Cooled Gas Turbine Blade Research Articles

CFD Analysis of Gas Turbine Blade Cooling with Staggered Holes

The role of turbine inlet temperature plays a pivotal role in determining overall turbine performance. As the temperature at the turbine inlet increases, it leads to an enhancement in the turbine's average efficiency. Gas turbines operate under extremely elevated temperatures, reaching levels of 1200°C to 1500°C. This extreme heat poses the risk of turbine blade melting and potential failure. Consequently, the need arises for a sophisticated cooling mechanism for the turbine blades. Optimally increasing the inlet temperature of the turbine blades can certainly elevate the turbine's overall efficiency. However, it's important to recognize that this increase in temperature directly impacts the durability of the turbine blades. To address this challenge, various innovative cooling technologies have been devised, including the implementation of distinct blade cooling hole configurations. These methodologies aim to ameliorate blade temperatures and consequently enhance turbine performance.Turbine blades are engineered using superalloys, renowned for their ability to withstand high thermal stress. Advanced software tools are harnessed to assess their performance under extreme temperatures. Computational Fluid Dynamics (CFD) software is commonly employed for analyzing turbine blades. A prevailing approach for blade temperature reduction entails employing diverse cooling systems, often utilizing air as a cooling agent. This directed airflow over the blades significantly curtails their temperatures. In this context, blade design featuring staggered hole layouts is applied, and both STATIC and CFD simulations, grounded in defined boundary conditions, are executed, yielding commensurate outcomes. Comparative analysis demonstrates that staggered hole configurations provide superior temperature distribution uniformity across blade surfaces, effectively reducing peak temperatures in contrast to blade configurations devoid of cooling holes..

Numerical Simulation of Heat Transfer and Flow Enhancement Through Multi-Row Film Holes for Gas Turbine Blade Cooling

Numerical Simulation of Heat Transfer and Flow Enhancement Through Multi-Row Film Holes for Gas Turbine Blade Cooling

Coupled aerothermal-mechanical analysis in single crystal double wall transpiration cooled gas turbine blades with a large film hole density

• Double wall transpiration cooling protects turbine blades against high heat flux. • A high film hole density can benefit both the aerothermal and mechanical response. • The wall spacing influences the cooling effectiveness and aerothermal response. • The wall thickness ratio influences thermal stresses and the mechanical response. • Stresses tend to apply along secondary crystallographic directions around features. The ambition for increasing gas turbine efficiency beyond current levels through the elevation of gas temperatures demands substantial progress in turbine blade cooling technology. Double wall transpiration cooling (DWTC) is an emerging technology which offers enhanced thermal protection at only modest coolant flows, thanks to a combination of impingement jets and densely packed arrays of film cooling holes. This paper presents a coupled aerothermal-mechanical investigation of a transpiration cooled double wall turbine blade design, by employing Computational Fluid Dynamics (CFD), heat transfer theory as well as stress analysis based on plate theory and Finite Element (FE) analysis. In comparison to previously explored systems with modest outer wall porosity, a system with high porosity is found to display enhanced cooling effectiveness and to reduce the temperature difference across the two walls that drives thermal stresses. This difference can be decreased further by reducing the wall spacing, H , or the inner-outer wall thickness ratio, t c / t h , at the cost of higher overall metal temperatures. A lower bound for H should be used to avoid undesirable poor coolant flow distribution and hot gas ingestion effects, whereas the t c / t h ratio does not impose any aerothermal constraints. Thermal stresses associated with a fixed temperature field are invariant with H but they vary drastically with t c / t h . From a design perspective, the above suggest that H is primarily determined by aerothermal requirements, whereas t c / t h by the mechanical performance. The single crystal orientation and elastic anisotropy of Nickel alloys are shown to have a profound impact on the stress concentration around cooling holes, with secondary crystallographic directions, such as 〈 110 〉 and 〈 111 〉 , playing a prominent role in the local stress state. Our study provides a framework for optimising the aerothermal and mechanical performance in a range of high temperature components and highlights key areas where more elaborate analysis is needed.

