Efficient Turbo-cast Cold Blade Cooling Design

The cooling technology of turbine blade is one of the most important key technologies for aero engine design. The progress in the cooling technology will directly promote the development of the discipline of heat transfer. Although film-impingement-convection composite cooling technology has been widely used on turbine blade, research activities on it are still very active, as to meet the increasingly need of high cooling effectiveness. Today, however, the turbine inlet temperature reaches 1800K～2000K for the series of aero engines whose thrust to weight ratio reaches 10, and further high turbine inlet temperature reaches 2100K～2300K (thrust to weight ratio 15～20). The hot gas temperature far exceeds the material permissible temperature of high temperature alloy. The multiple high efficiency cooled turbine vane merely fill the demand of the aero engine whose thrust to weight ratio reaches 10. In order to develop next generation higher performance aero engine, a new high efficiency cooling techniques of turbine blade have to be developed.In this paper, the new type cast-cooling turbine blade was designed using the foreign advanced cooling technologies of turbine blade and the thought way of high technology origin and large technology span for reference. The cooling technologies of film shower-cooling, film spill internal heat transfer coefficient enhancement and new type narrowness space impingement-film cooling were used in the turbine blade. Then several flow and heat transfer problems extracted from the cooling scheme were studied in details, which include: the distribution of film cooling effectiveness, the characteristics of flow and heat transfer in the internal cooling passage, and the temperature distribution of the blade. During the study the calculation method of film effectiveness and flow characteristics of internal cooling passage that currently used in engineering design was modified, and the accuracy of the modified method was verified experimentally.The main results of the study are as follows: 1) The modified calculation method of film cooling effectiveness and internal flow characteristics improved its accuracy. 2)Thermal network method can be used to calculate the flow characteristics of the very complex blade internal cooling passage. It can be used to calculate the mass flow rate distribution among impingement holes, film holes and trailing-edge slot, and the pressure and temperature distribution along the radial and chord direction. The agreement between calculated results and experiments results is quite good. 3)Under the design condition, an average cooling effectiveness in the cross section of half height of blade reached 0.7, and the temperature reduction was up to 860℃, and the cooling performance fulfills the design goal. The cooling techniques of turbine blade are close to the foreign advanced aero engine turbine blade cooling level (thrust to weight ratio reaches 15～20).All the results show that the cooling scheme for the cast-cooling turbine blade design in this thesis is reasonable and effective. It can meet the need to cool HP turbine blade of next generation aero engine.