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971.
推导出制导火箭弹为获得期望着角所需的脉冲发动机数量解析公式,并根据此公式对比研究了火箭弹分别采用同时点火和依次点火时俯仰角初始值、期望着角和脉冲发动机推力对脉冲发动机数量的影响。研究表明,同时点火方式优于依次点火方式;飞行末段姿态调整时间较长时,期望着角和初始俯仰角对发动机数量的影响很小;随着脉冲推力的增加,所需的脉冲发动机数量递减。 相似文献
972.
传递对准姿态匹配的优化算法 总被引:1,自引:0,他引:1
推导了4种传递对准姿态匹配算法,分析了这4种姿态匹配算法的优缺点,证明并验证了其中的“最优姿态匹配法”在姿态匹配算法中的最优性。首先介绍了传统的“姿态角匹配法”及其改进算法,即“姿态矩阵匹配法”,接着引入了量测失准角的概念,经过理论推导,提出了利用量测失准角进行传递对准姿态匹配的“量测失准角匹配法”。上述3种姿态匹配算法都是在子惯导安装角是小量的条件下推导而获得的,只能适用于安装角是小量的条件,具有一定的局限性。基于此,对“量测失准角匹配法”进行了完善,推导出了一种可在多挂点下使用的现代姿态匹配算法——最优姿态匹配法。从理论上证明了4种姿态匹配算法的相互关系。最后,采用“速度+姿态”匹配方案进行的传递对准仿真结果表明:4种姿态匹配算法具有相同的估计精度;推导的“最优姿态匹配法”在保证精度的同时,可应用于子惯导安装角是任意角度的情况,具有更广的应用范围。 相似文献
973.
974.
Nanosatellite constellation deployment using on-board magnetic torquer interaction with space plasma
Ji Hyun Park Shinji Matsuzawa Takaya Inamori In-Seuck Jeung 《Advances in Space Research (includes Cospar's Information Bulletin, Space Research Today)》2018,61(8):2010-2021
One of the advantages that drive nanosatellite development is the potential of multi-point observation through constellation operation. However, constellation deployment of nanosatellites has been a challenge, as thruster operations for orbit maneuver were limited due to mass, volume, and power. Recently, a de-orbiting mechanism using magnetic torquer interaction with space plasma has been introduced, so-called plasma drag. As no additional hardware nor propellant is required, plasma drag has the potential in being used as constellation deployment method. In this research, a novel constellation deployment method using plasma drag is proposed. Orbit decay rate of the satellites in a constellation is controlled using plasma drag in order to achieve a desired phase angle and phase angle rate. A simplified 1D problem is formulated for an elementary analysis of the constellation deployment time. Numerical simulations are further performed for analytical analysis assessment and sensitivity analysis. Analytical analysis and numerical simulation results both agree that the constellation deployment time is proportional to the inverse square root of magnetic moment, the square root of desired phase angle and the square root of satellite mass. CubeSats ranging from 1 to 3?U (1–3?kg nanosatellites) are examined in order to investigate the feasibility of plasma drag constellation on nanosatellite systems. The feasibility analysis results show that plasma drag constellation is feasible on CubeSats, which open up the possibility of CubeSat constellation missions. 相似文献
975.
作为实现导弹快速机动响应的关键部件,固体姿轨控发动机的性能需要通过动态多分力测试评价,但由于推力测试台结构复杂,对测试数据补偿提出了更高的要求。本文采用双模态阻尼补偿法开展固体姿轨控发动机推力补偿研究,通过脉冲激励试验获得了主要模态参数,并对脉冲激励和发动机冷流试验数据进行补偿,验证了双模态阻尼补偿方法的可行性和应用效果。结果表明:双模态阻尼补偿效果优于单模态补偿,可以有效恢复动态推力信号。所建立的双模态阻尼补偿,在姿轨控发动机研制中得到了应用。 相似文献
976.
977.
978.
979.
《Advances in Space Research (includes Cospar's Information Bulletin, Space Research Today)》2020,65(1):271-284
Space telescope ultrahigh precision pointing control requires the spacecraft platform to provide an ultra-quiet working environment. Vibration isolator rejection control and the multi-stage integrated control method is believed to be one of the best methods to improve the space telescope attitude control performance. In this paper, the fine dynamics model of multi-stage spacecraft systems is presented and the multi-stage integrated controller design techniques are provided. Effectiveness of the multi-stage integrated control approach is demonstrated by both the numerical simulation and experiment results. An integrated design and demonstrated experimental environment is developed for high-fidelity control performance assessment. The verification experiments for the space telescope attitude control and vibration control are carried out. The results show that the pointing accuracy and stability of the line-of-sight (LOS) for space telescope are improved at least one order by the multi-stage integrated control method. 相似文献
980.
This paper addresses the fixed-time adaptive model reference sliding mode control for an air-to-ground missile associated with large speed ranges, mismatched disturbances and un-modeled dynamics. Firstly, a sliding mode surface is developed by the tracking error of the state equation and the model reference state equation with respect to the air-to-ground missile. More specifically,a novel fixed-time adaptive reaching law is presented. Subsequently, the mismatched disturbances and the un-modeled dynamics are treated as the model errors of the state equation. These model errors are estimated by means of a fixed-time disturbance observer, and they are also utilized to compensate the proposed controller. Therefore, the fixed-time controller is obtained by an adaptive reaching law and a fixed-time disturbance observer. Closed-loop stability of the proposed controller is established. Finally, simulation results including Monte Carlo simulations, nonlinear six-DegreeOf-Freedom(6-DOF) simulations and different ranges are presented to demonstrate the efficacy of the proposed control scheme. 相似文献