EFFECTS OF ADDITIVELY MANUFACTURED SURFACE ROUGHNESS ON SMALL DIAMETER PLENUMS WITH MULTIPLE SIDE DISCHARGES FOR HIGH TEMPERATURE GAS TURBINE BLADE COOLING APPLICATIONS

Combined heat and power (CHP) applications have significant environmental and economic benefits that are consistent with the goals of the U.S. Department of Energy (DOE). One area that is currently being studied includes the potential benefits of CHP turbine operation at higher turbine inlet temperatures. Internal cooling concepts enabled by additive manufacturing (AM) are of primary interest. The effectiveness of internal cooling is hindered by many factors such as velocity distribution of the cooling air to the hot surface considering impingement cooling. To simulate cooling air exiting from a series of orifices for internal cooling in an airfoil, a straight smooth wall tubing with multiple side discharging orifices is used and compared to additively manufactured tubing (Ti6Al4V Grade 23) with orifice size and spacing as well as inner and outer diameters identical to the smooth wall tubing. Similar to flow discharging form perforated pipes, the flow discharged from individual orifices along the tubes in this study is found to be nonuniformly distributed, and the horizontal (axial direction) momentum can be observed from the experimental data. Discharge velocities have been measured with two-dimensional (2D) particle imaging velocimetry (PIV) and single element hot wire anemometry. Numerical analysis has also been conducted to predict the velocity distributions along the orifices in the smooth wall and additive manufacturing (AM) tubing, which are inherited with surface roughness. Numerical and measured results in this study are compared, presented, and discussed.

Mixed-dimensional geometric coupling of port-Hamiltonian systems

We propose a new interconnection relation for infinite-dimensional port-Hamiltonian systems that enables the coupling of ports with different spatial dimensions by integrating over the surplus dimensions. To show the practical relevance, we apply this interconnection to a model system of an actively cooled gas turbine blade. We also show that this interconnection relation behaves well with respect to a discretization in finite element space, ensuring its usability for practical applications.

Coupled Aerothermal-Mechanical Analysis in Single Crystal Double Wall Transpiration Cooled Gas Turbine Blades with a Large Film Hole Density

Coupled Aerothermal-Mechanical Analysis in Single Crystal Double Wall Transpiration Cooled Gas Turbine Blades with a Large Film Hole Density

On The Turbulent Flow Past Cambered Bodies with Blowing.(Dept.M)

This paper investigates, both experimentally and analytically, the heat and fluid flow around the trailing edge of a film cooled gas turbine blade. In the analysis, steady state, two dimensional, incompressible, recirculating turbulent flow is considered. The momentum and energy equations are derived and transformed to the generalized format of Gosman and et al (2). The kinetic energy and dissipation term of turbulence (i.e. the k-E, model) is con sidered in the turbulence computational programs. Small number of empirical constants were obtained from both the mathematical model and the conducted experimental work. These mathematical constants provide the most realistic predictions of heat transfer coefficients, when compared with the available experimental data. The investigation considers varying Reynolds number in the range of 6.25 x 104 to 1.29 x 105. There is a satisfactory agreement between the present experimental and theoritecal results, for the blowing parameter (M = Uc/U ) in the range from zero to one.

Heat Transfer Enhancement in Air Cooled Gas Turbine Blade Using

In this paper, an investigation of using corrugated passages instead of circular crosssection passages was achieved in conditions simulate the case in the gas turbine blade coolingusing ANSYS Fluent version (14.5) with Boundary conditions: inlet coolant air temperature of300 K with different air flow Reynolds numbers (191000, 286000 and 382000). Thesurrounding constant hot air temperatures was (1700 K). The numerical simulations was done bysolving the governing equations (Continuity, Reynolds Averaging Navier-stokes and Energyequation) using (k-ε) model in three dimensions by using the FLUENT version (14.5). Thepresent case was simulated by using corrugated passage of 3 m long, internal diameter of 0.3 m,0.01 m groove height and wall thickness of 0.01 m, was compared with circular cross sectionpipe for the same length, diameter and thickness. The temperature, velocity distributioncontours, cooling air temperature distribution, the inner wall surface temperature, and thermalperformance factor at the two passages centerline are presented in this paper. The coolant airtemperature at the corrugated passage centerline was higher than that for circular one by(12.3%), the temperature distribution for the inner wall surface for the corrugated passage islower than circular one by (4.88 %). The coolant air flow velocity seems to be accelerated anddecelerated through the corrugated passage, so it was shown that the thermal performance factoralong the corrugated passage is larger than 1, this is due to the fact that the corrugated wallscreate turbulent conditions and increasing thermal surface area, and thus increasing heat transfercoefficient than the circular case.

