Combustor Exit Flow Research Articles

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.

Parameterised Model of 2D Combustor Exit Flow Conditions for High-Pressure Turbine Simulations

An algorithm is presented generating a complete set of inlet boundary conditions for Reynolds-averaged Navier–Stokes computational fluid dynamics (RANS CFD) of high-pressure turbines to investigate their interaction with lean and rich burn combustors. The method shall contribute to understanding the sensitivities of turbine aerothermal performance in a systematic approach. The boundary conditions are based on a set of input parameters controlling velocity, temperature, and turbulence fields. All other quantities are derived from operating conditions and additional modelling assumptions. The algorithm is coupled with a CFD solver by applying the generated profiles as inlet boundary conditions. The successive steps to derive consistent flow profiles are described and results are validated against flow fields extracted from combustor CFD.

The Influence of Combustor Swirl on Pressure Losses and the Propagation of Coolant Flows at the Large Scale Turbine Rig (LSTR): Experimental and Numerical Investigation

The aerothermal interaction of the combustor exit flow on the first vane row has been examined at the Large Scale Turbine Rig (LSTR) at Technische Universität Darmstadt (Darmstadt, Germany). A baseline configuration of axial inflow and a variation of swirling combustor inflow have been studied. The nozzle guide vane (NGV) featured endwall cooling, airfoil film cooling and a trailing edge slot ejection as well as NGV-rotor wheel space purge flow. CO2 is injected for coolant flow tracing. The results are compared to five hole probe (5HP) measurements. The experiments for the baseline configuration are accompanied by numerical simulations using a passive scalar tracking method to validate the results and study the propagation of the coolant flow. The endwall coolant injection is detected to influence the pressure losses in the NGV. It has an impact on the Trailing Edge (TE) coolant ejection as well. For swirling combustor inflow, increased NGV pressure losses and increased mixing of Rear Inner Discharge Nozzle (RIDN) coolant and main flow is observed. An influence of the clocking position of the swirler to the vane is detected. Additional losses within the NGV row can be assigned to the swirler by means of flow tracing.

Influence of Combustor Swirl on Endwall Heat Transfer and Film Cooling Effectiveness at the Large Scale Turbine Rig

At the large scale turbine rig (LSTR) at Technische Universität Darmstadt, Darmstadt, Germany, the aerothermal interaction of combustor exit flow conditions on the subsequent turbine stage is examined. The rig resembles a high pressure turbine and is scaled to low Mach numbers. A baseline configuration with an axial inflow and a swirling inflow representative for a lean combustor is modeled by swirl generators, whose clocking position toward the nozzle guide vane (NGV) leading edge can be varied. A staggered double-row of cylindrical film cooling holes on the endwall is examined. The effect of swirling inflow on heat transfer and film cooling effectiveness is studied, while the coolant mass flux rate is varied. Nusselt numbers are calculated using infrared thermography and the auxiliary wall method. Boundary layer, turbulence, and five-hole probe measurements as well as numerical simulations complement the examination. The results for swirling inflow show a decrease of film cooling effectiveness of up to 35% and an increase of Nusselt numbers of 10–20% in comparison to the baseline case for low coolant mass flux rates. For higher coolant injection, the heat transfer is on a similar level as the baseline. The differences vary depending on the clocking position. The turbulence intensity is increased to 30% for swirling inflow.

Non-Uniformity of the Combustor Exit Flow Temperature in Front of the Gas Turbine

Abstract Various types of damages to gas-turbine components, in particular to turbine blades, may occur in the course of gas turbine operation. The paper has been intended to discuss different forms of damages to the blades due to non-uniformity of the exit flow temperature. It has been shown that the overheating of blade material and thermal fatigue are the most common reasons for these damages. The paper presents results from numerical experiments with use of the computer model of the aero jet engine designed for simulations. The model has been purposefully modified to take account of the assumed non-homogeneity of the temperature field within the working agent at the turbine intake. It turned out that such non-homogeneity substantially affects dynamic and static properties of the engine considered as an object of control since it leads to a lag of the acceleration time and to increase in fuel consumption. The summarized simulation results demonstrate that the foregoing properties of a jet engine are subject to considerable deterioration in pace with gradual increase of the assumed non-homogeneity of the temperature field. The simulations made it possible to find out that variations of the temperature field nonhomogeneity within the working agent at the turbine intake lead to huge fluctuation of the turbine rpm for the idle run.

