For the thermal design of combustion chambers and turbine blades in jet engines, a detailed knowledge of the combustion and of the heat loads to the walls is necessary. In general, high operating temperatures and reduced combustor size are striven for in order to increase engine efficiency and reduce weight. Consequently, the components are exposed to temperatures above the melting point of the materials and there is a growing risk of in-complete combustion within the combustion chambers. To study these effects, we set up a new test bench for fundamental investigation of chemical near-wall reactions at atmospheric pressure. First results of gaseous, non-premixed near-wall CH4/air and H2/air flames are presented. Optical methods such as two-line laser-induced fluorescence themometry and OH* chemiluminescence were applied. Further, the heat release to the wall was determined by means of inverse heat conduction calculation using the data of implemented thermocouples.

Gas turbine combustion chambers contain numerous small-scale features that help to dampen acoustic waves and alter the acoustic mode shapes. This damping helps to alleviate problems such as thermoacoustic instabilities. During computational fluid dynamics simulations (CFD) of combustion chambers, these small-scale features are often neglected as the corresponding increase in the mesh cell count augments significantly the cost of simulation while the small physical size of these cells can present problems for the stability of the solver. In problems where acoustics are prevalent and critical to the validity of the simulation, the neglected small-scale features and the associated reduction in overall acoustic damping can cause problems with spurious, non-physical noise and prevents accurate simulation of transients and limit cycle oscillations. Low-order dynamical systems (LODS) and artificial neural networks (ANNs) are proposed and tested in their ability to represent a simple two-dimensional acoustically forced simulation of an orifice at multiple frequencies. These models were built using compressible CFD, using OpenFOAM, of an orifice placed between two ducts. The acoustic impedance of the orifice has been computed using the multi-microphone method and compared to a commonly used analytical model. Following this, the flow field downstream of the orifice has been modelled using both a LODS and ANN model. Both methods have shown the ability to closely represent the simulated dynamical flows at much lower computational cost than the original CFD simulation. This work opens the possibility of models that can dynamically predict the flow through, for instance, acoustic liners, dilution ports and fuel injectors in real engines during thermoacoustic instabilities without having to mesh and simulate these small-scale features directly. Such models may also assist in the accurate simulation of flame quenching due to cooling flows or the design of effusion cooled aerodynamic surfaces such as nozzle guide vanes (NGVs) and turbine blades.

This paper describes the full-scale turbine section validation of the GT36 heavy duty gas turbine, which was conducted in a test Power Plant in Birr, Switzerland. The GT36 Test Power Plant is extensively instrumented with both standard and specialised instrumentation. In the turbine section, specialised instrumentation includes metal and gas thermocouples, thermal paint, pressure sensors, hot gas rakes, strain gauges, five-hole probes, pyrometers and tip timing sensors. Similar specialised instrumentation also exists for the compressor, combustor and the rotor sections. Three major test campaigns were conducted over an extended period, which consisted of both long and short duration tests, including a range of off-design tests. Within the turbine section, detailed transient and steady-state measurements were obtained of the stage inlet pressures and temperatures, airfoil surface pressures and metal temperatures. These measurements indicated that both the aerodynamic and cooling performances of the turbine blades and vanes are highly consistent and repeatable over a range of operating conditions. Detailed comparisons of the measured engine pressures and temperatures with predictions also indicated that there was generally a very good match in the Mach numbers and metal temperatures for all the turbine blades and vanes.

In gas turbine engines, the highest heat loads occur at the leading-edge areas of turbine blades and vanes. To protect the blades and vanes, a “showerhead” configuration of film cooling holes is often used for this location, in which several rows of holes are configured closely together to maximize film coverage. Typically, these film cooling holes are fed by impingement cooling jets, helping to cool the leading edge internally, but also changing the internal flow field. The effects of these internal flow fields on film cooling are not well known, and experimental research is very limited in its ability to analyze them. Because of this, computational fluid dynamic (CFD) simulations using RANS were used as a way to analyze these internal flow fields. To isolate the effects of the impingement jet, results were compared to a pseudo-plenum internal feed, and rotation in the hole was found to be a key factor in performance. Computational results from both coolant feed configurations were compared to experimental results for the same configurations. The CFD RANS results were found to follow the same trends as the experimental results for both the impingement-fed and plenum-fed cases, suggesting that RANS is able to accurately model some of the important physics associated with leading-edge film cooling.

