High Pressure Turbine Vane Research Articles

EBFOG: Deposition, Erosion, and Detachment on High-Pressure Turbine Vanes

Fouling and erosion are two pressing problems that severely affect gas turbine performance and life. When aircraft fly through a volcanic ash cloud, the two phenomena occur simultaneously in the cold as well as in the hot section of the engine. In the high-pressure turbine (HPT), in particular, particles soften or melt due to the high gas temperatures and stick to the wet surfaces. The throat area, and hence the capacity, of the HPT is modified by these phenomena, affecting the engine stability and possibly forcing engine shutdown. This work presents a model for deposition and erosion in gas turbines and its implementation in a three-dimensional Navier–Stokes solver. Both deposition and erosion are taken into account, together with deposit detachment due to changed flow conditions. The model is based on a statistical description of the behavior of softened particles. The particles can stick to the surface or can bounce away, eroding the material. The sticking prediction relies on the authors' Energy Based FOulinG (EBFOG) model. The impinging particles which do not stick to the surface are responsible for the removal of material. The model is demonstrated on a HPT vane. The airfoil shape evolution over the exposure time as a consequence of the impinging particles has been carefully monitored. The variation of the flow field as a consequence of the geometrical changes is reported as an important piece of on-board information for the flight crew.

Computational investigation of film cooling and secondary flow on turbine endwall with coolant injection from upstream interrupted slot

High pressure turbine vane surface and endwall regions are extensively cooled through discrete holes and leakage flow from combustor-turbine interface gap. For making better use of the limited amount of the leakage flow, this paper describes numerical investigation of endwall film cooling performance of a two-dimensional cascade with upstream interrupted slot injection. The geometry of the vane, size of the slot and mainstream parameters are all taken from a real engine high pressure turbine. The effects of varying blowing ratio, location of the upstream slot and coolant incidence angle on cooling effectiveness are studied. The ranges of the studied parameters are: blowing ratio 1.0, 1.7, 1.86; slot location Z/Cax = −0.1, −0.2, −0.3; coolant incidence angle −10°, 10°, 20°. The calculations are completed by solving three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with shear stress transport (SST) k-ω turbulence model, meanwhile, the turbulence model was validated by comparing the calculated results with the experiment data. The calculated results show an important influence of blowing ratio and axial position of the interrupted slot on film cooling effectiveness. Cooling Effectiveness is increased with increasing blowing ratio and decreasing distance between the slot and leading edge of the vane. To compare with BR = 1.7 cooling effectiveness can be improved significantly in the condition of BR = 1.86. Slot location of Z/Cax = −0.1 provides much higher cooling effectiveness than the rest two locations. The physical mechanism of the improvement is that in the condition of high blowing ratio high momentum coolant flow helps to reduce the strength of horseshoe vortex and therefore to limit its negative effect on the cooling effectiveness. The high momentum coolant jet impinges on leading edge of the vane and climbs the surface of it. The coverage of the coolant at the leading edge endwall and pressure side endwall junction becomes better. The contours of effectiveness for different slot location are similar to those of different blowing ratio. The momentum of the coolant jet increases with the slot moving close to the leading edge of the vane and creates similar effectiveness contour patterns of raising blowing ratio. The effect of varying coolant incidence angle on the cooling effectiveness is weaker than blowing ratio and slot location. Average cooling effectiveness is increased about 14% when the angle changed from 10° to 20°.

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.

DDES Analysis of the Wake Vortex Related Unsteadiness and Losses in the Environment of a High-Pressure Turbine Stage

In this work, the flows inside a high-pressure turbine (HPT) vane and stage are studied with a delayed detached eddy simulation (DDES) code. The fundamental nozzle/blade interaction is investigated with special attention paid to the development and transportation of the vane wake vortices. There are two motivations for this work. First, the extreme HPT operation conditions, including both transonic Mach numbers and high Reynolds numbers, impose a great challenge to modern computational fluid dynamics (CFD), especially for scale-resolved simulation methods. An accurate and efficient high-fidelity CFD solver is very important for a thorough understanding of the flow physics and the design of more efficient HPT. Second, the periodic wake vortex shedding is an important origin of turbine losses and unsteadiness. The wake and vortices not only cause losses themselves, but also interact with the shock wave (under transonic working condition), pressure waves, and have a strong impact on the downstream blade surface (affecting boundary layer transition and heat transfer). Based on one of our previous DDES simulations of a HPT vane, this work further investigates the development and length characteristics of the wake vortices, provides explanations for the length characteristics, and reveals the transportation of the wake vortices in the downstream rotor passages along with its impact on the downstream aero-thermal performance.

