Combustor-turbine Interface Research Articles

The Effect of Unsteady Inlet Boundary Conditions on the Aero-Thermal Behavior of High-Pressure Turbine Vanes – A Numerical Study Using Scale-Resolving Simulations

Abstract This paper presents the application of a novel method to prescribe unsteady boundary conditions to transient, scale-resolving computational fluid dynamics simulations of the high-pressure turbine in modern jet engines. The methodology is based on the compression of the interface data at the combustor–turbine interface, using proper orthogonal decomposition and Fourier series (PODFS). Doing so can reduce the stored data at the interface drastically. The capability of the PODFS method to produce realistic inlet boundary conditions was demonstrated in previous work. Here, the method is applied to a turbine case. The outlet data of a combustor simulation is used to create the PODFS boundary conditions for a scale-resolving simulation of a simplified first nozzle guide vane of the high-pressure turbine. This simulation is compared with simulations with steady-state boundary conditions to show the effect of unsteadiness in the inlet boundary condition on the aerodynamic and thermal behaviors of the turbine. While the aerodynamics show minor sensitivity against the way of applying the inlet boundary conditions, the thermal behavior of the vanes is strongly affected by the modeling of combustor unsteadiness.

Uncertainty quantification of turbine endwall aero-thermal performance under combustor-turbine interface louver coolant and cavity

The uncertainty characteristics have a significant influence on turbine cooling performance and its reliability, especially for the combustor-turbine interface region which faces more complicated flow interactions. Previous researches under the certainty framework are not able to reflect the realistic cooling characteristics of the interface region under real operation circumstances. Considering the random fluctuations of the major structure and flow parameters, uncertainty quantification of turbine endwall cooling performance under realistic combustor-turbine interface conditions is conducted in this paper using the polynomial chaos expansion method. The mix characteristic of combustor louver coolant and interface cavity leakage in the turbine endwall region is investigated, and the influences of parameter uncertainties on endwall cooling, vane phantom cooling, and cascade aerodynamic characteristics are analyzed. The results indicate that the uncertainties at the combustor-turbine interface largely influence the endwall cooling by modifying the strength and position of the secondary flows. Under the random fluctuations of interface geometry parameters, the local mean value of endwall cooling effectiveness can reach up to 0.25, and the standard deviation of endwall cooling effectiveness can reach 0.1. The geometrical parameters of interface width and misalignment height are the decisive factors for endwall cooling uncertainty. The random fluctuations of interface width and misalignment height separately cause great cooling effectiveness uncertainties on the pressure side endwall and suction side hot ring. The uncertainties in mainstream and structure parameters exhibit a reduced propensity to propagate onto vane phantom cooling. The decisive factors for vane phantom cooling uncertainty are interface width and misalignment height, and the major influenced area is the vane middle part. Four mainstream and structure parameters all make great contributions to the aerodynamic uncertainty at the strong secondary flow region and vane wake flow region. This paper can provide a theoretical basis for the design of combustor-turbine interface cooling structures to accommodate the actual turbine operation uncertainty conditions.

Film cooling effect of upstream jump coolant on turbine endwall with combustor-turbine interface misalignment

