Film Cooled Turbine Vane Research Articles

Numerical assessment of thermal insulation and stress responses in film-cooled turbine vane thermal barrier coatings under CMAS deposition conditions

Deposition of CaO-MgO-Al2O3-SiO2 (CMAS) significantly contributes to the spalling of thermal barrier coatings (TBCs) on turbine vanes. A thorough understanding of the thermodynamic properties during CMAS deposition is critical for advancing the lifetime design of TBCs. This study employs the CMAS gas thermal shock test to determine the deposition characteristics, which were then analyzed using the critical velocity model. The results of this analysis align closely with experimental outcomes. Based on the numerical simulation of the fluid-solid coupling method, we further explored the insulation efficiency and stress distribution in TBCs under CMAS deposition conditions. It was observed that CMAS predominantly accumulates on the pressure side and within the film pores of the vane TBCs, with minimal deposition on the suction side. Such deposition patterns result in an increased overall temperature of the vane, concurrently diminishing the TBCs' insulation efficiency. Specifically, CMAS deposition raised the maximum surface temperature of the vane by 100 K and decreased the peak insulation performance of the TBCs by 16 %. Additionally, the deposition induced higher stresses within both the TBCs and the underlying vane substrate, with a 7 % increase in the maximum principal stresses at the TBC surface and a 6 % increase in the substrate. Consequently, under CMAS deposition conditions, TBCs in regions of low insulation efficiency and high stress on turbine vanes are prone to cracking and subsequent spallation.

Impact of Predicted Combustor Outlet Conditions on the Aerothermal Performance of Film-Cooled High Pressure Turbine Vanes

Turbine inlet conditions in lean-burn aeroengine combustors are highly swirled and present nonuniform temperature distributions. Uncertainty and lack of confidence associated with combustor-turbine interaction affect significantly engine performance and efficiency. It is well known that only Large-eddy and scale-adaptive simulations (SAS) can overcome the limitations of Reynolds-averaged Navier–Stokes (RANS) in predicting the combustor outlet conditions. However, it is worth investigating the impact of such improvements on the predicted aerothermal performance of the nozzle guide vanes (NGVs), usually studied with RANS-generated boundary conditions. Three numerical modelling strategies were used to investigate a combustor-turbine module designed within the EU Project FACTOR: (i) RANS model of the NGVs with RANS-generated inlet conditions; (ii) RANS model of the NGVs with scale-adaptive simulation (SAS)-generated inlet conditions; (iii) SAS model inclusive of both combustor and NGVs. It was shown that estimating the aerodynamics through the NGVs does not demand particularly complex approaches, in contrast to situations where turbulent mixing is key. High-fidelity predictions of the turbine entrance conditions proved very beneficial to reduce the discrepancies in the estimation of adiabatic temperature distributions. However, a further leap forward can be achieved with an integrated simulation, capable of reproducing the transport of unsteady fluctuations generated from the combustor through the turbine, which play a key role in presence of film cooling. This work, therefore, shows how separate analysis of combustor and NGVs can lead to a poor estimation of the thermal loads and ultimately to a wrong thermal design of the cooling system.

Experimental investigation of wall thickness and hole shape variation effects on full-coverage film cooling performance for a gas turbine vane

The effects of wall thickness and hole shape variation on a full-coverage film cooled turbine vane are investigated in a stationary and linear cascade utilizing the pressure sensitive paint technique. The varied wall thickness produces hole length-to-diameter ratio (L/D) in a range from L/D = 2 to 5, and holes tested include simple angle hole, compound angle hole, and fan-shaped hole. Five rows of holes are provided on the pressure side while three rows of holes are provided on the suction side, with six rows of cylindrical holes drilled on the leading edge to construct showerhead film cooling. The tested blowing ratios for the showerhead, pressure side, and suction side range from 0.25 to 1.5, with a density ratio of 1.5. The freestream Reynolds number is 1.35 × 105, based on the axial chord length and the inlet velocity, with a freestream turbulence intensity level of 3.5% at the cascade inlet. The results indicate that the wall thickness variation produces significant influence on the pressure side film cooling effectiveness, while only marginal effect on the showerhead and suction side film cooling. Also observed is that the fan-shaped hole generates the highest film cooling effectiveness on pressure or suction side. Also discussed is the surface curvature effect, combining with effects of wall thickness and hole shape variations, on the film cooling effectiveness in comparison to the flat-plate data.

