Experimental investigation of the interaction between showerhead coolant jets and main flow

This paper presents the results of an experimental investigation into the thermal and aerodynamic behavior of coolant ejection at the leading edge of a highly loaded nozzle vane cascade. The leading-edge cooling scheme features four rows of cylindrical holes in a staggered configuration (showerhead). Pressure Sensitive Paints (PSP) technique was used to get the adiabatic film cooling effectiveness distribution, while Particle Image Velocimetry (PIV) and flow visualizations were used to investigate the mixing process taking place between coolant and main flow. PSP tests were conducted by using N2 (Density Ratio DR=1.0) as coolant at variable blowing ratio (BR=2.0 \u2013 4.0). Further tests were run by using CO2 (DR=1.5) at matching BR and momentum flux ratio (I) in order to investigate the effects of density ratio. The BR = 3.0 injection case was selected for the PIV investigation. Thermal and flow field data consistently show a shift in the position of stagnation line towards the suction side. Jet liftoff close to stagnation and a strong jet to jet as well as jet to mainstream interaction were also observed, resulting in a complex 3D flow characterized by high turbulence levels with a high degree of anisotropy. No coherent structures were detected, supporting the random nature of mixing process

Investigations of Heat Transfer and Fluid Flow in the Pocket Region of a Gas Turbine Engine and Cooling of a Turbine Blade

In the present work, heat transfer within gas turbine applications are investigated bothexperimentally and numerically. The main content concerns heat transfer and fluidflow over the pocket region and cooling of a turbine blade.A pocket cavity is generated at the junction part of the low pressure turbine (LPT) andthe outlet guide vane (OGV) in the rear part of a gas turbine engine. The heat transferdistribution and fluid flow over the pocket cavity have significant effects on theincoming flow of the OGV placed downstream. These pocket cavities are built withdifferent radii to find out improved heat transfer distributions and flow patterns. Theeffects of a pocket cavity on heat transfer and flow characteristics on the endwall witha symmetric vane are also investigated. The relative location between the pocketcavity and the symmetric vane is varied. In addition, the effect of incoming flowattack angle of the pocket cavity upstream of an OGV is investigated numerically.Liquid Crystal Thermography (LCT) is employed to measure the heat transfer of thetested surfaces. The results show that the smaller fillet radius provides a higher heattransfer peak value with a stronger recirculating flow inside the pocket cavity. When apocket cavity is placed upstream of the symmetric vane, the high heat transfer areasaround the symmetric vane are decreased. The attack angles of the incoming flow overthe pocket cavity affect the forming of horseshoe vortices in leading edge of the vaneand then affect the heat transfer distribution.Rib turbulators are widely employed in internal cooling passages of a turbine blade.Firstly, truncated ribs with various truncation types and arrangements are considered.Secondly, perforated ribs with differently shaped penetration holes and perforationratios are investigated. LCT is employed to measure surface temperature and deriveheat transfer coefficients over the ribbed surfaces in the tested channels. The turbulentflow details are presented by numerical calculations with an established turbulencemodel, i.e., the k-ω SST model. From the results, the truncated ribs can reduce thepressure loss penalty without reducing the heat transfer enhancement. By changing theconfigurations to staggered arrangements, the heat transfer can be further enhancedassociated with a moderate pressure drop. By using perforated ribs, the low heattransfer regions downstream of the rib rows are greatly improved.Endwall film cooling is a significant cooling method to protect the endwall regionwhere the flow structures are complex due to horseshoe vortices and generatedsecondary flows. This study firstly concentrates on film cooling holes arrangedupstream of the leading edge of a turbine vane. Several arrangements are designedaiming at improving the coolant coverage. Based on the calculated results, the filmcooling holes upstream the leading edge have cooling effects on both the vanesurfaces and the endwall. A case with two rows of compound angle holes in staggeredarrangement shows relatively high overall averaged cooling effectiveness independentof the blowing ratios. Then full-scale endwall film cooling is also investigated in thisstudy. The film holes arrangements are designed based on the pressure coefficientdistribution, streamline distribution and heat transfer distribution on the endwall. Withcompound angle holes, the design based on the pressure distribution forces the flowsto the suction side, which creates benefits for cooling the vane surfaces. The designbased on the streamline distribution has more uniform coolant coverage on theendwall. The design based on the heat transfer distributions has relatively largecoolant coverage and is effective in removing the high temperature region.Keywords: pocket cavity, symmetric vane, Liquid Crystal Thermography, truncatedribs, staggered arrangement, perforated ribs, secondary flows, endwall film cooling,leading edge, turbine vane, coolant coverag