The investigation of jump cooling performance on turbine endwall independently causes its actual cooling effectiveness to deviate from the design values. In this paper, the endwall jump cooling structure with combustor-turbine interface cavity and combustor liners is modeled. Taking the realistic combustor-turbine interface features into account, the effect of jump coolant on turbine endwall film cooling and heat transfer characteristics is numerically studied. Under different endwall misalignment modes, the aerodynamic interaction mechanisms of realistic combustor outflow profile, jump coolant jet and cascade secondary flows at turbine endwall region are revealed. The modification effects of combustor-turbine interface features on jump cooling of the endwall are investigated. The results show that the combustor outflow severely ingests into the combustor-turbine interface cavity after flowing across several steps. Two branches of cavity vortex are subsequently generated in the middle-pitch region of cascade to affect the jump coolant jet. At low coolant blowing ratio, the attachment and separation sides of cavity vortex individually result in high and low cooling regions, while the horseshoe vortex leads to a large wedge-shaped cooling region. At z/Cax < 0, the backward step leads to a higher cooling effectiveness than forward step by 0.2. The jump coolant only leads to phantom cooling effect on downstream part of the vane suction side surface. The net heat flux reduction (NHFR) between two cavity vortexes and at pressure side junction are individually 0<NHFR<0.2 and −0.2<NHFR < −0.1. At high blowing ratio, the jump coolant can cover the pressure side junction with the highest cooling effectiveness of 0.5. The jump coolant can flow across the horseshoe vortex and directly upwash the vane surface. The weak protection regions between two cavity vortexes and near the attachment line of the pressure side horseshoe vortex enlarges. Compared with the forward step, the backward step achieves a lower NHFR by up to 0.1 and a higher Nusselt number. This paper provides an in-depth analysis of the aero-thermal physics of endwall jump cooling concerning realistic combustor-turbine interface conditions.

Turbine Vane Passage Cooling Experiments With a Close-Coupled Combustor-Turbine Interface Geometry—Part I: Describing the Flow

Abstract Due to the proximity of the first-stage gas turbine vanes to the combustor, coolant introduced to the combustor walls interacts with the endwall film coolant and changes the vane passage flow physics. Recent results show that combustor coolant contributes significantly to cooling the endwall and vane surfaces. In this paper, the traditional combustor-turbine interface was modified to improve overall cooling performance. The performance of this new injection cooling scheme on passage fluid dynamics and surface cooling is assessed. The first of this two-part paper reports detailed experimental tests that document secondary flows and coolant transport throughout the vane passage for four combustor coolant flowrates. The experimental facility imitates combustor coolant injection and engine-level turbulence and has a modified transition duct design, called the “close-coupled combustor-turbine interface.” The “impingement vortex” seen in previous studies with combustor cooling appears as the dominant secondary flow. It is observed in the present study over a wide range of flowrates, confirming its tie to the combustor coolant flowrate and not the combustor-turbine interface geometry. It was found, however, that the location and size of the impingement vortex are affected by coolant flowrate. Part II of this paper discusses the impact of the observed secondary flows on cooling vane passage surfaces.

Aerothermal Performance and Soot Emissions of Reacting Flow in a Micro-Gas Turbine Combustor

Micro-gas turbines are used for power generation and propulsion in unmanned aerial vehicles. Despite the growing demand for electric engines in a world striving for a net zero carbon footprint, combustion gas turbines will continue to play a critical role. Hence, there is a need for improved micro-gas turbines that can meet stringent environmental regulations. This paper is the first part of a comprehensive study focused on understanding the fundamental performance and emission characteristics of a micro-gas turbine model, with the aim of finding ways to enhance its operation. The study used a multidisciplinary CFD model to simulate the reacting flow in the combustion chamber and validated the results against experimental data and throughflow simulations. The present work is one of the few work that attempts to address both the aerothermal performance and emissions of the gas turbine. The findings highlight that parameters such as non-uniform outlet pressure, fuel-to-air ratio, and fuel injection velocity can greatly influence the performance and emissions of the micro-gas turbine. These parameters can affect the combustion efficiency, the formation of hot spots at the combustor–turbine interface, and the soot emissions. The results provide valuable insights for optimizing the performance and reducing the emissions of micro-gas turbines and serve as a foundation for further research into the interaction between the combustor and the turbine.