Film Cooling Effectiveness in a Gas Turbine Engine: A Review

This study was carried out to extend database knowledge about the function of film cooling holes at the end of combustor and the inlet of turbine. Using the well-known Brayton cycle, raising the turbine inlet temperature is the key to obtain higher engine efficiency in gas turbine engines. However, high temperature of the combustor exit flow causes non-uniformities. These non-uniformities lead to the reduction of expected life of critical components. Therefore, an appropriate cooling technique should be designed to protect these parts. Film cooling is one of the most effective external cooling methods. Various film cooling techniques presented in the literature have been investigated. Moreover, challenges and future directions of film cooling techniques have been reviewed and presented in this paper. The aim of this review is to summarize recent development in research on film cooling techniques and attempt to identify some challenging issues that need to be solved for future research.

Experimental study of novel passive control methods to improve combustor exit temperature uniformity

The effect of non-uniform temperature flow which comes in contact with the turbine blades is of paramount importance since the thermal damage that occur to the blade during its operation, is governed by the non-uniformities present in the oncoming flow. This thermal damage leads to increased maintenance cost and reduced life-span of the turbine blades. This paper investigates the effectiveness of some novel passive control techniques to improve the temperature uniformity of the combustor exit flow to address the need to reduce the thermal damage and hence decrease the overall maintenance cost of the gas turbine system. The novel passive control techniques tested in this study include the use of streamlined body or guide vanes in the dilution zone of the combustor. For the guide vanes four different orientations were tested—0°, 30°, 60°, 90°. Extensive experimentation was conducted under different flow conditions. The deviation of the exit temperature from the equilibrium mixing temperature was used to compare the effectiveness of different passive control techniques. It was found that the 30o guide vanes gave the most uniform temperature flow under majority of cases considered. On an average, the flow with 30o guide vanes was about 15 % more uniform in temperature as compared to the staggered holes geometry with only 1 % higher pressure loss. The possible reason for this improvement is the combination of swirl and depth to which the dilution flow can enter the dilution zone and mix with the primary hot flow. The streamlined body came second with an improvement in pressure losses. Based on these experimental findings, the use of these guide vanes and streamlined body has potential in the gas turbine industry to deal with the high maintenance cost involved in these systems.

Film-Cooling Techniques at the End of Combustor and Inlet of Turbine in a Gas Turbine Engine: A Review

This study was carried out to extend database knowledge about the film cooling holes function at the end of combustor and inlet of turbine. Using the well-known Brayton cycle, rising the turbine inlet temperature, is the key to get higher engine efficiency in gas turbine engines. But the high temperature of the combustor exit flow causes non-uniformities. These non-uniformities lead to a reduction in the expected life of critical components. Therefore a cooling technique should be designed to protect these parts. There are two separate ways for Gas turbine cooling: internal cooling and external cooling. Film cooling is one of the most effective external cooling methods. In this system, a low temperature thin boundary layer such as buffer zone is formed and attached on the protected surface. In this study, a literature survey was done on the limited surveys, particularly since the first of 21th century.

Dynamic and Thermodynamic Analysis of Film-Cooling

This study was done to extend database knowledge about the film cooling action at the end of combustor and inlet of turbine. The well-known Brayton cycle plays a key role to get higher engine efficiency in gas turbine engines. But the high temperature of the combustor exit flow creates unpleasant environment. These surroundings lead to a reduction in the expected life of critical parts. There are two separate ways for Gas turbine cooling: internal cooling and external cooling. Film cooling is one of the most effective external cooling methods. In this system, a low temperature thin boundary layer such as buffer zone is formed and attached on the protected surface. In this study, a literature survey was done on the limited surveys that considered the effects of flow structure variations on film cooling, particularly from the first of 21th century.

Dynamic and Thermodynamic Analysis of Film-Cooling

This study was done to extend database knowledge about the film cooling action at the end of combustor and inlet of turbine. The well-known Brayton cycle plays a key role to get higher engine efficiency in gas turbine engines. But the high temperature of the combustor exit flow creates unpleasant environment. These surroundings lead to a reduction in the expected life of critical parts. There are two separate ways for Gas turbine cooling: internal cooling and external cooling. Film cooling is one of the most effective external cooling methods. In this system, a low temperature thin boundary layer such as buffer zone is formed and attached on the protected surface. In this study, a literature survey was done on the limited surveys that considered the effects of flow structure variations on film cooling, particularly from the first of 21th century.

