Combustor Liner Research Articles

Effect of Laser Shock Peening on High-Cycle Fatigue Performance of 1Cr18Ni9Ti/GH1140 Weld

The welded joints of 1Cr18Ni9Ti austenitic stainless steel and GH1140 nickel-based superalloy dissimilar materials used in certain types of aero-engine combustion liner components are prone to crack initiation during service, seriously affecting the service life of the combustion liner. In this study, laser shock peening (LSP) was applied to the dissimilar metal weld of 1Cr18Ni9Ti and GH1140, which are used in the combustion liner parts of aero engines. The effects of LSP on the residual stress, microhardness, microstructure and high-cycle fatigue performance of the weld were analyzed. The results show that the residual stress in the weld and heat-affected zones was converted from tensile residual stress to high amplitude compressive residual stress via LSP. Furthermore, the surface hardness of every region of the combustion liner weld was increased, especially in the weld zone, where an increase of 41.4% from 162 HV to 229 HV was observed. Simultaneously, with the introduction of grain refinement, gradient plastic deformation in the depth direction and the dislocation structure of the surface material, the high-cycle fatigue limit of the weld specimen was significantly increased and the fatigue limit of the 1Cr18Ni9Ti/GH1140 welded joint was improved by 65.39%, from 289 to 478 MPa.

Effects of lateral cooling hole configuration on a swirl-stabilized combustor

It is necessary to cool the combustor liner due to its harsh operating conditions. In the current study, wall temperatures of a swirling combustor with and without cooling holes are compared using a three-dimensional numerical simulation with conjugate heat transfer method. It is shown that the wall temperature can be reduced by 344–501 K by using cooling holes. Two lateral coolant injection designs are then proposed since the laterally injected coolant may interact with the swirling flow. Normal axial injection is also investigated to compare NOx emissions, cooling effectiveness and total pressure loss for various scenarios with different injection angles. The results show that when the incidence angle is 15°, NOx emissions remain low for all incidence configurations. When the lateral incidence is implemented, especially at small incidence angles such as 15° and 30°, the cooling effectiveness is maintained above 0.9. However, when the coolant incidence direction is opposite to the swirling flow and the incidence angle is 15° results in a high total pressure loss. At an incidence angle of 60°, the total pressure loss is about the same for all cases, at 3.2%.

Conjugate heat transfer characteristics of effusion cooling under realistic swirl flow in a three-sector gas turbine model combustor

• Convective heat transfer inside effusion holes contributes greatly to the heat removal even the film layer is damaged. • The full-coverage cooling film is developed and reachable after swirl flow-cooling jets intense interaction. • The swirl flow impinging location is not influenced by amount of main and cooling flows. · Increment of coolant. Conjugate heat transfer characteristics of the effusion cooling applied on the swirl-stabilized combustor liner is experimentally and numerically investigated in a three-nozzle gas turbine model combustor at non-reacting flow conditions. The interaction between the strong swirl main flow and multi-hole cooling jets inclined at 30° is stressed by varying the pressure drop across the nozzle ( Δ P m = 3–5%) and effusion plate ( Δ P e = 0–2%). The expanding swirl flow impinges on the liner wall and thus inhibits the injection of coolant via effusion holes locally, causing non-uniform cooling protection along both streamwise and lateral directions. The overall cooling effectiveness in the impingement region is much lower than other regions on the cooled plate. Despite of the detrimental effect of the main flow impingement on the formation of the effusion film, the cooling function doesn’t totally collapse due to the great capability of heat removal through convection heat transfer inside the cooling holes. Superposition effect induced full coverage of film on the wall is distinguished at further downstream. Enhancing the cooling air amount weakens the consequence of the mainstream impingement and lead the full coverage initial location to move upwards.

