Numerical Study On Blade Cooling Structure Modification Design Of High-pressure Turbine Rotor

With advantages such as high power, small size, light weight, quick start andacceleration, Gas Turbine has been widely used in aviation, marine and power industriessince it was invented in the early20th century. The fact that the gas turbine technologycan be involved in dual-use and take political, military, and economic factors intoconsideration，the developed countries strictly blockade the technology against China.Therefore，the design of high performance gas turbine can only rely on independentresearch and development of China’s engineering and technical personnel. It is wellknown that the combustor exit temperature is an important indicator to measure theefficiency of gas turbine cycle and it has been higher than the temperature limit ofmaterials nowadays, hence effective cooling technology of high temperaturecomponents must be included in the key to developing high-performance gas turbines.Since the high pressure rotor blades of gas turbine carry a huge aerodynamic load andwork in a very poor environment, it must be attached great importance to develop bettercooling structure to improve the cooling effects of gas turbine high temperaturecomponents.In this paper, the ANSYS CFX software is employed to study the first level of theF-class heavy-duty gas turbine rotor blades.The numerical simulation results offer abetter understanding of the first stage rotor blades of composite cooling structure andmore detailed analysis of the first stage rotor blades cooling structure and the coolingmechanism，which helps to explain the reasons of the high surface temperature and coldair uneven distribution in prototype blades. Therefore, a modification program of addingthe thermal barrier coating and removing the gas compensating hole is put forward towork out this problem.Since prototype rotor was modified, numerical simulation about the modificationof adding the thermal barrier coating is performed. Compared with the calculatingresults of the prototype, the modified rotor blade temperature is reduced to protect thesurface of metal blade, which results from the lower thermal conductivity of the thermalbarrier coating and leads to a higher temperature outside the thermal barrier coating.And the analysis of the second modification of removing the gas compensating holeshows that the cold air in the serpentine channel gets a more reasonable flow, and helpsto reduce the rotor metal blade temperature and the maximum temperature difference,which matches well with the desired cooling effect.Usually the high-pressure gas turbine rotor blades cooling experiments are finishedunder static condition. It’s necessary to verify the reliability of the results of staticexperiments when difficulty and economy factors are taken into account. Therefore, the impact of rotation on the first stage rotor blades cooling effect was studied bycomparing the aerodynamic performance and heat transfer performance of rotor bladesunder static and rotating conditions in this paper. Simulation results show thatcentrifugal force, Coriolis force, and their derivatives Bliss force of cold air in theserpentine channel under the rotating conditions can cause the turbulence intensity andthe heat transfer capacity to be increased. Consequently, as long as the coolingexperiments under static conditions meet the design requirements, it can be inferred thatthe rotor blades cooling experiments in rotating conditions can achieve better coolingeffect.