Conjugate Heat Transfer Simulation and Entropy Generation Analysis of Gas Turbine Blades

Abstract Reynolds-averaged Navier–Stokes (RANS) model-based conjugate heat transfer (CHT) method is so far popularly used in simulations and designs of internally cooled gas turbine blades. One of the important factors influencing the RANS-based CHT method's prediction accuracy is the choice of turbulence models for different fluid regions because the blade passage flow and internal cooling have considerably different flow features. However, most studies in the open literature adopted the same turbulence models in the blade passage flow and internal cooling. Another important issue is the comprehensive evaluation of the losses caused by the flow and heat transfer for both fluid and solid regions. In this study, a RANS-based CHT solver suitable for subsonic/transonic flows was developed based on OpenFOAM and then validated and used to explore suitable RANS turbulence model combinations for internally cooled gas turbine blades. Entropy generation, being able to weigh the losses caused by both flow friction and heat transfer, was used in the analyses of two vanes with smooth and ribbed cooling ducts to reveal the loss mechanisms. Findings indicate that the combination of the k–ω SST–γ–Reθ transition model for passage flow and the standard k–ε model for internal cooling provided the best agreement with measurement data. The relative error of vane surface dimensionless temperature was less than 3%. The variations of entropy generation with different internal cooling inlet velocities and temperatures indicate that reducing entropy generation was contradictory with enhancing heat transfer performance. This study, which provides a reliable computing tool and a comprehensive performance parameter, has an important application value for the design of advanced internally cooled gas turbine blades.

Mathematical modeling of coupled heat transfer on cooled gas turbine blades

The paper presents a physico-mathematical model for determining the heat transfer parameters between viscous gasdynamic flotations and cooled gas turbine blades made using the technology of composite permeable membranes (CPM). The mathematical model includes equations of dynamic and thermal boundary layers taking into account injection, three-dimensional transient heat conduction in CPM cooled frames, complex hydraulic flow and cooling inside the channels of a cooled blade with air taken from the power plant compressor. A complex problem was solved to determine the thermal state of microrocket engines, where a method was proposed for immersing a multiply-connected domain in a region of the simplest form. A new economical, absolutely stable method is proposed for numerically solving of spatial nonstationary heat transfer problems in multiply-connected domains. The results obtained showed greater cooling efficiency at lower cooling medium costs.

Pulsed micro-melting method for single-cristal turbine blade reduction technology

The paper is devoted to the study of the composition, microstructure and mechanical properties of heat-resistant single crystalline intermetallic alloy. In order to develop a resource-saving technology, the method of pulsed micro- surfacing has been tested, designed to restore the geometric dimensions of cooled gas turbine blades.

Geometric Optimization of a Gas Turbine Blade Cooling Passage using CFD

This work focuses on the rib-turbulated cooling which is a category of impingement cooling and aims at optimizing the geometry of rib-roughened cooling passage of a gas turbine blade. CFD analysis is carried out using Ansys/Fluent to solve the steady RANS equations. Computational domain consists of a long rectangular channel with the length of the channel being 9 times its height. Ratios of rib width, rib height and rib pitch to hydraulic diameter of the channel are taken as 0.1, 0.1 and 1.2, respectively. Numerical simulations are performed to analyze the performance of various rib shapes for Reynolds number, based upon the hydraulic diameter, in a range of 5000 to 50,000. Uniform heat flux of 800 W/m2 is applied to the ribbed wall. Incompressible air is used as the cooling fluid. Turbulent flow conditions are applied to the channel geometry with k-ω turbulence model. The effect of rib cross-section and rib pitch to rib width ratio on the heat transfer and friction factor is observed. The 2D CFD analysis revealed that the presence of ribs has significant effect on thermo-hydraulic performance of the cooling channel. Introducing square ribs in a smooth channel caused the Nusselt number to increase by two-folds. The highest value of Nusselt number was achieved by incorporating right-angle triangular ribs which caused the Nusselt number to increase by further 8%, as compared to the square ribs, and an increase in friction factor of 2.5%. The lowest value of friction factor was observed in semicircular ribs (2.95% less than the square ribs), however, the Nusselt number also decreased by 1.5%, as compared to square ribs. Decreasing rib pitch to rib width ratio increased both the Nusselt number and friction factor for all the cases. For square ribs, decreasing this ratio from 15 to 9 resulted in the rise of Nusselt number by 50% and increase in friction factor by 54%.