Effect of Hole Spacing on Deposition of Fine Coal Flyash Near Film Cooling Holes

Particulate deposition experiments were performed in a turbine accelerated deposition facility to examine the nature of flyash deposits near film cooling holes. Deposition on both bare metal and thermal barrier coating (TBC) coupons was studied, with hole spacing (s/d) of 2.25, 3.375, and 4.5. Sub-bituminous coal ash particles (mass mean diameter of 13 μm) were accelerated to a combustor exit flow Mach number of 0.25 and heated to 1183°C before impinging on a target coupon. The particle loading in the 1 h tests was 310 ppmw. Blowing ratios were varied in these experiments from 0 to 4.0 with the density ratio varied approximately from 1.5 to 2.1. Particle surface temperature maps were measured using two-color pyrometry based on the red/gree/blue (RGB) signals from a camera. For similar hole spacing and blowing ratio, the capture efficiency measured for the TBC surface was much higher than for the bare metal coupon due to the increase of surface temperature. Deposits on the TBC coupon were observed to be more tenacious (i.e., hard to remove) than deposits on bare metal coupons. The capture efficiency was shown to be a function of both the hole spacing and the blowing ratio (and hence surface temperature). Temperature seemed to be the dominant factor affecting deposition propensity. The average spanwise temperature downstream of the holes for close hole spacing was only slightly lower than for the large hole spacing. Roughness parameters Ra and Rt decreased monotonically with increased blowing ratio for both hole spacing analyzed. The roughness for s/d=3.375 was lower than that for s/d=4.5, especially at high blowing ratio. It is thought that these data will prove useful for designers of turbines using synfuels.

Effect of Particle Size and Trench Configuration on Deposition From Fine Coal Flyash near Film Cooling Holes

Particulate deposition experiments were performed in a turbine accelerated deposition facility to examine the effects of flyash particle size and trench configuration on deposits near film cooling holes. Deposition on two bare metal Inconel coupons was studied, with hole spacings (s/d) of 3.4 and 4.5. Two sizes of sub-bituminous coal ash particles were used, with mass mean diameter of 4 and 13 μm, respectively. The effect of a cooling trench at the exit of the cooling holes was also examined in this deposition facility. Experiments were performed at different angles of impaction. Particles were accelerated to a combustor exit flow Mach number of 0.25 and heated to 1183 °C before impinging on a target coupon. The particle loading in the 1-h tests was 160 ppmw. Blowing ratios were varied in these experiments from 0 to 4.0. Particle surface temperature maps were measured using two-color pyrometry based on RGB signals from a camera. Deposits generated from finer particles were observed to stick to the surface more tenaciously than larger particles. The capture efficiency measured for the small particles was lower than for the larger particles, especially at low blowing ratios. However, the finer particles exhibited a greater variation in deposition pattern as a function of hole spacing than seen with larger particles. The effect of trench configuration on deposition was examined by performing deposition tests with and without the trench for the same hole spacing and blowing ratio. The effects of trench configuration on capture efficiency, deposition pattern, and surface topography are reported. Deposition experiments at impingement angles from 45° to 15° showed changes in both deposit thickness and temperature. The trench increased cooling effectiveness, but did not change the particulate collection efficiency because the trench acted as a particulate collector.

Evolution of Surface Deposits on a High-Pressure Turbine Blade—Part II: Convective Heat Transfer

A thermal barrier coating (TBC)-coated turbine blade coupon was exposed to successive deposition in an accelerated deposition facility simulating flow conditions at the inlet to a first stage high pressure turbine (T=1150°C, M=0.31). The combustor exit flow was seeded with dust particulate that would typically be ingested by a large utility power plant. The turbine coupon was subjected to four successive 2h deposition tests. The particulate loading was scaled to simulate 0.02 parts per million weight (ppmw) of particulate over 3months of continuous gas turbine operation for each 2h laboratory simulation (for a cumulative 1year of operation). Three-dimensional maps of the deposit-roughened surfaces were created between each test, representing a total of four measurements evenly spaced through the lifecycle of a turbine blade surface. From these measurements, scaled models were produced for testing in a low-speed wind tunnel with a turbulent, zero pressure gradient boundary layer at Re=750,000. The average surface heat transfer coefficient was measured using a transient surface temperature measurement technique. Stanton number increases initially with deposition but then levels off as the surface becomes less peaked. Subsequent deposition exposure then produces a second increase in St. Surface maps of St highlight the local influence of deposit peaks with regard to heat transfer.

