Gas Turbine Heat Transfer: 10 Remaining Hot Gas Path Challenges

2006 ◽  
Author(s):  
Ronald S. Bunker
Keyword(s):  
Stress Reduction ◽  
Film Cooling ◽  
Gas Turbines ◽  
Secondary Flows ◽  
Internal Cooling ◽  
Thermal Gradients ◽  
Controlled Cooling ◽  
Hot Gas ◽  
Component Design ◽  
Gas Turbine Heat Transfer

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 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, micro cooling, 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.

Gas Turbine Heat Transfer: Ten Remaining Hot Gas Path Challenges

2006 ◽  
Vol 129 (2) ◽  
pp. 193-201 ◽  
Author(s):  
Ronald S. Bunker
Keyword(s):  
Stress Reduction ◽  
Film Cooling ◽  
Gas Turbines ◽  
Secondary Flows ◽  
Internal Cooling ◽  
Thermal Gradients ◽  
Controlled Cooling ◽  
Hot Gas ◽  
Component Design ◽  
Gas Turbine Heat Transfer

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.

The Influence of Turbulence and Reynolds Number on Multiple Slot Film Cooling Over the Suction Surface

2020 ◽  
Author(s):  
L. W. Soma ◽  
F. E. Ames ◽  
S. Acharya
Keyword(s):  
Film Cooling ◽  
Surface Film ◽  
Secondary Flows ◽  
Internal Cooling ◽  
Suction Surface ◽  
Film Cooling Effectiveness ◽  
Cooling Method ◽  
Cooling Section ◽  
Slot Film Cooling ◽  
Cooling Air

Abstract Developing robust film cooling protection on the suction surface of a vane is critical to managing the high heat loads which exist there. Suction surface film cooling often produces high levels of film cooling but can be influenced by secondary flows and some dissipation due to free-stream turbulence. Directly downstream from suction surface film cooling, heat loads are often significantly mitigated and internal cooling levels can be modest. One thermodynamically efficient way to cool the suction surface of a vane is with a counter cooling scheme. This combined internal/external cooling method moves cooling air in a direction opposite to the external flow through an internal convection array. The coolant is then discharged upstream where the high level of film cooling can offset the reduced cooling potential of the spent cooling air. The present suction surface film cooling arrangement combines a slot film cooling discharge on the near suction surface from an incremental impingement cooling method with a second from a counter cooling section. A second counter cooling section is added further downstream on the suction surface. The internal cooling plenums replicate the geometry of the cooling methods to ensure the fluid dynamics of the flow discharging from the slots are representative of the actual internal cooling geometry. These film cooling flows have been tested at blowing ratios of 0.5 and 1.0 for the initial slot and blowing ratios of 0.15 and 0.3 for the two downstream slots. The measurements have been taken at exit chord Reynolds numbers of 500,000, 1,000,000, and 2,000,000 with inlet turbulence levels ranging from 0.7% to 12.6%. Film cooling effectiveness measurements were acquired using both thermocouples and infrared thermography. The infrared thermography shows the influence of secondary flows on film cooling coverage near the suction surface endwall junction. The film cooling effectiveness results at varied blowing ratios, turbulence levels and Reynolds numbers document the impact of these major variables on suction surface slot film cooling. The results provide a consistent picture of the slot film cooling for the present three slot arrangement on the suction surface and they support the development of an advanced double wall cooling method.

Temperature-Based Optimization of Film Cooling in Gas Turbine Hot Gas Path Components

2013 ◽  
Author(s):  
Felipe A. C. Viana ◽  
Jack Madelone ◽  
Niranjan Pai ◽  
Genghis Khan ◽  
Sanghum Baik
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Optimization Design ◽  
Goal Attainment ◽  
High Efficiency ◽  
Cooling System ◽  
Computational Cost ◽  
Metal Temperature ◽  
Cooling Flow ◽  
Hot Gas

To achieve high efficiency, modern gas turbines operate at temperatures that exceed melting points of metal alloys used in turbine hot gas path parts. Parts exposed to hot gas are actively cooled with a portion of the compressor discharge air (e.g., through film cooling) to keep the metal temperature at levels needed to meet durability requirements. However, to preserve efficiency, it is important to optimize the cooling system to use the least amount of cooling flow. In this study, film cooling optimization is achieved by varying cooling hole diameters, hole to hole spacing, and film row placements so that the specified targets for maximum metal temperature are met while preserving (or saving) cooling flow. The computational cost of the high-fidelity physics models, the large number of design variables, the large number and nonlinearity of responses impose severe challenges to numerical optimization. Design of experiments and cheap-to-evaluate approximations (radial basis functions) are used to alleviate the computational burden. Then, the goal attainment method is used for optimizing of film cooling configuration. The results for a turbine blade design show significant improvements in temperature distribution while maintaining/reducing the amount of used cooling flow.