Thermal and Flow Field Investigations of a Micro-Tangential-Jet Film Cooling Scheme on Gas Turbine Components

Thermal and Flow Field Investigations of a Micro-Tangential-Jet Film Cooling Scheme on Gas Turbine Components Othman Hassan, Ph.D. Concordia University, 2013 Gas turbines play a major role in modern aerospace and in industrial power generation nowadays. Advanced gas turbines are designed to operate at increasingly higher inlet turbine gas temperature to increase their efficiency and specific power output. In order to enable this increase in the operating temperature, high-temperature resistant materials, Thermal-Barrier Coatings (TBCs), and advanced cooling techniques, are employed. Internal cooling, impingement cooling, and film cooling, are the typical cooling techniques that are being used nowadays for gas turbine engines cooling. For the past five decades, significant efforts have been implemented in the area of film cooling to design and investigate the performance of numerous cooling schemes at various operating conditions and geometries. However, the achieved effectiveness to date, especially over actual airfoil geometries, is still relatively low. Further efforts are essential to propose novel designs that are capable of providing the required cooling loads. The present study investigates the thermal performance and flow characteristics downstream a new film cooling scheme over a gas turbine vane and a flat plate. The state-of-the-art transient Thermochromic Liquid Crystal (TLC) technique has been employed for film cooling measurements, while the Particle Image Velocimetry (PIV) technique has been employed for flow field investigations. Validation of all measurement techniques were conducted and good agreement with literature works has been achieved. The Micro-Tangential-Jet (MTJ) scheme is a discrete-holes shaped cooling scheme with micro sized exit height that supplies the jet parallel to the surface. The MTJ scheme consists of two main parts, a circular supply micro-tube, and a shaped exit parallel to the vane surface. The shaped exit of the scheme starts with a circular cross section. Lateral expansion angles are then applied in both directions and a relatively constant height is maintained throughout the scheme yielding a squared exit. Due to the micro thickness of the jet, a deep penetration inside the main stream is achievable, while maintaining a tangential injection direction to the surface, thereby avoiding jet lift off. The film cooling performance of one row of MTJ scheme on the vane pressure side and another row on the suction side is investigated at different blowing ratios using the transient TLC technique. Comparisons with the film cooling performance of previously proposed shaped schemes are carried out to highlight the advantages and disadvantages of the new design. Mach number distributions over the airfoil surface are determined with and without the MTJ scheme to investigate the effect of the added material on the airfoil characteristics. A comprehensive analysis based on the current findings, previous efforts in the literature, and the flow field investigations using the PIV technique downstream the MTJ scheme is presented. Overall, the new design showed superior film cooling performance, compared to the best achieved results in literature. The effectiveness distribution downstream the MTJ scheme was characterized with superior lateral spreading over both pressure and suction surfaces. The measurements showed similarity in the characteristics of the 2-D film downstream the MTJ scheme and the one that accompanies the injection from continuous slot schemes. Moreover, the investigations showed that the presence of the MTJ scheme over the vane pressure or suction sides did not result in significant HTC augmentation, especially at blowing ratios less than unity. The MTJ scheme could be the first of a new generation of film cooling schemes over airfoil geometries

Fundamental Understanding of Interactions Among Flow, Turbulence, and Heat Transfer in Jet Impingement Cooling

