Reverse flow can annular combustor configuration becomes the inevitable option for industrial and marine gas turbine engine, due to its advantages over other configurations. The complexity associated with can annular configuration is optimum design of annular diffuser, as its flow field is dominated by downstream blockage created by transition duct geometry. In the present study, flow behavior in the annular diffuser has been analyzed by simulating realistic downstream combustor liner and transition duct geometry. Flow analysis has been carried out using ANSYS Fluent and turbulence has been modeled using Realizable k-ε model. The diffuser is designed based on G* method, for optimum pressure recovery. Six diffuser configurations have been analyzed by varying the inner wall profile. The effect of parameters on flow field within diffuser and dump region has been studied. Also, the static pressure recovery and total pressure loss coefficient of diffuser is calculated and compared. The results show that the profile of the inner wall and the dump region needs to be tailored to get optimum performance from diffuser.

Relight envelope of the combustor needs to be experimentally generated and established during the design and development of an aero gas turbine engine. Usually, during development stage of engine, compressor characteristics are not readily available at such low speeds and hence, it becomes difficult to specify the combustor inlet conditions such as pressure, temperature and Mach number during the engine light up studies. This paper compares the experimental test data generated on an annular combustor for windmill conditions during stand-alone mode and engine level tests under simulated flight conditions. The stand-alone combustor trials were conducted for the range of total pressure and temperature relevant to the flight altitude and Mach number range. During the engine level tests, combustor relight tests were conducted under simulated conditions (ISA+15) for altitudes ranging from 5.5 km to 10 km, flight Mach numbers in the range of 0.45 to 0.80. In this paper, effect of altitude and flight Mach number on the windmill spool speed, combustor pressure and temperature are studied.

Swirl cups (hybrid atomizers) are being widely employed in aero gas turbine engine combustors for their established merits in terms of achieving satisfactory atomization over the entire combustor operating regime. Even though several investigators have worked on development of these swirl cups, there is a scanty data reported in literature relevant to their design. In the present study, flow behavior in a swirl cup assembled in a confined chamber similar to a gas turbine combustor has been analyzed. Flow analysis has been carried out using ANSYS Fluent and turbulence has been modeled using Realizable k-ϵ model. Six swirl cup configurations have been analyzed; mass flow ratio between primary and secondary swirler and venturi converging area ratio have been varied. The effect of these parameters on downstream flow field has been studied by analyzing the profiles of axial, tangential and radial velocities downstream of swirl cup. The size and shape of the recirculation zone has been analyzed and reported for all configurations. Also, the mass flow recirculated by swirl cup has been estimated and compared amongst the configurations analyzed. Data thus generated is very useful in designing such swirl cups of gas turbine combustors.

Over the years, the requirements of higher specific thrust and lower specific fuel consumption have been necessitating a continual increase in the maximum temperature and pressure in gas turbine engines. However, such an increase has a direct impact on the structural integrity of various modules of the engine; combustor being one of the severely affected modules. This makes the combustor designer’s task of achieving the targeted life of liner, the hottest component of combustor, a challenging one. Estimation of liner metal temperature, thereby arriving at the combustor life, is an essential part of the design process. In the present study, CHT analysis of a radial annular combustor has been carried out. RANS based analysis of a sector combustor with periodicity in flow and geometry has been performed at realistic engine operating conditions using ANSYS Fluent. Predicted liner metal temperatures have been compared with the measured data and a close agreement has been noted between them, the maximum variation being ± 10%.

Achieving minimum pressure loss and exit temperature pattern factors are the two important, but conflicting performance goals of a gas turbine combustor. In this study, Taguchi method was used to design CFD (Computational Fluid Dynamics) analysis based experiments in order to arrive at optimized liner hole sizes (primary and dilution) of an annular gas turbine combustor. The CFD experiments were steady & three-dimensional Reynolds Averaged Navier-Stokes (RANS) based analyses performed on a 20-degree sector combustor model. Nine CFD analyses were performed as per the Taguchi L-9 orthogonal array, based on which, an optimal hole configuration was arrived. An additional CFD run was carried out to verify the predicted result. The optimized hole size combination reduces the Circumferential Pattern Factor (CPF) by 28% and pressure loss by 0.06% (difference between absolute values of percentage pressure losses) in comparison to the baseline configuration.