CAJ | 학술논문

本文利用实验与数值模拟的方法研究航空发动机涡轮叶栅在改变节距时，其总压损失、出口气流角及叶栅表面静压系数的变化。通过实验和数值计算的结果对比，验证了数值计算的准确性。数值模拟结果显示：相对节距珋t在0.5～1.0范围内变化，叶栅的总压损失在珋t＝0.9时达最小；叶栅入口攻角从-10°逐渐增加至＋10°时，总压损失呈现递增的趋势；随着相对节距每增加0.1，气流出口角相应增加大约2°；随着叶栅相对节距的增加，附面层增厚、脱离，叶型损失增大。
본문이용실험여수치모의적방법연구항공발동궤와륜협책재개변절거시，기총압손실、출구기류각급협책표면정압계수적변화。통과실험화수치계산적결과대비，험증료수치계산적준학성。수치모의결과현시：상대절거류t재0.5～1.0범위내변화，협책적총압손실재류t＝0.9시체최소；협책입구공각종-10°축점증가지＋10°시，총압손실정현체증적추세；수착상대절거매증가0.1，기류출구각상응증가대약2°；수착협책상대절거적증가，부면층증후、탈리，협형손실증대。
In this paper , cascade pitch is changed to find out the trend of total pressure , exit flow angle and surface static pressure co-efficient by experiment and numerical simulation .Contrast experiment result with numerical simulation , prove that numerical simulation is correct .The result of simulation show:Relative pitches of cascade are changed from 0.5 to 1.0 and find out that total pressure loss has a best result when relative pitch t=0.9;When angle of attack is increased from -10°to +10°, the total pressure loss also in-creases;Flow exit angle increases 2°with each 0.1 increase in relative pitch;With the increasing of relative pitch , boundary layer be-come thicker and separate from the cascade , profile loss increases .