Multi-step optimizations of leading edge and downstream film cooling configurations on a high pressure turbine vane

Multi-step optimizations of leading edge and downstream film cooling configurations on a high pressure turbine vane are conducted using full three-dimensional optimization method. Firstly, a full three-dimensional optimization platform for film-cooled turbines, which consists of the parametric modeling, automatic mesh generation, the CFD numerical calculation and the optimization strategy is established. On the basis of it, the optimization platform is used to optimize the configurations of multi-rows film cooling holes for C3X gas turbine vane cooled with leading edge showerhead film cooling, and the layout of rows of film cooling holes on pressure side and suction side is optimized for the further step optimization. The sensitivity analysis for film cooling performance is used to decide the final design variables, and the adiabatic film cooling effectiveness is adopted to evaluate the overall performance of film-cooled turbine vane. The results show that the average film cooling effectiveness of the optimized leading edge region, pressure surface and suction surface are increased by 8.4%, 1.1% and 5.5% compared with the original case, while the average temperature of the leading edge region is decreased by 19 K.

Aero-thermal optimization on multi-rows film cooling of a realistic marine high pressure turbine vane

Aero-thermal optimizations on multi-rows film cooling of a realistic marine high pressure turbine vane are performed in this paper. Firstly, a full three-dimensional optimization platform for film-cooled turbines, which consists of the parametric modeling, automatic mesh generation, the CFD numerical calculation and the optimization strategy is established. The design variables are selected based on the considerations of hole-to-hole interactions and invalid coverage of cooling film. Two objective functions of aerodynamic efficiency and adiabatic film cooling effectiveness are adopted to evaluate the overall performance of film-cooled turbine vane. Bothe single objective and multi-objective optimization are adopted to optimize the configurations of multi rows film cooling holes, and the Multi-Island Genetic Algorithm and The response surface approximation with the Non-dominated Sorting Genetic Algorithm (NSGA-II) are selected to conduct the optimizations. The objective of the present study is to use a numerical optimization method to find the optimal cooling structure configurations that yields better cooling performance with an acceptable aerodynamic penalty. The results show that the film cooling effectiveness of multi-objective optimization is increased by 11.4%, and average temperature of the blade surface is reduced by 75K.

Microstructural Degradation of Thermal Barrier Coatings on an Integrated Gasification Combined Cycle (IGCC) Simulated Film-Cooled Turbine Vane Pressure Surface Due to Particulate Fly Ash Deposition

Research is being conducted to study the degradation of thermal barrier coatings (TBC) employed on IGCC turbine hot section airfoils due to particulate deposition from contaminants in coal syn-thesis gas (syngas). West Virginia University (WVU) had been working with US Department of Energy, National Energy Technology Laboratory (NETL) to simulate deposition on the pressure side of an IGCC turbine first stage vane. To simulate the contaminant deposition, several TBC coated, angled film-cooled test articles were subjected to accelerated coal fly ash, which was injected into the flow of a combustor facility with a high pressure (approximately 4 atm) and a high temperature (1560 K) environment. To investigate the degradation of the TBCs due to particulate deposition, non-destructive tests were performed using scanning electron microscopy (SEM) evaluation and energy dispersive X-ray spectroscopy (EDS) examinations. The SEM evaluation was used to display the microstructure change within the layers of the TBC system directly related to the fly ash deposition. The SEM micrographs showed that deposition-TBC interaction made the YSZ coating more susceptible to delamination and promoted a dissolution-reprecipitation mechanism that changed the YSZ morphology and composition. The EDS examination provided elemental maps of the shallow infiltration depth of the fly ash and chemical composition spectrum results which showed yttria migration from the YSZ into the deposition.

Evaluation of CFD simulations of film cooling performance on a turbine vane including conjugate heat transfer effects

Computational simulations using a steady RANS approach with the k–ω SST turbulence model were performed to complement experimental measurements of overall cooling effectiveness and adiabatic film effectiveness for a film cooled turbine vane airfoil. The vane included a single row of holes located on the suction side of the airfoil. The simulated geometry also included the internal impingement cooling configuration. Internal and external boundary conditions were matched to experiments using the same vane model. To correctly simulate conjugate heat transfer effects, the experimental vane model was constructed to match the Biot number for engine conditions. Computational predictions of the overall cooling effectiveness and adiabatic film effectiveness were compared to experimental measurements. The CFD predictions showed that the k–ω SST RANS model over-predicted local adiabatic film effectiveness for an attached jet, while performance was under-predicted for a detached jet. The corresponding predictions of overall cooling effectiveness were also over and under-predicted. Further, it was shown that the adiabatic wall temperature was not the correct driving temperature for heat transfer, especially in the case of a detached jet.