Lean-Burn Combustor Simulator for an Engine-Component Test Facility: An Experimental and Computational Study

Abstract The computational and experimental assessment of a lean-burn low-NOx combustor simulator for an engine component test facility is presented. The Engine Component Aero-Thermal (ECAT) facility is a full-scale engine-parts facility, designed for the study of the aero-thermal performance of fully cooled high-pressure nozzle guide vanes (NGVs). The facility operates with non-dimensionally matched engine conditions in terms of Reynolds number, Mach number, and coolant-to-mainstream pressure ratio. The combustor simulator is designed to replicate lean-burn conditions of swirl and temperature distortion upstream of the nozzle guide vanes. The purpose is to allow the study of flow capacity, aerodynamic performance (with film cooling), and thermal performance (overall effectiveness) in the presence of engine-realistic inlet distortions. The ECAT combustor simulator design builds on the work of Hall and Povey, who developed a full-scale low-speed atmospheric-pressure combustor simulator for the preliminary design of similar simulators later implemented in the Oxford Turbine Research Facility and ECAT. This pilot facility produced non-dimensional lean-burn combustor-exit conditions closely matched to target profiles representative of modern aero engines. This design was modified and scaled for compatibility with the annulus-line and higher-pressure operating conditions of the ECAT facility. In the present study, detailed experimental measurements with multi-hole probes and thermocouples (pressure profile and temperature field) are presented and compared to results from Reynolds-Averaged Navier–Stokes Simulations. Additional simulations were performed to understand how the elevated back-pressure and vane potential field affect the non-dimensional profiles of pressure loss, residual swirl, and temperature at the combustor–turbine interface. This is perhaps the most comprehensive study to date of a combustor simulator in an engine-scale research facility, providing a unique insight into the known challenges of simulator design, scaling issues when moving from low to high Reynolds number, and limitations of computational fluid dynamics in this flow environment. The results, which will serve as boundary conditions to understand the impact of this flow on components in future studies in this facility, demonstrate the fidelity with which lean-burn target conditions can be replicated in a non-reacting environment.

Turbine nozzle endwall aero-thermal characteristics under combustor louver coolant with interface cavity

Targeting on combustor-turbine interface cooling system of combustor louver coolant and interface cavity, the endwall film cooling and heat transfer characteristics as well as the vane phantom cooling performance of a nozzle guide vane are investigated by solving the Reynolds-Averaged Navier-Stokes equations, and the secondary flow structure at endwall region are analyzed. The results show that most of the louver coolant leads to the core cooling region on the endwall after recirculating in the cavity with the cavity vortex, while a minority of the louver coolant escapes from the cavity vortex and forms the outer endwall cooling region limited by the horseshoe vortex. Under a certain mainstream flow condition, the endwall cooling performance and net heat flux reduction (NHFR) only depend on the mass flow ratio (MFR) of louver coolant instead of blowing ratio (BR). With increasing louver BR, the coolant mainly leads to the increase of the width and film cooling effectiveness of the core region at high MFR, but it leads to increasing cooling effectiveness of the outer region at low MFR. When the louver MFR increases from 0.25 % to 1.91 %, the averaged endwall cooling effectiveness at z/Cax = 0.1 is increased from 0.04 to 0.21. The NHFR of the core and outer cooling region reaches 0.3 and 0.1, respectively. The pressure and suction side horseshoe vortexes cause two protection failure regions, and its NHFR further decreases with increasing BR. At high MFR, the heat transfer in the low Nusselt number regions resulted by cavity vortex separation is further weakened with the increase of BR, and the region is enlarged toward two sides of the cascade. When MFR = 1.91 %, the Nusselt number of the high heat transfer region at z/Cax = 0.7 decreases from 2200 to 2000, and the averaged NHFR of fore part endwall reaches up to 0.2. This paper provides guidance for endwall cooling configuration design by taking the compound effect of the combustor-turbine interface cooling scheme into account.

