Advanced Cooling Technique for Gas Turbine Combustors

A key problem associated with aircraft turboshaft engines is the need to provide adequate cooling for the combustor liner. The general approach which has been taken toward this problem is to utilize a portion of incoming combustor air for film-cooling of the liner. This proposal deals with a cooling technique for combustor liners which does not utilize film cooling. Instead, a special liner design is employed which enables incident heat to be absorbed by incoming combustor air. The innovative combustor liner cooling concept provides a more uniform temperature distribution in the combustor exit gas, and permits effective liner cooling even when film-cooling techniques are inadequate. Project objectives are: to establish design characteristics of the combustor liner, to assess the feasibility of the innovative liner cooling technique, and to identify research and development needs. Design and performance studies will be performed, a conceptual liner cooling system design will be prepared, and key problem areas will be evaluated. It is expected that results will verify the technical attractiveness of the new liner cooling concept, and its applicability to military and civilian turboshaft and other gas turbine applications.