This paper presents a CFD-based numerical analysis on the potential benefits of non-radial blading turbine for low speed-low pressure applications. Electric turbocompounding is a waste heat recovery technology consisting of a turbine coupled to a generator that transforms the energy left over in the engine exhaust gases, which is typically found at low pressure, into electricity. Turbines designed to operate at low specific speed are ideal for these applications since the peak efficiency occurs at lower pressure ratios than conventional high speed turbines. The baseline design consisted of a vaneless radial fibre turbine, operating at 1.2 pressure ratio and 28,000rpm. Experimental low temperature tests were carried out with the baseline radial blading turbine at nominal, lower and higher pressure ratio operating conditions to validate numerical simulations. The baseline turbine incidence angle effect was studied and positive inlet blade angle impact was assessed in the current paper. Four different turbine rotor designs of 20, 30, 40 and 50° of positive inlet blade angle are presented, with the aim to reduce the losses associated to positive incidence, specially at midspan. The volute domain was included in all CFD calculations to take into account the volute-rotor interactions. The results obtained from numerical simulations of the modified designs were compared with those from the baseline turbine rotor at design and off-design conditions. Total-to-static efficiency improved in all the non-radial blading designs at all operating points considered, by maximum of 1.5% at design conditions and 5% at off-design conditions, particularly at low pressure ratio. As non-radial fibre blading may be susceptible to high centrifugal and thermal stresses, a structural analysis was performed to assess the feasibility of each design. Most of non-radial blading designs showed acceptable levels of stress and deformation.

Over the last decade, new concepts have evolved to promote a significant azimuthal flow in annular combustors of gas turbines. The benefits are better flame propagation at ignition, positive flame-flame interaction, and better interaction with the burnt gases. Other advantages in terms of size, congestion and conditioning of the turbine inlet flow are also significant. The technical challenges reported by the literature are often related to the higher thermal stress of the flames on the walls compared to a conventional frame. Other sources of inspiration for this work are the principles of burnt gas recirculation, sequential combustion and flameless combustion. This contribution focuses on a novel tangential burner arrangement inspired by the previous references. It offers a synthesis of key features and properties of the latter and goes even further. Here, a significant part of the burnt gases produced by one burner intentionally enters the inlet of the next burner, and so on along the azimuthal direction. This takes advantage of the closed loop aspect of an annular combustor when considering the toroidal direction. It also proposes a solution to the thermal load problem. We named this principle Recursive Sequential Combustion (RSC). While the flame alignment is organised along the generatrix of the combustor’s annulus, one difficulty lies in the design of the lateral feeds of reactants and the lateral exit of the exhaust gases. A double-spiral combustor design is proposed, which has similarities with the Swiss Roll Combustor concept. It directs the flow in the toroidal direction, as well as it creates the favourable conditions for a dynamically stabilised premixed flame centred along the torus’ generatrix at some distance from the walls. This design maximises the interaction between the fresh reactants and the burnt gases. The technical challenge is to find the right balance in terms of momentum flux of the incoming and outgoing flows to keep the flame in the middle of the torus. If this concept is successful, a lean flame could be operated with an unmatched trade-off between stability, flexibility and low-emissions (including soot). The paper reports about the RSC concept, the design, and the early results.

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.

Due to increase in the power generation from renewable sources, steam and gas turbines will be required to adapt for more flexible operations with frequent start-ups and shut-downs to provide load levelling capacity. During shut-down regimes, mixed convection takes place with natural convection dominance depending on the operating conditions in turbine cavities. Buoyant flows inside the turbine that are responsible for non-uniform cooling leading to thermal stresses and compromise clearances directly limits the operational flexibility. Computational fluid dynamics (CFD) tools are required to predict the flow field during these regimes since direct measurements are extremely difficult to conduct due to the harsh operating conditions. Natural convection with the presence of cross-flow -mixed convection has not been extensively studied to provide detailed measurements. Since the literature lacks of research on such flows with real engine representative operating conditions for CFD validation, the confidence in numerical predictions is rather inadequate. This paper presents a novel experimental facility that has been designed and commissioned to perform very accurate unsteady temperature and flow field measurements in a simplified turbine casing geometry. The facility is capable of reproducing a wide range of Richardson, Grashof and Reynolds numbers which are representative of engine realistic operating conditions. In addition, high fidelity, wall resolved LES with dynamic Smagorinsky subgrid scale model has been performed. The flow field as well as heat transfer characteristics have been accurately captured with LES. Lastly, inadequacy of RANS for mixed type of flows has been highlighted.

