脉冲设备二维高超进气道压缩面激波—边界层干扰显示技术及应用研究
发布时间：2018-08-05 09:23
【摘要】：在吸气式高超声速飞行器研究中,激波-边界层干扰问题是一个研究多年但远未解决的难题。其中,高超进气道外压缩面上的激波-边界层干扰通过影响波系结构、’边界层状态而对进气道性能产生重要影响。然而,在众多激波-边界层干扰研究中,关注高超进气道外压缩面上情况较少,特别是多级压缩拐角和前缘半径对激波-边界层干扰特性及进气道性能的影响规律,尚缺乏系统的研究。在激波边界层干扰(包括进气道外压缩面)研究中,既要研究波系分布,也要关注边界层行为.。脉冲型高超声速试验设备因能够模拟高超声速总温和总压试验条件,是常规风洞的重要补充,’两者结合,可覆盖更全面的马赫数、雷诺数和壁温比范围,也是开展激波边界层干扰研究的重要试验平台。由于试验时间极短,脉冲型高超声速试验设备上尚缺少成熟的、可显示边界层分离情况的油流试验技术；同时由于高超声速流场中的高温和高速测试环境,常规的边界层厚度测试技术如皮托探针无法使用,测试高超声速边界层厚度存在巨大困难。针对上述问题,本文在脉冲型高超声速风洞上,发展了两类高超进气道激波边界层干扰研究测试技术。首先,在激波风洞上发展了基于聚焦纹影技术的高超声速边界层厚度测量方法,结合理论和数值模拟结果,提供了从聚焦纹影照片中提取密度边界层厚度数据的方法,证明了基于密度梯度的边界层厚度近似密度边界层厚度,密度边界层与温度边界层厚度相同,.对于空气速度边界层厚度略小于密度边界层厚度,在应用中改进了聚焦纹影系统光源和成像系统。其次,在脉冲燃烧风洞上发展了油流显示技术,结合脉冲燃烧设备运行特点,剖析了脉冲类设备上发展油流显示技术的难点所在；分析了油膜受力情况和运动机理,给出了油膜运动速度的影响因素,提出了瞬时脉冲型风洞上膜式油流试验准则；提供了合适的显示油粘度范围‘,发展了油膜喷涂方法,解决了高温试验气体自发光问题,设计了实时摄像系统；在极短的有效试验时间内,采集了油流图谱序列,油膜的动态发展过程验证了上述关系准则作为边界层厚度测量方法和油流显示技术的应用,结合其它测试和数值模拟手段,初步研究了多级压缩拐角和前缘半径对进气道性能和激波边界层干扰特性的影响规律。在脉冲燃烧风洞上名义马赫数6和单位雷诺数5.4×106/m条件下,得到的油流图谱序列表明,两级压缩楔拐角处初始分离角度介于100～150之间,三级压缩楔模型第二拐角处流动未分离。在激波风洞上名义马赫数6和单位雷诺数3.4×107/m条件下,研究了三级平面压缩高超声速进气道前缘半径的影响,获得了壁面静压、壁面热流、激波几何特征、边界层密度厚度分布数据。壁面静压从拐点上游开始爬升,在拐点后一定距离达到压强平台；前缘半径增加,压强平台值减小,达到压强平台需要的距离增加,’意味着拐角处激波-边界层干扰区范围扩大,尖前缘情况下激波-边界层干扰区范围是上游边界层厚度的5倍,3mm钝前缘情况下达到8倍。壁面热流从拐点上游开始逐渐爬升,在拐点后达到一个峰值,然后下降直到下一个压缩拐角；随着前缘半径的增加,拐角热流峰值显著减小,热流峰值位置向上游移动。边界层厚度分布在拐角上游达到上游最大值,过干扰区后边界层厚度比入流边界层厚度显著减小,在下一个压缩面上重新增长；随着前缘半径的增加,边界层厚度增加。使用FLUNET软件选用不同湍流模型得到的壁面静压都与试验结果吻合较好,但壁面热流分布、边界层密度厚度分布差异较大。随着前缘半径的增加,第一级激波角逐渐增加,第二级和第三级激波角逐渐减小,拐角处第二、三道激波根部弯曲程度和影响区域增大。
[Abstract]:In the study of air breathing hypersonic vehicles, the shock boundary layer interference problem is a difficult problem for many years, but the shock boundary layer interference on the hyperactive outer compression surface has an important influence on the performance of the intake port by influencing the structure of the wave system and the state of the boundary layer. However, in a number of shock wave boundary layers. In the perturbation study, it is not necessary to study the influence of the high intake air compression surface on the outer compression surface, especially the influence of the multistage compression corner and the radius of the front edge to the shock boundary layer interference and the inlet performance. Boundary layer behavior, pulse type hypersonic test equipment is an important supplement to conventional wind tunnel because of its ability to simulate high hypersonic total pressure test conditions. The combination of the two can cover a more comprehensive Maher number, Reynolds number and the range of wall temperature ratio. It is also an important test platform for the study of shock wave boundary layer interference. An oil flow test technique which can display boundary layer separation conditions is still lacking on the flush hypersonic test equipment. At the same time, due to the high temperature and high speed testing environment in the hypersonic flow field, the conventional boundary layer thickness testing technology, such as the pitot probe, is difficult to test the thickness of the high supersonic boundary layer. In this paper, we develop two kinds of high supersonic wave boundary layer interference in high supersonic wind tunnel. Firstly, the method of measuring the thickness of hypersonic boundary layer based on the focused schlieren technique is developed on the shock wind tunnel, and the density is extracted from the focused schlieren photos with the theoretical and numerical simulation results. The method of boundary layer thickness data shows that the thickness of the boundary layer based on the density gradient is approximately the thickness of the density boundary layer, the density boundary layer is the same as the thickness of the temperature boundary layer. The thickness of the air velocity boundary layer is slightly less than the thickness of the density boundary layer, and the light source and imaging system of