The flow physics of impinging jet is very complex and is not fully understood yet. The flow field in an impingement problem comprised of three different distinct regions: a free jet with a potential core, a stagnation region where the velocity goes to zero as the jet impinges onto the wall and a creation of wall jet region where the boundary layer grows radially outward after impinging. Since impingement itself is a broad topic, effort is being made in the current study to narrow down on three particular geometric configurations (a narrow wall, an array impingement configuration and a curved surface impingement configuration) that shows up in a typical gas turbine impingement problem in relation to heat transfer. Impingement problems are difficult to simulate numerically using conventional RANS models. It is worth noting that the typical RANS model contains a number of calibrated constants and these have been formulated with respect to relatively simple shear flows. As a result typically these isotropic eddy viscosity models fail in predicting the correct heat transfer value and trend in impingement problem where the flow is highly anisotropic. The common RANS-based models over predict stagnation heat transfer coefficients by as much as 300% when compared to measured values. Even the best of the models, the v^2-f model, can be inaccurate by up to 30%. Even though there is myriad number of experimental and numerical work published on single jet impingement; the knowledge gathered from these works cannot be applied to real engineering impingement cooling application as the dynamics of flow changes completely. This study underlines the lack of experimental flow physics data in published literature on multiple jet impingement and the author emphasized how important it is to have experimental data to validate CFD tools and to determine the suitability of Large Eddy Simulation (LES) in industrial application. In the open literature there is not enough study where experimental heat transfer and flow physics data are combined to explain the behavior for gas turbine impingement cooling application. Often it is hard to understand the heat transfer behavior due to lack of time accurate flow physics data hence a lot of conjecture has been made to explain the phenomena. The problem is further exacerbated for array of impingement jets where the flow is much more complex than a single round jet. The experimental flow field obtained from Particle Image Velocimetry (PIV) and heat transfer data obtained from Temperature Sensitive Paint (TSP) from this work will be analyzed to understand the relationship between flow characteristics and heat transfer for the three types of novel geometry mentioned above. There has not been any effort made on implementing LES technique on array impingement problem in the published literature. Nowadays with growing computational power and resources CFD are widely used as a design tool. To support the data gathered from the experiment, LES is carried out in narrow wall impingement cooling configuration. The results will provide more accurate information on impingement flow physics phenomena where experimental techniques are limited and the typical RANS models yield erroneous result The objective of the current study is to provide a better understanding of impingement heat transfer in relation to flow physics associated with it. As heat transfer is basically a manifestation of the flow and most of the flow in real engineering applications is turbulent, it is very important to understand the dynamics of flow physics in an impingement problem. The work emphasis the importance of understanding mean velocities, turbulence, jet shear layer instability and its importance in heat transfer application. The present work shows detailed information of flow phenomena using Particle Image Velocimetry (PIV) in a single row narrow impingement channel. Results from the RANS and LES simulations are compared with Particle Image Velocimetry (PIV) data. The accuracy of LES in predicting the flow field and heat transfer of an impingement problem is also presented the in the current work as it is validated against experimental flow field measured through PIV. Results obtained from the PIV and LES shows excellent agreement for predicting both heat transfer and flow physics data. Some of the key findings from the study highlight the shortcomings of the typical RANS models used for the impingement heat transfer problem. It was found that the stagnation point heat transfer was over predicted by as much as 48% from RANS simulations when compared to the experimental data. A lot of conjecture has been made in the past for RANS\u27 ability to predict the stagnation point heat transfer correctly. The length of the potential core for the first jet was found to be ~ 2D in RANS simulations as oppose to 1D in PIV and LES, confirm the possible underlying reason for this discrepancy. The jet shear layer thickness was underpredicted by ~ 40% in RANS simulations proving the model is not diffusive enough for a flow like jet impingement. Turbulence production due to shear stress was over predicted by ~130% and turbulence production due to normal stresses were underpredicted by ~40 % in RANS simulation very close to the target wall showing RANS models fail where both strain rate and shear stress plays a pivotal role in the dynamics of the flow. In the closing, turbulence is still one of the most difficult problems to solve accurately, as has been the case for about a century. A quote below from the famous mathematician, Horace Lamb (1849-1934) express the level of difficulty and frustration associated with understanding turbulence in fluid mechanics. I am an old man now, and when I die and go to heaven there are two matters on which I hope for enlightenment. One is quantum electrodynamics, and the other is the turbulent motion of fluids. And about the former I am rather optimistic. Source: http://scienceworld.wolfram.com/biography/Lamb.html This dissertation is expected to shed some light onto one specific example of turbulent flows