Effects of Syngas Particulate Fly Ash Deposition on the Mechanical Properties of Thermal Barrier Coatings on Simulated Film-Cooled Turbine Vane Components

Research is being conducted to study the effects of particulate deposition from contaminants in coal synthesis gas (syngas) on the mechanical properties of thermal barrier coatings (TBC) employed on integrated gasification combined cycle (IGCC) turbine hot section airfoils. West Virginia University (WVU) had been working with US Department of Energy, National Energy Technology Laboratory (NETL) to simulate deposition on the pressure side of an IGCC turbine first stage vane. To model the deposition, coal fly ash was injected into the flow of a combustor facility and deposited onto TBC coated, angled film-cooled test articles in a high pressure (approximately 4 atm) and a high temperature (1560 K) environment. To investigate the interaction between the deposition and the TBC, a load-based multiple-partial unloading micro-indentation technique was used to quantitatively evaluate the mechanical properties of materials. The indentation results showed the Young’s Modulus of the ceramic top coat was higher in areas with deposition formation due to the penetration of the fly ash. This corresponds with the reduction of strain tolerance of the 7% yttria-stabilized zirconia (7YSZ) coatings.

Experimental Simulation of Contaminant Deposition on a Film-Cooled Turbine Vane Pressure Side With a Trench

An important issue in the use of coal- or biomass-derived synthetic gaseous (syngas) fuels is the deposition of contaminants on film-cooled turbine surfaces, which alter cooling and aerodynamic performance and increase material degradation. The current study applied a new experimental technique that simulated the key physical aspects of contaminant deposition on a film-cooled turbine vane. The depositing contaminants were modeled in a wind tunnel facility with a spray of molten wax droplets of a size range that matched the Stokes number of the contaminant particles in engine conditions. Most experiments were performed using a vane model with a thermal conductivity selected such that the model had the same Biot number of an actual engine airfoil, resulting in a cooler surface temperature. Some experiments were performed using an approximately adiabatic model for comparison. The film cooling design consisted of three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the pressure side. Two designs of pressure side body film cooling holes were considered: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. The results showed thin deposits formed in the trench, with the thickest deposits on its downstream wall between coolant jets. Adiabatic film effectiveness levels were essentially unchanged by the presence of deposits for either film configuration. Deposit formation was strongly influenced by the model surface temperature with cooler surfaces inhibiting deposition. There was evidence of a threshold surface temperature above which deposits became significantly thicker.

A Novel Technique for Assessing Turbine Cooling System Performance

A new experimental technique for the accurate measurement of steady-state metal temperature surface distributions of modern heavily film-cooled turbine vanes has been developed and is described in this paper. The technique is analogous to the thermal paint test but has been designed for fundamental research. The experimental facility consists of an annular sector cascade of high pressure (HP) turbine vanes from a current production engine. Flow conditioning is achieved by using an annular sector of deswirl vanes downstream of the test section, being both connected by a three-dimensionally contoured duct. As a result, a transonic and periodic flow, highly representative of the engine aerodynamic field, is established: Inlet turbulence levels, mainstream Mach and Reynolds numbers, and coolant-to-mainstream total pressure ratio are matched. Since the fully three-dimensional nozzle guide vane (NGV) geometry is used, the correct radial pressure gradient and secondary flow development are simulated and the cooling flow redistribution is engine-realistic. To allow heat transfer measurements to be performed, a mainstream-to-coolant temperature difference (up to 33.5°C) is generated by using two steel-wire mesh heaters, operated in series. NGV surface metal temperatures are measured (between 20°C and 40°C) by wide-band thermochromic liquid crystals. These are calibrated in situ and on a per-pixel basis against vane surface thermocouples, in a heating process that spans the entire color play and during which the turbine vanes can be assumed to slowly follow a succession of isothermal states. Experimental surface distributions of overall cooling effectiveness are presented in this paper. By employing resin vanes of the same geometry and cooling configuration (to implement adiabatic wall thermal boundary conditions) and the transient liquid crystal technique, surface distributions of external heat transfer coefficient and film cooling effectiveness can be acquired. By combining these measurements with those from the metal vanes, the results can be scaled to engine conditions with a good level of accuracy.

