Combustor Liner Research Articles

Isolating influences of varying pitch from the effects of non-zero compound angles on effusion cooling

Abstract Effusion cooling is the state-of-the-art cooling technology for gas turbine hot-gas path components. Typically, effusion cooling holes across the entire combustor liner are aligned with the combustor axis, rendering a nominal zero compound angle between highly directional miniature effusion cooling jets and the main flow direction. The pitch of effusion cooling holes is optimised accordingly. However, the swirling main flow results in a non-zero compound angle and an effectively different pitch from the design. The directional effect of effusion cooling as a result of swirling main flow on the adiabatic film cooling effectiveness (AFE) is a combined effect of a non-zero compound angle and a varied pitch. The current experimental study aims to investigate the isolated effects of compound angle on AFE by excluding the influences of varying pitch. With an improved understanding of the sole effects of non-zero compound angles on AFE, the roles that a varied pitch plays in modifying AFE are further discussed to guide future effusion cooling designs under swirling main flow conditions. Binary pressure sensitive paint (PSP) was used to determine AFE experimentally.

Effect of laser drilling process on combustor liner effusion hole quality

Abstract Effusion cooling is one of the significant cooling technologies in combustor liners in terms of cooling efficiency and weight reduction. However, effusion cooling technology is difficult to manufacture. In fact this technology requires laser-drilling of thousands of tiny holes with shallow angles on a sheet metal with a thickness generally varying between 0.5 to 1.5 mm. In addition, the use of thermal barrier coating is common in gas turbine engines and is one more challenge for the drilling process. In order to obtain more efficient gas turbine engines, the inlet temperature keeps increasing in the last decades, which induces the combustion chamber to operate in a hotter environment. Therefore, efficient cooling technology is needed, even if it is hard to manufacture. For laser drilling, several parameters have to be explored to obtain acceptable holes. This study includes the microstructure investigations of the holes produced with different laser parameters and the optimal laser parameters determined according to the microstructure of six different effusion cooling hole configurations. The results show that laser process differences affect the metal substrate microstructure and thermal barrier coating structure. Drilling method, peak power, number of pulses, gas type and pressure value have a significant effect on the hole geometry and its microstructure.

Isothermal Flow Field Characterization of a Full-Scale Sector Combustor at Elevated Pressures

Abstract An experimental investigation in a sector (20 deg) of full-scale annular gas turbine combustor is performed. The sector combustor is optically accessible for the flow and flame visualization of the primary and exit zones of the combustor. The distinctive feature of the experimental setup is that it preserves the geometrical details of an annular combustor that includes the casing, dome and combustor liner. The combustor design features a series of primary and secondary dilution holes with multiple film cooling strips on the outer and inner liner. In the present study, the combustor is operated at inlet Mach numbers of 0.02–0.3 at operating absolute pressures of 1–5 bar. Static pressure measurements are performed at multiple locations in the rig to characterize the pressure drop across the combustor. Two-dimensional particle image velocimetry (PIV) is performed to measure the velocity fields of the primary and exit zones of the combustor simultaneously. The results show the presence of a central recirculation zone (CRZ), high-velocity annular jets, and a pair of dilution jets in the primary zone of the combustor. The steady-state flow structures are invariant of inlet Mach number and pressures. The relationship between the relative pressure drop across the combustor and the combustor inlet condition is obtained. Mass flowrate and momentum flux are calculated for the flow through the swirler, central recirculation zone, the primary dilution jets, and the exit zone. The paper shows how the flow structures in a realistic combustor change with variations in global combustor parameters.

