Santosh Abraham

Experimental and Numerical Investigation of Convective Heat Transfer in a Gas Turbine Can Combustor

Experiments and numerical computations are performed to investigate the convective heat transfer characteristics of a gas turbine can combustor under cold flow conditions in a Reynolds number range between 50,000 and 500,000 with a characteristic swirl number of 0.7. It is observed that the flow field in the combustor is characterized by an expanding swirling flow, which impinges on the liner wall close to the inlet of the combustor. The impinging shear layer is responsible for the peak location of heat transfer augmentation. It is observed that as Reynolds number increases from 50,000 to 500,000, the peak heat transfer augmentation ratio (compared with fully developed pipe flow) reduces from 10.5 to 2.75. This is attributed to the reduction in normalized turbulent kinetic energy in the impinging shear layer, which is strongly dependent on the swirl number that remains constant at 0.7 with Reynolds number. Additionally, the peak location does not change with Reynolds number since the flow structure in the combustor is also a function of the swirl number. The size of the corner recirculation zone near the combustor liner remains the same for all Reynolds numbers and hence the location of shear layer impingement and peak augmentation does not change.

Effect of Endwall Contouring on a Transonic Turbine Blade Passage: Part 1—Aerodynamic Performance

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.Copyright

Effect of Endwall Contouring on a Transonic Turbine Blade Passage: Part 2—Heat Transfer Performance

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.”Copyright

Measurement of Aerodynamic Losses for Turbine Airfoil Cascades With Varying Pitch, Operating Under Transonic Conditions

The paper presents detailed experimental results of the midspan total pressure losses, secondary flow field, and static pressure measurements on two linear, high-turning turbine cascades at transonic conditions. The airfoils in the two cascades being studied are identical and their aerodynamic loading levels are varied by increasing the pitch of one cascade by 25% with respect to the other. The turbine cascades are referred to as B1-SP and B1-IP. Cascade B1-IP, with its increased pitch, has a Zweifel coefficient increased by 25%. The airfoils have a turning angle of ∼127°. Measurements are made at design and off-design conditions, at exit Mach numbers ranging from 0.71 to 0.95. The exit span of the airfoils are increased relative to the inlet span with the intention of obtaining a ratio of inlet Mach number to exit Mach number that is representative to that encountered in a real engine. This results in one end wall diverging from inlet to exit at a 13 degree angle, which simulates the required leading edge loading as seen in an engine. The objective of this study is to investigate the variation in airfoil loading distribution and the effect it has on aerodynamic performance in terms of pressure losses. Detailed loss measurements, both in the pitchwise as well as spanwise directions, are conducted at 0.1 Cax and 1.0 Cax locations downstream of the trailing edge. Results from 3D viscous numerical simulations have been used to assist the interpretation of experimental results.Copyright

Experimental and Numerical Investigations of a Transonic, High Turning Turbine Cascade With a Divergent Endwall

The paper presents detailed measurements of midspan total pressure loss, secondary flow field, static pressure measurements on airfoil surface at midspan, near hub and near the end walls in a transonic turbine airfoil cascade. Numerous low-speed experimental studies have been carried out to investigate the performance of turbine cascades. Profile and secondary loss correlations have been developed and improved over the years to include the induced incidence and leading edge geometry and to reflect recent trends in turbine design. All of the above investigations have resulted in better understanding of flow field in turbine passages. However, there is still insufficient data on the performance of turbine blades with high turning angles operating at varying incidences angles at transonic Mach numbers. In the present study, measurements were made at +10, 0 and −10 degree incidence angles for a high turning turbine airfoil with 127 degree turning. The exit Mach numbers were varied within a range from 0.6 to 1.1. Additionally, the exit span is increased relative to the inlet span resulting in one end wall diverging from inlet to exit at 13 degree angle. This was done in order to obtain a ratio of inlet Mach number to exit Mach number which is representative to that encountered in real engine and simulates the blade and near end wall loading that is seen in an engine. 3D viscous compressible CFD analysis was carried out in order to compare the results with experimentally obtained values and to further investigate the design and off-design flow characteristics of the airfoil under study. All aerodynamic measurements were compared with CFD analysis and a reasonably good match was observed.Copyright

