With gas temperatures far exceeding the melting point of nickel-base alloys, advanced cooling schemes are essential to meet the desired mission life of turbine airfoils. Naturally, combustion systems produce gas-temperature non-uniformity in the exiting flowfield. Downstream turbine components must be tolerant to the maximum anticipated gas temperatures. On the other hand, excessive use of cooling air reduces engine efficiency and compromises combustor durability. Throughout gas turbine design history it has been the desire of Turbine Aerodynamicists to be able to compute combustor hot streak migration and mixing through multiple turbine airfoil stages. Typically, hot streak migration studies have been performed using (a) mixing-plane models between rotating and stationery domains or (b) unsteady simulations in which the flowpath annulus is represented by a segment containing airfoil counts that are integer multiples in each blade row or (c) Non-Linear Harmonic methods. With the development of highly-parallelized Computational Fluid Dynamic (CFD) codes driving high performance computer clusters simulation of combustor hot streak migration through multiple High Pressure (HP) turbine stages using an unsteady, 360° (full-annulus) model can be achieved. To this end, Honeywell, in collaboration with Numeca Corporation, has performed a study to evaluate the state-of the art for computation of the effect on second-stage HP turbine nozzle metal temperatures of combustor hot streaks migrated through the first-stage of a two-stage HP turbine.

Modern direct-drive turbofan engines typically have the fan turbine designed at significantly higher diameter than the gas producer turbine. Furthermore, the gas turbine industry is being pushed to shorten engine length with the goal of reducing weight. This results in a need to design very aggressive inter-turbine-ducts (ITD’s) that have high endwall slopes. The gas turbine design cycle typically begins with conceptual design where many engine configuration iterations are made. During conceptual design, there usually is little firm geometric definition or time for detailed CFD studies on aggressive ITD’s. This can cause a large amount of risk to the engine development schedule and cost if the space allocated for the ITD during conceptual design is found to be insufficient later in the design cycle. Therefore, simple analytical tools for accurately assessing the risk of an ITD in conceptual design are important. The gas turbine industry is familiar with the Sovran and Klomp annular diffuser performance chart [1] as a conceptual design tool for assessing ITD’s. However, its applicability to modern gas turbine ducts with high endwall slope is limited. The location of the maximum pressure recovery for a given length, the Cp* line, considers only two geometric parameters: area ratio and normalized length. The chart makes no distinction of risk of flow separation regarding the level of slope or the pitch-wise turning in the duct. However, intuition would suggest that a high slope duct would have more risk of separation than an equivalent area ratio duct with low slope. Similarly, a duct that turns the flow from axial to radial would be expected to be riskier than a pure axial duct. To help assess the interaction of duct slope and pitch-wise turning with area ratio and length, an analytical Design of Experiments (DOE) was run using approximately sixty different duct configurations. The DOE was carried out using 3D, steady CFD analysis. The results of the DOE are presented with insights provided into how the Cp* line may shift as a function of duct slope. Of particular interest is that slope by itself does not work particularly well as a risk indicator. However, a combination of new area ratio-length and slope-length parameters was found to segregate ducts between separated and non-separated cases.

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