Heat removal in axial flow high pressure gas turbine

The demand for high power in aircraft gas turbine engines as well as industrial gas turbine prime mover promotes increasing the turbine entry temperature, the mass flow rate and the overall pressure ratio. High turbine entry temperature is however the most convenient way to increase the thrust without requiring a large change in the engine size. This research is focused on improving the internal cooling of high pressure turbine blade by investigating a range of solutions that can contribute to the more effective removal of heat when compared with existing configuration. The role played by the shape of the internal blade passages is investigated with numerical methods. In addition, the application of mist air as a means of enhanced heat removal is studied. The research covers three main area of investigation. The first one is concerned with the supply of mist on to the coolant flow as a mean to enhancing heat transfer. The second area of investigation is the manipulation of the secondary flow through cross-section variation as a means to augment heat transfer. Lastly a combination of a number of geometrical features in the passage is investigated. A promising technique to significantly improve heat transfer is to inject liquid droplets into the coolant flow. The droplets which will evaporate after travelling a certain distance, act as a cooling sink which consequently promote added heat removal. Due to the promising results of mist cooling in the literature, this research investigated its effect on a roughened cooling passage with five levels of mist mass percentages. In order to validate the numerical model, two stages were carried out. First, one single-phase flow case was validated against experimental results available in the open literature. Analysing the effect of the rotational force, on both flow physics and heat transfer, on the ribbed channel was the main concern of this investigation. Furthermore, the computational results using mist injection were also validated against the experimental results available in the literature. Injection of mist in the coolant flow helped achieve up to a 300% increase in the average flow temperature of the stream, therefore in extracting significantly more heat from the wall. The Nusselt number increased by 97% for the rotating leading edge at 5% mist injection. In the case of air only, the heat transfers decrease in the second passage, while in the mist case, the heat transfer tends to increase in the second passage. Heat transfer increases quasi linearly with the increase of the mist percentage when there is no rotation. However, in the presence of rotation, the heat transfers increase with an increase in mist content up to 4%, thereafter the heat transfer whilst still rising does so more gradually. The second part of this research studies the effect of non-uniform cross- section on the secondary flow and heat transfer in order to identify a preferential design for the blade cooling internal passage. Four different cross-sections were investigated. All cases start with square cross-section which then change all the way until it reaches the 180 degree turn before it changes back to square cross-section at the outlet. All cases were simulated at four different speeds. At low speeds the rectangle and trapezoidal cross-section achieved high heat transfer. At high speed the pentagonal and rectangular cross-sections achieved high heat transfer. Pressure loss is accounted for while making use of the thermal performance factor parameter which accounts for both heat transfer and pressure loss. The pentagonal cross-section showed high potential in terms of the thermal performance factor with a value over 0.8 and higher by 33% when compared to the rectangular case. In the final section multiple enhancement techniques are combined in the sudden expansion case, such as, ribs, slots and ribbed slot. The maximum heat enhancement is achieved once all previous techniques are used together. Under these circumstances the Nusselt number increased by 60% in the proposed new design

