Understanding the interaction between the combustor and turbine subsystems of a gas turbine engine is believed to be key in developing focused strategies for improving turbine performance. Past studies have approached the problem starting with an existing turbine rig with inlet conditions provided by "representative" hardware which attempts to mimic some key flow features exiting the combustor. In this paper, experiments are performed which center around complete engine hardware of the combustor, including engine geometry turbine nozzle guide vanes (NGVs) to solely represent the upstream impact of the complete turbine. This domain ensures that the traditional interface between combustor and turbine is sufficiently encompassed and not compromised by obfuscating or limiting effects due to approximating combustor hardware. The full-annular experimental measurements include all components of the velocity and pressure fields at various planar sections perpendicular to the primary flow direction. These include detailed, two-dimensional measurements both upstream and downstream of the NGVs. The combustor is a classic rich-burn design. Passive scalar (CO2) tracing measurements are peiformed to gain insight into the flow responsible for the temperature fields in the coupled system, including the impact of the NGVs on the upstream flow at the conventional combustor-turbine inteiface. CFD simulations are used to develop a complete picture of the combustor-turbine interface and the coupling between the two subsystems. The complementary experimental and simulation datasets are together intended to provide a benchmark for future, more traditional turbine rig tests and turbine CFD simulations where inlet conditions are at the exit plane of the combustor.

Turbulence measurements are made in a novel gas turbine rig facility recently used to study combustor-turbine interactions in jet engines [1]. The rig is capable of numerous area traverses surrounding engine turbine nozzle guide vanes (NGVs). The rig is unique in that complete engine hardware of the annular combustion subsystem is used to simulate the upstream flow entering the turbine. The rig operates at cold, near-atmospheric conditions. The turbulence measurements include both the turbulence intensities and lengthscales and span an area over a single combustor sector. Axial measurement planes include locations both upstream and downstream of the real engine hardware NGVs. The upstream plane corresponds to a conventional combustor-turbine interface plane. In [1], pressure, velocity, and passive scalar mixing measurements were presented along with RANS CFD predictions. Here, in addition to the newly measured turbulence quantities, large-eddy simulations (LES) are performed for the complete, coupled combustor-turbine system. Good agreement between rig data and CFD is seen at the combustor-turbine interface, with LES yielding improved predictions over RANS. For the flow through the NGV passages, vortex visualizations of the simulated flowfields show significant differences to the classic, commonly accepted picture of Langston [2] and others [3]. The difference is attributed to the high turbulence levels created by the combustor. The impact of the limitations of the combustor-turbine rig on these findings is discussed.

Thermal protection of gas turbine combustors relies heavily upon the delivery of a carefully managed film of coolant air to the hot-side of the combustor liner. Furthermore, improvements in engine sfc and the trend to ever more aggressive engine cycles means greater emphasis is being placed upon more efficient use of the proportion of combustion system air made available for cooling. As a result, there is a requirement to better understand the development of cooling films deposited onto the hot-side of the liner through complex effusion arrays. This study, therefore, is concerned with the prediction and measurement of adiabatic film effectiveness of a number of engine-representative designs. A RANS based CFD approach is used to predict film effectiveness in which computational cost is minimised by solving first for a single coolant passage to provide high fidelity, near-exit boundary conditions to the effusion arrays. Equivalent measurements are made for each test case using a Pressure Sensitive Paint (PSP) technique in which the oxygen-quenched fluorescence properties of the paint are employed together with a Nitrogen gas cooling simulant to determine adiabatic film effectiveness. This study demonstrates that whist the model under-predicts the mixing of the coolant with the main-stream flow, and hence the film development over the surface, the approach works well at quantifying the relative performance of each design.

Legislation controlling the permitted levels of pollutant emissions from aircraft gas turbines has been an increasingly important design driver for the combustion system for some time, particularly with respect to oxides of nitrogen. This has lead to many suggestions for radical departures from the geometry of the classical combustor configuration involving, for example, lean premixed module technology, or staging (axially or radially) of combustor pilot and main zones. The optimum operation of any combustor also requires, however, appropriate and efficient distribution of compressor delivery air to the various flametube features (fuel injectors, dilution ports, for cooling and for air bleed purposes). Radial staging, leading to double annular combustor configurations, poses a particularly difficult challenge. The radial depth of the combustor increases to a level where the external aerodynamics of the combustor involves large flow turning after the pre-diffuser. Careful design is then needed to achieve acceptable levels of loss coefficient in the outer annulus. If these aspects are not properly addressed then inadequate penetration and mixing in the combustor interior can result, rendering low emissions performance impossible. This paper will report on the design, instrumentation and operation of a fully annular isothermal test facility, which has been developed specifically to enable this important issue of external flow quality in double annular combustor systems to be assessed. Representative inlet conditions to the combustion system are generated using a single stage axial compressor; modular construction enables quick and inexpensive changes to components of the combustor (pre-diffuser, cowl shape, liner port locations and geometrical details). Computerised rig control and data acquisition allows the collection of large amounts of high quality data. In addition to the calculation of overall system performance, it is then possible to identify flow mechanisms and loss-producing features in various zones and suggest appropriate modifications.

The increasing radial depth of modern combustors poses a particularly difficult aerodynamic challenge for the pre-diffuser. Conventional diffuser systems have a finite limit to the diffusion that can be achieved in a given length and it is, therefore, necessary for designers to consider more radical and unconventional diffuser configurations. This paper will report on one such unconventional diffuser; the hybrid diffuser which, under the action of bleed, has been shown to achieve high rates of diffusion in relatively short lengths. However, previous studies have not been conducted under representative conditions and have failed to provide a complete description of the relevant flow mechanisms making optimization difficult. Utilizing an isothermal representation of a modern gas turbine combustor an experimental investigation was undertaken to study the performance of a hybrid diffuser compared to that of a conventional, single-passage, dump diffuser system. The hybrid diffuser achieved a 53% increase in area ratio within the same axial length generating a 13% increase in the pre-diffuser static pressure recovery coefficient which, in turn, produced a 25% reduction in the combustor feed annulus total pressure loss coefficient. A computational investigation was also undertaken in order to investigate the governing flow mechanisms. A detailed examination of the flow field, including an analysis of the terms within the momentum equation, demonstrated that the controlling flow mechanisms were not simply a boundary layer bleed but involve a more complex interaction between the accelerating bleed flow and the diffusing mainstream flow. A greater understanding of these mechanisms enabled a more practical design of hybrid diffuser to be developed that not only simplified the geometry but also improved the quality of the bleed air making it more attractive for use in component cooling.