J. D. Piggush

Bio: J. D. Piggush is an academic researcher from University of Minnesota. The author has contributed to research in topics: Turbine & Leakage (electronics). The author has an hindex of 4, co-authored 4 publications receiving 153 citations. Previous affiliations of J. D. Piggush include Pratt & Whitney.

Turbine Endwall Aerodynamics and Heat Transfer

Abstract: This review addresses recent literature on turbine passage aerodynamics and endwall heat transfer; articles that describe the endwall flow and cooling problems are summarized, recent activity on improving endwall aerothermal design is discussed, improved cooling schemes are proposed, and methods for managing secondary flows to allow more effective cooling are suggested. Much attention is given to aerodynamic losses associated with secondary flows developed near the endwalls. The endwall region flowfield is influenced by the stagnation zones established as the endwall approach flow boundary layer meets the airfoil leading edges, by the curvature of the passages, by the steps and gaps on the endwall surface ahead of and within the passage, by the leakage and coolant flows introduced through the endwall surface ahead of and within the passage, by the tip leakage flows between the blades and shroud in the rotor endwall region, and by many more effects. Recent combustor redesigns have flattened the turbine inlet temperature profile and have raised the turbine inlet temperatures. This, coupled with a continued need to improve engine durability and availability, has spurred strong interest in thermal control of the turbine endwall regions. Thus, much of the literature presented herein is focused on endwall cooling and, in particular, the effects of near-endwall secondary flows on endwall cooling.

Flow Measurements in a First Stage Nozzle Cascade Having Endwall Contouring, Leakage, and Assembly Features

Abstract: This work supports new gas turbine designs for improved performance by evaluating the use of endwall contouring in a cascade that is representative of a first stage stator passage. Contouring accelerates the flow, reducing the thickness of the endwall inlet boundary layer to the turbine stage and reducing the strength of secondary flows within the passage. Reduction in secondary flows leads to less mixing in the endwall region. This allows improved cooling of the endwall and airfoil surfaces with injected and leakage flows. The present paper documents component misalignment and leakage flow effects on the aerodynamic losses within a passage having one contoured and one straight endwall. Steps on the endwall and leakage flows through the endwall can lead to thicker endwall boundary layers, stronger secondary flows and possibly additional vortex structures in the passage. The paper compares losses with steps of various geometries and leakage of various flow rates to assess their importance on aerodynamic losses in this contoured passage. In particular, features associated with the combustor-to-turbine transition piece and the slashface gap, a gap between two vane segments on the vane platform, are addressed. An n-factorial study is used to quantify the importance of such effects on aerodynamic losses.Copyright © 2005 by ASME

Heat transfer measurements in a first stage nozzle cascade having endwall contouring, leakage and assembly features

Abstract: This work supports new gas turbine designs for improved performance by evaluating sealing flow effects in a cascade representative of a contoured first stage stator passage. Contouring accelerates the flow, reducing the thickness of the endwall inlet boundary layer to the turbine stage and reducing the strength of secondary flows within the passage. Injected flows, used to seal gaps and cool surfaces, may affect endwall boundary layers, increase secondary flows and possibly create additional vortex structures in the passage. The present paper documents injected flow effects on the endwall heat transfer within a passage with one contoured and one straight endwall. The paper discusses heat transfer distributions measured with different leakage flow rates. In particular, leakage is from the gap between the combustor and turbine sections and from the gap at the assembly joint on the vane platform between two vanes.© 2005 ASME

Flow Measurements in a First Stage Nozzle Cascade Having Leakage and Assembly Features: Effects of Endwall Steps and Leakage on Aerodynamic Losses

Abstract: This work supports new gas turbine designs for improved performance by evaluating endwall leakage and assembly features in a cascade that is representative of a first stage stator passage. The present paper documents component misalignment and leakage flow effects on the aerodynamic losses within a passage having one contoured and one straight endwall. Steps on the endwall and leakage flows through the endwall can lead to thicker endwall boundary layers, stronger secondary flows and possibly additional vortex structures in the passage. The paper compares losses with steps of various geometries and leakage of various flow rates to assess their importance on aerodynamic losses in this contoured passage. In particular, features associated with the combustor-to-turbine transition piece and the slash-face gap, a gap between two vane segments on the vane platform, are addressed.Copyright © 2005 by ASME

