摘要
声爆精确预测及低声爆设计方法已成为新一代军民用超声速飞机研制过程中必须解决的关键难题之一。基于计算流体力学(CFD)、波形参数法以及MARK-VII方法构建了高精度声爆预测方法,利用该方法对低声爆静音锥的设计展开研究。研究结果表明,静音锥的设计存在临界长度,静音锥长度小于临界长度时,静音锥产生的激波在传播过程中与机头弓形激波合并,静音锥无法起到降低声爆的作用;静音锥长度大于临界长度时,声爆水平也会略有上升。静音锥临界长度随飞行高度和飞行马赫数的变化而变化,可以根据实际飞行状态采用可伸缩设计,达到最佳的低声爆效果。多级静音锥利用多道弱激波取代机头强弓形激波,其声爆水平较单级静音锥也更低,同样,多级静音锥每一级的长度也要达到临界长度。不同静音锥头部形状产生的脱体激波形状不同,脱体距离也不同,导致阻力系数以及静音锥壁面温度有所不同,但静音锥头部形状对远场声爆信号的影响并不明显。采用静音锥的低声爆方案与原始方案比较,声爆水平得到大幅降低,阻力系数略有上升。
High fidelity sonic boom prediction and low sonic boom design method is one of the key technologies of next gen- eration environment-friendly supersonic aircraft which has a direct bearing on the feasibility of commercial operation and op- eration economics. This paper developes a high fidelity sonic boom prediction method based on computational fluid dynamic （CFD）, the wave form parameter method and the MARK-VII method. The design of a quiet spike is studied by using the high fidelity sonic boom prediction method. It is found that the critical length of the quiet spike is important. If the length of a quiet spike is shorter than its critical length, the shock wave of the spike will coalesce with the nose shock, which can lead to in- validation of the quiet spike. On the other hand, if the length of the spike is longer than the critical length, the sonic boom will increase too. The critical length of a quiet spike varies with the flight altitude and flight Mach number; therefore the length of a quiet spike changes with flight condition. Multi-order quiet spikes can produce multi-weak shocks to replace the strong nose shock, which can suppress the sonic boom level more effectively. At the same time, the length of each order of spike should be equal to the critical length. Different nose shapes of a quiet spike can produce different detached shocks and shock detachment distances, which can lead to different drag coefficients and temperatures. But the influence of nose shape on far field sonic boom is not obvious. The sonic boom of the quiet spike layout is lower than the original layout, but the drag coefficient is slightly increased.
出处
《航空学报》
EI
CAS
CSCD
北大核心
2013年第5期1009-1017,共9页
Acta Aeronautica et Astronautica Sinica
基金
西北工业大学博士论文创新基金(CX201232)~~
关键词
超声速飞机
激波
计算流体力学
声波传播
气动噪声
supersonic aircraft
shock wave
computational fluid dynamics
acoustic wave propagation
aerodynamic noise