the focused schlieren system are improved in application. Secondly, the pulse combustion is used. The oil flow display technology is developed in the wind tunnel. The difficulties in the development of oil flow display technology on pulse equipment are analyzed in combination with the characteristics of the pulse combustion equipment. The stress situation and movement mechanism of the oil film are analyzed. The influence factors of the motion velocity of the oil film are given, and the test criterion of the film type oil flow on the instantaneous pulse blast tunnel is put forward. The oil film spraying method is properly displayed and the oil film spraying method is developed. The self luminescence problem of the high temperature test gas is solved. The real-time camera system is designed. The oil flow map sequence is collected during the very short effective test time. The dynamic development process of the oil film proves that the above relationship criterion is used as the boundary layer thickness measurement method and the oil flow display. The effect of the multistage compression corner and the radius of the front edge on the performance of the intake port and the interference characteristics of the shock wave boundary layer is preliminarily studied with the use of other testing and numerical simulation methods. Under the condition of the nominal Maher number 6 and the unit Reynolds number of 5.4 x 106/m on the pulse combustion wind tunnel, the obtained oil flow map sequence shows that the two stage compression wedge abduction is shown. The initial separation angle at the corner is between 100~150, and the flow of the three stage compression wedge is not separated at second corners. Under the condition of the nominal Maher number 6 and the unit Reynolds number 3.4 x 107/m on the shock wind tunnel, the influence of the three level plane compression on the radius of the front edge of the high supersonic inlet is studied, and the wall static pressure, the wall heat flow, the shock geometry feature and the edge are obtained. The boundary layer density thickness distribution data. The wall static pressure begins to climb up the inflection point and reaches the pressure platform at a certain distance after the inflection point; the radius of the front edge increases and the pressure platform value decreases, which can increase the distance needed by the pressure platform, 'means that the disturbance area of the shock boundary layer is enlarged at the corner and the shock boundary layer in the front edge is the interference area. The range is 5 times the thickness of the upstream boundary layer, and the 3mm blunt front edge reaches 8 times. The wall heat flow gradually rises from the upstream of the inflection point, reaches a peak after the turning point, and then drops to the next compression corner. With the increase of the radius of the front edge, the peak heat flow peak decreases significantly, the peak position of the heat flow moves upstream. Boundary layer thickness. The thickness of the upper boundary layer is significantly lower than the inflow boundary layer thickness, and the thickness of the boundary layer is increased again. With the increase of the radius of the front edge, the thickness of the boundary layer increases with the increase of the radius of the front edge. The wall static pressure obtained by using different turbulence models using FLUNET software is in good agreement with the test results, but the wall wall is in good agreement with the test results. The distribution of surface heat flow and the thickness distribution of the boundary layer density vary greatly. With the increase of the radius of the front edge, the first shock angle increases gradually, the second and third magnitude shock angles gradually decrease, and the bending degree and the affected area of the second, third shock waves at the corner are increased.
