飞翼模型高速风洞PIV试验研究

Experimental research on flying-wing by PIV in high speed wind tunnel

对小展弦比飞翼标模在2.4米跨声速风洞中创新开展了PIV试验。对空风洞进行了测速校核,并对小展弦比飞翼标模开展了二维、三维涡迹PIV测试,试验马赫数为0.4~0.9。测试结果表明,2.4m风洞PIV试验数据具有较高的准确度,M≤0.8时空风洞测速结果与理论值相差不超过1%,M=0.9时相差不超过2%。小展弦比飞翼标模测试结果显示,M数增大使机翼尾涡涡量和切向速度增大,涡核向内展向方向移动。前缘涡与上翼面分离具有密切关系:当M=0.8、α≤12°时,翼梢测试截面的前缘涡尚未破裂,上翼面未发生显著的流动分离;当α≥13°时,前缘涡破碎时机提前,当地后1/2弦长区域产生了比较明显的流动分离。

A PIV experimental investigation has been made in 2.4m transonic wind tunnel. The item,include 2D and stereo PIV test,and the resluts show that the deviation of the PIV results from calibration results are less than 1% when M≤0.8 and less than 2% when M=0.9. The low-aspect-ratio flying-wing model test shows that the vorticity and tangential velocity of wake increases and the vortex core moves closer as Mach number increases.The leading edge vortex has direct relationship with the flow seperation.The leading edge vortex doesn′t break in test region while M=0.8 andα≤12°and there is no significant flow seperation visible.Whenα≥13°,the leading edge vortex breaks,in advance,and significant flow seperation can be observed after 1/2 chord length of the wing.