Unsteady phenomena at the combustor-turbine interface

The combustor-turbine interface in a gas turbine is characterised by complex, highly unsteady flows. In a combined experimental and large eddy simulation (LES) study including realistic combustor geometry, the standard model of secondary flows in the nozzle guide vanes (NGV) is found to be oversimplified. A swirl core is created in the combustion chamber which convects into the first vane passages. Four main consequences of this are identified: variation in vane loading; unsteady heat transfer on vane surfaces; unsteadiness at the leading edge horseshoe vortex, and variation in the position of the passage vortex. These phenomena occur at relatively low frequencies, from 50–300 Hz. It seems likely that these unsteady phenomena result in non-optimal film cooling, and that by reducing unsteadiness designs with greater cooling efficiency could be achieved. Measurements were performed in a high speed test facility modelling a large industrial gas turbine with can combustors, including nozzle guide vanes and combustion chambers. Vane surfaces and endwalls of a nozzle guide vane were instrumented with 384 high speed thin film heat flux gauges, to measure unsteady heat transfer. The high resolution of measurements was such to allow direct visualisation in time of large scale turbulent structures over the endwalls and vane surfaces. A matching LES simulation was carried out in a domain matching experimental conditions including upstream swirl generators and transition duct. Data reduction allowed time-varying LES data to be recorded for several cycles of the unsteady phenomena observed. The combination of LES and experimental data allows physical explanation and visualisation of flow events.

A New Experimental Approach for Heat Transfer Coefficient and Adiabatic Wall Temperature Measurements on a Nozzle Guide Vane With Inlet Temperature Distortions

AbstractModern gas turbines lean combustors allow to limit NOx pollutant emissions by controlling the flame temperature, while maintaining high turbine inlet temperatures. On the other hand, their adoption presents other challenges, especially concerning the combustor–turbine interaction. Turbine inlet conditions are generally characterized by severe temperature distortions and swirl degree, which, in turn, is responsible for very high turbulence intensities. Several past studies have focused on the description of the effects of these phenomena on the behavior of the high pressure stages of the turbine, both considering them as separated aspects, and, in very recent years, accounting for their combined impact. Nevertheless, very limited experimental results are available when it comes to evaluate the heat transfer coefficient (HTC) on the nozzle guide vane (NGV) external surface, since relevant temperature distortions present a severe challenge for the commonly adopted measurement techniques. The work presented in this paper was carried out on a non-reactive, annular, three-sector test rig, made by a combustor simulator and a NGV cascade. Making use of three real hardware burners of a Baker Hughes heavy-duty gas turbine, operated in similitude conditions, it can reproduce a representative swirling flow, with temperature distortions at the combustor–turbine interface plane. This test apparatus was exploited to develop an experimental approach to retrieve reliable HTC and adiabatic wall temperature distributions simultaneously, in order to overcome the known limitations imposed by temperature gradients on state-of-the-art methods for HTC calculation from transient tests. A non-cooled mockup of a NGV doublet, manufactured using low thermal diffusivity plastic material, was used for the tests, carried out using infra-red thermography with a transient approach. In the authors’ knowledge, this presents the first experimental attempt of measuring a NGV HTC in the presence of relevant temperature distortions and swirl.

Clocking and Potential Effects in Combustor–Turbine Stator Interactions

Investigations of combustors and turbines separately have been carried out for years by research institutes and aircraft engine companies, but there are still many questions about the interaction effect. In this paper, a prediction of a turbine stator’s potential effect on flow in a combustor and the clocking effect on temperature distribution in a nozzle guide vane are discussed. Numerical simulation results for the combustor simulator and the nozzle guide vane (NGV) of the first turbine stage are presented. The geometry and flow conditions were defined according to measurements carried out on a test section within the framework of the EU FACTOR (full aerothermal combustor–turbine interactions research) project. The numerical model was validated by a comparison of results against experimental data in the plane at a combustor outlet. Two turbulence models were employed: the Spalart–Allmaras and Explicit Algebraic Reynolds Stress models. It was shown that the NGV potential effect on flow distribution at the combustor–turbine interface located at 42.5% of the axial chord is weak. The clocking effect due to the azimuthal position of guide vanes downstream of the swirlers strongly affects the temperature and flow conditions in a stator cascade.