This work is an experimental study of film cooling effectiveness on a blade tip in a stationary, linear cascade. The cascade is mounted in a blowdown facility with controlled inlet and exit Mach numbers of 0.29 and 0.75, respectively. The free stream turbulence intensity is measured to be 13.5 % upstream of the blade’s leading edge. A flat tip design is studied, having a tip gap of 1.6%. The blade tip is designed to have 15 shaped film cooling holes along the near-tip pressure side (PS) surface. Fifteen vertical film cooling holes are placed on the tip near the pressure side. The cooling holes are divided into a 2-zone plenum to locally maintain the desired blowing ratios based on the external pressure field. Two coolant injection scenarios are considered by injecting coolant through the tip holes only and both tip and PS surface holes together. The blowing ratio (M) and density ratio (DR) effects are studied by testing at blowing ratios of 0.5, 1.0, and 1.5 and three density ratios of 1.0, 1.5, and 2.0. Three different foreign gases are used to create density ratio effect. Over-tip flow leakage is also studied by measuring the static pressure distributions on the blade tip using the pressure sensitive paint (PSP) measurement technique. In addition, detailed film cooling effectiveness is acquired to quantify the parametric effect of blowing ratio and density ratio on a plane tip design. Increasing the blowing ratio and density ratio resulted in increased film cooling effectiveness at all injection scenarios. Injecting coolant on the PS and the tip surface also resulted in reduced leakage over the tip. The conclusions from this study will provide the gas turbine designer with additional insight on controlling different parameters and strategically placing the holes during the design process.

The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Re θ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

The effects that leading-edge impingement coolant feeds have on the external flowfield and on film cooling performance in the showerhead have not been studied thoroughly in the literature. To isolate the influence of the impingement feed, experimental adiabatic effectiveness and off-the-wall thermal field measurements were made using a shaped hole geometry fed by an ideal plenum coolant feed and by an engine-realistic impingement coolant feed. The impingement configuration exhibited around 10% higher adiabatic effectiveness levels than the plenum configuration did — a finding in agreement with the few studies isolating this effect. CFD RANS simulations of the impingement and the pseudo-plenum configurations from a companion study were consulted to investigate the root cause of this difference in performance because the experimental data alone did not sufficiently explain it. In the impingement feed simulation, flow remained better attached throughout the hole (both at the inlet and at the diffuser) due to a rotation caused by the impingement flow, leading to better attachment on the exterior surface. This was most significant for the suction side holes at higher blowing ratios wherein the pseudo-plenum caused much more severe separation in the holes than the impingement configuration did.

Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.

Impingement cooling is one of the powerful cooling methods in high-temperature devices. For the gas turbine applications, impingement cooling is commonly applied in the transition piece of a combustor and in the leading edge, suction and pressure sides of a turbine blade/vane. In the suction side and pressure side, impingement cooling is applied as a form of an array jet. However, due to the small gap between the jet hole and target surface, the wall jet faces a crossflow inside of the gap. This crossflow has an adverse effect on the jets and deteriorates the heat/mass transfer performance. Therefore, several studies have been conducted to minimized the crossflow effect. The present study also investigated the effect of crossflow reduction in the gap by having a castellated hole plate. The heat/mass transfer was measured using the naphthalene sublimation method. Heat/mass transfer data are compared among three different cases. One is the baseline case which is simple array impinging jets. Others are the castellated cases with and without rib structures on the target wall. Jet-to-jet spacing (s/d) and jet-to-target spacing (z/d) are selected as geometrical variables. Also, the experiments were conducted for the Reynolds numbers (based on jet hole diameter) of 5,000, 15,000 and 30,000. The baseline case was named as B case, the castellated case without rib structure as C case and with rib structure as CR case. Both C and CR cases showed the crossflow reduction effect and resulted high and similar Nusselt number values.