Aerothermodynamics of Impingement and Film Cooling in a Gas Turbine Blade

The service life of gas turbine engine turbine blades depends on the blade’s material, service temperature and total stress. In high-performance gas turbines, film cooling is widely used to reduce the blade service temperature. Often impingement cooling is also employed to target the stagnation point heat transfer for internally-cooled gas turbine blades. A novel thermal wind tunnel was designed to study the combined effect of the impingement and film cooling on blunt airfoils. The hot exhaust plume of a micro-jet is used as the source of high-temperature gas flow in the thermal wind tunnel. An ejector nozzle was designed and integrated with the hot jet to provide a thermally controlled test section environment in the research facility. Measurements of freestream parameters such as gas speed, turbulence intensity and gas temperature were made. An airfoil that utilizes leading-edge (internal) impingement as well as film cooling holes on its suction surface was designed and fabricated. A cooling sleeve is used inside the airfoil to guide the impingement jets on the leading edge and to supply the coolant to the film holes. The surface temperature distribution is measured by an array of eight thermocouples flush-mounted on the airfoil surface downstream of the film holes. The initial ranges of blowing parameters (Mb) investigated were between 5 and 6. Numerical simulation using a commercially available Reynolds-Averaged Navier-Stokes (RANS) software was used and validated by the experimental measurements. The numerical simulations for the airfoil consisted of two thermal wall boundary conditions, the adiabatic and conjugate heat transfer (CHT) models. The adiabatic model focuses on the effect of film cooling on an adiabatic wall. The conjugate heat transfer model represents the solid and fluid heat transfer exchange, conduction and convection. Verification and validation was completed to ensure accurate aerothermodynamic simulations. The experimental and numerical data showed a close comparison for the suction surface temperatures and cooling effectiveness. A broader range of characteristic parameters (blowing parameter, turbulence intensity (Tu) and density ratio) were studied to show their impact on film cooling effectiveness parameter. The effects from the blowing parameter are reported for different Mb of 0.53 to 5.95 with two turbulent intensities, 5% and 20%. The adiabatic film effectiveness parameter showed two unique trends: low Mb with low Tu or high Mb with high Tu both exhibited improved film cooling effectiveness. Jet detachment is also detected at Mb ~ 1.5 for the current film cooling set up. The study of turbulence intensity effects was completed in the range of 5% to 25 % for two density ratios of 1.65 and 1.99. The turbulence intensity study showed that higher Tu caused the adiabatic film effectiveness to decrease by an average 18%. The density ratio (DR) in the film cooling is studied to explore the real turbine environment. The velocity ratio and turbulence intensity is held at a constant of 0.64 and 20%, respectively, for a range of the density ratio: 1.49 to 1.99. The results show that coolant density would cause the adiabatic film effectiveness to increase an average of 12% from the baseline (DR: 1.65) to the representative engine condition (DR: 1.99)

Turbine end-wall aerothermal management with engineered surface structure

Motivated by the enlarged design space and additional flexibility offered by the latest advances in manufacturing techniques, especially Additive Manufacturing (AM), this thesis investigates a novel turbine end-wall aerothermal management method with the engineered surface structures, through closely coupled experi-mental and numerical studies. A 90-degree turning duct and a linear turbine cascade test section were employed for the experimental research in a low-speed wind tunnel. Duct and turbine end-wall heat transfer and cooling effectiveness were measured by transient Infrared Thermography. PIV measurement was conducted to obtain the exit flow field. Computational fluid dynamics (CFD) simulations were performed using ANSYS FLUENT to compliment the experimental findings. The flow solver uses the finite volume method to solve the three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations. The k-ω shear stress transport (SST) turbulence model was validated and chosen for all the numerical studies. The secondary flow control principle of the engineered surface structure in the simplified duct is revealed through a detailed investigation of the flow produced by multiple small surface structures. The CFD and PIV measurement results consistently show that addition of the engineered surface structure on end-wall can effectively reduce the magnitude of streamwise vorticity associated with the secondary flow and alleviate its lift-off motion. For turbine cascade applications, it can be observed that the strength of the passage vortex is effectively reduced, and the passage vortex loss core moves closer to the end-wall with the addition of the engineered surface structure. The purge air cooling enhancement by the engineered surface structure is then studied. The purge air cooling flow becomes more attached to the end-wall and covers a larger wall surface area with the added end-wall rib structures. Both experimental and numerical results reveal a consistent trend on improving film cooling effectiveness and net heat flux reduction (NHFR). This novel concept was success-fully demonstrated in a more realistic turbine cascade case. Enhanced cooling effectiveness and net heat flux reduction were obtained from both experimental data and CFD analysis. The additional surface features were proved to be effective in reducing the passage vortex and providing more coolant coverage without introducing additional aerodynamic loss. The overall Net Heat Load Reduction for the 90-degree turning duct and the turbine cascade is increased by 11% and 2% respectively