Numerical prediction of heat transfer characteristics on a turbine nozzle guide vane under various combustor exit hot‐streaks

AbstractThe precise prediction of the aerothermal conditions at the outlet of modern combustors of aero engines is of primary interest in the aerothermal design of high‐pressure turbines. The use of high swirling flows to stabilize the flame in the new generation of lean‐burn combustors (LBCs) leads to shorter combustor length and high degrees of unsteady temperature. These aspects generate high levels of turbulence and complex temperature nonuniformity patterns at the combustor–turbine interface. Therefore, the nozzle guide vane (NGV) located at the interface exhibits a high inlet temperature gradient called a hot‐streak. The transport mechanism of the hot‐streak across the NGV needs to be decrypted to increase turbine service life and enhance its performance. In this context, the present study aims to examine numerically the effect of three new distorted hot‐streak gradients on the heat transfer characteristics across a turbine NGV. An innovative LBC is used to generate the three inlet hot‐streaks, which are highly swirled and distorted similar to real aero‐engine conditions. The use of such a generation approach avoids any simplification associated with uniform or circular contours that are widely used by researchers. A solid NGV cascade developed by NASA is adopted for the investigations. The results show significant differences in the thermal field on the cascade at each distorted thermal condition. The transport of the hot‐streak behaves differently under axial and swirl conditions. The presence of a swirl induces flow mixing and results in thermal load increasing on the cascade surfaces (NGV, hub, and shroud).

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwall—Part 1: Aerodynamics

Abstract The first-stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics is not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first-stage nozzle guide vane passage. The test section involves an engine-representative combustor–turbine interface geometry, combustor coolant flow, and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwall—Part 2: Thermal Measurements

Abstract Flow over gas turbine endwalls is complex and highly three-dimensional. As boundaries for modern engine designs are pushed, this already-complex flow is affected by aggressive application of film cooling flows that actively interact. This two-part study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first-stage nozzle guide vane passage (NGV). The approach flow conditions represent flow exiting a low-NOx combustor. The test section includes geometric and cooling details of a combustor-turbine interface in addition to endwall film cooling flows injected upstream of the passage. The first part of this study describes in detail, the passage aerodynamics as affected by injection of cooling flows. It reveals a system of secondary flows, including the newly discovered Impingement Vortex, which redefines our understanding of the aerodynamics of flow in a modern, film-cooled, first-stage vane row. The second part investigates, through thermal measurements, the distribution, mixing, and disruption of cooling flows over the endwall. Measurements are made with and without active endwall film cooling. Descriptions are made through adiabatic surface effectiveness measurements and correlations with in-passage velocity (presented in part one) and thermal fields. Results show that the newly discovered impingement vortex has a positive effect on coolant distribution through passage vortex suppression and by carrying the coolant to hard-to-cool regions in the passage, including the pressure surface near the endwall.

Effects of Endwall Film Coolant Flowrate on Secondary Flows and Coolant Mixing in a First Stage Nozzle Guide Vane

AbstractModern gas turbines are subjected to very high thermal loading. This leads to a need for aggressive cooling to protect components from damage. Endwalls are particularly challenging to cool due to a complex system of secondary flows near them that wash and disrupt the protective coolant films. This highly three-dimensional flow not only affects but is also affected by the momentum of film cooling flows, whether injected just upstream of the passage to intentionally cool the endwall or as combustor cooling flows injected further upstream in the engine. This complex interaction between the different cooling flows and passage aerodynamics has been recently studied in a first stage nozzle guide vane. The present paper presents a detailed study on the sensitivity of aero-thermal interactions to endwall film cooling mass flow to mainstream flow ratio. The test section represents a first stage nozzle guide vane with a contoured endwall and endwall film cooling injected just upstream of it. The test section also includes an engine-representative combustor–turbine interface geometry with combustor cooling flows injected at a constant rate. The approach flow conditions represent flow exiting a low-NOx combustor. Adiabatic surface thermal measurements and in-passage velocity and thermal field measurements are presented and discussed. The results show the dynamics of passage vortex suppression and the increase of impingement vortex strength as MFR changes. The effects of these changes of secondary flows on coolant distribution are presented.