Gas Turbine Heat Transfer: Ten Remaining Hot Gas Path Challenges

Abstract: The advancement of turbine cooling has allowed engine design to exceed normal material temperature limits, but it has introduced complexities that have accentuated the thermal issues greatly. Cooled component design has consistently trended in the direction of higher heat loads, higher through-wall thermal gradients, and higher in-plane thermal gradients. The present discussion seeks to identify ten major thermal issues, or opportunities, that remain for the turbine hot gas path (HGP) today. These thermal challenges are commonly known in their broadest forms, but some tend to be little discussed in a direct manner relevant to gas turbines. These include uniformity of internal cooling, ultimate film cooling, microcooling, reduced incident heat flux, secondary flows as prime cooling, contoured gas paths, thermal stress reduction, controlled cooling, low emission combustor-turbine systems, and regenerative cooling. Evolutionary or revolutionary advancements concerning these issues will ultimately be required in realizable engineering forms for gas turbines to breakthrough to new levels of performance. Herein lies the challenge to researchers and designers. It is the intention of this summary to provide a concise review of these issues, and some of the recent solution directions, as an initial guide and stimulation to further research.

Turbine Endwall Aerodynamics and Heat Transfer

Abstract: This review addresses recent literature on turbine passage aerodynamics and endwall heat transfer; articles that describe the endwall flow and cooling problems are summarized, recent activity on improving endwall aerothermal design is discussed, improved cooling schemes are proposed, and methods for managing secondary flows to allow more effective cooling are suggested. Much attention is given to aerodynamic losses associated with secondary flows developed near the endwalls. The endwall region flowfield is influenced by the stagnation zones established as the endwall approach flow boundary layer meets the airfoil leading edges, by the curvature of the passages, by the steps and gaps on the endwall surface ahead of and within the passage, by the leakage and coolant flows introduced through the endwall surface ahead of and within the passage, by the tip leakage flows between the blades and shroud in the rotor endwall region, and by many more effects. Recent combustor redesigns have flattened the turbine inlet temperature profile and have raised the turbine inlet temperatures. This, coupled with a continued need to improve engine durability and availability, has spurred strong interest in thermal control of the turbine endwall regions. Thus, much of the literature presented herein is focused on endwall cooling and, in particular, the effects of near-endwall secondary flows on endwall cooling.

Film-Cooling Effectiveness on a Rotating Blade Platform

Abstract: Film cooling effectiveness measurements under rotation were performed on the rotor blade platform using a pressure sensitive paint (PSP) technique. The present study examines, in particular, the film cooling effectiveness due to purging of coolant from the wheel-space cavity through the circumferential clearance gap provided between the stationary and rotating components of the turbine. The experimental investigation is carried out in a new three-stage turbine facility, recently designed and taken into operation at the Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. This new turbine rotor has been used to facilitate coolant injection through this stator-rotor gap upstream of the 1st stage rotor blade. The gap was inclined at 25° to mainstream flow to allow the injected coolant to form a film along the passage platform. The effects of turbine rotating conditions on the blade platform film cooling effectiveness were investigated at three speeds of 2550 rpm, 2000 rpm and 1500 rpm with corresponding incidence angles of 23.2°, 43.4° and 54.8° respectively. Four different coolant-to-mainstream mass flow ratios varying from 0.5% to 2.0% were tested at each rotational speed. Aerodynamic measurements were performed at the 1st stage stator exit using a radially traversed five-hole probe to quantify the mainstream flow at this station. Results indicate that film cooling effectiveness increases with an increase in the coolant-to-mainstream mass flow ratios for all turbine speeds. Higher turbine rotation speeds show more local film cooling effectiveness spread on the platform with increasing magnitudes.Copyright © 2006 by ASME

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

Abstract: This paper is focused on the effect of film hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film cooling holes are divided into four rows and loosely distributed on the platform. Four film cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film cooling effectiveness distributions on the platform was obtained using pressure sensitive paint (PSP) technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.Copyright © 2008 by ASME

The Effects of Varying the Combustor-Turbine Gap

Abstract: To protect hot turbine components, cooler air is bled from the high pressure section of the compressor and routed around the combustor where it is then injected through the turbine surfaces. Some of this high pressure air also leaks through the mating gaps formed between assembled turbine components where these components experience expansions and contractions as the turbine goes through operational cycles. This study presents endwall adiabatic effectiveness levels measured using a scaled up, two-passage turbine vane cascade. The focus of this study is evaluating the effects of thermal expansion and contraction for the combustor-turbine interface. Increasing the mass flow rate for the slot leakage between the combustor and turbine showed increased local adiabatic effectiveness levels while increasing the momentum flux ratio for the slot leakage dictated the coverage area for the cooling. With the mass flow held constant, decreasing the combustor-turbine interface width caused an increase in uniformity of coolant exiting the slot, particularly across the pressure side endwall surface. Increasing the width of the interface had the opposite effect thereby reducing coolant coverage on the endwall surface.