To be able to set uniform inlet boundary conditions in simulation, there must be an inlet extension at the first guide vane. In the inlet extension, turbulence experiences strong numerical dissipation, which has not been paid attention to. In the current paper, the influence of the numerical dissipation of turbulence on accuracy in predicting heat transfer was discussed. Two cases, where the numerical dissipation of turbulence was neglected, were analysed. In the first case, wrong conclusion about effect of turbulence scale on heat transfer was drawn: blade heat transfer increases with inlet turbulence scale under the same inlet turbulence intensity. The mechanism for the wrong conclusion is that turbulence with larger scale numerically dissipates more slowly in the inlet extension so that turbulence intensity at blade leading edge is greater under turbulence with larger scale, it is the turbulence intensity not turbulence scale itself really affects heat transfer. In the second case, when the numerical dissipation of turbulence is neglected and turbulence parameters at measuring plane of inlet are directly as input for turbulence boundary condition, flow transition is postponed downstream and heat transfer error is greater, however, when the numerical dissipation of turbulence is considered and turbulence parameters at measuring plane are regard as benchmark and matched by adjusting parameters of inlet turbulence boundary condition, the result shows better agreement with experiment. Thus, the correct way to set turbulence boundary condition is to match turbulence parameters at measuring plane by adjusting parameters of inlet turbulence boundary.

The power generation industry is targeting heavy duty gas turbine to increase power and efficiency. Hot gas temperature and massflow are continuously being increased. It brings new challenges for the design of cooling systems for turbine blades and vanes. Up to date most of studies of heat transfer in internal cooling channels were in the range of Reynolds numbers below 80,000 for cooling air flow, for example, experimental series done by J. Chin Han et al. since 1985. Actually the range of Reynolds numbers is increased with the increase of total massflow. Extrapolation of available data is not reliable while local distribution of heat transfer coefficients becomes critical in terms of thermal stresses. Only few recent studies deal with the range of Reynolds number above 80,000 , for example, in 2009 J. Chin et. al showed results for 45° angled ribs provided only area averaged values for heat transfer coefficient over one pitch and in 2003 R. Bunker showed local distribution for 45° angled ribs only. Within current study the experimental measurements of local heat transfer and friction in ribbed cooling channel were performed for Reynolds numbers in range of 100,000 – 180,000 , what fits the parameters of modern and perspective heavy duty gas turbines. Using thermochromic liquid crystal technology the following rib configurations were tested: angled 45°, 60°, 90° and chevron 45°, 60°; pitch to height ratio of 10; rib turbulator height-to-channel hydraulic diameter ratio of 0.083. Maximum averaged heat transfer value was provided by 60° angled ribs. Comparison of local distribution of heat transfer coefficients for considered configurations was performed. Minimum non-uniformness of heat transfer coefficient was provided by chevron ribs, having maximum friction factor. Conjugated thermal-hydraulic analysis for cooled vane for heavy duty gas turbine was performed in order to quantify the effect of local heat transfer coefficient distribution in ribbed cooling channel. Metal temperature calculation was performed for two cases of air side thermal boundary condition application for wall surface between rib-turbulators: averaged value of heat transfer coefficient and detailed local distribution. Comparison of calculated metal temperature for 2 cases shows that usage of locally distributed air side heat transfer coefficient is important and should increase the accuracy of temperature prediction by 50°C. Consideration of local distribution of heat transfer coefficient is important for cooling design of modern heavy duty gas turbine in order to provide acceptable thermal gradients and consequently reach lifetime targets.

