| Gas turbine engine used in aviation can improve plane thrust, save space, reduce aircraft volume, and improve plane motor performance. With the development of aero engine technology, the temperature before turbine was gaining height. The degree of T3*in the most advanced aircraft engine has been raised to about2000K. It is therefore most necessary to adopt rational cooling strategies to reduce the heat load of turbine vanes and blades, which is important guarantee to extend the life of gas turbine. Film cooling technology, as an important cooling method, has received attention from more and more researchers. Based on the previous research work about film cooling, different film cooling geometry has been studied in this paper. Additionally, a new-type of film cooling geometry has been used on the tip of HPT blade of1+1/2counter-rotating turbine. Detailed numerical research work has been carried out to obtain the heat transfer and cooling performance of the new-type of film coolig geometry. Leading edge film cooling has also been studied in the paper. The main content in this paper is listed as follows:1. In order to further improve the film cooling effectiveness of double-jet film-cooling (DJFC) geometry, a diffused double-jet film-cooling geometry is presented in this paper. Numerical investigation of this geometry and cylindrical double-jet geometry is carried out. This paper is focus on the effect of vertical distance of the two holes on film cooling effectiveness. Computational results show that both of the two DJFC holes can lead to anti-kidney vortex structure in all cases. Diffused double-jet geometry is able to provide larger lateral film coverage area on the platform, more symmetrical vortex structure and better cooling effect. It is found that the optimal vertical distance of diffused double-jet geometry is1.5D, while it is ID for cylindrical geometry. Additionally, the laterally averaged film cooling effectiveness of the diffused double-jet hole with vertical distance of1.5D is higher than the cylindrical hole with vertical distance of ID at all blowing ratios investigated.2. A low speed wind tunnel has been constructed to verify the numerical research work of diffused double-jet film-cooling geometry. The experimental results indicate that blowing ratio has little effect on the laterally averaged film cooling effectiveness as the distance of the two single holes of diffused double-jet film-cooling geometry is1.5D.3. A new-type of converging-slot hole on1+1/2counter-rotating turbine blade tip has been designed and studied. Baseline flat tip geometry with cylindrical holes was considered. The research work found that the blade tip heat load reduced by20.8%of1+1/2counter-rotating turbine as the new-type of converging-slot hole has been used on the blade tip at blowing ratio M=1.0, tip clearance k=0.7mm. Moreover, this new-type of cooling geometry has been detailed studied at different tip clearance and different blowing ratios. Detailed heat transfer and film cooling performance of this new-type of cooling geometry have been obtained..4. In this paper,3D unsteady numerical simulation of leading edge of1+1/2counter-rotating turbine LPT has been carried out to obtain detailed flow field and temperature field. And then, the unsteady characteristics of leading edge film cooling can be received.
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