Effect of Airfoil Shape and Turning Angle on Turbine Airfoil Aerodynamic Performance at Transonic Conditions

Performance data for high turning gas turbine blades under transonic Mach numbers is significantly lacking in literature. Performance of three gas turbine airfoils with varying turning angles at transonic flow conditions was investigated in this study. Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface in a linear cascade setting were measured. Airfoil curvature and true chord were varied to change the loading vs. chord for each airfoil. Airfoils A, D and E are designed to operate at different velocity triangles. Velocity triangle requirements (inlet/exit Mach number and gas angles) come from 1D and 2D models that include calibrated loss systems. One of the goals of this study was to use the experimental data to confirm/refine loss predictions for the effect of various Mach numbers and gas turning angles. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. The airfoil turning angle ranges from 120° to 138°. A realistic inlet/exit Mach number ratio, that is representative of that seen in a real engine, was obtained by reducing the inlet span with respect to the exit span of the airfoil, thereby creating a quasi 2D cascade. In order to compare the experimental results and study the detailed flow characteristics, 3D viscous compressible CFD analysis was also carried out.Copyright

Effect of Turbine Airfoil Shape on Aerodynamic Losses for Turbine Airfoils Operating Under Transonic Conditions

Profile and secondary loss correlations have been developed and improved over the years to include the induced incidence and leading edge geometry and to reflect recent trends in turbine design. All of these investigations have resulted in better understanding of the flow field in turbine passages. However, there is still insufficient data on the performance of turbine airfoils with high turning angles operating at varying incidence angles at transonic Mach numbers. The paper presents detailed aerodynamic measurements for three different turbine airfoils with similar turning angles but different aerodynamic shapes. Midspan total pressure loss, secondary flow field, and static pressure measurements on the airfoil surface in the cascades are presented and compared for the three different airfoil sets. The airfoils are designed for the same velocity triangles (inlet/exit gas angles and Mach number). Airfoil curvature and true chord are varied to change the loading vs. chord. The objective is to investigate the type of loading distribution and its effect on aerodynamic performance (pressure loss). Measurements are made at +10, 0 and −10 degree incidence angles for high turning turbine airfoils with ∼127 degree turning. The cascade exit Mach numbers were varied within a range from 0.6 to 1.1. In order to attain a ratio of inlet Mach number to exit Mach number that is representative to that encountered in a real engine, the exit span is increased relative to the inlet span. This results in one end wall diverging from inlet to exit at a 13 degree angle, which simulates the required leading edge loading as seen in an engine. 3D viscous compressible CFD analysis was carried out in order to compare the results with experimentally obtained values and to further investigate the flow characteristics of the airfoils under study.Copyright

Investigation of Effect of End Wall Contouring Methods on a Transonic Turbine Blade Passage

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.© 2011 ASME

Film Cooling Study of Novel Orthogonal Entrance and Shaped Exit Holes

Film cooling effectiveness depends on several geometrical parameters like location on the airfoil, exit shape, orientation and arrangement of the holes. The focus of this investigation is to propose and explore a new film cooling hole geometry. The adiabatic film cooling effectiveness is determined experimentally, downstream of the exit of the film cooling holes on a flat plate using a steady state IR thermography technique. Coolant holes that are perpendicular to the direction of flow detach from the surface and enhance the heat transfer coefficient on the turbine blade without providing any coolant coverage, while angled holes along the mainstream direction result in superior film cooling effectiveness and lower heat transfer to the surface. The objective of this study is to examine the external cooling effects using coolant holes that are a combination of both angled shaped holes as well as perpendicular holes. The inlet of the coolant hole is kept perpendicular to the direction of flow to enhance the internal side heat transfer coefficient and the exit of the coolant hole is expanded and angled along the mainstream flow to prevent the coolant jet from lifting off from the blade external surface. A total of six different cases with variations in exit shape geometry are investigated at different blowing ratios (BR varying from 0.5 to 2.0). Results suggest that the film cooling effectiveness values obtained from these geometries are comparable with those of conventional angled holes. With the added advantage of enhanced heat transfer coefficient on the coolant channel internal side, as proven earlier by Byerley [3], overall superior cooling is accomplished. Furthermore this shaped hole can be made using the same technology being presently used in the industry.Copyright