Scramjet尾喷管几何调节方案的计算与实验研究

CFD and Experimental Investigation for an Adjustable Scramjet Nozzle

高超声速飞行器飞行接力点和巡航结束点尾喷管冷、热态俯仰力矩差较大,给飞行器的飞行姿态控制造成严重影响。为了减小喷管冷、热态俯仰力矩差,提出了在喷管上膨胀面末端增加移动板进行调节的方案,并进行了详细的三维数值模拟和相应的风洞缩比冷流实验研究。计算结果表明,Ma=4.5时,调节移动板伸出400mm,喷管冷、热态力矩差最大减小21.74%,推力系数损失1.64%;Ma=6.5时,调节移动板喷管冷、热态力矩差可降低77.59%,而推力系数只减小1.35%,调节收益非常明显。最后通过将喷管各调节状态下的冷流缩比实验壁面压力数据与计算结果的对比,证明了该调节方案的计算方法及其结果是可靠的,同时得出该调节方案可以有效地降低冷、热态力矩差的结论。

The difference of pitch moment under cold and hot conditions can be changed greatly when the hypersonic vehicle works at relay point and cruising end point, which will cause the big problem of the vehicle control. In order to decrease the difference of cold/hot pitch moment, an adjustment method with a slide wall at the end of nozzle up expansion wall was proposed, and was investigated by using CFD and cold flow experiment of scaled wind tunnel, and it was proven to be effective to decrease the difference of cold/ hot pitch moment. The computed results show that in Mach 4.5, the difference between the cold and hot pitch moment is reduced by 21.74% at the maximum and the thrust coefficient penalty is only 1.64% when the sliding wall is extended 400mm, as the same sliding wall out length in Mach 6.5, the difference be- tween the cold and hot pitch moment decreases by 77.59% and the thrust coefficient decreases by 1.35% , which is proven to be highly efficient in Mach 6.5. Finally the computed results are compared with experi- mental data.