for the element of its capacitance ∼10 nF. The output amplitudes were expressed as a linear function of the momentum at collision. Empirical formulas obtained from on-ground experiments could be used for the calibration of the detector. The pentagonal element was potential to measure the momentum during collision from the output amplitude. A set of electrodes on the surface was used to confirm the measurement of the coordinates at collision. A possible application of this pentagonal element on a real-time dust detector was discussed.     

多孔材料在整流罩内中高频降噪的应用与优化研究

基于JCA模型，采用有限元仿真和声学试验获取了三聚氰胺泡沫和吸音棉的声学参数，采用统计能量分析（SEA）方法建立了某卫星-整流罩的系统级模型。通过与混响室试验结果进行对比，验证了模型的正确性；并通过仿真和试验获取了两种材料对整流罩内噪声环境的降噪效果，分析了厚度、敷设率和敷设位置对降噪效果的影响。结果表明，采用结合Biot理论的SEA方法能有效预测整流罩内中高频噪声环境及多孔材料降噪效果；总降噪量与厚度、敷设率和敷设位置关系紧密，采用多孔材料降噪时应综合考虑以上因素。     

火星动力下降自主导航与制导技术研究进展

基于火星定点软着陆探测任务的需求，阐述了火星着陆动力下降导航与制导技术的必要性。首先分析了火星着陆动力下降段导航与制导问题，并总结了动力下降段导航与制导面临的挑战与难点，然后综述了动力下降段导航与制导技术的研究现状。最后针对未来复杂地形区精确着陆，提出了实现火星动力下降段高精度自主导航与制导需要解决的关键问题。     

从盖茨比中看菲茨杰拉德

弗兰西斯·司各特·菲茨杰拉德是现代美国著名作家,被称作爵士时代的代言人.他的多数作品描绘的是二十世纪二十年代美国年轻一代的期望、不满和失望.1925年出版的<了不起的盖茨比>是菲茨杰拉德写作生涯的顶点.书中主人公盖茨比从多个角度反映了作者的生活和精神历程,从而被认为是作者的缩影.     

Numerical studies of a non-linear aeroelastic system with plunging and pitching freeplays in supersonic/hypersonic regimes

The flutter and post flutter of a two-dimensional double-wedge lifting surface with combined freeplay and cubic stiffness nonlinearities in both plunging and pitching degrees-of-freedom operating in supersonic/hypersonic flight speed regimes have been analyzed. In addition to the structural nonlinearities, the third-order piston theory aerodynamics is used to evaluate the unsteady non-linear aerodynamic force and moment. Such model accounts for stiffness and damping contributions produced by the aerodynamic loads. Responses involving limit cycle oscillation and chaotic motion are observed over a large number of parameters that characterizes the aeroelastic system. Results of the present study show that the freeplay in the pitching degree-of-freedom and soft/hard cubic stiffness in the pitching and plunging degrees-of-freedom have significant effects on the LCOs exhibited by the aeroelastic system in the supersonic/hypersonic flight speed regimes. The simulations also show that the aeroelastic system behavior is greatly affected by physical structural parameters, such as the radius of gyration and the frequency ratio, especially in post-flutter regimes, when accounting for all system nonlinearities. It has been shown that at high Mach numbers the non-linear aerodynamic stiffness yields detrimental effects from the aeroelastic point of view, while the damping one do not.