Nickel-base superalloys (NBSAs) are extensively utilized as the design materials to develop turbine blades in gas turbines due to their excellent high-temperature properties. Gas turbine blades are exposed to extreme loading histories that combine high mechanical and thermal stresses. Both directionally solidified (DS) and single crystal NBSAs are used throughout the industry because of their superior tensile and creep strength, excellent low cycle fatigue (LCF), high cycle fatigue (HCF), and thermomechanical fatigue (TMF) capabilities. Directional solidification techniques facilitated the solidification structure of the materials to be composed of columnar grains in parallel to the <001> direction. Due to grains being the sites of failure initiation the elimination of grain boundaries compared to polycrystals and the alignment of grain boundaries in the normal to stress axis increases the strength of the material at high temperatures. To develop components with superior service capabilities while reducing the development cost, simulating the material’s performance at various loading conditions is extremely advantageous. To support the mechanical design process, a framework consisting of theoretical mechanics, numerical simulations, and experimental analysis is required. The absence of grain boundaries transverse to the loading direction and crystallographic special orientation cause the material to exhibit anisotropic behavior. A framework that can simulate the physical attributes of the material microstructure is crucial in developing an accurate constitutive model. The plastic flow acting on the crystallographic slip planes essentially controls the plastic deformation of the material. Crystal Visco-Plasticity (CVP) theory integrates this phenomenon to describe the effects of plasticity more accurately. CVP constitutive models can capture the orientation, temperature, and rate dependence of these materials under a variety of conditions. The CVP model is initially developed for SX material and then extended to DS material to account for the columnar grain structure. The formulation consists of a flow rule combined with an internal state variable to describe the shearing rate for each slip system. The model presented includes the inelastic mechanisms of kinematic and isotropic hardening, orientation, and temperature dependence. The crystallographic slip is accounted for by including the required octahedral, cubic, and cross slip systems. The CVP model is implemented through a general-purpose finite element analysis software (i.e., ANSYS) as a User-Defined Material (USERMAT). Uniaxial experiments were conducted in key orientations to evaluate the degree of elastic and inelastic anisotropy. The temperature-dependent modeling parameter is developed to perform non-isothermal simulations. A numerical optimization scheme is utilized to develop the modeling constant to improve the calibration of the model. The CVP model can simulate material behavior for DS and SX NBSAs for monotonic and cyclic loading for a range of material orientations and temperatures.

The need for better aeroengine efficiency and fuel burn savings can be addressed by increasing turbine cycle temperatures. New cooling technologies such as emerging double skin transpiration (DST) systems, are essential for preserving the integrity of metallic parts if operating temperatures are to be increased beyond current levels. Their implementation, however, requires complex architectures with detailed features such as inclined film holes that raise local stresses. Our aim is to accelerate the implementation of DST systems in turbomachines, by providing an understanding of their implications on thermal stresses and the creep-fatigue failure processes. By using geometric, temperature field and nickel alloy material property idealizations, we generate both theoretical and Finite Element (FE) solutions for the thermal stress field and the critical cyclic strain range. Theoretical stress analysis and a local approach to failure lead to life predictions that are in reasonable agreement with inelastic cycle-by-cycle FE analysis, suggesting that analytical approaches can be useful for immediately identifying the significance of the various geometric features and thermal loading parameters. Film hole locations are predicted to fail under low cycle fatigue, suggesting that optimizing hole inclination, hole shape and other important geometric features is critical.

