Thermal design of a cooled micro gas turbine
Author(s)Evans, Simon William, 1977-
Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics.
Alan H. Epstein.
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One of the major challenges associated with designing a micro gas turbine engine is the problem of heat transfer. The demonstration version of the engine deals with this problem by transferring excess heat from the turbine, to the compressor wall, through the rotor shaft. This is necessary to keep the turbine wall within its temperature constraints. The resulting heat transfer into the compressor flow however reduces the compressor performance to the point that the cycle will no longer close. A film cooled turbine has thus been pursued as a means of keeping the turbine within its temperature constraints and at the same time reducing heat transfer to the compressor. The thermal design of this cooled micro gas turbine has involved the design of the thermodynamic cycle, a secondary flow system to carry compressor discharge air to the turbine for cooling, and conceptual design of a turbine and rotor shaft to match the compressor. The analysis leading to this design identified turbine wall temperature, turbine exit radius and shaft area as three tools for increasing the power of the turbine, required to close the cycle. The design converged upon revealed that a very high cooling effectiveness is required to close the cycle, if the turbine wall is to be limited to 950K. This high effectiveness is calculated according to an empirical model established with data from full size engines, and thus represents an extrapolation of data with its attendant risks. A comparative model was developed as a regression of CFD results produced for the engine geometry. This model predicts adiabatic cooling effectiveness values too low to close the cycle. From the cycles studied, the recommended cycle configuration includes a 10mm diameter turbine with 1600K at rotor inlet. 41% of compressor inlet air is required to cool the turbine wall to 950K, and shaft area required to be 0.1% of a solid 6mm diameter shaft, i.e. 0.079mm2. The resulting cycle breaks even with a compressor pressure ratio of 2.46 and efficiency of 43%. Turbine efficiency is 63%. This solution shows that closure of the cycle is possible. It however suggests that further design study and technology development is needed to generate useful levels of engine performance.
Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 2001.Includes bibliographical references (p. 169-170).
DepartmentMassachusetts Institute of Technology. Dept. of Aeronautics and Astronautics.
Massachusetts Institute of Technology
Aeronautics and Astronautics.