Investigating the Influence of Mainstream Turbulence and Large Vortices on Gas Turbine Trenched Film Cooling Flows Using Planar Temperature-Velocity Imaging

Abstract Film cooling is a critical gas turbine technology to reduce heat loads on the combustion chamber and the turbine vanes and blades. To improve film cooling performance, trenched cooling configurations where the holes are connected by a cross-stream slot in the surface have been developed. In this study, for the first time, we experimentally examine the influence of mainstream turbulence and large vortices on trenched film cooling flows. We apply an imaging technique based on thermographic phosphor particles seeded into the flow to measure 2D time-resolved gas temperature and velocity distributions in film cooling flows established in a closed-loop heated wind tunnel. Two trenched configurations (straight and an optimized zigzag design) are compared with ordinary effusion holes at momentum ratios (I=3.5,5.7, and 8.3) relevant to gas turbine combustors. Furthermore, two turbulence generation methods were used to create realistic mainstream turbulence levels (Tu=16.4% and 22.7%) and large vortices, so film cooling flows at typical wind tunnel turbulence (Tu=5.3%) can be compared with combustor-relevant conditions. By comparing the film position derived from the time-average temperature fields, the optimized trench performs best at low momentum ratios and turbulence levels but the performance rapidly drops when the momentum ratio rises or the turbulence level is increased. The straight trench performs best across all conditions studied and this configuration is therefore recommended for combustor liner cooling where main flow turbulence levels are high. In general, increased turbulence intensity reduces the effective length of the cooling film for all geometries. The straight trench, however, produces a more stable cooling film than typical effusion holes or the optimized trench configuration. Both phase-locked and time-resolved data indicate that, in the presence of a dominating frequency in the turbulent main flow field produced using a vortex generator, there are instances where the cooling films are strongly disturbed in the streamwise direction and hot gas is locally in contact with the surface.

Exploration on Nonaxisymmetric Flow Phenomenon in a Slinger Injector

The slinger is an important configuration of an aeroengine combustion chamber. It is well matched to smaller gas turbines that are more sensitive to cost than larger and more sophisticated models. The slinger combustor is inherently low cost itself, by eliminating fuel injectors and much of the fuel tubing and manifolding and also reducing costs in the requirements of the main fuel pump by operating at fuel delta pressures significantly lower than conventional fuel systems. The slinger's inherent drawbacks of increased combustion liner surface area and less atomization are less of a concern in a lower temperature, lower cost small turbines. The slinger's internal flow is of great significance to fuel atomization. This paper conducts experiments and numerical simulations on the slinger fuel injector. Numerical analysis and high-speed photography of the fuel slinger were applied to investigate the film flow pattern inside the slinger and the liquid phase distribution inside the combustion chamber to predict the flow field in an aeroengine combustion chamber. The transient Reynolds-averaged Navier–Stokes method, the volume of fluid method, and the discrete phase method were adopted in the simulation, and a high-speed camera and the rotational rig were used to perform the experiments. The nonaxisymmetric flow was present in the slinger and the combustor space. In the simulation results, the total mass-flow rate varies with time. Each hole's mass-flow rate value in the slinger is also different at the same time. The Sauter mean diameter (SMD) difference for each injection orifice is relatively small compared with the SMD difference caused by a rotation rate. The spatial distribution of the spray is also uneven as shown in the result of different single discharge channels' mass-flow rates. The experimental photos confirmed the simulation outcomes. The general theoretical analysis was made that these nonaxisymmetric phenomena were driven by the combined action of centrifugal force, surface tension, and instability phenomenon.

Effect of Spanwise Hole-to-Hole Spacing on Overall Cooling Effectiveness of Effusion Cooled Combustor Liners for a Swirl-Stabilized Can Combustor

Abstract One of the most effective ways to cool the combustor liner is through effusion cooling. Effusion cooling (also known as full-coverage effusion cooling) involves uniformly spaced holes distributed throughout the combustor liner wall. Effusion cooling configurations are preferred for their high effectiveness, low-pressure penalty, and ease of manufacturing. In this article, experimental results are presented for effusion cooling configurations for a realistic swirl driven can combustor under reacting (flame) conditions. The can combustor was equipped with an industrial engine swirler and gaseous fuel (methane), subjecting the liner walls to engine representative flow and combustion conditions. In this study, three different effusion cooling liners with spanwise spacings of r/d = 6, 8, and 10 and streamwise spacing of z/d = 10 were tested for four coolant-to-main airflow ratios. The experiments were carried out at a constant main flow Reynolds number (based on combustor diameter) of 12,500 at a total equivalence ratio of 0.65. Infrared thermography (IRT) was used to measure the liner outer surface temperature, and detailed overall effectiveness values were determined under steady-state conditions. The results indicate that decreasing the spanwise hole-to-hole spacing (r/d) from ten to eight increased the overall cooling effectiveness by 2–5%. It was found that reducing the spanwise hole-to-hole spacing further to r/d = 6 does not affect the cooling effectiveness implying the existence of an optimum spanwise hole-to-hole spacing. Also, the minimum liner cooling effectiveness on the liner wall was found to be downstream of the impingement location, which is not observed in the existing literature for experiments done under nonreacting conditions.

