The Study Of Film Cooling And Heat Transfer Characteristics On The Turbine Shroud

The turbine shroud is one of the main components of the turbine cascade.It is the part of the turbine casing that directly contacts the gas.It is subjected to the strong scouring of the high-temperature and high-speed airflow.And it has a large thermal load and is easily ablated,so it must be efficiently cooled.The turbine shroud of foreign advanced aero-engines adopts advanced air film cooling technology,but the domestic understanding of the film cooling characteristics and mechanism under this special flow-structure condition is not deep enough,and the basic research is very insufficient,which restricts the design of high-efficiency film cooling structure for engine turbine shroud in china.Based on the above background,this paper studies the heat transfer characteristics and flow mechanism of the gas film cooling structure of the turbine shroud unit of the turbine by establishing the turbine shroud experiment and numerical research method.In this study,by simulating the unsteady turbine shroud working environment under blade rotation,it is found that the main factor affecting the film cooling characteristics of the turbine shroud is the gas incident angle of the blade,velocity and the turbine shroud film injection is brought by the blade rotation of the narrow channel height effect.Therefore,under the similar conditions of flow parameters,and the mainstream of the narrow angle of incident angle,the experimental research method of the uniform film cooling characteristics of the turbine shroud is proposed.And the flow field structure and working environment of the turbine shroud film injection are relatively real,the characteristics of the film cooling are obtained.In the experimental study of circumferential film hole in turbine shroud,the distribution of cooling effectiveness of single row hole on turbine shroud surface under different blowing ratio,Reynolds number,hole spacing and angle was compared by simulating the working environment of turbine shroud with incident angle and main channel under narrow slot.The main conclusions are as follows: under circumferential degrees,similar conclusions can be drawn.The distribution area of high cooling effectiveness near the exit of the film hole is larger,the spread distribution is wider,and the film cooling effectiveness value is higher than that near the outlet of the film hole.At the high blowing ratio,the cooling effectiveness of the film hole near the exit of the hole decreases obviously,and the effect of the Reynolds number on the turbine shroud film hole increases with the increase of the Reynolds number.At the high blowing ratio,the cooling effectiveness increases at first and then decreases with the Reynolds number increasing.The cooling effectiveness decreasing with the increase of the hole spacing.Comparing the film holes with different circumferential dip angles,the cold effect of the small dip angle film hole is the highest.With the increase of the hole inclination angle,the cold film effect value decreases,and the film cooling effect gradually deteriorates.In the experimental study of the down-flow direction and back-flow direction of the turbine shroud,the distribution of the cooling effectiveness of the single exhaust film on the outer ring surface under the basic parameters of the film hole,such as the blowing rat io and the Reynolds number of the aperture,were compared under two different flow direction angles.The main conclusions are as follows: the film holes have better film coverage under small blowing ratio,and there are more high cooling efficiency areas near the outlet of the holes,and the distribution of cooling effectiveness is wider.With the blowing ratio increasing,the distribution of the high cooling effectiveness area near the exit of the hole becomes narrower and the flow direction becomes shorte r.At the high blowing ratio,the phenomenon of air-conditioning breaking away from the wall,the film flow direction becomes worse,and the cooling effectiveness decreases.With the blowing ratio increasing,the film cooling effectiveness decreases with the Reynolds number increasing,the comparison of the two holes shows that the film cooling effectiveness decreases with the Reynolds number increasing.Comparing the flow holes of two different angles,under the small air blowing ratio,especially near the design point,the cooling efficiency under the dip angle of the downstream hole is higher,and the cooling effect is better.In the numerical study of the unsteady turbine shroud film cooling with the blade rotation,compared with the time averaged cooling effectiveness and the heat transfer coefficient ratio with the blade rotation,the distribution of the time-averaged result is mainly affected by the flow in the channel and the complex flow in the tip,the rotation of the blade.The hole angle has less influence on the time-average results.And the time-average heat transfer ratio is larger at a smaller angle and is significantly enhanced.In the study of multi-row holes,according to the film cooling effectiveness and heat transfer characteristics of the turbine shroud under different hole arrangement,the difference in film cooling and heat transfer characteristics under the hole layout is not obvious change.The impact diaphragm is added into the secondary flow chamber.As the blowing ratio increases,the impinging airflow flows back into the impact wall in the confined space of the secondary flow chamber of the turbine shroud,so that the cold air side heat change increases.Comparing the film cooling effectiveness and heat transfer characteristics o f the turbine shroud with impact inlet or large chamber inlet,it can be seen that the relative velocity of the exit of the shroud under the impact inlet is larger,and the gas membrane is out of the wall surface.The cooling effectiveness is smaller than that of the non-impacted case.And the heat transfer coefficient is significantly different at different blow ratios.