The Effect of Hot Streaks on a High Pressure Turbine Vane Cascade with Showerhead Film Cooling

Hot streak migration in a linear vane cascade with showerhead film cooling was experimentally and numerically investigated at isentropic exit Mach number of Ma2is = 0.40, with an inlet turbulence intensity level of Tu1 = 9%. Two tangential positions of the hot streak center were taken into account: 0% of pitch (hot streak is aligned with the vane leading edge) and 45% of pitch. After demonstrating that computations correctly predict hot streak attenuation through the vane passage with no showerhead blowing, the numerical method was used to investigate hot streak interaction with showerhead film cooling, at blowing ratio of BR = 3.0, corresponding to a coolant-to-mainstream mass flow ratio of MFR = 1.15%. The effects of mixing and coolant interaction on the hot streak reduction were interpreted under the light of the superposition principle, whose accuracy was within 12% on the leading edge region, in the central section of the vane span.

Non-ideal effects on the typical trailing edge shock pattern of ORC turbine blades

At the trailing edge of supersonic high-pressure turbine vanes, a typical shock pattern, the so-called fish-tail shocks, originates due to the flow rotation imposed by its finite thickness. In addition, shock and shock/fan systems can arise in case of a post-expanded channel design or at off-design conditions. ORC turbine stator blades are particularly prone to this phenomena since they are designed to provide a high outlet Mach number, especially at the first stage. In the close proximity of the saturation curve, near the critical point, molecularly complex organic fluids for ORC applications may exhibit a number of non-ideal gasdynamic effects, including a remarkable dependency of the shock properties on the upstream thermodynamic state of the fluid, at a fixed upstream Mach number. The influence of thermodynamic conditions on the shock pattern is assessed as a function of the flow deviation and compared against the ideal gas case, for which the shock properties depends on the upstream Mach number only. Non-ideal effects are investigated here using siloxane vapor MDM (Octamethyltrisiloxane, C8H24O2Si3), as an exemplary organic fluid. The present results can be arguably extended to most vapors currently employed in ORC applications.

Experimental Investigation of Aerodynamic Performance of a Turbine Cascade with Trailing-Edge Injection

AbstractTrailing-edge mixing flows associated with coolant injection are complex and result in significant aerodynamics losses. A series of tests on a high-pressure turbine vane cascade with traili...

Effects of Coolant Density, Specific Heat Capacity, and Biot Number on Turbine Vane Cooling Effectiveness

This paper describes the effects of coolant-to-mainstream density ratio and specific heat capacity flux ratio (the product of blowing ratio and specific heat capacity ratio) on the overall cooling effectiveness of high pressure (HP) turbine vanes. Experimental measurements have been conducted at correct engine-matched conditions of Mach number, Reynolds number, turbulence intensity, and coolant-to-mainstream momentum flux ratio. Vanes tested were fully cooled production parts from an engine currently in service. A foreign gas mixture of SF6 and Ar was selected for injection as coolant in the facility so that density and blowing ratios were also matched to the engine situation. The isentropic exponent of the foreign gas mixture coincides with that of air. Full-coverage surface maps of overall cooling effectiveness were acquired by an infrared (IR) thermography technique at a range of mainstream-to-coolant temperature ratios. Measurements were subsequently scaled to engine conditions by employing a new theory based on the principle of superposition and a recovery and redistribution temperature demonstrated in previous papers. It is shown that the two aerodynamically matched situations of air- and foreign-gas-cooled experiments give virtually the same effectiveness trends and patterns. Actual levels differ, however, on account of specific heat capacity flux ratio differences. The effect is described and quantified by a one-dimensional analytical model of the vane wall. Differences in Biot number with respect to engine conditions are discussed as they also influence the scaling of turbine metal temperatures.

