摘要
通过孔缝结构把温度较低的气体冷却介质输送到需要热防护的壁面上,并沿高温气流的流向形成冷气薄膜,将固体壁面与高温气流隔离开来,防止壁面因直接接触高温气流而超温损坏的技术称为气膜冷却,该技术有效地应用于燃气轮机燃烧室和透平叶片等高温零部件。本研究对某航空燃气涡轮发动机单级跨声速涡轮进行了全气膜冷却设计,采用源项模型对冷气喷射进行仿真,通过数值模拟得到了不同冷气喷射条件下的涡轮流场,研究了冷气喷射对涡轮气动特性的影响,并分析了叶片表面气膜冷却效率。结果表明,随着冷气喷射量的增加,涡轮膨胀比和气动效率均有所下降,降低幅度约2%,而通流能力的增强以及冷气部分动能的有效利用使得涡轮输出功率有一定程度的提高。通过分析叶片表面气膜冷却效率发现当冷气/主流流量比为0.09时获得了最佳气膜冷却效果。
Increasing gas temperature at inlet of the turbine is useful to improve the output power and the efficiency of the aero engine thermal cycle, thereby increasing the thrust-to-weight ratio and re-ducing the fuel consumption rate. It is the main technical approach to improve engine perfor-mance. However, with the increase of gas temperature, the thermal load of turbine blade will in-crease dramatically, which poses a severe challenge to its mechanical strength, working stability and service life. As an important cooling method, film cooling has always been one of the research hotspots of high-performance aero engines. The use of the least cooling medium to obtain the best cooling effect to meet the normal working needs of the blades, and to minimize the influence of the cooling medium on the aerodynamic performance of the turbine, is the significance of the film cooling research. Full film cooling was designed for the single stage transonic turbine of an aero gas turbine engine. The coolant injections were simulated by source term model, and the turbine flow fields under different conditions of coolant injections were obtained by numerical simulation. The influence of coolant injections on aerodynamic characteristics and the film cooling effectiveness of turbine blade surface were studied. It shows that, the expansion ratio and aerodynamic efficiency of the turbine decrease by about 2% with the increase of injections, and the output power increases to a certain extent because of enhanced through-flow capability and partial utilization of the kinetic energy from coolant. The film cooling efficiency on the blade surface indicates that the optimal film cooling effect is obtained when the coolant-to-mainstream flow ratio is 0.09.
出处
《应用物理》
CAS
2020年第3期222-230,共9页
Applied Physics
基金
中国民航大学实验技术创新基金项目(2017SYCX11)
中央高校基本科研业务费项目(3122014D015)。