Within the research area of sCO 2 power cycles, the direct-fired oxy-combustion cycle is of great interest because the inherent carbon capture makes it a desirable option for clean power generation. In order to better understand the combustion kinetics of the direct-fired cycle in sCO 2 , a laboratory-scale combustion test rig was designed. The design and thermal-mechanical analysis of the combustor test rig for steady-state open-loop combustion testing is presented. The test rig is designed for operating conditions of 265 bar and full-scale oxy-combustion temperatures. The test rig incorporates a laser igniter and modular fully premixed injector to enable testing with step- or swirl-style flame holders. The rig also employs an existing dense-phase sCO 2 pump to provide a closed-loop high-flow cooling circuit to enable cost-effective short-lead manufacturing of liner and pressure casing components. Finally, the test rig’s multi-zone copper alloy C18150 liner design minimizes thermal stresses and allows for independent control of dilution flow in each zone to vary combustion temperatures. Planned measurements include exit temperature profile and chamber dynamic pressures. The sensitivity of CO oxidation and other reaction rates to temperature will be indirectly evaluated using post-combustion gas sampling for comparison with combustion simulations.

Stator and rotor blades in intermediate-pressure steam turbines gradually deteriorate during operation because of solid particle erosion. In addition to that, turbine blades unexpectedly crack because of metal fatigue or thermal stress deformation. As eroded blades increase the aerodynamic losses and cracked blades may induce rupture of the blade, the periodic maintenance, repair, and overhaul of steam turbines is essential. Eroded or cracked blades should be replaced with new ones or repaired for further use. Cutback treatment is one of the repair methods wherein the deteriorated trailing edge on a turbine blade is removed to avoid further cracking and blade fracturing. The use of cutback blades can reduce the replacement cost; however, that may affect the steam flow and the turbine’s performance. In this study, we numerically investigated the effect of the blade deterioration on the performance of a three-stage intermediate-pressure steam turbine using a numerical method that was developed at Tohoku University. Further various cutback lengths were considered for the deteriorated first-stage stator-blade trailing edge. The obtained numerical results indicate that the cutback first-stage stator blades certainly affected the steam flow in the turbine, resulting in a negative influence on the torque obtained from the adjacent rotor blades, which depends on the cutback length. However, the torque decrement can be mitigated by arranging the cutback and non-cutback stator blades alternately in a row.

The SOFC performance and lifetime highly depend on the operation condition, especially the SOFC operation temperature. The temperature fluctuation causes thermal stress in electrodes and electrolyte ceramics. On the other hand, it also needs to maintain a sufficiently high temperature to enable the efficient transport of oxygen ions across the electrolyte. Therefore, it is necessary to design an effective SOFC temperature management system to guarantee safe and efficient operation. In this paper, a two-side temperature control method is proposed to avoid the temperature difference between anode and cathode. Therefore, the SOFC thermal management system includes two control loops. The anode inlet temperature and cathode inlet temperature are controlled by blowers adjusting the recirculated flow rate. In addition, the control performance of the proposed SOFC thermal management system is compared with one-side temperature control systems. The results show that both anode control loop and cathode control loop are essential to get a better control performance. The SOFC would operate with less efficiency with only anode temperature control. On the other hand, the safety problem would occur with only cathode temperature control. The temperature gradient would be more than the upper limit at a part load condition. Therefore, the SOFC thermal management strategy with anode and cathode temperature control loops is feasible for the SOFC-GT system.