Rib-turbulators have been widely used in turbine blade cooling designs to enhance internal heat transfer. However, application of the ribs in blade cooling channels inevitably causes large pressure losses and lower heat transfer areas behind the ribs and at the joint parts of the ribbed wall with sidewalls in comparison to smooth channels. Semi-attached rib having two rectangular holes at the bottom of the ribs at the joint regions of the ribbed wall and two sidewalls is able to reduce friction losses and alleviates lower heat transfer in the channels. This paper presented a method of combining computational fluid dynamic simulation with an optimization method to optimize the performance of a channel with semi-attached ribs and to discuss the effects of hole geometric variables on both pressure losses and heat transfer enhancement. A fully automatic multi-objective genetic algorithm based on the Non-dominated Sorted Genetic Algorithm-II is applied for a one-pass squared channel with 60° inclined semi-attached ribs. Four geometrical variables are widths and heights of the two rectangular holes, and two objective functions are the maximum of area-averaged Nusselt number on the rib roughend wall and the minimum of friction factor. The Kriging model is constructed by using the Latin hypercube sampling method to choose 15 experimental points in the design space. The results showed that the optimization method helps to increase the channel performance and to eliminate nearly all low heat transfer areas by appropriately adjusting the heights and widths of the holes. Compared to solid ribs, the optimal semi-attached rib increases the area-averaged Nu number on the rib-roughened walls by 9.45%, whereas the friction factor in the channel is decreased by 6.95%, which corresponds to an increase of 12.13% in thermal performance. In the optimal model, the existence of large holes on the right side of the optimal semi-attached ribs moves the vortex core of the secondary flow downstream of the ribs from the right side of the solid-ribbed channel to the middle, which causes a large increase in Nu number on the middle and left side of the rib roughened walls and only a moderate reduction in Nu number on the right part of the walls. In the whole range discussed, heat transfer coefficient on the rib-roughened walls increased with the height of the right holes, but it increases with the width of the hole only in the range of 8–10mm. For the holes on the left, heat transfer coefficient deteriorates as the height and width increase. Increasing width and height of the holes on both sides of the semi-attached ribs reduces the channel friction losses.

For an optimum performance design of low-pressure turbine (LPT) blades, it is crucial to understand the generation mechanism of profile loss properly. As the profile loss is usually taken to be the loss generated inside the blade boundary layer due to viscous effects, most of the efforts for the performance optimization have concentrated on the reduction in the boundary layer loss using the flow parameters that represent the loss generation in the boundary layers. Kodama and Funazaki [1] investigated the generation mechanism of profile loss from a view point of blade drag forces, friction drag force and pressure drag force, and suggested that the loss due to pressure drag is dominant in the profile loss of a typical LPT blade. The loss due to pressure drag is not a boundary layer loss that is generated in the boundary layers, but a mixing loss that is generated downstream of the trailing edge. It is necessary to clarify a key flow parameter to the loss due to pressure drag for an effective performance optimization. This paper aims at investigating the flow parameter that is a measure of the profile loss. In the investigation, the profile loss is broken down into the loss components which are expressed by the boundary layer integral parameters at the trailing edge. Then the loss components are categorized into the loss due to friction drag or the loss due to pressure drag. The loss level of each component is evaluated by using the results of steady Reynolds Averaged Navier-Stokes (RANS) simulations to assess the contribution to the total profile loss. The evaluations are conducted for two kinds of blade profiles at three different Reynolds numbers. It is found that the largest contributor to the loss due to pressure drag, consequently to the total profile loss, is the loss associated with a mixing of accelerated free stream flow by the flow blockage at the trailing edge plane. The loss level is simply determined by the flow blockage. This suggests that the flow blockage at the trailing edge plane is the most important flow parameter for an optimum performance design of LPT blades.

The clearance that exists between the casing and turbine blade tips is one of the key drivers of efficiency in gas turbine engines. For this reason, engine manufacturers utilize precise manufacturing techniques and may employ clearance control systems to minimize tip clearances to reduce associated losses. Despite these efforts, turbines typically exhibit some nominal casing ovality or rotor-casing eccentricity, and changes to blade tip clearance during operation commonly occur due to thermal and mechanical stresses. The present study investigates non-axisymmetric tip clearance effects by creating a rotor-casing eccentricity in a one-stage axial test turbine operating in a continuous-duration mode at engine relevant conditions with engine representative hardware. A magnetic levitation bearing system was leveraged to move the turbine shaft to vary the rotor-casing eccentricity without test section disassembly. The results of this study indicate that rotor-casing eccentricity does not affect overall turbine efficiency over the range that was tested, but does locally influence efficiency and the rotor exit flow field. Comparisons of flow angle and secondary flow kinetic energy agreed with previous studies and existing analytical methods, respectively. Collectively, these results indicate that tip clearance can be studied locally on an eccentric rotor.