Linear Stability Prediction of Vortex Structures on High Pressure Turbine Blades

Velocity profiles are extracted from time- and span-averaged direct numerical simulation data, describing the flow over a high-pressure turbine vane linear cascade near engine-scale conditions with reduced inlet disturbance levels. Based on these velocity profiles, local as well as non-local linear stability analysis of the boundary-layer over the suction side of the vane is carried out in order to characterise a linearly unstable region close to the trailing edge. The largest growth rates are found for oblique modes, but those are only slightly more unstable than 2D modes, which describe the locations and frequencies of most unstable modes very well. The frequencies of the most unstable linear modes predict with good accuracy the predominant frequencies found in the direct numerical simulations (DNS) close to the trailing edge.

Scaling of Turbine Metal Temperatures in Cooled Compressible Flows—Experimental Demonstration of a New Theory

Experimental measurements of overall cooling effectiveness conducted on a high-pressure turbine vane in a warm rig flow are scaled to engine conditions in this paper. A new theory for the scaling of turbine metal temperatures in cooled compressible flows has been applied, based on the principle of superposition, and demonstrated analytically and numerically in a previous paper. The analysis employs a definition of overall cooling effectiveness based on a new recovery and redistribution temperature, which makes it independent of the temperature boundary conditions of the hot and cold flow streams. This enables the vane external wall temperatures to be scaled to engine conditions by varying, in a fixed aerodynamic field, the mainstream-to-coolant temperature ratio. Experimental validation of the theory is provided in this article. Measurements were conducted in the Annular Sector Heat Transfer Facility, which employs fully cooled nozzle guide vanes, production parts of a civil aviation engine currently in service. Mainstream Mach and Reynolds numbers, inlet turbulence intensity, and coolant-to-mainstream total pressure ratio (and thus momentum flux ratio) are all matched to engine conditions. Full-coverage overall cooling effectiveness distributions, acquired by infrared thermography, are presented for a range of mainstream-to-coolant temperature ratios between 1.05 and 1.22 and subsequently scaled to engine conditions by an iterative procedure. In reducing to practice the principles of the new scaling theory, it is demonstrated that direct validation of turbine cooling system performance is possible in experiments at lower than engine temperatures.

Experimental and numerical investigations of hole injection on the suction side throat of transonic turbine vanes in a cascade with trailing edge injection

Trailing-edge mixing flows associated with coolant injection are complex, in particular at transonic flows, and result in significant aerodynamics losses. The objective of this paper is to evaluate the impacts of hole injection near the suction side throat on shock wave control and aerodynamic losses. A series of tests and calculations on effects of hole injection on the suction-side throat of a high-pressure turbine vane cascade with and without trailing-edge injection were conducted. Wake traverses with a five-hole probe and tests of pressure distributions on the turbine profile were taken for total injection mass flow ratios of 0% and 1.2% under test Mach numbers of 0.7, 0.78, and 0.87. Meantime, numerical predictions are carried out for exit isentropic Mach numbers of 0.7, 0.78, 0.87, and 1.1 and hole-injection mass flow ratios of 0%, 0.17%, 0.3%, and 0.89%. Numerical predictions show a reasonable agreement with the experimental data, and wake total pressure losses and flow angles as well as pressure distributions on the turbine profile were compared to calculations without hole injection, indicating a significant effect of hole injection on the profile wake development and its blockage effect on the shock-wave flow in the vane cascade passage. At subsonic flows, the hole injection on the suction side throat thickens the suction-side boundary layer, and increases the flow mixing, thus causing increased wake losses and flow angles. At transonic flows, while the trailing-edge injection reduces the strength of the shock wave at the trailing-edge pressure side, the hole injection on the suction side throat alters the local pressure fields, and then tends to enhance the shock-wave at the trailing-edge pressure-side; however, it seems to reduce the strength of the shock-wave at the trailing-edge suction side.

Tensile and Low Cycle Fatigue Behaviour of Cast Superalloy Supercast 247 A in Polycrystalline Condition

Indigenously developed SUPERCAST 247 A is a nickel base cast super alloy, nearly equivalent to CM 247 LC. It is intended for the manufacture of critical aero-engine hot end components that include high pressure turbine blades and vanes, Integral Rotors and Stators of Jet fuel starter of an indigenous aircraft. The present study intends to evaluate the temperature dependent tensile and low cycle fatigue properties of this alloy in the range of 25 to 650°C as a function of microstructure that is obtainable in two specific heat treatment conditions. Gamma prime size and its morphology varied with the heat treatment conditions. Its size is finer in Solutionising + Single Aging (SSA) condition as compared to that in Solutionising + Double Aging (SDA) condition which reflected in high tensile strength values of the material in SSA condition. The alloy in SSA condition exhibited good fatigue resistance at room temperature whereas the alloy in SDA condition showed good fatigue resistance at 650°C.