THE EFFECT OF THERMAL CYCLING ON DEPOSITION IN AN IMPINGEMENT/EFFUSION COOLING GEOMETRY

Abstract Experiments were conducted in a high-pressure deposition facility to study the effect of engine cycling on the successive buildup of deposits as aircraft shut down and power up. The test facility simulates the pressure and temperature environment of a combustor liner cooling circuit. The coolant flow temperature reaches 894K (1150 °F) and discharges into a 17atm (250 psi) cavity pressure at a nominal pressure ratio of 1.027. AFRL05 test dust with a 0-10micron size distribution is added to the coolant . The cooling circuit consists of a double-walled impingement/effusion cooling plate with nominal hole sizes on the order of 0.5mm. To simulate cycling, the facility is brought up to the desired operating conditions where the first batch of dust is delivered (2-8g). The facility is then ramped down to ambient conditions. Following a “dwell” period of approximately 24 hours, another batch of dust is delivered once the facility is brought back up to the desired operating condition. Test data were acquired for one, two, and four cycles with different dust mass delivered. Values such as discharge coefficient, pressure ratio, Reynolds Number, and temperature are used to evaluate the effects of dust deposition on the impingement/effusion plate setup. Dust capture efficiency is shown to be insensitive to cycling whereas flow blockage is negatively impacted by increased number of cycles for the same total mass delivered. The sloughing of deposit structures during cooling circuit cool down is postulated to be responsible for the observed behavior.

Simulating conjugate heat and mass transfer by the advanced perforated wall model for effusion cooling

Full-scale Computational Fluid Dynamics (CFD) numerical simulations can accurately predict the heat flux and metal wall temperature of gas turbine combustor liners, but they require considerable computational resources. In present work, an Advanced Perforated Wall model (A-PWM) has been developed to resolve the flow and heat transfer behaviors in this confined region. The geometric structure of the effusion cooling perforated plate is simplified, while the fluid-solid coupling heat transfer processes are still fully considered in details. The three-dimensional Reynolds Average Navier–Stokes simulations have been carried out for both a single-hole cooling plate and a multi-hole perforated cooling plate by using A-PWM. For simulations of the single-hole cooling plate, the cooling effectiveness in the downstream of the jet flow with the x/D ranging from 2 to 8 predicted by A-PWM shows good agreement with conjugate heat transfer model and experimental results. For multi-hole perforated cooling plate, the differences of cooling effectiveness between the prediction results by A-PWM and experimental results are just within 5%, confirming the developed model’s accuracy in predicting the effusion cooling. In addition, the A-PWM can also be employed to analyze the temperature distribution inside the solid zone. In simulations of both the single-hole cooling plate and the multi-hole perforated cooling plate, the use of the A-PWM model can lead to a reduction of over 35% in grid count compared to the conjugate heat transfer model, significantly reducing the computational time.

Enhancing thermal stability and corrosion resistance of carbon-carbon composites with iridium coatings deposited by electron beam physical vapor deposition

Abstract Carbon–carbon (C−C) composites are extensively used in high-temperature environments such as Combustor Liners and Turbine Blades in jet engines and Throat Inserts, Nozzle Extensions and Exit Cones in rocket engines due to their excellent thermal stability and mechanical properties. However, at temperatures exceeding 800 °C, these composites require additional protection to prevent degradation. This study aims to investigate the behavior of C−C composites when coated with high-purity metallic iridium using Electron Beam Physical Vapor Deposition (EBPVD). The research problem focuses on enhancing the high-temperature performance and corrosion resistance of C−C composites for aerospace applications. The methodology involves depositing a uniform 5.6 microns thick iridium coating on C−C substrates and characterizing the coating’s microstructure, hardness, and corrosion resistance. FESEM micrographs reveal that the iridium coating adheres uniformly to the substrate without any seepage, and XRD analysis confirms an FCC crystal structure with a densely packed grainy surface. Corrosion tests were conducted using a BIOLOGIC electrochemical workstation in a sodium chloride environment indicate a corrosion rate of 0.00307 mm year−1. The Nyquist, Bodo plots, and Taffel plots were constructed for the better understanding of the corrosion mechanism. While the OCP was constructed to understand the stability and the corrosion resistance of the C−C samples. Microhardness of the coating, measured under a normal applied load of 0.20 N, is 702 HV. The coated samples also could withstand thermal shocks between −40 °C and 1500 °C for 40 h without observable damage or color change. These findings demonstrate the potential of iridium-coated C−C composites to maintain structural integrity and performance in extreme aerospace environments, significantly impacting the field by providing a reliable protective solution for high-temperature applications.