Optimization of Gas Turbine Blade Cooling System

At present, Gas turbines play an essential responsibility in different areas such as jet, generating power and various commercial and industrial sectors. Melting point of the turbine blade may causes the hotness levels which go rapidly raise. Likewise, heavy crack may cause because of Turbine Inlet Temperature (TIT) at turbine blades for the period of expansion procedure of turbine sector. Hence, a highly developed blade cooling system is required for safe operation of turbines. The proposed system deals with the serpentine rip - roughened passage with micro pin fin cooling system and it has been analyzed corresponding to serpentine cooling system. It increases the heat transfer enhancement. Therefore, very warm gases in and around the turbine blade may have a stream temperature at 1500K. On the other side, cool air disclosed to the blade core duct and an entry temperature may have been 650K. The proposed systems with 2D and 3D model were developed by using CATIA. The 3D design is then analyzed using CFD. Further, the corresponding results of serpentine rip - roughened passage and micro pin fin arrangement in serpentine rip-roughened passage are compared and the details are presented.

Heat Transfer in a Square Ribbed Channel: Evaluation of Turbulent Heat Transfer Models

This paper presents the results of integral heat transfer measurements taken in a square ribbed cooling channel configuration for evaluating heat transfer and turbulent flow characteristics in convective cooled gas turbine blades and draws a comparison with numerical results. The heated section of the channel is either smooth or equipped with 45 ∘ crossed ribs on two opposite walls. The first part of the paper describes the instrumentation and experimental setup in detail. The second part compares the numerical calculations with the experimentally determined results. The turbulent heat transfer is calculated using two common algebraic models and three implemented explicit algebraic models, each time in combination with an explicit algebraic Reynolds stress model. The numerical calculations show that the use of higher-order models for the turbulent heat flux provides a higher accuracy of the heat transfer prediction for both configurations. The best model is able to predict almost all results within the experimental uncertainties.

Effect of Helix Angle and Cross Section of Helicoidal Ducts in Gas Turbine Blade Cooling

Gas turbines play a prominent role in aircraft propulsion and land-based power generation systems. To improve the turbine blade efficiency and power, present-day gas turbines are subjected to higher inlet temperatures. However, this has an adverse effect as it increases the thermal loads on gas turbine blades which weaken the structure. Modern gas turbines are operated at temperatures beyond 1500°C. Therefore, power output from a gas turbine blade is directly dependent on the cooling, so that turbine blades are not damaged from such high thermal loads. In this paper, an effort has been made to numerically analyze the effect of helix angle on circular helicoidal shape geometry in gas turbine blade cooling. Different cross-section of helicoidal geometries discussed in this paper is elliptical, circular, square and triangular. It is observed that reduction in helix angle results in improved cooling of a gas turbine blade. It is also inferred that elliptical helicoidal ducts enhance cooling as compared to other helicoidal geometries due to a larger convective area and higher heat dissipation as compared to other cross-section of helicoidal ducts. Copyright © 2019 Praise Worthy Prize S.r.l. - All rights reserved.

A Three-Dimensional Conjugate Approach for Analyzing a Double-Walled Effusion-Cooled Turbine Blade

A double-wall cooling scheme combined with effusion cooling offers a practical approximation to transpiration cooling which in turn presents the potential for very high cooling effectiveness. The use of the conventional conjugate computational fluid dynamics (CFD) for the double-wall blade can be computationally expensive and this approach is therefore less than ideal in cases where only the preliminary results are required. This paper presents a computationally efficient numerical approach for analyzing a double-wall effusion cooled gas turbine blade. An existing correlation from the literature was modified and used to represent the two-dimensional distribution of film cooling effectiveness. The internal heat transfer coefficient was calculated from a validated conjugate analysis of a wall element representing an element of the aerofoil wall and the conduction through the blade solved using a finite element code in ANSYS. The numerical procedure developed has permitted a rapid evaluation of the critical parameters including film cooling effectiveness, blade temperature distribution (and hence metal effectiveness), as well as coolant mass flow consumption. Good agreement was found between the results from this study and that from literature. This paper shows that a straightforward numerical approach that combines an existing correlation for film cooling from the literature with a conjugate analysis of a small wall element can be used to quickly predict the blade temperature distribution and other crucial blade performance parameters.

