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101.
102.
A parameterized geometry design method for inward turning inlet compatible waverider 总被引:1,自引:1,他引:1
Intensive studies have been carried out on generations of waverider geometry and hypersonic inlet geometry. However, integration efforts of waverider and related air-intake system are restricted majorly around the X43A-like or conical flow field induced configuration, which adopts mainly the two-dimensional air-breathing technology and limits the judicious visions of developing new aerodynamic profiles for hypersonic designers. A novel design approach for integrating the inward turning inlet with the traditional parameterized waverider is proposed. The proposed method is an alternative means to produce a compatible configuration by linking the off-the-shelf results on both traditional waverider techniques and inward turning inlet techniques. A series of geometry generations and optimization solutions is proposed to enhance the lift-to-drag ratio. A quantitative but efficient aerodynamic performance evaluation approach (the hypersonic flow panel method) with lower computational cost is employed to play the role of objective function for opti- mization purpose. The produced geometry compatibility with a computational fluid dynamics (CFD) solver is also verified for detailed flow field investigation. Optimization results and other numerical validations are obtained for the feasibility demonstration of the proposed method. 相似文献
103.
为验证一种双楔顶压、侧板中置的侧压式进气道基本性能,设计了一套进口面积为110mm×91mm的双流道试验模型,并在300mm马赫数6的高焓脉冲风洞中进行了吹风实验。实验测量了进气道和隔离段内的沿程静压分布和隔离段进出口截面的皮托压力分布,分析了进气道内的典型流场特征,获得了进气道的基本性能参数,并以马赫数的测量为例阐述了流场不均匀性对测量结果可能造成的影响。实验结果表明,马赫数6来流条件下,该侧压式进气道流量系数为0.83,隔离段出口平均马赫数为2.57,总压恢复系数为0.296,增压比为23.7,表明这种侧压式进气道的气动布局方式能够获得较好的总体性能。 相似文献
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为了使高超声速冲压发动机在宽飞行条件下同时具有高比冲、高推力系数、高推重比,在讨论多模态冲压发动机的不同工作模态特性基础上,提出了改进进气道/燃烧室/尾喷管参数协调状态的技术途径。在固定几何的条件下,采用一体化设计内流通道,并巧妙地调节加热规律,使得在不同飞行条件下采用不同的优化工作模态,从而防止进气道出现亚 声速溢流或过度超临界,防止尾喷管产生膨胀过度或不足,防止燃烧室内的过度高温高压,并使冲量增量最大。此外,就国内外在研制过程中曾出现过经验教训及应引起关注的技术创新点进行了讨论。 相似文献
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在前期开展再入飞行器RCS(雷达散射截面)特性试验和理论研究的基础上,对典型临近空间高超声速飞行器RCS特性开展了研究,分析了绕流和尾迹对飞行器本体RCS特性的影响。研究表明等离子体流场在头身部绕流、近尾尾迹和部分远尾尾迹的最大电子密度将远高于电离层最高电子密度,更高于典型天波超视距雷达工作频段对应的临界电子密度。因而等离子体尾迹将会对3~30 MHz频段电磁波产生较强的散射,使得等离子体尾迹的RCS远远大于飞行器本体的RCS。利用临近空间高超声速飞行器尾迹RCS的这一特点,有可能实现对临近空间高超声速飞行器的超视距探测。 相似文献
110.
Experimental investigation on aero-heating of rudder shaft within laminar/turbulent hypersonic boundary layers 总被引:1,自引:0,他引:1
The aero-heating of the rudder shaft region of a hypersonic vehicle is very harsh, as the peak heat flux in this region can be even higher than that at the stagnation point. Therefore, studying the aero-heating of the rudder shaft is of great significance for designing the thermal protection system of the hypersonic vehicle. In the wind tunnel test of the aero-heating effect, we find that with the increase of the angle of attack of the lifting body model, the increasement of the heat flux of the rudder shaft is larger under laminar flow conditions than that under turbulent flow conditions. To understand this, we design a wind tunnel experiment to study the effect of laminar/turbulent hypersonic boundary layers on the heat flux of the rudder shaft under the same wind tunnel freestream conditions. The experiment is carried out in the ?2 m shock tunnel(FD-14 A) affiliated to the China Aerodynamics Research and Development Center(CARDC). The laminar boundary layer on the model is triggered to a turbulent one by using vortex generators, which are 2 mm-high diamonds. The aero-heating of the rudder shaft(with the rudder) and the protuberance(without the rudder) are studied in both hypersonic laminar and turbulent boundary layers under the same freestream condition. The nominal Mach numbers are 10 and 12, and the unit Reynolds numbers are2.4 × 10~6 m~(-1) and 2.1 × 10~6 m-1. The angle of attack of the model is 20°, and the deflection angle of the rudder and the protuberance is 10°. The heat flux on the model surface is measured by thin film heat flux sensors, and the heat flux distribution along the center line of the lifting body model suggests that forced transition is achieved in the upstream of the rudder. The test results of the rudder shaft and the protuberance show that the heat flux of the rudder shaft is lower in the turbulent flow than that in the laminar flow, but the heat flux of the protuberance is the other way around,i.e., lower in the laminar flow than in the turbulent flow. The wind tunnel test results is also validated by numerical simulations. Our analysis suggests that this phenomenon is due to the difference of boundary layer velocities caused by different thickness of boundary layer between laminar and turbulent flows, as well as the restricted flow within the rudder gap. When the turbulent boundary layer is more than three times thicker than that of the laminar boundary layer, the heat flux of the rudder shaft under the laminar flow condition is higher than that under the turbulent flow condition. Discovery of this phenomenon has great importance for guiding the design of the thermal protection system for the rudder shaft of hypersonic vehicles. 相似文献