The present study focuses on the very high-lift T106C cascade with passing wakes and aims to validate the γ - Re θ ¯ model of Menter-Langtry used to predict laminar-turbulent transition based on unsteady Reynolds-Averaged Navier-Stokes simulations. The comparison to experimental data provided by Von Karman Institute, shows that the transition model is able to capture the influence of passing wakes on transition phenomenon. Like the experiments, the simulations show a reduction of the time-averaged separation bubble length and of the overall losses in the presence of passing wakes. For this numerical study, four other wakes have been generated in order to study the influence of wake parameters on the transition onset, on the laminar separation bubble formation and on the turbine cascade performances. For a given averaged turbulence intensity and total pressure deficit, thinner wakes seem to have a more positive effect on boundary layer, reducing the separation and the overall losses.

Axial turbine stages for gas generator cycle type rocket engines typically employ highly-loaded supersonic stator vanes. Consequently, the flow pattern downstream of the vanes is characterized by shock waves which induce high-frequency excitation on the subsequent rotor. For this reason, the optimal design of the stator is crucial in the context of development of the next generation of high-performance rocket engines, where reusability is a principal design criterion. A thorough comprehension of the loss mechanisms combined with the adoption of automated optimisation techniques can therefore enable new stator designs that may provide large benefits in terms of overall turbine performance and lifespan. The scope of this study stems from these considerations and its objective is twofold, namely i) the shape optimisation of a supersonic stator for rocket engines and ii) the investigation of the loss mechanisms in supersonic axial turbine stator vanes at on- and off-design conditions. The investigation is performed on stator vanes that are under development for the first turbine stage of a gas generator cycle type rocket engine. The stator vanes are therefore optimised in order to reduce the profile losses by exploiting a novel adjoint optimisation framework for turbomachinery implemented in the open-source code SU2. The effect on the resulting flow field and loss sources is finally investigated. Results show that entropy based loss coefficient gains of 6% can be achieved via shape optimisation and that the fluid-dynamic performance of these vanes is less sensitive to changes in pressure ratio compared to the performance provided by the baseline configuration. Eventually, shock-waves remain the primary loss source.

The engine-cycle performance of jet engines can be improved by more efficient cooling systems, either by reducing the required cooling air or by intensifying the cooling efficiency with the same amount of cooling mass flow. However, the multitude of geometrical design parameters and the strong multidisciplinary aspect of cooling mass flow consumption optimization make designing the cooling systems extremely challenging. Integrating probabilistic methods into the thermal design process enables the automated evaluation of multiple design variants which contributes to the development of more efficient systems. In the present study, the sensitivity of a multi-pass cooling system to geometric variations is investigated. The cooling air flow, solved using a 1D, correlation based flow solver, is iteratively coupled with the 3D-FE thermo-mechanical analysis of the blade. The geometry of the cooling system is varied using the Harmonic-Spline-Deformation parametric, which has been extended to modify the wall thickness enabling to perform a geometrical-holistic analysis. Furthermore, the Elementary-Effects-Method (EEM) and the Monte-Carlo-Simulation (MCS) are compared to identify the most influential parameters and analyze their complex interactions. It is shown that the cooling system’s performance is mostly affected by the shape and position of the first web. Furthermore, MCS proves to be robust towards changes in design space while simultaneously enabling a more detailed analysis of the system behavior compared to EEM.