Evolution of Surface Deposits on a High-Pressure Turbine Blade—Part I: Physical Characteristics

Turbine blade coupons with three different surface treatments were exposed to deposition conditions in an accelerated deposition facility. The facility simulates the flow conditions at the inlet to a first stage high-pressure turbine (T=1150°C, M=0.31). The combustor exit flow is seeded with dust particulate that would typically be ingested by a large utility power plant. The three coupon surface treatments included: (1) bare polished metal; (2) polished thermal barrier coating with bondcoat; and (3) unpolished oxidation resistant bondcoat. Each coupon was subjected to four successive 2h deposition tests. The particulate loading was scaled to simulate 0.02 parts per million weight (ppmw) of particulate over 3months of continuous gas turbine operation for each 2h laboratory simulation (for a cumulative 1year of operation). Three-dimensional maps of the deposit-roughened surfaces were created between each test, representing a total of four measurements evenly spaced through the lifecycle of a turbine blade surface. From these measurements the surface topology and roughness statistics were determined. Despite the different surface treatments, all three surfaces exhibited similar nonmonotonic changes in roughness with repeated exposure. In each case, an initial buildup of isolated roughness peaks was followed by a period when valleys between peaks were filled with subsequent deposition. This trend is well documented using the average forward facing roughness angle in combination with the average roughness height as characteristic roughness metrics. Deposition-related mechanisms leading to spallation of the thermal barrier coated coupons are identified and documented.

High-Pressure Turbine Deposition in Land-Based Gas Turbines From Various Synfuels

Ash deposits from four candidate power turbine synfuels were studied in an accelerated deposition test facility. The facility matches the gas temperature and velocity of modern first-stage high-pressure turbine vanes. A natural gas combustor was seeded with finely ground fuel ash particulate from four different fuels: straw, sawdust, coal, and petroleum coke. The entrained ash particles were accelerated to a combustor exit flow Mach number of 0.31 before impinging on a thermal barrier coating (TBC) target coupon at 1150°C. Postexposure analyses included surface topography, scanning electron microscopy, and x-ray spectroscopy. Due to significant differences in the chemical composition of the various fuel ash samples, deposit thickness and structure vary considerably for each fuel. Biomass products (e.g., sawdust and straw) are significantly less prone to deposition than coal and petcoke for the same particle loading conditions. In a test simulating one turbine operating year at a moderate particulate loading of 0.02 parts per million by weight, deposit thickness from coal and petcoke ash exceeded 1 and 2mm, respectively. These large deposits from coal and petcoke were found to detach readily from the turbine material with thermal cycling and handling. The smaller biomass deposit samples showed greater tenacity in adhering to the TBC surface. In all cases, corrosive elements (e.g., Na, K, V, Cl, S) were found to penetrate the TBC layer during the accelerated deposition test. Implications for the power generation goal of fuel flexibility are discussed.

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

Pronounced nonuniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of endwall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the effects of these temperature nonuniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimize the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and endwall heat transfer rate for a high-pressure (HP) nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two nonuniform temperature profiles. The temperature profiles were nondimensionally similar to profiles measured in an engine. A difference of one-half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and endwall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualization study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short-duration engine-sized turbine facility. Mach number, Reynolds number, and gas-to-wall temperature ratios were correctly modeled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

Combustor Turbine Interface Studies—Part 2: Flow and Thermal Field Measurements

Most turbine inlet flows resulting from the combustor exit are nonuniform in the near-platform region as a result of cooling methods used for the combustor liner. These cooling methods include injection through film-cooling holes and injection through a slot that connects the combustor and turbine. This paper presents thermal and flow field measurements in the turbine vane passage for a combustor exit flow representative of what occurs in a gas turbine engine. The experiments were performed in a large-scale wind tunnel facility that incorporates combustor and turbine vane models. The measured results for the thermal and flow fields indicate a secondary flow pattern in the vane passage that can be explained by the total pressure profile exiting the combustor. This secondary flow field is quite different than that presented for past studies with an approaching flat plate turbulent boundary layer along the upstream platform. A counter-rotating vortex that is positioned above the passage vortex was identified from the measurements. Highly turbulent and highly unsteady flow velocities occur at flow impingement locations along the stagnation line.

Heat transfer measurements on an intermediate-pressure nozzle guide vane tested in a rotating annular turbine facility, and the modifying effects of a non-uniform inlet temperature profile

In modern gas turbine engines there exist significant temperature gradients in the combustor exit flow. These gradients arise because both fuel and dilution air are introduced within the combustor as discrete jets. The effects of this non-uniform temperature field on the aerodynamics and heat transfer rate distributions of nozzle guide vanes and turbine blades is difficult to predict, although an increased understanding of the effects of temperature gradients would enhance the accuracy of estimates of turbine component life and efficiency. Low-frequency measurements of heat transfer rate have been conducted on an annular transonic intermediate-pressure (IP) nozzle guide vane operating downstream of a high-pressure (HP) rotating turbine stage. Measurements were conducted with both uniform and non-uniform inlet temperature profiles. The non-uniform temperature profile included both radial and circumferential gradients of temperature. Experiments were conducted in the isentropic light piston facility at QinetiQ Pyestock, a short-duration engine-size turbine facility with 1.5 turbine stages, in which Mach number, Reynolds number and gas—wall temperature ratios are correctly modelled. Experimental heat transfer results are compared with predictions performed using boundary layer methods.