Gas turbine combustor CFD study and single-objective DoE optimization

A fast RANS method for methane-air non-premixed combustion is proposed, using a k-ω SST model for turbulence, a transport species model for chemistry and eddy dissipation model for turbulence-chemistry interaction. The method is validated with experimental data from Sydney Swirl Flow and Flame Database. The same method is then used to simulate a two-dimensional axisymmetric combustor liner geometry, and finally a Design Exploration is proposed to optimize the Pattern Factor at the combustor outlet, obtaining the response surface for the entry variables, namely the Swirl Number and mass flows for each row of holes. An optimal combustor configuration is chosen. The results are then discussed and compared to similar works, showing the viability of the simplified CFD simulation and DoE optimization methodology.

CFD analysis of combustion liner with different geometry holes for effusion film cooling

The advanced gas turbine is designed to operate at higher temperature to increase thermal efficiency. Since the gas temperature exceed the allowable material temperature, so cooling techniques of turbine components are more important. Effusion film cooling is an external and highly sophisticated cooling technology, which protects the liner by injecting the secondary cooler fluid at particular locations over the liner surface exposed to the higher gas temperatures. The injected fluid forms as a thin layer (film) on the liner surface and protects the liner surface from higher gas temperatures. In elliptical geometry of holes in plate for effusion cooling is studied for the samples made of adiabatic and conductivity material. To increase the life of combustion chamber it is essential to know the amount of cooling air to be used for cooling of the liner. For effusion film cooling, geometry of the hole is one of the key parameters which influences the coolant mass flow rate and it spread over the liner surface thus affecting the cooling effectiveness. The commercially available FloEFD software is used for calculating the surface temperature and effusion cooling effectiveness distribution throughout the sample plate and result values are in good agreement and this has given confidence to usage of the commercial software FloEFD for predicting the film effectiveness. After result validation finally concluded that span wise elliptical hole (Config-2) have more effectiveness than circular holes.

Investigation of the acoustics of full-scale perforated liners in gas turbine combustors

In conventional gas turbine combustors, the combustion chamber linings are perforated and used for cooling. To cool the liner evenly, bias flows are introduced into the combustor at rates that depend on the operating condition. It has been found that airflow through the liner not only provides cooling but also improves sound absorption and acoustic instability. This experiment reports a unique comprehensive investigation of the influences of single and double-layer cylindrical full-scale gas turbine combustor liner sound absorption properties based on the no flow and non zero bias flows. In particular it is shown that combustor liner porosity (determined by orifice diameter and axial pitch distance) has an important influence on non-zero bias flow in that it increases the peak absorption or dissipation compared with that which occurs in the absence of flow. It is shown that the main influence of bias flow is to increase absorption compared with no flow above 600 Hz and to decrease the transmission loss measured in the absence of flow below 600 Hz but to increase it above 600 Hz. Internal resonance in the combustion liner test section influences both absorption and transmission loss spectra near 600 Hz. To create higher damping, and decrease in acoustic instability during the combustion process, gas turbine combustors require mapping between the inner and outer liner perforation to increase efficiency and lower the hydrocarbon emission. The calculated pressure ratio versus mass flow and combined discharge coefficient effect explain the non-linear distribution of the absorptive and dissipative energy measured at the gas turbine combustor.

