Effects of Combustor Exit Flow on Turbine Performance and Endwall Heat Transfer

Strict limits for pollutant emissions by aircraft engines strive the development of new combustion concepts. Lean combustion decreases the peak temperatures during the combustion process as the stoichiometric mixture of fuel and air is avoided. Thereby, the production of NOx can be reduced. However, a highly swirling flow inside the combustion chamber is required to enhance the mixing of fuel and air and to stabilize the flame front in a zone of recirculating flow. This swirl persists downstream at the turbine inlet. It is hardly attenuated by radial injection of dilution air as it is the case in conventional combustion chambers. The recirculation zone causes a redistribution of mass towards the endwalls, temperature and pressure peaks and a very high level of turbulence.
In contrast, moderately low levels of turbulence are assumed during blade design of high pressure turbines. An axial inflow direction is applied as well as constant values of pressure and temperature. The current thesis investigates the effect of realistic turbine inlet boundary conditions in comparison to the design point using computational fluid dynamics simulations. Swirling inflow and the resulting pressure distribution as well as turbulence and hot streaks from the combustion chamber are included. The complexity of the inflow condition is successively increased over three different test cases: First, a cascade is investigated for different swirl intensities. Second, an annular 1.5 stage turbine is analyzed applying different clocking positions and swirl orientations. Third, the 2.5 stage high pressure turbine of a real jet engine is simulated at varying temperature levels.