【学位授予单位】：中国空气动力研究与发展中心
【学位级别】：博士
【学位授予年份】：2015
【分类号】：V211.74
本文编号：2165354
[Abstract]:In the study of air breathing hypersonic vehicles, the shock boundary layer interference problem is a difficult problem for many years, but the shock boundary layer interference on the hyperactive outer compression surface has an important influence on the performance of the intake port by influencing the structure of the wave system and the state of the boundary layer. However, in a number of shock wave boundary layers. In the perturbation study, it is not necessary to study the influence of the high intake air compression surface on the outer compression surface, especially the influence of the multistage compression corner and the radius of the front edge to the shock boundary layer interference and the inlet performance. Boundary layer behavior, pulse type hypersonic test equipment is an important supplement to conventional wind tunnel because of its ability to simulate high hypersonic total pressure test conditions. The combination of the two can cover a more comprehensive Maher number, Reynolds number and the range of wall temperature ratio. It is also an important test platform for the study of shock wave boundary layer interference. An oil flow test technique which can display boundary layer separation conditions is still lacking on the flush hypersonic test equipment. At the same time, due to the high temperature and high speed testing environment in the hypersonic flow field, the conventional boundary layer thickness testing technology, such as the pitot probe, is difficult to test the thickness of the high supersonic boundary layer. In this paper, we develop two kinds of high supersonic wave boundary layer interference in high supersonic wind tunnel. Firstly, the method of measuring the thickness of hypersonic boundary layer based on the focused schlieren technique is developed on the shock wind tunnel, and the density is extracted from the focused schlieren photos with the theoretical and numerical simulation results. The method of boundary layer thickness data shows that the thickness of the boundary layer based on the density gradient is approximately the thickness of the density boundary layer, the density boundary layer is the same as the thickness of the temperature boundary layer. The thickness of the air velocity boundary layer is slightly less than the thickness of the density boundary layer, and the light source and imaging system of the focused schlieren system are improved in application. Secondly, the pulse combustion is used. The oil flow display technology is developed in the wind tunnel. The difficulties in the development of oil flow display technology on pulse equipment are analyzed in combination with the characteristics of the pulse combustion equipment. The stress situation and movement mechanism of the oil film are analyzed. The influence factors of the motion velocity of the oil film are given, and the test criterion of the film type oil flow on the instantaneous pulse blast tunnel is put forward. The oil film spraying method is properly displayed and the oil film spraying method is developed. The self luminescence problem of the high temperature test gas is solved. The real-time camera system is designed. The oil flow map sequence is collected during the very short effective test time. The dynamic development process of the oil film proves that the above relationship criterion is used as the boundary layer thickness measurement method and the oil flow display. The effect of the multistage compression corner and the radius of the front edge on the performance of the intake port and the interference characteristics of the shock wave boundary layer is preliminarily studied with the use of other testing and numerical simulation methods. Under the condition of the nominal Maher number 6 and the unit Reynolds number of 5.4 x 106/m on the pulse combustion wind tunnel, the obtained oil flow map sequence shows that the two stage compression wedge abduction is shown. The initial separation angle at the corner is between 100~150, and the flow of the three stage compression wedge is not separated at second corners. Under the condition of the nominal Maher number 6 and the unit Reynolds number 3.4 x 107/m on the shock wind tunnel, the influence of the three level plane compression on the radius of the front edge of the high supersonic inlet is studied, and the wall static pressure, the wall heat flow, the shock geometry feature and the edge are obtained. The boundary layer density thickness distribution data. The wall static pressure begins to climb up the inflection point and reaches the pressure platform at a certain distance after the inflection point; the radius of the front edge increases and the pressure platform value decreases, which can increase the distance needed by the pressure platform, 'means that the disturbance area of the shock boundary layer is enlarged at the corner and the shock boundary layer in the front edge is the interference area. The range is 5 times the thickness of the upstream boundary layer, and the 3mm blunt front edge reaches 8 times. The wall heat flow gradually rises from the upstream of the inflection point, reaches a peak after the turning point, and then drops to the next compression corner. With the increase of the radius of the front edge, the peak heat flow peak decreases significantly, the peak position of the heat flow moves upstream. Boundary layer thickness. The thickness of the upper boundary layer is significantly lower than the inflow boundary layer thickness, and the thickness of the boundary layer is increased again. With the increase of the radius of the front edge, the thickness of the boundary layer increases with the increase of the radius of the front edge. The wall static pressure obtained by using different turbulence models using FLUNET software is in good agreement with the test results, but the wall wall is in good agreement with the test results. The distribution of surface heat flow and the thickness distribution of the boundary layer density vary greatly. With the increase of the radius of the front edge, the first shock angle increases gradually, the second and third magnitude shock angles gradually decrease, and the bending degree and the affected area of the second, third shock waves at the corner are increased.
【学位授予单位】：中国空气动力研究与发展中心
【学位级别】：博士
【学位授予年份】：2015
【分类号】：V211.74
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