Thermodynamics of CMC bladed marine gas turbine-LM 2500: Effect of cycle operating parameters

Recent developments in ceramic-matrix composites and their successful use in combustor liners and shrouds have generated interest among researchers to adopt these materials in rotating gas turbine blades, especially in the first stages of high-pressure turbines where gas temperatures are highest. CMC blades have the potential of being retrofitted to replace superalloy turbine blades in operating gas turbines. In this paper, a comparative study on the thermodynamic performance of a marine gas turbine engine, LM 2500, featuring directionally solidified nickel superalloy blades versus novel CMC blades in the high-pressure turbine sections has been reported. Mathematical modeling of the gas turbine cycle components has been developed and then coded in C++ language. The effects of turbine inlet temperature on thermodynamic efficiencies, coolant mass flow rates, and work ratios on the two systems have been analyzed. Finally, the exergy analysis of the systems’ components has been done to identify the benefits of adopting CMC blades in the LM 2500 system. It has been observed that when compared to the directionally solidified bladed turbine system, the projected first law efficiency of CMC bladed LM 2500 gas turbines can be enhanced over 7% (from 34.17% to 41.21%). The projected work ratio can be improved by over 16% (from 0.49 to 0.57) at the turbine inlet temperature of 1725 K.

A Comprehensive Review on the Application of 3D Printing in the Aerospace Industry

Additive manufacturing is a process wherein a three-dimensional object is created layer-by-layer. It offers adaptability to the geometrical complexity and customizability of the design, which is difficult to manufacture using conventional manufacturing. The aerospace industry is one of the sectors that first adopted additive manufacturing, particularly three-dimensional printing (3D printing) in the production of aircraft parts such as rocket engine components, oil fuel tanks, environmental control system ducting, combustor liner, custom cosmetic aircraft interior components, and unmanned aerial vehicle (UAV) components. The aircraft's most common materials used in the 3D printing prototype parts are acrylonitrile butadiene styrene (ABS) thermoplastic, carbon-fiber and thermoplastic composite, and nylon 12 using selective laser sintering, fused deposition modeling, or composite filament co-extrusion technology. One of the aerospace industry's challenges is ensuring the efficiency and quality of aircraft structural parts that typically require complicated manufacturing due to their complexity and variability of function. Additive manufacturing is seen to respond to this challenge by developing and prototyping 3D printed parts and exploring practical 3D printing technologies.

Comparison of adiabatic and conjugate heat transfer models on near-wall region flows and thermal characteristics of angled effusion cooling holes

This paper aims to evaluate the accuracy and efficiency of different numerical simulation approaches on cooling performance for the angled effusion cooling holes applied in a micro gas turbine (MGT) combustor liner. The geometric structure processing methods of a 3-D multi-perforated plate are performed for conjugate heat transfer (CHT) and adiabatic models. Numerical approaches are verified by experimental data in terms of centerline film cooling effectiveness. The wall cooling film effectiveness, film thickness and discharge coefficient of the multi-perforated plate are studied and discussed under the condition of whether there is solid heat conduction. Results show that the adiabatic model cannot predict the veritable wall cooling effectiveness distribution on the inner wall surface and overestimates the discharge coefficient by nearly 16.5%. However, the accuracy of the adiabatic model in predicting the near wall flow field structure is not lower than that of the CHT model. Both of the two model can predict a film thickness of 4.1 mm at the same sampling position. Further studies indicate that the increases of flow direction β have a significant impact on the improvement of wall cooling effectiveness distribution in the spanwise direction. The Ux magnitude predicated by adiabatic model is increasingly close to that of CHT model. The predication results also illustrate a relatively fixed difference of 0.1 between these two models.

