共查询到19条相似文献,搜索用时 281 毫秒
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超声速高超声速风洞测力数据衔接性的研究 总被引:1,自引:0,他引:1
本文叙述半锥角θc=1 0°尖锥模型和HB 2标模在FD 0 7风洞的气动力测量结果。通过与FD 0 6跨超声速风洞及国内外其它风洞实验数据的比较 ,讨论从超声速到高超声速的不同风洞设备中气动力测量数据随 Ma数的衔接变化 ,分析新建的FD 0 7风洞气动力测量的准确性和可靠性。 相似文献
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吸气式高超声速飞行器内外流同时测力试验 总被引:2,自引:0,他引:2
为研究吸气式高超声速飞行器新型测力试验技术的可行性,在马赫数Ma=6、高度H=27km模拟飞行环境下开展了风洞试验。以轴对称吸气式高超声速飞行器无舵翼简化构型为研究对象,通过合理的试验系统设计解决内外流解耦关键技术,采用两个环式六分量天平在同一车次上测量模型内外流气动力载荷。试验结果表明:内外流气动力天平测量数据和全模气动力数据准确反映了内外流解耦、内外流窜流、进气道起动/不起动以及溢流影响等物理现象;从物理上实现了内外流解耦;内流气动力对全机气动性能贡献大;内外流解耦、进气道起动时,气动力数据准确、重复性好;窜流产生的传力属于内力。试验证明,同时测力试验技术可行,为解决内外流气动数据源于不同试验所致的不确定性问题提供了有效手段。 相似文献
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介绍了带环量控制尾梁的直升机模型在风洞中和旋翼下的测压实验。根据测量数据计算出尾梁上气动力随环量控制参数的变化情况,并绘制成曲线。主要研究动量系数和缝隙几何参数对圆柱尾梁上气动力的影响。 相似文献
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为了解决小不对称再入体滚动气动力测量问题,北京空气动力研究所研制开发了以空气轴承为核心的滚转气动力测量技术,利用空气了轴承自身旋转阻尼非常小的特点,使模型做自由滚转运动,一个特殊设计的非接触的光学测量系统测出模型的转角随时间的历程,用参数拟合的方法得到滚转力矩的大小和方向。为验证该项实验技术的正确性与可靠性,在φ500高超声速风洞中对4个模型进行了吹风实验,吹风马赫数为5,测量滚转力矩系数C10和滚转阻力矩系数C1p。实验结果表明,该文方法数据合理,并较其它方法更具有鲁棒性。 相似文献
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采用结冰风洞实验方法,在0.3m×0.2m结冰风洞第二实验段对圆柱模型的结冰/除冰过程进行了气动参数测量实验。建立了电加热圆柱模型和适合低温高湿环境的五分量外式微量天平,获得了结冰气象环境下圆柱模型结冰/除冰过程的气动力/力矩随时间的变化规律。喷雾对载荷和动压的影响可以忽略,单位时间内模型受到喷雾的最大水平力、最大动压增量分别为0.6%和0.2%。基于结冰风洞低温高湿环境的测力实验技术可以捕捉结冰/除冰过程的气动力/力矩变化。结冰过程中,圆柱模型阻力系数随时间不断增大,呈现出近似线性增长趋势,而升力系数、俯仰力矩系数、偏航力矩系数、滚转力矩系数的变化可忽略不计。除冰过程中,前缘冰壳滑动改变了姿态,会造成阻力系数、偏航力矩系数、滚转力矩系数等迅速变化,其对气动性能的影响难以预测。 相似文献
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旋翼翼型低Ma数动态失速特性计算 总被引:3,自引:1,他引:2
低Ma数下,翼型前缘涡强度的增加和移动使法向力系数产生很大的超调量,Beddoes通过增加一项与延迟后后缘分离点有关项来模拟该特性,并改进Leishman-Beddoes二维翼型动态失速模型。在此基础上,本文在非定常法向力系数中引入一阶延迟,推迟失速判断点,得到修正后模型;而后,通过计算NACA0012、OA207翼型在低Ma数下的非定常气动力,并与实验结果进行对比,验证了模型在计算翼型低Ma数下非定常气动力的准确性,并分析了折合频率、迎角平均值、振幅对计算结果的影响。 相似文献
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基于符号计算的风洞试验测量不确定度评估 总被引:1,自引:0,他引:1
基于符号计算进行风洞试验测量不确定度分析可以实现实验数据处理公式及误差灵敏度系数的自动推导,采用该方法对ZSDD-1导弹标模风洞试验结果进行了定量的试验不确定度评估,计算得到的气动力系数精度极限与重复性试验得出的试验精度吻合良好,气动力系数偏离极限计算值通常是其精度极限的3-4倍,其不确定度大约是其精度极限的4倍。笔者所述分析方法和分析程序为定量评估风洞试验数据的可靠性提供了一种有效手段。 相似文献
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跨声速压气机动叶平面叶栅实验 总被引:1,自引:0,他引:1
对某跨声速压气机动叶根部叶型平面叶栅流场在不同冲角和进口等熵马赫数下进行了详细的测量,得到了冲角特性曲线和叶片表面及端壁静压.结果表明:负冲角及零冲角时,叶栅出口总压损失系数随进口等熵马赫数增加变化不大,而在正冲角时变化较大.相同进口等熵马赫数下,负冲角和零冲角时,叶片负荷较高;正冲角时,由于气流分离严重,叶片负荷下降,叶栅出口总压损失系数升高.随冲角由负冲角向正冲角增加,气流落后角逐渐增大,叶栅出口总压损失系数先减小后增大,最小值为0.034.冲角相同时,随进口等熵马赫数增加,叶栅出口总压损失系数总体呈增加趋势. 相似文献
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An aerodynamic force and moment measurement was conducted in JF12 long-testduration detonation-driven shock tunnel of Institute of Mechanics,Chinese Academy of Sciences.The test duration of JF12 is 100–130 ms.The nominal Mach number is 7.0 and the exit diameter of the contoured nozzle is 2.5 m.The total enthalpy is 2.5 MJ/kg which duplicates the hypersonic flight conditions of Mach number 7.0 at 35 km altitude.The test model is the standard aerodynamic force model of 10° half-angle sharp cone.The length of the test model is 1500 mm and the weight is 57 kg.The aerodynamic forces were measured with a six-component strain balance.The angles of attack were set to be à5°,0°,5°,10° and 14°,respectively.The experimental results show that in the 100–130 ms test duration,the signals of strain balance have 3–4 complete vibration cycles.So,the aerodynamic forces and moments can be obtained directly by averaging the signals of balance without acceleration compensation.The force measurement error of repeatability of JF12 is less than 2%.The aerodynamic force coefficients of JF12 are in good agreement with those of conventional hypersonic wind tunnels.For this test model at Mach number 7.0 and total enthalpy of 2.5 MJ/kg,the real-gas effects on aerodynamic force characteristics are not very evident. 相似文献