As gas turbine inlet temperature continuously increases, blade trailing edge suffers an extremely high thermal load due to the thin structure and constraint of internal convective cooling arrangements. To overcome these difficulties, pressure-side cutback, which is strengthened with multiple internal structures and land extensions, is widely used in blades to protect trailing edge from high thermal stresses. Due to the geometrical complexity and strong interactions between coolant flow and mainstream, sophisticated heat transfer and flow patterns exist in the cutback region, which presents a great challenge for the trailing edge cutback design. To understand the heat transfer and aerodynamic performance in blade with trailing edge cutback, CFD method has become an efficient tool which provides deep insights into the flow mechanisms and heat transfer characteristics in the detailed region. To accurately resolve the flow and heat transfer performance in a turbine blade with trailing edge cutback, structured grids are preferred because of higher resolution in flow/heat transfer prediction than unstructured grids, especially in boundary layers. However, for a blade with landed trailing edge cutback, few researchers tried to employ structured grids to predict aero-thermal performance due to the geometrical complexity. In this paper, the Background-Grid Based Mapping (BGBM) method proposed in Part I of this study was adopted to generate multi-block structured grids for a gas turbine vane with landed trailing edge cutback. With the coordinate transformation strategies, multi-block structured grids for the vane with landed trailing edge cutback were generated conveniently. With the generated structured grids, flow and heat transfer performance in vane were investigated using RANS (Reynolds-Averaged Navier-Stokes) equations solutions combined with transitional turbulence model. Effects of land extensions on the heat transfer and aerodynamic performance were analyzed, as well as the effects of inflow turbulence intensity, mainstream Reynolds number and ejection rate. The results show that heat transfer coefficients on vane surface, total pressure loss coefficient and energy loss coefficient in vane are all increased with the increase of inflow turbulence intensity. However, heat transfer coefficients on cutback and trailing edge surface are not sensitive to inflow turbulence intensity. At the same inflow turbulence intensity, the aerodynamic loss in vane is decreased with increasing the Reynolds number of mainstream. The increase of ejection rate significantly increases the heat transfer coefficients on cutback surface. Compared with the vane without land extensions, heat transfer coefficients and pressure coefficients on vane surface are reduced and the heat transfer coefficients on cutback surface are increased for the vane with land extensions. In the case of Re = 2.0 × 10 6 , the area-averaged heat transfer coefficient on landed cutback is 14.46% higher than the cutback without lands. Compared with the experimental data, predictions with structured grids based on BGBM method are more agreeable than those with unstructured grids.

Tip leakage flow in high speed turbine induce significant thermal loads and give rise to intense thermal stresses on blade tip, while increasing inlet pressure tends to accelerate leakage velocity beyond transonic regime. The present research quantifies heat transfer and film cooling effect on a squealer tip with three film cooling layouts, three coolant mass flow rates and a relative casing movement. The results indicate that area-averaged HTC of PS layout is higher than that of CAM layout by 6.9% and that of SS layout by 5.7% when coolant flow rate equals to 0.6% mainstream flow rate. By comparison, it is clearly observed that area of the high heat transfer coefficient regions are significantly enlarged when the flow rate of coolant is increased. With relative casing movement, a significant high HTC stripe parallel to pressure side rim is formed. In case of the PS layout, heat transfer coefficient is reduced by 7.3% with casing movement. While in case of CAM layout and SS layout, heat transfer coefficient increased by 4.8% and 2.3% with casing movement, respectively. Detailed flow patterns with three film cooling layouts are also illustrated.

A conjugate heat transfer study is carried out to obtain temperature and thermal stress field of a film-cooled superalloy with multi-layer thermal barrier coatings (TBCs). The aim is to understand the effects of the blowing ratio and ceramic top coating (TC) thickness on temperature and thermal stress which have an influence on component reliability and life. Results reveal that the distribution of film cooling effectiveness gets more uniform as TC thickness decrease because thick TC with low thermal conductivity prevents heat conduction in the axial and spanwise directions. In the upstream of the film cooling hole, the cooling effect is enhanced nonlinearly with the increase of the blowing ratio since the flow separation in the cooling tube affects the heat transfer enhancement. The insulation performance is improved by about 10 K for every 0.1D increase in TC thickness and the cooling effect is improved by about 20 K when the blowing ratio is increased from 0.5 to 1.0 at the leading edge of the film-cooling tube. The influence of jet lift-off and hotgas entrainment on the insulation effect is greater than TC thickness. The stress is concentrated at the leading edge of the film cooling hole and interfaces of TBCs. The maximum Von-Mises stress (761 MPa) on the interfaces is not at the leading or trailing sides of the film-cooling tube, it is about ± 45° from the centerline of the BC/SUB interface. The debonding stress at TC/BC interface and BC/SUB interface are about 26 MPa and 175 MPa respectively. The normal stress near the film-cooling tube on the BC/SUB interface is 5 – 7 times the one at TC/BC interface. Therefore, the interface crack is more likely to initiate at the BC/SUB interface, and the crack may keep growing and cause the spalling of TBC.

