American Institute of Aeronautics and Astronautics 092407 1 Performance of Choked Unsteady Ejector-Nozzles for use in Pressure-Gain Combustors Jonathan JH Heffer and Robert J Miller, Whittle Laboratory University of Cambridge Engineering Department If the conventional steady flow combustor of a gas turbine is replaced with a device which achieves a pressure gain during the combustion process then the thermal efficiency of the cycle is raised. All such ‘Pressure Gain Combustors’ (e.g. PDEs, pulse combustors or wave rotors) are inherently unsteady flow devices. For such a device to be practically installed in a gas turbine it is necessary to design a downstream row of turbine vanes which will both accept the combustors unsteady exit flow and deliver a flow which the turbine rotor can accept. The design requirements of such a vane are that its exit flow both retains the maximum time-mean stagnation pressure gain (the pressure gain produced by the combustor is not lost) and minimises the amplitude of unsteadiness (reduces unsteadiness entering the downstream rotor). In this paper the exit of the pressure gain combustor is simulated with a cold unsteady jet. The first stage vane is simulated by a one-dimensional choked ejector nozzle with no turning. The time-mean and rms stagnation pressure at nozzle exit is measured. A number of geometric configurations are investigated and it is shown that the optimal geometry both maximizes time mean stagnation pressure gain (75% of that in the exit of the unsteady jet) and minimizes the amplitude of unsteadiness (1/3 of that in the primary jet). The structure of the unsteady flow within the ejector nozzle is determined computationally. Introduction If a stagnation pressure gain could be achieved across the combustion process in a gas turbine then its entropy rise would be reduced compared with a conventional combustor, and the exergy and availability of the exit flow would be increased. Sir William Hawthorne in the conclusions of his 1994 IGTI scholar lecture [1], captured this concisely when he said ‘the largest loss of thermodynamic availability occurs in the combustion chamber. What is needed is a work producing combustion chamber.’ If such a combustion chamber could be practically realized then the rise in turbine inlet exergy and availability would result in a step increase in gas turbine thermal efficiency. Three types of pressure gain combustor have been reported in the literature, pulse combustors, pulse detonation engines (PDEs) and wave rotors. Each has relative merits and disadvantages, however all produce a highly unsteady exit flow which has a higher exergy than would be possible with a conventional steady flow combustor. If a practical engine using such a combustor is to achieve a higher thermal efficiency, then the downstream turbine must be able to extract a significant proportion of the additional exergy as shaft work. To achieve this, a first stage turbine vane must be designed which accepts the unsteady combustor exit flow and delivers a flow to the downstream rotor which retains the maximum possible time mean exergy and stagnation pressure. It is also a requirement of the first turbine vane that the exit flow has the minimum variation in stagnation pressure. This is because unsteadiness in the nozzle exit flow will propagate into a downstream rotor and may adversely affect the efficiency. Three conceptual methods of joining a pressure gain combustor to a downstream turbine stage have been reported in the literature. Schematics of the three are shown in figure 1. The papers are all concerned with ensuring that the increase in work output from a downstream turbine would be maximised. In the first method, fig 1a, a plenum is located between combustor exit and turbine inlet. Such a geometry was reported by Gemmen et al [2]. They found that the plenum reduced unsteadiness in the exit flow but that the dissipation of the exit flow in the plenum resulted in a stagnation pressure gain of only 1%. These tests show that dumping the flow into a plenum at combustor exit would significantly reduce the benefit of pressure gain combustion. 47th AIAA Aerospace Sciences Meeting Including The New Horizons Forum and Aerospace Exposition 5 - 8 January 2009, Orlando, Florida AIAA 2009-1063 Copyright © 2009 by J Heffer and R Miller. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.