发动机喷管羽流对近场声爆特性影响的风洞试验技术  

Wind tunnel test techniques for exhaust nozzle plume effects on near-field sonic boom

作　　者：刘中臣 钱战森[1,2,3] 李雪飞 冷岩 郭大鹏 LIU Zhongchen;QIAN Zhansen;LI Xuefei;LENG Yan;GUO Dapeng(AVIC Aerodynamics Research Institute,Shenyang 110034,China;Key Laboratory of Hypersonic Aerodynamic Force and Heat Technology,AVIC Aerodynamics Research Institute,Shenyang 110034,China;Aviation Key Laboratory of Science and Technology on High Speed and High Reynolds Number Aerodynamic Force Research,AVIC Aerodynamics Research Institute,Shenyang 110034,China)

机构地区：[1]中国航空工业空气动力研究院,沈阳110034 [2]中国航空工业空气动力研究院高超声速气动力/热技术重点实验室,沈阳110034 [3]中国航空工业空气动力研究院高速高雷诺数气动力航空科技重点实验室,沈阳110034

出　　处：《航空学报》2023年第2期136-149,共14页Acta Aeronautica et Astronautica Sinica

基　　金：国家自然科学基金(11672280)。

摘　　要：超声速飞行所引发的声爆问题是困扰新一代环保型超声速客机发展的关键技术难题,发动机喷管羽流对全机声爆特性尤其是后激波特性具有重要影响。设计了单喷管喷流试验模型及声爆试验装置,评估了风洞试验段洞壁反射激波对模型近场压力测量的影响,重点针对通气支臂对喷管羽流的支撑干扰问题进行了分析与优化。基于中国航空工业空气动力研究院FL-60风洞,开展了发动机喷管羽流对旋成体单喷管模型近场声爆特性影响试验技术研究,试验来流马赫数2.0、落压比(NPR)范围1~20.39。研究结果表明,通过对来流马赫数、通气支臂外形、喷流模型长度、通气支臂与模型的相对位置等参数的综合优化,消除了通气支臂带来的支撑干扰对喷管羽流的影响,确保在风洞试验段受限空间内模型近场压力测量不受洞壁反射和通气支臂波系的影响;喷管羽流主要对模型尾部的近场压力特征产生影响,在来流马赫数一定的条件下,提高喷管NPR使喷流状态从过膨胀到欠膨胀,喷管唇口激波逐渐增强、位置逐渐向上游移动,抑制了喷管船尾膨胀波的发展。The sonic boom caused by supersonic flight is a key technical problem hindering the development of a new generation of environmentally friendly supersonic airliners.The exhaust nozzle plume has an important effect on the sonic boom characteristics of the whole aircraft,particularly the rear shock characteristics.We design the single nozzle test model with jet and sonic boom test setup,and evaluate the effects of the reflected shock wave from the tunnel wall on the near-field pressure measurement of the model in the wind tunnel test section.The ventilation strut interference on the nozzle plume is analyzed and optimized.Based on the FL-60 wind tunnel of AVIC Aerodynamics Research Institute,the test techniques for the exhaust nozzle plume effects on the near-field sonic boom characteristics of the axisymmetric nozzle model are researched.The test free stream Mach number is 2.0,and the Nozzle Pressure Ratio(NPR)ranges from 1 to 20.39.The test results demonstrate that the comprehensive optimization of the free stream Mach number,the shape of the ventilation strut,the length of the axisymmetric nozzle model,and the relative position between the ventilation strut and the model eliminates the effect of the ventilation strut interference on the nozzle plume,ensuring that the near-field pressure measurement of the model in the confined space of the wind tunnel test section is not affected by the reflected shock wave from the tunnel wall and the ventilation strut.The nozzle plume mainly affects the near-field pressure signatures at the tail of the model.When the free stream Mach number is constant,the nozzle lip shock wave gradually increases and moves to the upstream with the NPR increasing and the nozzle exhaust flow changing from over-expansion to under-expansion,thereby effectively suppressing the development of the nozzle boat-tail expansion wave.

关 键 词：声爆 喷管羽流 落压比 支撑干扰 风洞试验 

分 类 号：V211.7[航空宇航科学与技术—航空宇航推进理论与工程]