Unsteady Analysis of a Time-Dependent Hot Streak in a Turbine Stage

Experimental studies have shown that combustor temperature nonuniformities, or hot can significantly affect the heating of first-stage rotor blading in turbines. In addition, it has been observed that these hot streaks are not constant in time, but rather fluctuate with instabilities within the combustor. The timedependent variations in the combustor generate not only positive temperature differences from the free stream, but negative as well (i.e., streaks). The unsteady cyclic thermal loading of the blading can degrade turbine durability and efficiency. To investigate this phenomenon, twodimensional unsteady Navier-Stokes simulations have been performed for the 1-1/2 stages of a high-pressure turbine operating in subsonic flow with unsteady hot streaks introduced at the inlet. BPF fm N P RS. 5.5. T TO NOMENCLATURE Blade passing frequency Normalized hot streak frequency Rotor blade passing frequency Static Pressure Pressure surface Suction surface Static temperature SUBSCRIPTS amp Amplitude of fluctuation Base Nominal value c Cold flow HS Hot streak t Stagnation quantity 0 Stator-1 inlet quantity 3 Stator-2 exit quantity INTRODUCTION In the drive for improved fuel economy and cycle efficiency combustor exit temperatures have risen to a point where they can locally exceed the maximum allowable metal temperatures of the blading by 50 to 100°F. Experimental data taken from gas turbine combustors have shown large temperature gradients in both the radial and circumferential directions [1]. These gradients vary both spatially and temporally, and result from the combined effects of the combustor core flow, combustor bypass flow, and combustor surface cooling flow. These combustor hot can typically have temperatures 1.2 to 2.5 times that of the free stream flow [2]. It has also been experimentally observed that the combustor exit flow not only contains positive temperature gradients but negative as well, resulting in The positive temperature gradients are caused by the combustion process, while the negFree stream static temperature w/o H&tive gradients arise because of the combustor Phase * Associate Professor, Senior Member AIAA. ^Graduate Research Assistant. Copyright ©2001 by The American Institute of Aeronautics and Astronautics, Inc. All Rights Reserved. coolant flows. Roback and Bring [3. 4] showed that cold streaks tend to migrate to the rotor suction surface, whereas hot streaks will tend to migrate towards the rotor pressure surface. This fundamental condition exists because the incidence and velocity variations associated with a American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization. AIAA-2001-0525 cold streak are opposite those associated with a hot streak. While there have been numerous experimental (e.g., Refs. [1, 3, 4, 5]) and numerical studies (e.g., Refs. [6, 7, 8, 9]) of the effects of hot streaks, less effort has been given to the study of the combined effects of unsteady cold and hot streaks. In order to achieve a better understanding of the combined effects of hot streak/cold streak cogeneration a numerical study has been devised in which a streak with time-dependent temperature variations is introduced into the inlet flow of a 1-1/2 stage axial turbine. Using the geometry of the United Technologies Research Center (UTRC) Large Scale Rotating Rig (LSRR), simulations have been performed in which timevarying temperature profiles were introduced at the entrance to the first-stage stator and traced throughout the stage. The study involved varying the frequency of the time-dependent temperature variations.

Flowfield Analysis of a Scramjet Combustor with a Coaxial Fuel Jet

A calculation procedure for predicting turbulent mixing and burning.and skin friction and wall heat transfer in a scramjet combustor with a central fuel jet from a gas generator has been developed. All the important physical and chemical processes have been modeled, including the upstream influence of heat release on the initial conditions and the pressure distribution in the duct. Calculations for a representative engine with Shelldyne H fuel at Mach 4 and 7 indicate that a combustor 1 to 2 m long is sufficient to insure complete heat release but that substantial nonuniformity of the combustor exit flow would still exist. The skin friction and wall heat transfer are both shown to be very sensitive to the local pressure gradient and to the impingement of combustion products and turbulence from the central jet mixing and burning zone. However, the responses of these two quantities are not of the same importance, and they are not in phase.