Modern Micro Gas Turbines must be capable to operate at different load points, in order to fulfill the demand of Combined Heat and Power for which they are designed. The combustion chamber structures are therefore subjected to regularly variable thermal loads, yet remaining physically constrained at the rest of the structure. Hence, they experience variable metal temperatures, temperature gradients and thermal stresses which can lead to thermal failure. Typical failure mechanisms in combustion chambers are fatigue and creep. Oxidation can also play an important role. In the present study, Computational Fluid Dynamics methods are used for validation of flow prediction, combustion and heat transfer of an atmospheric combustion chamber. Convective and radiative heat losses towards the ambient are specifically taken into account, leading to better agreement with experimental data from preceding studies. The comparison is presented in this paper. The real-scale machine-operated combustion chamber is then tested at its nominal high-pressure conditions. In this respect, the hereby validated numerical models are employed to simulate the high-pressure operation. A fully coupled Thermo-Structural analysis is performed in order to account for heat fluxes within the solid materials. By so doing, wall temperature distributions can be obtained. The results from Fluid Dynamics simulations serve as input for a Finite Elements Analysis, which provides equivalent stress, strain and deformation distributions by using a linear elastic mechanical model. Such distributions highlight the most critical areas, allowing a first estimate of the components’ life according to thermo-mechanical fatigue. The additional influence of Creep and Oxidation is currently under development at DLR Stuttgart and will be presented in subsequent works.

Conventional power plants are obliged to compensate for the fluctuations in power generation, due to the rising amount of renewable energies, to ensure grid stability. Consequently, steam turbines are more frequently facing load variation and startup/shut-down cycles leading to an increase of thermal stress induced by phase change phenomena. The review of existing test facilities providing measurement data of heat transfer coefficients influenced by multiphase phenomena, such as surface wettability and dry-out, revealed the necessity for a new measurement application. This paper presents the design of the Experimental Multi-phase Measurement Application “EMMA” to generate the required conditions in combination with an academic turbine housing geometry. The performed investigations are focused on the local distribution of heat transfer coefficients (HTC) and the surface wettability affected by phase change phenomena. Two main film formation mechanisms can be observed, depending on the thermal gradient between the fluid and the wall. These are a) saturated/superheated steam in contact with a sub-cooled wall leading to film-wise/drop-wise condensation and b) primary condensed wet steam droplets depositing on a superheated wall, leading to evaporation. Both, the liquid film and the local heat transfer are measured simultaneously. An overview of applicable thickness measurement methods for transparent liquid films is given and the applied optical measurement system is further described. Moreover the HTC measurement methods are presented considering the occurring case of phase change.

Solid oxide fuel cells integrated with gas turbine (SOFC-GT) systems are considered among the most promising power generation units, not only because of the high efficiency, low emissions and carbon capture ability, but also the flexibility to use different kinds of fuels such as natural gas, syngas and biogas directly. In the case of natural gas, Previous researches have demonstrated that solid oxide fuel cells possess the ability to utilize natural gas directly by reforming it inside the anode because of the high operating temperature. But the major problem of internal reforming is that it increases the temperature gradient at the leading edge of fuel cell which may lead to high thermal stress and damage the cells. On the other side, external reforming requires an additional reformer outside of fuel cell, which may increase the investment costs. Also, the amount of air needed to cool the fuel cell is doubled, compared with internal reforming. A full comparison between internal reforming and external reforming of the pressurized SOFC is needed for the hybrids application. In this paper, a real time equilibrium reformer model based on minimization of Gibbs free energy was built to couple with 1D real time solid oxide fuel cell model. An internal on-anode reforming SOFC stack configuration for hybrid SOFC-GT system application was compared with external reforming configurations with 800K, 900K and 1000K reforming temperatures. The results show that internal reforming provides better performance of SOFC stack in the case of high fuel utilization. However, the external reforming showed a higher stack efficiency and smaller stack size compared with on-anode reforming when keeping a relatively lower SOFC stack fuel utilization, necessarily for high hybrid efficiency. Results indicated that external and internal reforming of fuel needs to be optimized depending on different design conditions of the entire hybrid system in terms of efficiency and investment cost. This paper shows that the hybrid system provides the opportunities for thermal integration on performance and efficiency improvement over fuel cell anode reforming.