Impact of a Combustor–Turbine Interface on Turbine Vane Aerodynamics and Film Cooling

Abstract Optimizing the aerothermal performance of the combustor–turbine interface is an important factor in enhancing the efficiency of heavy-duty gas turbines. Also, it is a key requirement to fulfill the lifetime in this hottest area of the gas turbine. Typically transition pieces of can combustors induce a highly nonuniform swirling flow at the turbine inlet. In order to better understand the impact of the nonuniform combustor flow at the first stage vanes, a combined experimental and numerical study was carried out. The experimental facility consisted of a high-speed linear cascade with four vane passages, including an upstream transition piece, which was representative of a heavy-duty gas turbine can combustor–turbine interface geometry. The experiments were conducted at engine representative Mach numbers, and film cooling effectiveness measurements were performed at three different blowing ratios. Computational fluid dynamics (CFD) Reynolds-averaged Navier–Stokes simulations were undertaken using a commercial flow solver. The numerical model was first validated with the experimental data, using inlet traverse five-hole probe measurements, pressure taps along the airfoil perimeter, and oil flow visualization results. The investigation shows that the position of the vane relative to the combustor transition piece has a significant impact on the vane aerodynamics and also film cooling behavior. This understanding was important for a robust first vane aerothermal design of the GT36.

Uncertainty Propagation Analyses of Lean Burn Combustor Exit Conditions for a Robust Nozzle Cooling Design

Abstract Knowing the flow conditions at the combustor turbine interface is a key asset for an efficient cooling design of high-pressure turbines. However, measurements and numerical predictions of combustor exit conditions are challenging due to the extreme temperatures and complex flow patterns in modern combustors. Even the time-averaged flow fields at the combustor exit which are commonly used as inlet condition for simulations of the turbine are therefore subject to uncertainty. The goal of this paper is to illustrate how aleatory uncertainties in the magnitude and position of residual swirl and hot spots at the combustor exit affect uncertainties in the prediction of cooling and heat load of the first nozzle guide vane. Also, it is identified which of these uncertain parameters have the greatest impact. An iso-thermal test rig and an engine realistic setup with lean burn inflow conditions are investigated. The analysis combines a parameterized model for combustor exit flow fields with uncertainty quantification methods. It is shown that the clocking position of turbine inlet swirl has a large effect on the formation of secondary flows on the vane surface and thus affects the uncertainty of thermal predictions on the hub and vanes.

Effect of inflow swirl on the vane endwall aerothermal performance with endwall misalignment at transonic condition

Effect of turbine inlet swirl condition (swirl intensity and orientation) on endwall heat transfer performance with combustor-turbine misalignment at transonic condition was numerically investigated, and coolant injection from combustor-turbine interface was also considered. Simulations with two mainstream swirl numbers (SN) of 0.6 and 1.0, two swirl orientations, two endwall misalignment modes and slot coolant injection were conducted by solving the three-dimensional RANS equations coupled with the Reynolds Stress Turbulence Model (RSM). Computational results indicate that for anticlockwise swirled condition, the averaged Nu number of the high heat transfer attachment region upstream of the leading edge decreases from 4400 to 3100 with the increasing SN from 0.6 to 1.0. Negative incidence is caused at hub followed by a low heat transfer region at pressure side junction. Enhanced horseshoe vortex and mainstream direct attachment enlarge the hot ring (uncooled area around the vane leading edge) when SN = 0.6. As for clockwise swirled condition, positive incidence is caused at hub and can have some effect on reducing the passage crossflow. The pressure side low heat transfer region enlarges under aligned mode but shrinks under misaligned mode compared with anticlockwise condition. Averaged Nu distributes more uniformly along axial direction. The hot ring can only be observed when SN = 0.6 under misaligned condition. For aligned mode, endwall heat transfer at x/Cax < 1.0 is 30–50% higher for SN = 0.6 than SN = 1.0. When coolant is introduced at the combustor-turbine interface, the heat transfer strength at attachment region is largely reduced. Under aligned condition, coolant tends to eject out between two swirl cores. The majority of the coolant is constrained in the recirculation region by cavity vortices under misaligned mode. Averaged Nu number at the attachment region is 4450 and 3600 under aligned and misaligned mode, respectively. But the averaged Nu number downstream of the passage throat is less than 2000 and distributes uniformly.

