摘要
进气道是飞行器动力装置的重要组成部分,准确测量进气道流量系数是进气道风洞试验的重要内容。对来流马赫数Ma=4.5,5.0和6.0状态下皮托管进气道开展流量系数测量研究,通过对比理论值和实测值,获取各状态流量系数修正系数。试验结果表明,随着来流马赫数增加,进气道流量系数与理论值偏差较明显,并逐渐增大。超声速风洞试验通常认为测量截面总温与来流总温相等,通过对测量截面总温与来流总温偏差以及测量截面流场畸变情况的分析,判断测量偏差主要是由测量截面总温等于来流总温的假设导致的。在高超声速风洞试验中,由于模型壁面热交换的存在,测量截面总温低于来流总温,进气道流量系数测量时需要进行总温修正,以提高流量测量精度。
Inlet is a critical component of aircraft propulsion system and accurate measurement of its flow coefficient is important in inlet wind tunnel experiments. Flow coefficient measurement investigations using Pitot tube inlet were carried out at free stream Mach number 4.5, 5.0, 6.0, respectively. Correction of flow coefficients in all kinds of cases were acquired by comparing the experimental values and theoretical values. The experimental results indicate that the difference compared with the theoretical values gradually rises up along with the increase in the incoming flow Mach number. According to the engineering experience in supersonic wind tunnel experiments, the temperature at measurement cross section is equal to the incoming flow total temperature. This hypothesis is also used in the current experiments. However, it is proved that this hypothesis is the main cause resulting in measure- ment bias by analyzing the deference between the two kinds of total temperature as mentioned above and the flow field distortion at measurement cross section. In hypersonic wind tunnel, the total temperature at measurement cross section is lower than that of the incoming flow because of the existence of heat transform on model wall surface. In order to increase the precision of the inlet flow coefficient measurement, the correction of total temperature is also required.
出处
《推进技术》
EI
CAS
CSCD
北大核心
2013年第4期470-476,共7页
Journal of Propulsion Technology
关键词
超声速进气道
高超声速风洞
试验
流量系数
Supersonic inlet
Hypersonic wind tunnel
Experiment
Flow coefficient