The main function of labyrinth seal is to control leakage flow in clearance that involves with rotating and stationary parts. Therefore, the effective of clearance gap in labyrinth seal is critical to sealing, heat transfer and vibration characteristics. However, due to the mechanical expansions, vibrations, thermal stress, misalignment of seal components in transient periods of startup, shutdown and hot restart, the stationary and rotating parts of the labyrinth seal are likely to contact each other, causing wear damages in labyrinth fin. Mushrooming damages are often occurred in the rubbing events when labyrinth fin is made of soft material compared with the opposite component. To investigate how mushrooming damage affects the leakage performance of labyrinth seal, many numerical and experimental studies have been carried out in last decades. However, little attention has been paid on the influence of labyrinth fin mushrooming on aerodynamic performance of turbine stages. In this study, using the RANS equations solution methods, the effect of labyrinth fin mushrooming on isentropic efficiency, leakage rates, outlet flow angles, reaction degrees, profile static pressure distributions and flow fields in turbine stages were investigated at three different mushrooming radii and three effective clearances. It shows the leakage rate is increased with increasing the mushroom radius and effective clearance. At the same effective clearance, as the mushrooming radius increases from 0.2mm to 0.4mm, the leakage rate is increased by about 0.19–0.32%, and the overall isentropic efficiency is decreased by 0.78%. At the same mushrooming radius, as the effective clearance increases from 1mm to 1.4mm, the leakage rate is increased by 0.21–0.31%, and the overall isentropic efficiency is decreased by 0.65%. As mushroom radius and effective clearance increase, the secondary flows near hub and shroud are intensified and developed along axial direction, causing pronounced aerodynamic loss in turbine stages.

Better life assessment of hot-components of an aero-engine can help improve its reliability and service life, while, reducing associated maintenance cost. Accurate prediction of Thermo-Mechanical Fatigue (TMF) is one of the crucial aspects of life prediction. Therefore, fully resolved simulation methodologies have gained attention as an ingredient for solving TMF problems owing to their potential for providing comprehensive insights into a system having hot components undergoing transient loading during operation. The present work focuses on a multi-physics simulation-based approach for the life-prediction of a representative gas-turbine combustor liner with an objective of providing a complete framework for TMF analysis of an actual aero-engine combustor liner. The presented methodology consists of a coupling between Computational Fluid Dynamics (CFD) and Finite Element Method (FEM). Thermal loads on the representative aero-engine combustor are predicted using Conjugate Heat Transfer (CHT) modeling in the CFD analyses for different operating conditions suitable for a flight cycle. A load cycle is then constructed using these thermal loads and is transferred to the structural analysis to evaluate the stresses in the liner. Results are obtained regarding spatially varying thermal expansion resulting in inelastic strains as governed by temperature and rate dependent material behavior. Stress and plastic strain history information from the structural analysis are processed to predict the life of different regions of the combustor liner. Different simulation methods for conjugate heat-transfer, load-cycle, material property extraction, thermal-stresses, and fatigue are evaluated, and an overall methodology involving accuracy and reasonable computational cost is proposed. The proposed methodology is numerically verified, and the verification results are presented in this work.