Performance of Public Film Cooling Geometries Produced Through Additive Manufacturing

2020 ◽  
Vol 142 (5) ◽  
Author(s):  
Jacob C. Snyder ◽  
Karen A. Thole
Keyword(s):  
Additive Manufacturing ◽  
Film Cooling ◽  
Gas Turbines ◽  
Nickel Alloys ◽  
Cooling Effectiveness ◽  
Powder Bed ◽  
The Public ◽  
Hot Gas ◽  
Cooling Hole ◽  
Cooling Technology

Abstract Film cooling is an essential cooling technology to allow modern gas turbines to operate at high temperatures. For years, researchers in this community have worked to improve the effectiveness of film cooling configurations by maximizing the coolant coverage and minimizing the heat flux from the hot gas into the part. Working toward this goal has generated many promising film cooling concepts with unique shapes and configurations. However, until recently, many of these designs were challenging to manufacture in actual turbine hardware due to limitations with legacy manufacturing methods. Now, with the advances in additive manufacturing, it is possible to create turbine parts using high-temperature nickel alloys that feature detailed and unique geometry features. Armed with this new manufacturing power, this study aims to build and test the promising designs from the public literature that were previously difficult or impossible to implement. In this study, different cooling hole designs were manufactured in test coupons using a laser powder bed fusion process. Each nickel alloy coupon featured a single row of engine scale cooling holes, fed by a microchannel. To evaluate performance, the overall cooling effectiveness of each coupon was measured using a matched Biot test at engine relevant conditions. The results showed that certain hole shapes are better suited for additive manufacturing than others and that the manufacturing process can cause significant deviations from the performance reported in the literature.

Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
M. D. Barringer ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Secondary Flows ◽  
Pressure Side ◽  
Surface Deposition ◽  
Hot Gas ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes ◽  
Effect Of Surface

Film cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film cooling along the pressure side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the pressure side of the airfoil. Deposits having a semielliptical shape were placed along the pressure side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets toward the vane endwall junction on the pressure side. Results also indicated that there was a steady decrease in adiabatic effectiveness levels with a sequential increase in the number of rows with the deposits.

Overall Effectiveness and Flowfield Measurements for an Endwall With Nonaxisymmetric Contouring

2015 ◽  
Vol 138 (3) ◽  
Author(s):  
Amy Mensch ◽  
Karen A. Thole
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Transfer Model ◽  
Internal Cooling ◽  
Blowing Ratio ◽  
Passage Vortex ◽  
Endwall Contouring ◽  
Overall Effectiveness

Endwall contouring is a technique used to reduce the strength and development of three-dimensional secondary flows in a turbine vane or blade passage in a gas turbine. The secondary flows locally affect the external heat transfer, particularly on the endwall surface. The combination of external and internal convective heat transfer, along with solid conduction, determines component temperatures, which affect the service life of turbine components. A conjugate heat transfer model is used to measure the nondimensional external surface temperature, known as overall effectiveness, of an endwall with nonaxisymmetric contouring. The endwall cooling methods include internal impingement cooling and external film cooling. Measured values of overall effectiveness show that endwall contouring reduces the effectiveness of impingement alone, but increases the effectiveness of film cooling alone. Given the combined case of both impingement and film cooling, the laterally averaged overall effectiveness is not significantly changed between the flat and the contoured endwalls. Flowfield measurements indicate that the size and location of the passage vortex changes as film cooling is added and as the blowing ratio increases. Because endwall contouring can produce local effects on internal cooling and film cooling performance, the implications for heat transfer should be considered in endwall contour designs.

scholarly journals Effects of Crossflow in an Internal-Cooling Channel on Film Cooling of a Flat Plate Through Compound-Angle Holes

2015 ◽  
Author(s):  
Zachary T. Stratton ◽  
Tom I-P. Shih ◽  
Gregory M. Laskowski ◽  
Brian Barr ◽  
Robert Briggs
Keyword(s):  
Flat Plate ◽  
Film Cooling ◽  
Gas Flow ◽  
Density Ratio ◽  
Internal Cooling ◽  
Cfd Simulations ◽  
Hot Gas ◽  
Adiabatic Effectiveness ◽  
Cooling Passage ◽  
Internal Cooling Passage

CFD simulations were performed to study the film cooling of a flat plate from one row of compound-angles holes fed by an internal-cooling passage that is perpendicular to the hot-gas flow. Parameters examined include direction of flow in the internal cooling passage and blowing ratios of 0.5, 1.0, and 1.5 with the coolant-to-hot-gas density ratio kept at 1.5. This CFD study is based on steady RANS with the shear-stress transport (SST) and realizable k-ε turbulence models. To understand the effects of unsteadiness in the flow, one case was studied by using large-eddy simulation (LES). Results obtained showed an unsteady vortical structure forms inside the hole, causing a side-to-side shedding of the coolant jet. Values of adiabatic effectiveness predicted by CFD simulations were compared with the experimentally measured values. Steady RANS was found to be inconsistent in its ability to predict adiabatic effectiveness with relative error ranging for 10% to over 100%. LES was able to predict adiabatic effectiveness with reasonable accuracy.