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Research team on airfoil 《航空学报》1981,2(1):10-20
在西北工业大学跨音速翼型风洞(又名57风洞)中,对弦长分别为50、100、125毫米的三个RAE 104翼型模型进行了压力分布测量。结果表明:当上下壁开闭比选用2%时,风洞的堵塞干扰基本消除。 采用实侧壁、多层网板而不进行抽气及多层网板进行抽气这三种侧壁状态,对RAE 104翼型模型所做实验结果表明:有多层网板而不抽气,使升力系数大大低于无干扰值;在M相似文献
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《Aerospace Science and Technology》2001,5(3):167-178
The Types III and IV interference flows, as defined by Edney, and corresponding heat transfer distributions were investigated experimentally. The model consists of a cylindrically blunted plate and a wedge serving as an oblique shock generator. The ‘thin wall’ technique was used for heat transfer measurements on the cylinder surface. These experiments were carried out in the TsAGI short duration wind tunnel UT-1 at Mach numbers 6 and 16 in air and at Mach number 6.6 in carbon dioxide. The Reynolds number based on the plate bluntness diameter was varied in the range 2.2×104 to 1.6×106. Tests of the cylinder alone (without the wedge) at Mach number 6 and for different Reynolds numbers revealed an influence of incoming disturbances on the stagnation line heat transfer. The influence of the impinging shock location on the interference heat transfer was carefully investigated. Systematic calculations of inviscid flow at Mach number 6 were also performed. Estimations of the maximum interference heat transfer rate, based on these calculations and a boundary layer approach, compare well with the data. Influence of the specific heat ratio on the interference flow was studied. These experiments and calculations revealed important features of interference flow patterns and heat transfer distributions. 相似文献
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气动声学相似理论与试验研究 总被引:2,自引:0,他引:2
本文研究了气动声学的一般性相似原理,为气动声学的模型实验研究方法建立理论基础。由此得到了四个关于气动声场相似的相似准则数,导出了适合于流动分析的相似条件。对非定常流动中所产生的脉动力谱进行的相似分析,证实当气动力学完全相似时,以Strouhal数作为无量纲频率参数的脉动力谱函数也完全相似,此时,在几何相似点上测得的声谱在以St数作为频率坐标的频谱图上完全相同。风扇噪声的试验证实了气动声学的相似律。 相似文献
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For the purpose of establishing and validating aerodynamic performance predictions at transonic Mach numbers, a wind tunnel test was conducted in the High-Speed Tunnel (HST) of the German-Dutch Wind Tunnels. The test article is the aerodynamic validation model from the Chinese Aeronautical Establishment, which is a full-span scale model representation of a business jet aircraft. The wind tunnel test comprised of parallel deployments of balance, pressures, infrared thermography, and model marker measurement techniques. Dedicated investigations with a dummy support were conducted as well, in order to derive and correct for the interference that the support system imposed on the overall model loads. This enabled the establishment of a comprehensive dataset in which the steady overall model loads, the wing load distribution, the state of the wing boundary layer, and the aeroelastic wing shape were quantified for conditions up to and beyond the cruise Mach number of 0.85. 相似文献
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给出了前缘后掠65°、双弧形剖面的细长梯形翼背风面流动显示结果。实验Mach数为1.10,1.53,2.53,3.01和4.01,攻角范围为5°~25°。应用蒸汽屏、纹影和油流技术拍摄了空间和表面流型照片。蒸汽屏显示表明:在机翼背风面三角形区域的空间流型随法向攻角αN(在垂直于前缘的平面内流速与弦线间的夹角)和法向Mach数MaN(来流Mach数在垂直于前缘平面内的分量)变化,并可在αN和MaN为坐标的平面上划分出7种流型存在的区域。侧缘区有侧缘分离涡形成;后缘有尾涡拖出。从纹影照片与横截面上的蒸汽屏照片对照可获得机翼锥面激波位置随Mach数的变化;以及激波-诱导分离线位置随Mach数和攻角变化曲线。机翼表面油流谱显示出了主再附线、二次分离线、二次再附线和侧缘涡区。显示出的流型与其他有关实验和数值计算结果比较符合得很好 相似文献