Research On High Pressure Turbine Blade Tip Aerodynamic And Heat Transfer

Gas turbine plays a key role in the aviation, marine, electric power, oil and other energy application fields. Turbine, as one of the high speed revolution core components, its performance directly affects the overall performance of gas turbine. The existence of clearance between blade of high pressure turbine and the stationary casing, not only results in the turbine aerodynamic performance degradation, makes the blade tip exposed in high temperature gas and more likely to cause erosion and abrasion in the tip area appears. Effective protection strategies for the blade tip can reduce leakage losses, diminish the tip surface temperature and increase the service life of the rotor. At the beginning of this paper, the two-dimensional groove gap model is established to analysis the flow and heat transfer by numerical simulation. Then a linear cascade experiment of tip clearance is made to discuss the clearance leakage flow control methods. At last the high pressure turbine blade tip aerodynamic and heat transfer are fully studied by numerical simulation.In the two-dimensional groove gap model, the effects of the gap size, rim height, groove width, pressure ratio on the clearance internal flow, total pressure loss and leakage flow rate are researched. The influence of wall motion on the clearance flow is studied. Then the film cooling in the internal of groove with the up wall moving is analyzed. It analyzes the influence of the relative location, angle, blowing ratio of the cooling hole on the wall Nusselt number. The results show that the smaller gap size, greater groove width and rim height, higher pressure ratio lead to a greater internal loss. Under the same pressure ratio, a greater total loss corresponds to a smaller leakage flow rate. Wall movement makes the groove internal loss increase and leakage flow rate reduce. When the cooling hole located in the upstream of the leakage flow impingement region, the wall Nusselt number can be greatly reduced by the film cooling.The linear cascade experiment is made to study the effects of tip rim and tip injection on the tip leakage flow. The goal is to explore the blade tip structure and jet location to the leakage flow control effect. The tip leakage flow distribution is analyzed by numerical calculation. Results show than the partition opening of tip rim increases the leakage flow rate at opening location. The nearer to the pressure side is the tip injection position, the greater impact on passage vortex. The tip rim and tip injection can reduce the total pressure loss at the exit of cascade to achieve effectively tip leakage flow control under small gap size, but invalid approaches under large gap size.Numerical calculations for rotor passage are made to analysis the effects of blade tip construction and film cooling on the aerodynamic and heat transfer. Research on rim height at five kinds of gap size suggests that rim height increasing will lead to the total energy loss decreased first and then increased. The more significant control effects of leakage flow by tip groove would be achieved at larger gap size. The start and end position more closer to the leading and trialing edge, the rotor aerodynamic loss is smaller. In the types of rim opening at the rear of blade tip, the size of the rim opening at pressure side increasing leads to the losses decreased first then increased, while rim opening at suction side works well at all opening sizes. Film cooling at flat tip can effectively decrease the blade tip temperature at the downstream region of cooling holes, but has little cooling effect at upstream and front region of blade tip. Finally, fluid-structure coupling results show that the temperature of the tip leading edge rim and the suction side rim is very high without blade tip cooling. The cold air tip leakage which comes from leading and pressure side cooling holes makes the pressure side rim temperature lower than the rest of blade tip. Case 4 which arrange radial holes in the leading edge and impingement holes to tip rib at bottom of groove shows a significant reduction in blade surface high temperature area. The highest temperature of blade surface is lower than any other computation schemes. And it realizes the effective cooling of blade tip.