Exit survey measurements comparing exit losses for a constant inlet height cascade and a contracting inlet cascade have been acquired in a low speed wind tunnel facility. The measurements were taken for both a low and a high turbulence level across a four to one range in chord Reynolds number (500,000, 1,000,000, and 2,000,000). The high intensity turbulence has been generated using a simulated aero-combustor for both cascades. Exit survey measurements have been acquired for the two cascades using a five-hole cone probe at stations representing an exit location of 1/4 axial chord downstream from the trailing edge plane. Measurements detail total pressure loss, turning angle, and secondary velocities. At the low turbulence condition, the contracting inlet cascade shows a significant reduction in losses compared with the constant height inlet cascade. The contracting inlet cascade also features an aft loaded vane profile which may have some impact on the development of secondary flows. The constant cross-section cascade uses a fully loaded vane profile. The difference in losses between the two cascades is also significant at the high turbulence condition. However, at the high turbulence condition, the losses for the contracting inlet cascade are greater than the constant cross-section cascade. These increased losses are believed to be due to the more aggressive turbulence of the contracting inlet case. The contracting inlet effectively moves the vane leading edge plane forward into the exit of the combustor where the vane passage is subjected to more aggressive turbulence levels.

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The growing trend to achieve a higher Turbine Inlet Temperature (TIT) in the modern gas turbine industry requires, in return, a more efficient and advanced cooling system design. Therefore, a complete study of heat transfer is necessary to predict the thermal loadings in the turbine vane/blade. To estimate the metal temperatures, it is important to simulate the external hot gas flow condition, the conduction in the blade material, and the internal coolant flow characteristics accurately and simultaneously. As a result, proposing novel, quicker, and more convenient ways to study the heat transfer behavior of gas turbine blades is of absolute necessity. In the current work, a predictive model for the gas turbine blade cooling analysis in the form of a computer program has been developed to answer this need. The program is capable of estimating distribution of coolant mass flow rate, internal pressure and metal temperature of a turbine blade based on external and internal boundary conditions. The simultaneous solutions result from the coupled equations of mass and energy balance. The model is validated by showing its accuracy to predict the temperature distributions of a NASA E3 blade with an uncertainty of less than +/−10%. Later, this paper documents the overall analysis for a set of different boundary conditions with the same blade model (E3) and demonstrates the capability of the program to extend for other cases as well.

The growing trend to achieve a higher turbine inlet temperature (TIT) in the modern gas turbine industry requires a more efficient and advanced cooling system design. Therefore, a complete study of heat transfer is necessary to predict the thermal loadings on the gas turbine vanes and blades. In the current work, a predictive model for the gas turbine blade cooling analysis has been developed. The model is capable of calculating the distribution of coolant mass flow rate (MFR) and metal temperatures of a turbine blade using the mass and energy balance equations at given external and internal boundary conditions. Initially, the performance of the model is validated by demonstrating its capability to predict the temperature distributions for a NASA E3 blade. The model is capable of predicting the temperature distributions with reasonable accuracy, especially on the suction side (SS). Later, this paper documents the overall analysis for the same blade profile but at different boundary conditions to demonstrate the flexibility of the model for other cases. Additionally, guidelines are provided to obtain external heat transfer coefficient (HTC) distributions for the highly turbulent mainstream.

Turbine vanes are typically assembled as a section containing single or double airfoil units in an annular pattern. First stage guide vane assembly results in two common mating interfaces: a gap between combustor and vane endwall and another resulted from the adjacent sections, called slashface. High pressure coolant could leak through these gaps to reduce the ingestion of hot gas and achieve certain cooling benefit. As vane endwall region flow field is already very complicated due to highly three-dimensional secondary flows, then a significant influence on endwall cooling can be expected due to the gap leakage flows. To determine the effect of leakage flows from those gaps, film cooling effectiveness distributions were measured using pressure sensitive paint (PSP) technique on the endwall of a scaled up, midrange industrial turbine vane geometry with the multiple rows of discrete film cooling (DFC) holes inside the passages. Experiments were performed in a blow-down wind tunnel cascade facility at the exit Mach number of 0.5 corresponding to Reynolds number of 3.8 × 105 based on inlet conditions and axial chord length. Passive turbulence grid was used to generate free-stream turbulence (FST) level about 19% with an integral length scale of 1.7 cm. Two parameters, coolant-to-mainstream mass flow ratio (MFR) and density ratio (DR), were studied. The results are presented as two-dimensional film cooling effectiveness distribution on the vane endwall surface with the corresponding spanwise averaged values along the axial direction.