Conjugate heat transfer with Large Eddy Simulation for gas turbine components

CHT (Conjugate Heat Transfer) is a main design constraint for GT (gas turbines). Most existing CHT tools are developed for chained, steady phenomena. A fully parallel environment for CHT has been developed and applied to two configurations of interest for the design of GT. A reactive Large Eddy Simulations code and a solid conduction solver exchange data via a supervisor. A flame/wall interaction is used to assess the precision and the order of the coupled solutions. A film-cooled turbine vane is then studied. Thermal conduction in the blade implies lower wall temperature than adiabatic results and CHT reproduces the experimental cooling efficiency. To cite this article: F. Duchaine et al., C. R. Mecanique 337 (2009).

Three Component Velocity Field Measurements in the Stagnation Region of a Film Cooled Turbine Vane

The showerhead region of a film-cooled turbine vane in a gas turbine engine involves a complex interaction between mainstream flow and coolant jets. This flow field was studied using three component laser Doppler velocimeter measurements in a simulated turbine vane test facility. Measurements were focused around the stagnation row of holes. Low and high mainstream turbulence conditions were used. The spanwise orientation of the coolant jets, typical for showerhead coolant holes, had a dominating effect. Very high levels of turbulence were generated by the mainstream interaction with the coolant jets. Furthermore, this turbulence was highly anisotropic, with the spanwise component of the turbulent fluctuations being twice as large as the other components. Finally, there was an interaction of the high mainstream turbulence with the coolant injection resulting in increased turbulence levels for the spanwise velocity component, but had little effect on the other velocity components.

Thermal Field and Flow Visualization Within the Stagnation Region of a Film-Cooled Turbine Vane

Abstract To develop quality computational codes for the film cooling of a turbine vane, a detailed understanding is needed of the physical mechanisms of the mainstream-coolant interactions. In this study flow visualization, thermal profiles, and laser Doppler velocimetry measurements were used to define the thermal and velocity fields of the film cooled showerhead region of a turbine vane. The showerhead consisted of six rows of spanwise oriented coolant holes, and blowing ratios ranged from 0.5 to 2.0. Performances with low and high mainstream turbulence levels were tested. Coolant jets from the showerhead were completely separated from the surface even at relatively low blowing ratios. However, the interaction of the coolant jets from laterally adjacent holes created a barrier to the mainstream flow, resulting in relatively high adiabatic effectiveness.

A Three-Dimensional Coupled Internal/External Simulation of a Film-Cooled Turbine Vane

A three-dimensional Navier–Stokes simulation has been performed for a realistic film-cooled turbine vane using the LeRC-HT code. The simulation includes the flow regions inside the coolant plena and film cooling holes in addition to the external flow. The vane is the subject of an upcoming NASA Lewis Research Center experiment and has both circular cross-sectional and shaped film cooling holes. This complex geometry is modeled using a multiblock grid, which accurately discretizes the actual vane geometry including shaped holes. The simulation matches operating conditions for the planned experiment and assumes periodicity in the spanwise direction on the scale of one pitch of the film cooling hole pattern. Two computations were performed for different isothermal wall temperatures, allowing independent determination of heat transfer coefficients and film effectiveness values. The results indicate separate localized regions of high heat flux in the showerhead region due to low film effectiveness and high heat transfer coefficient values, while the shaped holes provide a reduction in heat flux through both parameters. Hole exit data indicate rather simple skewed profiles for the round holes, but complex profiles for the shaped holes with mass fluxes skewed strongly toward their leading edges. [S0889-504X(00)02802-6]

LEADING EDGE FILM-COOLING EFFECTS ON TURBINE BLADE HEAT TRANSFER

A three-dimensional Navier-Stokes code has been used to study the effect ofspanwise pitch of shower-head holes and coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat transfer coefficient on a film-cooled turbine vane. The mainstream is akin to that under real engine conditions with stagnation temperature 1900 K and stagnation pressure 3 MPa. It is found that with the ratio of coolant to mainstream mass flow being fixed, reducing the spanwise pitch for shower-head holes (P) from 7.5d to 3.0d, where d is the hole diameter, increases the average effectiveness considerably over the blade surface. However, when P / d = 7.5, increasing the coolant mass flow increases the effectiveness on the pressure surface but reduces it on the suction surface owing to coolant jet lift-off. For P / d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant-to-mainstream-mass-flow ratios analyzed. In all cases, adiabatic effectiveness and heat transfer coefficient are highly three-dim...