Numerical Simulation of CM247SX Single Crystal High Pressure Turbine Vane

Numerical simulation of single crystal high pressure turbine vane is carried out with an objective to investigate the effects of process parameters such as (i) process temperature, (ii) mould withdrawal rate, on thermal profiles, mushy zone, cooling rate and temperature gradient of the casting. The grain structure modelling is performed by Cellular Automation Finite Element (CAFE) module of the commercial finite element based software. Scanning electron microscopy of the experimentally cast single crystal vane is characterized. The simulation studies have shown that the increased process temperature and mold withdrawal rate led to increased temperature gradient and cooling rate of the vane. Conditions for the formation of undesirable stray grains is determined.

Film Cooling Effectiveness Downstream of Trailing Edge Slots Including Cutback Surface Protuberances

The trailing edge of the high pressure turbine blade and vane presents significant challenges to the turbine cooling engineer. The current research has focused specifically on the effect of cutback surface protuberance, or “land”, shapes on film cooling effectiveness. A set of six different land geometries has been investigated in a large scale model of the trailing edge pressure side ejection slot exit. Slot height and width and lip height was maintained. Pressure sensitive paint was used to measure adiabatic film cooling effectiveness at five blowing ratios ranging from 0.6 to 1.4 in increments of 0.2. High-resolution full surface distributions of film cooling effectiveness both on the cutback surface and the top of the lands were recorded. It was found that tapering the lands did not significantly increase effectiveness on the lands and slightly reduced effectiveness near the lands. Using a diffuser shape improved average effectiveness greatly and gave the best overall performance up to the end of the lands except at the lowest blowing ratio of 0.6, where having no lands was slightly better.

An Energy-Based Fouling Model for Gas Turbines: EBFOG

Fouling is a major problem in gas turbines for aeropropulsion because the formation of aggregates on the wet surfaces of the machine affects aerodynamic and heat loads. The representation of fouling in computational fluid dynamics (CFD) is based on the evaluation of the sticking probability, i.e., the probability a particle touching a solid surface has to stick to that surface. Two main models are currently available in literature for the evaluation of the sticking coefficient: one is based on a critical threshold for the viscosity, and the other is based on the normal velocity to the surface. However, both models are application specific and lack generality. This work presents an innovative model for the estimation of the sticking probability. This quantity is evaluated by comparing the kinetic energy of the particle with an activation energy which describes the state of the particle. The sticking criterion takes the form of an Arrhenius-type equation. A general formulation for the sticking coefficient is obtained. The method, named energy-based fouling (EBFOG), is the first “energy”-based model presented in the open literature able to account any common deposition effect in gas turbines. The EBFOG model is implemented into a Lagrangian tracking procedure, coupled to a fully three-dimensional CFD solver. Particles are tracked inside the domain, and equations for the momentum and temperature of each particle are solved. The local geometry of the blade is modified accordingly to the deposition rate. The mesh is modified, and the CFD solver updates the flow field. The application of this model to particle deposition in high-pressure turbine vanes is investigated, showing the flexibility of the proposed methodology. The model is particularly important in aircraft engines where the effect of fouling for the turbine, in particular the reduction of the high pressure (HP) nozzle throat area, influences heavily the performance by reducing the core capacity. The energy-based approach is used to quantify the throat area reduction rate and estimate the variation in the compressor operating condition. The compressor operating point as a function of the time spent operating in a harsh environment can be in this way predicted to estimate, for example, the time that an engine can fly in a cloud of volcanic ashes. The impact of fouling on the throat area of the nozzle is quantified for different conditions.

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.