Computational study of film cooling and flowfields on a stepped vane endwall with a row of cylindrical hole and interrupted slot injections

Endwall is one of the critical regions to perform film cooling for gas turbine because of not only the increase of turbine inlet temperature but also the complicated secondary flow. Typical endwall film cooling configuration includes leakage slot at combustor-vane interface and discrete cooling holes on the wall. This paper mainly studied the effect of endwall step on the cooling performance of downstream cooling holes. The combustor-turbine interface in this paper is a kind of backward-facing step, causing an interrupted slot. The detailed characteristic of the interrupted slot has been investigated in the previous study. Three dimensional Reynolds-averaged Navier-Stokes equations with shear stress turbulence model (SST) k-ω was solved to perform the calculations of endwall film effectiveness. The adiabatic film effectiveness on the flat endwall was compared with that on the stepped endwall with and without the interrupted slot. Results showed a significant impact of the endwall step on the adiabatic film cooling effectiveness. Coolant jet downstream of the step is split into two parts due to the effects of step vortex (SV) and cross flow. Detailed flowfields in spanwise, pitchwise and axial planes explain the formation and development of secondary flow near the cooling holes. For a range of blowing ratio from 0.6 to 1.2, increasing the blowing ratio of cooling holes deteriorates the adiabatic film cooling effectiveness upstream and downstream of the cooling holes due to the jet liftoff. It is found that cooling effectiveness is strongly dependent on the axial position of the hole. Placing the cooling holes downstream at leading edge plane of the cascade little effect of the step vortex on the film effectiveness and, subsequently, relatively uniform coolant coverage was observed on the endwall.

Adiabatic Effectiveness on High-Pressure Turbine Nozzle Guide Vanes Under Realistic Swirling Conditions

In modern lean-burn aero-engine combustors, highly swirling flow structures are adopted to control the fuel-air mixing and to provide the correct flame stabilization mechanisms. Aggressive swirl fields and high turbulence intensities are hence expected in the combustor-turbine interface. Moreover, to maximize the engine cycle efficiency, an accurate design of the high-pressure nozzle cooling system must be pursued: in a film-cooled nozzle, the air taken from last compressor stages is ejected through discrete holes drilled on vane surfaces to provide a cold layer between hot gases and turbine components. In this context, the interactions between the swirling combustor outflow and the vane film cooling flows play a major role in the definition of a well-performing cooling scheme, demanding for experimental campaigns at representative flow conditions. An annular three-sector combustor simulator with fully cooled high-pressure vanes has been designed and installed at THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion-cooled liners, and six film-cooled high-pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central airfoil aligned with the central swirler. In this experimental work, adiabatic film effectiveness measurements have been carried out in the central sector vanes, in order to characterize the film-cooling performance under swirling inflow conditions. The pressure-sensitive paint (PSP) technique, based on heat and mass transfer analogy, has been exploited to catch highly detailed 2D distributions. Carbon dioxide has been used as coolant in order to reach a coolant-to-mainstream density ratio of 1.5. Turbulence and five-hole probe measurements at inlet/outlet of the cascade have been carried out as well, in order to highlight the characteristics of the flow field passing through the cascade and to provide precise boundary conditions. Results have shown a relevant effect of the swirling mainflow on the film cooling behavior. Differences have been found between the central airfoil and the adjacent ones, both in terms of leading edge stagnation point position and of pressure and suction side film coverage characteristics.

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°.