On-Design Component-Level Multiple-Objective Optimization of a Small-Scale Cavity-Stabilized Combustor

Abstract This work presents an on-design component-level multiple-objective optimization of a small-scaled uncooled cavity-stabilized combustor. Optimization is performed at the maximum power condition of the engine thermodynamic cycle. The computational fluid dynamics simulations are managed by a supervised machine learning algorithm to divide a continuous and deterministic design space into nondominated Pareto frontier and dominated design points. Steady, compressible three-dimensional simulations are performed using a multiphase realizable k–ε RANS and nonadiabatic flamelet/progress variable combustion model. Conjugate heat transfer through the combustor liner is also considered. There are fifteen geometrical input parameters and four objective functions viz., maximization of combustion efficiency, and minimization of total pressure losses, pattern factor, and critical liner area factor. The baseline combustor design is based on engineering guidelines developed over the past two decades. The small-scale baseline design performs remarkably well. Direct optimization calculations are performed on this baseline design. In terms of Pareto optimality, the baseline design remains in the Pareto frontier throughout the optimization. However, the optimization calculations show improvement from an initial design point population to later iteration design points. The optimization calculations report other nondominated designs in the Pareto frontier. The Euclidean distance from design points to the Utopic point is used to select a “best” and “worst” design point for future fabrication and experimentation. The methodology to perform computational fluid dynamics optimization calculations of a small-scale uncooled combustor is expected to be useful for guiding the design and development of future gas turbine combustors.

FILM COOLING STUDIES ON COMBINED THREE-DIMENSIONAL SLOT AND EFFUSION JET CONFIGURATION OF AN ANNULAR COMBUSTOR LINER

A detailed numerical investigation has been carried out to study the film cooling characteristics of a combined three dimensional slot and effusion jet configuration of a combustor liner. The effect of slot jet geometry, effusion jet pattern and coolant mass flow ratio on film cooling performance is identified. Mass flow ratio is the flow parameter that determines the flow transition between pure slot and effusion film. The three-dimensional slot coupled with the staggered effusion array is studied at a range of blowing ratios (BR), i.e., BR= 0.6-3. The enhancement in adiabatic effectiveness obtained with combined film cooling over pure three-dimensional slot cooling is reported in terms of laterally averaged and area averaged effectiveness on the film cooled surface. The most commonly studied two-dimensional planar slot is seen to have a cooling behavior different from the three-dimensional slots employed in an actual combustor, where the discrete nature of three-dimensional slot jets resulted in a lower film cooling effectiveness. Additionally, the adiabatic film cooling effectiveness has a non-linear variation with the coolant mass flow ratio (MFR) at low BRs (<2), and a monotonic decrease with MFR at high BRs (>2). Within the low blowing ratio range, the maximum improvement in the area average adiabatic effectiveness is 14% at BR=1.8, whereas it is 34% at high blowing ratio, BR=3. The standard deviation of effectiveness, a measure of uniformity in film cooling, is also lower with the combined film cooling configuration. A maximum reduction of 32% in standard deviation is noticed for BR=2.4 at MFR=1.

Effect of blowing ratio on adiabatic effectiveness for effusion cooling in gas turbine combustor liners

Effect of blowing ratio on adiabatic effectiveness for effusion cooling in gas turbine combustor liners

Comparisons of Cooling Performance and Flow Characteristics of a Combustor Liner Plate with Compound Angle and Simple Angle Effusion Holes

Comparisons of Cooling Performance and Flow Characteristics of a Combustor Liner Plate with Compound Angle and Simple Angle Effusion Holes

Numerical Investigation of Effusion Cooling in Gas Turbine Combustor Liner

In this study, a numerical analysis was performed to find the aerothermal characteristics of the effusion cooled gas turbine combustor liner. The analyzed geometric model is a scale model of an actual combustor liner. The study aims to investigate the effect of different blowing ratios by validating an experimental test setup. In experimental studies on effusion cooling, the sidewall effect is a serious problem that can distort the results. Numerical analyses provide advantages in visualizing temperature and velocity contours in different sections of physical model. The counter rotating vortex pairs, the horseshoe vortex and the recirculation zone are the main flow features of the jet mixture. At a blowing ratio of 3.35, numerical analyzes gave the highest value of cooling effectiveness. Although the blowing ratio slightly changes the cooling effectiveness in experimental data, it has been found that the effect of blowing ratio is more pronounced on the numerical results, especially at high blowing ratios.

