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Due to the proximity of the first-stage gas turbine vanes to the combustor, coolant introduced to the combustor walls interacts with the endwall film coolant and changes the vane passage flow physics. Recent results show that combustor coolant contributes significantly to cooling the endwall and vane surfaces. In this paper, the traditional combustor-turbine interface was modified to improve overall cooling performance. The performance of this new injection cooling scheme on passage fluid dynamics and surface cooling is assessed. The first of this two-part paper reports detailed experimental tests that document secondary flows and coolant transport throughout the vane passage for four combustor coolant flowrates. The experimental facility imitates combustor coolant injection and engine-level turbulence and has a modified transition duct design, called the “close-coupled combustor-turbine interface.” The “impingement vortex” seen in previous studies with combustor cooling appears as the dominant secondary flow. It is observed in the present study over a wide range of flowrates, confirming its tie to the combustor coolant flowrate and not the combustor-turbine interface geometry. It was found, however, that the location and size of the impingement vortex are affected by coolant flowrate. Part II of this paper discusses the impact of the observed secondary flows on cooling vane passage surfaces.

The first-stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics is not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first-stage nozzle guide vane passage. The test section involves an engine-representative combustor–turbine interface geometry, combustor coolant flow, and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.