航空发动机
航空髮動機
항공발동궤
AERO ENGINE
2011年
3期
50-54
,共5页
涡轮静子端壁%换热%全气膜冷却%进口吹风比
渦輪靜子耑壁%換熱%全氣膜冷卻%進口吹風比
와륜정자단벽%환열%전기막냉각%진구취풍비
turbine vane endwall%heat transfer%film cooling%inlet blowing ratio
对跨声速涡轮静子端壁气膜冷却进行了数值研究。研究发现涡轮静子端壁存在几个强换热区域：叶片前缘马蹄涡及前缘马蹄涡区域、吸力面马蹄涡分支覆盖区域、通道中靠近压力面侧和尾缘附近及尾缘后区域。针对端壁区域复杂的换热分布,设计了1种新型端壁全气膜冷却布置。通过数值研究对比了在不同进口吹风比情况下的壁面Nu、壁面气膜冷却效果和壁面热负荷。结果表明：存在最佳的进口吹风比,即在前缘Minlet=1.0时,尾缘Minlet=4.0时,端壁区域冷却效果最好。
對跨聲速渦輪靜子耑壁氣膜冷卻進行瞭數值研究。研究髮現渦輪靜子耑壁存在幾箇彊換熱區域：葉片前緣馬蹄渦及前緣馬蹄渦區域、吸力麵馬蹄渦分支覆蓋區域、通道中靠近壓力麵側和尾緣附近及尾緣後區域。針對耑壁區域複雜的換熱分佈,設計瞭1種新型耑壁全氣膜冷卻佈置。通過數值研究對比瞭在不同進口吹風比情況下的壁麵Nu、壁麵氣膜冷卻效果和壁麵熱負荷。結果錶明：存在最佳的進口吹風比,即在前緣Minlet=1.0時,尾緣Minlet=4.0時,耑壁區域冷卻效果最好。
대과성속와륜정자단벽기막냉각진행료수치연구。연구발현와륜정자단벽존재궤개강환열구역：협편전연마제와급전연마제와구역、흡력면마제와분지복개구역、통도중고근압력면측화미연부근급미연후구역。침대단벽구역복잡적환열분포,설계료1충신형단벽전기막냉각포치。통과수치연구대비료재불동진구취풍비정황하적벽면Nu、벽면기막냉각효과화벽면열부하。결과표명：존재최가적진구취풍비,즉재전연Minlet=1.0시,미연Minlet=4.0시,단벽구역냉각효과최호。
The numerical study was conducted on the endwall film-cooling of transonic turbine vane.Several regions of high heat transfer rate were presented on turbine vane endwall.They are leading edge region where the horseshoe vortex occurs,suction side region where horseshoe vortex happens,channel region near the pressure side,and region near the trailing edge and downstream of trailing edge.According to the complex heat transfer distribution on the endwall,a new full film cooling holes on the endwall was designed.The endwall Nu number,film cooling effectiveness and thermal loads on the endwall were compared under different inlet blowing ratio,the result shows that there is the best inlet blowing ratio,the film-cooling effectiveness on the endwall is best at M■=1.0 on the leading edge and M■=4.0 on the trailing edge.