Experimental investigation of film cooling effectiveness on gas turbine blades

The hot gas temperature in gas turbine engines is far above the permissible metal temperatures. Advanced cooling technologies must be applied to cool the blades, so they can withstand the extreme conditions. Film cooling is widely used in modern high temperature and high pressure blades as an active cooling scheme. In this study, the film cooling effectiveness in different regions of gas turbine blades was investigated with various film hole/slot configurations and mainstream flow conditions. The study consisted of four parts: 1) effect of upstream wake on blade surface film cooling, 2) effect of upstream vortex on platform purge flow cooling, 3) influence of hole shape and angle on leading edge film cooling and 4) slot film cooling on trailing edge. Pressure sensitive paint (PSP) technique was used to get the conduction-free film cooling effectiveness distribution. For the blade surface film cooling, the effectiveness from axial shaped holes and compound angle shaped holes were examined. Results showed that the compound angle shaped holes offer better film effectiveness than the axial shaped holes. The upstream stationary wakes have detrimental effect on film effectiveness in certain wake rod phase positions. For platform purge flow cooling, the stator-rotor gap was simulated by a typical labyrinth-like seal. Delta wings were used to generate vortex and modeled the passage vortex generated by the upstream vanes. Results showed that the upstream vortex reduces the film cooling effectiveness on the platform. For the leading edge film cooling, two film cooling designs, each with four film cooling hole configurations, were investigated. Results showed that the shaped holes provide higher film cooling effectiveness than the cylindrical holes at higher average blowing ratios. In the same range of average blowing ratio, the radial angle holes produce better effectiveness than the compound angle holes. The seven-row design results in much higher effectiveness than the three-row design. For the trailing edge slot cooling, the effect of slot lip thickness on film effectiveness under the two mainstream conditions was investigated. Results showed thinner lips offer higher effectiveness. The film effectiveness on the slots reduces when the incoming mainstream boundary layer thickness decreases

Analyse expérimentale et simulation CFD en vue de l'étude du refroidissement des aubes Rotor/Stator : cas d'une turbine à gaz

128 p. : ill. ; 30 cmLes performances aérothermiques d'un système de refroidissement interne du bord de fuite d'une aube de turbine à gaz sont évaluées expérimentalement et numériquement dans les conditions stationnaires et rotatives. La géométrie étudiée est un modèle à échelle 30 : 1 représentative d'une conduite sans et avec perturbateurs avec une ligne de 7 vannes élargies. Six géométries sont testées par le moyen de la technique TLC pour un nombre de Reynolds entre 10000-40000 et un nombre de Rotation jusqu'à 0.23. En outre, l'analyse CFD est basée sur ANSYS-Fluent et un modèle de turbulence k À- SST tout en considérant un écoulement d'air iso-thermique stationnaire à l'intérieur de la géométrie étudiée pour les conditions fixes et rotatives. Les résultats sont présentés sous forme de cartes 2D illustrant le coefficient d'échange thermique sur la surface en dépression, en plus de corrélations pour le nombre de Nusselt moyenné en fonction de Re, Pr, Ro et une fraction de la hauteur de l'aube. Les résultats obtenus sont d'un grand intérêt pour les concepteurs des systèmes de refroidissement pour les aubes de turbines à ga