Influence of pin-fin height and diameter on flow and cooling characteristics of three-layer porous laminates: An experimental study

ABSTRACT Pin-fins play a very important role in heat-transfer enhancement in engine combustor liners. In this engineering-scale experimental study, the flow and cooling characteristics of two- and three-layer porous laminates with pin-fins were compared. The influence of the diameter and height of the pin-fin on the flow and cooling characteristics of three-layer porous laminates was also investigated. Adding a third (impingement) layer to the two-layer porous laminate in the study increased the heat-transfer coefficient by 66.00% and the cooling efficiency by 5.90%; both performance factors were further improved as the pin-fin diameter and height in the three-layer laminate were reduced. An empirical expression for the Nusselt number of a three-layer porous laminate was obtained, with a prediction error below 5%.

Thermal Analysis and Creep Lifetime Prediction Based on the Effectiveness of Thermal Barrier Coating on a Gas Turbine Combustor Liner Using Coupled CFD and FEM Simulation

Thermal barrier coating (TBC) plays a vital role in the gas turbine combustor liner (CL) to mitigate the internal heat transfer from combustion gas to the CL and enhance the parent material lifetime of the CL. This present study examined the thermal analysis and creep lifetime prediction based on three different TBC thicknesses, 400, 800, and 1200 μm, coated on the inner CL using the coupled computational fluid dynamics/finite element method. The simulation method was divided into three models to minimize the amount of computational work involved. The Eddy Dissipation Model was used in the first model to simulate premixed methane-air combustion, and the wall temperature of the inner CL was obtained. The conjugate heat transfer simulation on the external cooling flows from the rib turbulator, impingement jet, and cross flow, and the wall temperature of the outer CL was obtained in the second model. The thermal analysis was carried out in the third model using three different TBC thicknesses and incorporating the wall data from the first and second model. The effect of increasing TBC thickness shows that the TBC surface temperature was increased. Thereby, the inner CL metal temperature was decreased due to the TBC thickness as well as the material properties of Yttria Stabilized Zirconia, which has low thermal conductivity and a high thermal expansion coefficient. With the increase in TBC thickness, the average temperature difference between the TBC surface and the inner metal surface increased. In contrast, the average temperature difference between the inner and outer metal surfaces remained nearly constant. The von Mises equivalent stress, based on the material property and thermal expansion coefficient, was determined and used to find the creep lifetime of the CL using the Larson–Miller rupture curve for all TBC thickness cases in order to analyze the thermo-structure. Except in the C-channel, the increasing TBC thickness was found to effectively increase the CL lifespan. Furthermore, the case without TBC was compared with the damaged CL with cracks due to thermal stress, which was prevented by increasing TBC thickness shown in this present study.

Liquid ammonia spray combustion in two-stage micro gas turbine combustors at 0.25 MPa; Relevance of combustion enhancement to flame stability and NOx control

The development of liquid ammonia spray combustion technology is expected to reduce the cost and size of gas turbines otherwise fueled with gaseous ammonia. However, successful studies on this topic are scarce due to the difficulty in flame stabilization. In this study, the flame stability and emissions control in liquid ammonia spray combustion is studied for the first time for input thermal power up to 230 kW using novel two-stage gas turbine combustors with and without slot film cooling at 0.25 MPa. The experiments were conducted in a micro gas turbine combustor test facility originally developed for gaseous ammonia combustion. The results show that the injection of liquid ammonia spray into a methane-air flame encouraged rapid cooling of the combustor liner head, enhancing flame blowoff. The blowoff limit based on ammonia fraction in the fuel showed a strong dependence on the combustor temperature, hence the combustor without slot film cooling extended the blowoff limit allowing the successful stabilization of pure liquid ammonia spray flame. Result of emissions analysis showed that two-stage rich-lean combustion is effective in controlling emissions from the flames, however, requires further enhancement of the flame in the primary combustion zone. An increase in combustor inlet temperature and a decrease in combustor wall heat loss, both of which enhanced the combustion of the liquid spray, resulted in a decrease in the emissions of NO, NO2, N2O, NH3, HCN and CO from liquid ammonia co-fired with methane. Conversely, pure liquid ammonia spray combustion resulted in higher emissions of NH3, NO and N2O than did the co-combustion with methane, encouraged by inhomogeneity in droplets distribution and larger heat transfer to the liquid droplets.