Tip leakage flow is inevitable due to the tip clearance over rotor blades in turbines. This phenomenondeteriorates blade aerodynamic performance and induces severe thermal damage to the tip surface.Introduction of cooling jets to the tip can effectively controls the tip leakage flow and improves the tip heat transfer. Therefore, this paper aims to optimize film cooling holes on a flat tip of a subsonic cascade and an topology-optimized tip of a transonic cascade. A design variable is a material parameter defined at each grid node along the blade camber line. This idea is based on the topology optimization method. The objective is to minimize blade energy loss and maximize tip heat transfer intensity. A response surface optimization based on the design of experiment (DOE) analysis is employed, and a multi-objective Genetic Algorithm is used to get Pareto optimum solutions. During the DOE process, a CFD method using injection source terms is integrated for numerical simulations to reduce computational costs. Optimized tip film cooling holes are finally re-constructed. The influence of the newly designed tip cooling holes configuration on blade aero-thermal performance is evaluated via CFD simulations using body-fitted mesh. Results show that compared with the uniform arrangement of cooling film holes along the axial direction, all the optimized film cooling holes can improve both blade aerodynamic performance and tip heat transfer performance.

The first-stage rotor squealer tip is a key area in the gas turbine for both aerodynamic performance and heat transfer characteristics, which should be carefully designed. However, harsh operating conditions near the rotor squealer tip can cause the geometry of the squealer tip to degrade, and manufacturing inaccuracies can also cause the squealer tip geometry to deviate from the ideal design. In this work, an uncertainty quantification (UQ) method is proposed using the non-intrusive polynomial chaos expansion method and Smolyak sparse grids. Then coupled with three dimensional (3D) Reynolds-Averaged Navier-Stokes (RANS) solutions, an uncertainty quantification procedure is carried out for aerodynamic and heat transfer performance of GE-E3 rotor blade squealer tip. A parameter sensitivity analysis using the Sobol Indice method is carried out to identify the key parameters for aerothermal performance of the squealer tip. Wherein, the inlet total temperature and the blowing ratio are considered as flow condition uncertainty parameters and tip clearance is considered as geometrical uncertainty parameters. The uncertainty analysis results show that under the influence of the uncertain geometry and operating conditions, the heat flux of squealer tip basically conforms to the normal distribution and the statistical mean value of it increased by 13.56% relative to the design value and the probability of it deviating from the design value by 10% is as high as 65.68% The statistical average of the squealer tip film cooling effectiveness is reduced by 29.52% compared to the design value, and the probability of it deviating from the design value by 10% is as high as 91.83%. The result of sensitivity analysis reveals that the uncertainty of the aerodynamic characteristics of the squealer tip is almost entirely caused by the tip clearance which accounts for 88.02% of the variance of the leakage flow rate. The inlet total temperature has almost no effect on the uncertainty of the aerodynamic performance. However, it is the dominant variable for the uncertainty of the heat transfer performance considering that its variance indexes for tip heat flux Q Tip and film cooling effectiveness η are 84.87% and 24.87% respectively. Compared with the main effects, the influence of the interaction effects among the variables on the squealer tip aerothermal performance is almost negligible.

The performance of turbomachines is often dependent on the unsteady flow fields they naturally produce, owing primarily to row-to-row interactions from both moving and stationary components, as well as the unsteady nature of the turbulent flow. When it comes to computational fluid dynamics, a disparity exists between steady state and transient simulation as far as accuracy is concerned, albeit the computational cost of transient simulation on fully complex industrial hardware can be overwhelming. This study bridges the gap by presenting a harmonic balance conjugate heat transfer simulation approach in Simcenter STAR-CCM+, to model the unsteady flow phenomena while also providing accurate temperature predictions throughout the gas turbine blade solid bodies. The harmonic balance method used is a mixed time domain and frequency domain technique, which is suitable for periodic unsteady flows and is much less expensive than transient simulation. With this method, the impact of capturing these unsteady flow structures, such as the wake interactions and secondary cooling flows, is quantified on the resulting metal temperature distribution. Such is investigated and characterized throughout an industrial gas turbine blade with fully complex internal cooling passages, as well as film cooling for the external blade surface. Comparisons to steady simulation and transient simulation are also made to quantify the relative fidelity of each approach. Regarding the final resulting blade heat transfer, analysis is also provided to differentiate between important sources: the unsteadiness in the primary gas path flow and the classical unsteady nature of turbulence. Often these effects are lumped together when analyzing the resulting heat transfer, which is incorrect and can be better understood with more detailed analysis.