Efficient uncertainty quantification of stochastic heat transfer problems by combination of proper orthogonal decomposition and sparse polynomial chaos expansion

To increase the contribution and reliability of computational fluid dynamics efforts in design process of industrial equipments, it is necessary to quantify the effects of uncertainties on the system performance. Due to exponentially increment of the computational cost with number of uncertain variables for uncertainty quantification using classical polynomial chaos expansion methodology, reducing the required number of samples for uncertainty quantification is a real engineering challenge. In this paper, the proper orthogonal decomposition method based on the multifidelity approach is combined with the full and sparse polynomial chaos expansions for efficient uncertainty quantification of complex heat transfer problems with large number of random variables. The conjugate conduction heat transfer in NASA C3X cooled gas turbine blade with geometrical uncertainties and the convective heat transfer in ribbed passage with the stochastic wall heat flux boundary condition are considered as the test cases. Results of uncertainty quantification analysis in both test cases showed that proposed multi-fidelity approaches are able to produce the statistical quantities with much lower computational cost compare to the classical regression-based polynomial chaos method. It is shown that the combination of the proper orthogonal decomposition with the sparse polynomial chaos gives a computational gain at least 2 times greater than combination of the proper orthogonal decomposition with the full polynomial chaos expansion.

Verification of Thermal Models of Internally Cooled Gas Turbine Blades

Numerical simulation of temperature field of cooled turbine blades is a required element of gas turbine engine design process. The verification is usually performed on the basis of results of test of full-size blade prototype on a gas-dynamic test bench. A method of calorimetric measurement in a molten metal thermostat for verification of a thermal model of cooled blade is proposed in this paper. The method allows obtaining local values of heat flux in each point of blade surface within a single experiment. The error of determination of local heat transfer coefficients using this method does not exceed 8% for blades with radial channels. An important feature of the method is that the heat load remains unchanged during the experiment and the blade outer surface temperature equals zinc melting point. The verification of thermal-hydraulic model of high-pressure turbine blade with cooling allowing asymmetrical heat removal from pressure and suction sides was carried out using the developed method. An analysis of heat transfer coefficients confirmed the high level of heat transfer in the leading edge, whose value is comparable with jet impingement heat transfer. The maximum of the heat transfer coefficients is shifted from the critical point of the leading edge to the pressure side.

A Lumped Thermodynamic Model of Gas Turbine Blade Cooling: Prediction of First-Stage Blades Temperature and Cooling Flow Rates

Turbine inlet temperatures (TIT) of 1500–2000 K have become a sort of standard for most modern advanced applications. First-stage blades are obviously the most exposed components to such hot gases, and thus they need proper cooling. In the preliminary design of the blades and their cooling system, designers must rely on simple models that can be further refined at a later stage, in order to have an approximate but valuable set of guidelines and to reach a feasible first-order configuration. In this paper, a simple lumped thermodynamic model of blade cooling is proposed. It is based on mass/energy balances and heat transfer correlations, and it predicts a one-dimensional temperature profile on the blade external surface along the chord for a given gas temperature profile, as well as the required cooling air flow rates to prevent blade material from creep. The greatest advantage of the model is that it can be easily adapted to any operating condition, process parameter, and blade geometry, which makes it well suited to the last technological trends, namely, the investigation of new cooling methods and alternative coolants instead of air. Therefore, the proposed model is expected to be a useful tool in the field of innovative gas turbine cycle analysis, replacing more computationally intensive and very time-consuming models.

Single wavelength and ratio pyrometry reflection errors in temperature measurement of gas turbine blade

Radiation thermometers are the instruments most opted for when measuring surface temperatures of high speed rotating parts non-intrusively in many industrial applications. However, reflection of ambient radiation from the measured target goes against the operating principles of most of these thermometers thus introducing errors to the temperature measurements. Reflection errors introduced into single wavelength and ratio pyrometer temperature measurements of a rotating cooled gas turbine blade were investigated in this work through simulation. Errors for the two pyrometers based on varied target surface emissivity, view factor and ambient temperature were calculated. At high ambient temperatures, high view factors and low surface emissivity, errors exceeding 70% and 30% were realized for single wavelength and ratio pyrometers respectively. Thus temperature measured on reflective targets surrounded by surfaces at much higher temperature should be corrected for reflection errors if its reliability is to be guaranteed.