The paper presents a detailed experimental and numerical study on the effect of endwall contouring in a quasi 2D cascade, operating at transonic conditions. Aerodynamic performance of two contoured endwalls are studied and compared with a baseline (planar) endwall. The first contoured endwall was generated with the goal of reducing secondary losses (Aero-Optimized contoured endwall) and the second endwall was generated with the objective of reduced overall heat transfer to the endwall (HT-optimized contoured endwall). Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface were measured. The cascade exit Mach numbers range from 0.71 to 0.95 and the turning angle of the airfoil is ∼127°. The inlet span of the airfoils was reduced with respect to the outlet span with the intention of obtaining a realistic inlet/exit Mach number that is observed in a real engine. 3D viscous compressible CFD analysis was carried out to study the detailed behavior of the complex flow structures that develop as a result of endwall contouring. A 3% reduction in area averaged losses was achieved at 0.1 Cax downstream of the trailing edge and a 17% reduction in mixed out losses was achieved at 1.0 Cax downstream location with the Aero-Optimized contoured endwall.
Contouring of turbine endwalls has been widely studied for aerodynamic performance improvement of turbine passages. However, it is equally important to investigate the effect of contouring on endwall heat transfer, because a substantial increase in endwall heat transfer due to contouring will render the design impractical. In this paper, the effect of contouring on endwall heat transfer performance of a high-turning HP-turbine blade passage, operating under transonic exit Mach number conditions, is reported. Three endwall geometries were experimentally investigated at three different passage exit Mach numbers, 0.71, 0.88(design) and 0.95, for their heat transfer performance. One endwall is a non-contoured baseline endwall and the other two are contoured endwall geometries. One of the contoured endwall geometry was generated with the goal of reduction in stagnation pressure losses and the other was generated with the goal of reduced overall heat transfer through the endwall. The experiments were carried out in Virginia Tech’s transient, blow down, transonic linear cascade facility. Endwall surface temperatures were measured using infrared thermography technique. Local heat transfer coefficient values were calculated using the measured temperatures. The heat transfer coefficient values were then related to the endwall geometries using a camera matrix model. The measurement technique and the methodology for the post-processing of the heat transfer coefficient data have been presented in detail. Details of the flow behavior for these endwalls were obtained using CFD simulations and have been used to assist the interpretation of the experimental results. In this study, the heat transfer performance of the contoured endwalls in comparison to the non-contoured baseline case is presented. Both the contoured endwalls demonstrated a significant reduction in the overall average heat transfer coefficient values. The surface heat transfer coefficient distributions also indicated a reduction in the level of hot spots for most of the endwall surface. However, increase in the heat transfer coefficient values was observed especially in the area near the leading edge. The results indicate that, in addition to a probable improvement in aerodynamic performance, endwall contouring may also be used to improve the heat transfer performance of turbine passages. Additionally, aerodynamic behavior of these endwalls is discussed in detail in the companion paper GT2012-68425, “Effect of endwall contouring on a transonic turbine endwall: Part 1 – Aerodynamic performance.”
Effect of turbine endwall contouring on its aerodynamic performance has been widely studied, but only a few studies are available in the open literature investigating its effect on heat transfer performance; especially at transonic exit Mach number conditions. In this paper, we report a study of effect of contouring on endwall heat transfer performance of a high-turning high-pressure (HP) turbine blade passage operating under transonic exit conditions. The paper describes comparison of heat transfer performance of two contoured endwall geometries, one aerodynamically optimized (AO) and the other heat transfer optimized (HTO), with a baseline, noncontoured geometry. The endwall geometries were experimentally investigated at Virginia Tech's transient, blow down, transonic linear cascade facility at three exit Mach numbers, Mex= 0.71, 0.88(design) and 0.95, for their heat transfer performance. Endwall surface temperatures were measured using infrared (IR) thermography and local heat transfer coefficient (HTC) values were calculated using measured temperatures. A camera matrix model-based data postprocessing technique was developed to relate the two-dimensional images captured by IR camera to three-dimensional endwall contours. The measurement technique and the methodology for postprocessing of the heat transfer coefficient data have been presented in detail. Discussion and interpretation of experimental results have been augmented using aerodynamic CFD simulations of the geometries. Both the contoured endwalls demonstrated a significant reduction in the overall average heat transfer coefficient values of the order of 10%. The surface Stanton number distributions also indicated a reduction in the level of hot spots for most of the endwall surface. However, at some locations an increase was also observed, especially in the area near the leading edge (LE). The results indicate that the endwall contouring could significantly improve heat transfer performance of turbine passages.
End wall contouring has been widely studied during past two decades for secondary loss reduction in turbine passages. Recent non-axisymmetric end wall contouring methods have shown more promise for loss reduction as compared to the axisymmetric end wall contouring methods used in initial studies. The end wall contouring methods have shown definite promise, especially, for the turbine passages at low design exit Mach numbers. A class of methods exists in the literature where the end wall surface is defined by using a combination of two curves. These curves specify surface topology variation in streamwise and pitchwise directions. Another class of methods depends on surface contour optimization, in which the modification of surface contours is achieved by changing the control point locations that define the surface topology. A definitive, passage design parameter based method of contouring is still not available. However, a general guideline for the trend of contour variation, along pitchwise and streamwise direction, can certainly be extrapolated from the existing literature. It is not clear, however, whether such a trend can be fitted to any blade profile to achieve, least of all a nonoptimum but a definite, reduction in losses. Moreover, almost all of the existing studies have focused on end wall contouring of passages with low exit Mach numbers. Some researchers, indeed, have used blades designed for high turning and high exit Mach number. However, such studies were done at Mach number well below the intended design condition. A study of effect of end wall contouring on a high turning blade with high design exit Mach number is not available in open literature. The present study investigates the effect of application of three different types of end wall contouring methods through numerical simulation, on a high turning transonic turbine blade passage. The main contouring method is based on total loss reduction criterion which is described here in detail. The contouring methodology described here avoids the deficiency of current commercial mesh generation software in context of automated meshing and provides a robust end wall optimization methodology. The geometry that gives minimum SKE values is compared with this loss optimized geometry. Additionally, a normalized contoured surface topology was extracted from a previous study that has similar blade design parameters and this surface was fitted to the turbine passage under study in order to investigate the effect of such trend based surface fitting. This contour geometry has also been compared with the other two contour geometries. Aerodynamic response of these geometries has been compared in detail with the baseline case without any end wall contouring. A comparison of shape and location of end wall contours on aerodynamic performance has been provided. The results indicate that end wall contouring for transonic turbine blades may not result in as significant gains at design conditions as those claimed for low speed turbine passages in previous studies.
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