Robust Optimization of the HPT Blade Cooling and Aerodynamic Efficiency

2016 ◽  
Author(s):  
Kirill A. Vinogradov ◽  
Gennady V. Kretinin ◽  
Kseniya V. Otryahina ◽  
Roman A. Didenko ◽  
Dmitry V. Karelin ◽  
...  
Keyword(s):  
Robust Optimization ◽  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
Operational Parameters ◽  
Cooling Effectiveness ◽  
Aerodynamic Efficiency ◽  
Hot Gas ◽  
Blade Tip

Constant rise of hot gas temperature is crucial for the creation of modern gas-turbines engines requiring considerable improvement of cooling configurations. A high pressure turbine blade is one of the most crucial and loaded details in gas-turbine engines. A HPT blade is affected by different operational deviations: stochastic fluctuations of inlet parameters and difference in operational parameters for manufactured engines. Combination of these factors makes the task of uncertainty quantification and robust optimization of the HPT blade relevant in modern science. The authors make an attempt to implement robust optimization to the HPT blade of the gas-turbine engine. The two most important areas of the cooling blade (the leading edge (LE) and the blade tip) were taken into account. The operational and the aleatoric uncertainties were analyzed. These uncertainties represent the fluctuations in the operational parameters and the random-unknown conditions such as the boundary values and or geometrical variations. Industrial HPT blade with a serpentary cooling system and film cooling at the LE was considered. Results of many engine tests were applied to construct probability density function distributions for operational uncertainties. More than 100 real gas-turbines were examined. The following operational uncertainties were reviewed: inlet hot gas pressure and temperature together with cooling air pressure. The tip gap was used as geometrical variation. Conjugate Heat Transfer computations were carried out for the temperature distribution obtained. Geometrical variations of the LE film cooling rows and the tip gap are variables in the robust optimization process. The authors developed a special technology for full parameterization of the LE film-cooling rows only by two parameters. A surrogate model technique (the response surface and the Monte-Carlo method) was applied for the uncertainty quantification and the robust optimization processes. The IOSO technology was employed as one of the robust optimization tools. This technology is also based on the widespread application of the response surface technique. Robust optimal solution (the Pareto set) between cooling effectiveness of the leading edge and the blade tip and aerodynamic efficiency was obtained as the result. At chosen point from the Pareto set (angle point) we calculated necessary levels of robust criteria characterized LE and blade tip cooling effectiveness and kinetic energy losses.

A Method for Identifying and Visualizing Foreign Particle Motion Using Time Resolved Particle Tracking Velocimetry

2009 ◽  
Author(s):  
N. D. Cardwell ◽  
P. P. Vlachos ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Particle Tracking Velocimetry ◽  
Internal Cooling ◽  
Particle Collisions ◽  
Two Phase ◽  
Time Resolved ◽  
Image Velocimetry ◽  
Particle Ingestion ◽  
Film Cooling Holes

Gas turbines for aircraft are designed for operation with a clean inlet air flow. This ideal operational condition is often violated during take-off and landing, where the probability of particle ingestion is high with sand and dirt being the most commonly observed foreign particles. Current research on particle ingestion has identified several mechanisms that contribute to performance degradation in the turbine: erosion of internal and external surfaces; and flow blockages of film-cooling holes and internal cooling passages. The focus of the study given in this paper is to present a method that identifies the motion of foreign particles within an internal ribbed passage. The method uses a high-resolution, flowfield interrogation method known as Time-Resolved Digital Particle Image Velocimetry (TRDPIV). Observations from the two-phase flows showed that particle collisions occurred more frequently on the upstream surface of the ribs, especially in the inlet region. Results from these collisions included substantial particle breakup and a particle rebounding phenomenon between the upper and lower walls. Comparisons are made to LES predicted particle trajectories indicating some agreement, but also phenomena that are not predicted due to the inherent assumption of the modeling.

Effects of Endwall Film Coolant Flowrate on Secondary Flows and Coolant Mixing in a First Stage Nozzle Guide Vane

2021 ◽  
Vol 143 (3) ◽  
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Test Section ◽  
Film Cooling ◽  
Gas Turbines ◽  
Thermal Loading ◽  
Guide Vane ◽  
Secondary Flows ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract Modern gas turbines are subjected to very high thermal loading. This leads to a need for aggressive cooling to protect components from damage. Endwalls are particularly challenging to cool due to a complex system of secondary flows near them that wash and disrupt the protective coolant films. This highly three-dimensional flow not only affects but is also affected by the momentum of film cooling flows, whether injected just upstream of the passage to intentionally cool the endwall or as combustor cooling flows injected further upstream in the engine. This complex interaction between the different cooling flows and passage aerodynamics has been recently studied in a first stage nozzle guide vane. The present paper presents a detailed study on the sensitivity of aero-thermal interactions to endwall film cooling mass flow to mainstream flow ratio. The test section represents a first stage nozzle guide vane with a contoured endwall and endwall film cooling injected just upstream of it. The test section also includes an engine-representative combustor–turbine interface geometry with combustor cooling flows injected at a constant rate. The approach flow conditions represent flow exiting a low-NOx combustor. Adiabatic surface thermal measurements and in-passage velocity and thermal field measurements are presented and discussed. The results show the dynamics of passage vortex suppression and the increase of impingement vortex strength as MFR changes. The effects of these changes of secondary flows on coolant distribution are presented.

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