Pressure surface film cooling from discrete holes can often be challenging due to higher than optimum coolant to surface pressure ratios, effects of high levels of flow field turbulence, and the potential for clogging. Double wall cooling methods can be designed to collect spent cooling air and distribute the film cooling downstream through a slot. Incremental impingement is a new internal cooling method designed for cooling the leading edge region and pressure surface. Internally, incremental impingement includes high solidity pedestals to conduct heat and transmit thermal stresses due to temperature variations between cold and hot side surfaces. Subsequently, the flow is collected downstream from the last row of pedestals and discharged through a slot. Experimental and computational research from mesh slots, which have dense arrays of pedestals upstream from the discharge, and slots downstream from high solidity pedestal arrays have shown that turbulence and vorticity generated inside a film cooling plenum can have a significant impact on downstream film cooling. This impact of plenum flow disturbances is in addition to the film cooling dissipation caused by external flow field turbulence. Incremental impingement, in addition to high solidity pedestals, has impingement jets integrated behind the last row of pedestals which may cause further disruption to the film discharge and flow field interaction. The present measurements document the film cooling effectiveness distributions downstream from a slot located at 62% arc along the pressure surface of a vane. The plenum has been designed to include high solidity pedestals and impingement jets consistent with an incremental impingement geometry. Blowing ratios of 0.4, 0.7 and 1.0 have been investigated at vane exit chord Reynolds numbers of 500,000, 1,000,000 and 2,000,000 at density ratios a little over 1. These conditions have been run at 5 independent turbulence levels ranging from 0.7% to over 17%. The results provide a consistent picture of pressure surface slot film cooling downstream from incremental impingement.

An experimental study is carried out to analyze effects of swirl on the structure and stabilization of methane non-premixed oxygen-enriched flames above a co-axial injector in which the two streams are eventually swirled. The mean position of the flame and the liftoff height above the injector lips are investigated with OH* chemiluminescence images. The oxygen enrichement, the momentum flux ratio between the two co-flows, the swirl level inside the central jet and the swirl level in the annular jet are varied over a large range of operating conditions. It is found that, in the absence of swirl in the central stream, the flame is always attached to the lips of the internal injection tube. As the inner swirl level increases, the flame front located at the lips of the internal injection tube disappears. When the annular swirl level is high enough to create a central recirculating bubble, the flame detaches from the nozzle rim and remains lifted at a finite distance above the injector. Increasing the oxygen concentration shifts this transition to smaller momentum flux ratios and smaller annular swirl levels. The lift-off distance can be finely tuned and depends on the inner and outer swirl levels, and on the momentum flux ratio between the two coaxial streams. It is shown that this feature neither depends on the confinement of the injector nor on the thermal stress exerted by the hot burnt gases on the injector back plane. About 1000 configurations were investigated that could be classified into only four distinct stabilization modes, in which the flame structure was shown to follow a similar pathway when the momentum flux ratio between the two streams, the swirl level in the central and external streams and the quarl angle of the annular stream are varied. It is finally shown how these limits are altered when the oxygen concentration in the annular oxidizer stream is varied from air to oxygen-enriched operation.

Leakage flow between the rotating turbine blade tip and the fixed casing causes high heat loads and thermal stress on the tip and near the tip region. For this study, new squealer tips called partial cavity tips, which combine the advantages of plane and squealer tips, were suggested, and the effects of the cavity shape on the tip heat transfer coefficient and film cooling effectiveness were investigated experimentally in a low speed linear cascade. The suggested blade tips had a flat surface near the leading edge and a squealer cavity from the mid-chord to trailing edge region to achieve the advantages of both blade tip types. The heat transfer coefficient was measured via the 1-D transient heat transfer technique using an IR camera, and the film cooling effectiveness was obtained via the pressure sensitive paint (PSP) technique. Results showed that the heat transfer coefficient and film cooling effectiveness on the partial cavity tips strongly depended on the cavity shape. Near the leading edge, the heat transfer coefficients for the partial cavity tip cases were lower than that for the squealer tip case. However, the heat transfer coefficient on the cavity surface was higher for the partial cavity tip cases. The D10 tip showed a similar distribution of film cooling effectiveness to that of the PLN tip near the leading edge and the DSS tip near the mid-chord region. However, the overall averaged film cooling effectiveness of the DSS tip was higher than that of the D10 tip.