Effect of Low-NOX Combustor Swirl Clocking on Intermediate Turbine Duct Vane Aerodynamics With an Upstream High Pressure Turbine Stage—An Experimental and Computational Study

Flow in an intermediate turbine duct (ITD) is highly complex, influenced by the upstream turbine stage flow structures, which include tip leakage flow and nonuniformities originating from the upstream high pressure turbine (HPT) vane and rotor. The complexity of the flow structures makes predicting them using numerical methods difficult, hence there exists a need for experimental validation. To evaluate the flow through an intermediate turbine duct including a turning vane, experiments were conducted in the Oxford Turbine Research Facility (OTRF). This is a short duration high speed test facility with a 3/4 engine-sized turbine, operating at the correct nondimensional parameters for aerodynamic and heat transfer measurements. The current configuration consists of a high pressure turbine stage and a downstream duct including a turning vane, for use in a counter-rotating turbine configuration. The facility has the ability to simulate low-NOx combustor swirl at the inlet to the turbine stage. This paper presents experimental aerodynamic results taken with three different turbine stage inlet conditions: a uniform inlet flow and two low-NOx swirl profiles (different clocking positions relative to the high pressure turbine vane). To further explain the flow through the 1.5 stage turbine, results from unsteady computational fluid dynamics (CFD) are included. The effect of varying the high pressure turbine vane inlet condition on the total pressure field through the 1.5 stage turbine, the intermediate turbine duct vane loading, and intermediate turbine duct exit condition are discussed and CFD results are compared with experimental data. The different inlet conditions are found to alter the flow exiting the high pressure turbine rotor. This is seen to have local effects on the intermediate turbine duct vane. With the current stator–stator vane count of 32-24, the effect of relative clocking between the two is found to have a larger effect on the aerodynamics in the intermediate turbine duct than the change in the high pressure turbine stage inlet condition. Given the severity of the low-NOx swirl profiles, this is perhaps surprising.

Damage mechanisms in an EB-PVD thermal barrier coating system during TMF and TGMF testing conditions under combustion environment

Modern high pressure turbine blades and vanes of aero-engines are cooled and thermally protected using a thermal barrier coating (TBC) system. TBC systems are submitted to very complex service loadings, with a combination of mechanical and thermal fatigue under complex thermal gradients, in addition to the contribution of oxidation to the damage processes.The high temperature durability under combustion environment of the AM1/NiAlPt/EB-PVD TBC system used for high temperature blades and vanes has been investigated performing out of phase load-controlled thermo-mechanical fatigue (TMF) and thermal-gradient-mechanical fatigue (TGMF) tests in the 500–1100°C temperature range. This characterization has been performed using the MAATRE burner where massive or hollow samples can be mechanically loaded under hot gas flow conditions. This unique test rig allows for gas testing temperatures up to 1600°C and internal cooling of hollow samples.The damage mechanisms under different TMF conditions are presented and the impact of a through thickness thermal gradient (induced by means of internal cooling) is analyzed. Moreover, a prior thermal over-aging of the whole system is investigated to mimic service degradation. It is shown in this work that depending on the tensile stress applied at low temperature, a change in the main degradation mechanism of the TBC system is observed. Moreover, the thermal over-aging before TMF tests favors cycling ratcheting in compression during thermo-mechanical cycling due to the degradation of the substrate microstructure.

Direct Numerical Simulations of a High-Pressure Turbine Vane

In this paper, we establish a benchmark data set of a generic high-pressure (HP) turbine vane generated by direct numerical simulation (DNS) to resolve fully the flow. The test conditions for this case are a Reynolds number of 0.57 × 106 and an exit Mach number of 0.9, which is representative of a modern transonic HP turbine vane. In this study, we first compare the simulation results with previously published experimental data. We then investigate how turbulence affects the surface flow physics and heat transfer. An analysis of the development of loss through the vane passage is also performed. The results indicate that freestream turbulence tends to induce streaks within the near-wall flow, which augment the surface heat transfer. Turbulent breakdown is observed over the late suction surface, and this occurs via the growth of two-dimensional Kelvin–Helmholtz spanwise roll-ups, which then develop into lambda vortices creating large local peaks in the surface heat transfer. Turbulent dissipation is found to significantly increase losses within the trailing-edge region of the vane.

Development of the design scheme of cooling for a nozzle vane of high pressure turbine of gas turbine engine

The design schemes of cooling for the nozzle blades of high-temperature gas turbines are considered. The results of the thermal and hydraulic tests for the cooling systems of the nozzle vanes are presented. The prospective cooling system of the nozzle vanes of low and high pressure turbines is developed and presented. The test results for the vane with proposed design are presented.