Numerical Investigations on the Effects of Dome Cooling Air Flow on Combustion Characteristics and Emission Behavior in a Can-Type Gas Turbine Combustor

To meet the requirements of achieving higher efficiency and lower NOx pollution, the flame temperature in gas turbine combustors increases continually; thus, the effusion-cooling technology has been used to ensure the combustor liner remains within the allowed temperature, by which the combustion characteristics and emission behavior are possibly influenced. In order to investigate the effects of dome cooling air flow on combustion characteristics and NOx emissions, three-dimensional combustion simulations for a swirl-stabilized can-type gas turbine combustor are carried out in this work by using the computational fluid dynamics (CFD) method. Through adjusting the ratio of the dome cooling air flow and the dilution cooling air flow, the characteristics of flow field, temperature distribution and NOx emissions under each work condition are analyzed. At different ratios of the dome-cooling air flow to the total air flow, the flow velocity field in the region near the center of the combustion chamber is not changed much, while the velocity field near the chamber wall shows a more significant difference. The temperature in the outer recirculation zone within the combustion chamber is effectively reduced as the dome cooling air flow increases. By analyzing the distribution characteristics of the concentration of OH*, it is demonstrated that the dome cooling air flow does not have a direct effect on the reaction of combustion. It is also found that as the ratio of the dome cooling air flow to the total air flow increases from 0 to 0.15, the value of the NOx emissions drops from 28.4 to 26.3 ppmv, about a 7.4% decrease. The distribution of the NOx generation rate in the combustion chamber does not vary significantly with the increasing dome cooling air flow. Furthermore, by calculating the residence time in different stages, when the the ratio of the dome cooling air flow to the total air flow varies from 0 to 0.15, the residence time in the pilot stage decreases obviously, from 42 ms to 18 ms. This means that reduction in residence time is the main factor in the decrease of NOx emissions when the dome cooling air flow increases.

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.

Comparison of Response Surface Based Preliminary Design Methodologies for a Gas Turbine Combustor

Preliminary design of gas turbine combustor is a multi-objective optimization problem. The methodology to be used at the preliminary design stage depends on the freedom of design choices available. In this article, we explore three preliminary design methodologies for gas turbine combustor - M1: combustion liner design for a given casing; M2: combustion liner design without the casing and M3: coupled design of combustion liner and casing. A workflow for the automated design space exploration of gas turbine combustor using response surface methodology is presented. Computational fluid dynamics studies along with central composite design for design of experiments and genetic aggregation for response surface generation are used to quantify the combustor performance in design space. Comparison of three different design methodologies (M1, M2 and M3) is made to show how the choice of design methodology changes the available design space and limits/expands combustor performance. Candidate optimal designs and associated trade-offs from the optimization study are also presented. This study can aid combustor design engineers choose the most suitable preliminary design methodology for their specific use case.

Spectral effects of radiating gases and soot on an experimental aero-engine combustion chamber using full-spectrum k-distribution method

Radiative heat transfer is the main mode of heat transfer and has been proven to be important in the aero-engine combustion. Those radiating gases and soot resulted from combustion may display very strong spectral, or “nongray” behavior, which is difficult to both characterize and evaluate. In this work, the full-spectrum k-distribution (FSK) method is employed to resolve the spectral behavior of both radiating gases and soot during combustion in an experimental aero-engine. Different FSK-based spectral treatments are combined to illustrate how the spectral behavior affects the aero-engine combustion. Those weighted-sum-of-gray-gases (WSGG) results previously done by the authors (Wang, et al. 2024) are also partially compared to the FSK results in this work. Results show that different FSK-based spectral treatments result in small differences for temperature distribution of the flow field. However, since different spectral models are found to impose a considerable effect on radiative heat transfer rate, it is inferred that spectral behavior may be deeply coupled with the generation of pollutants. Results also show that the wall temperature of the combustor liner is highly affected by the spectral model: ignoration of soot radiation underestimates the wall temperature while assuming soot to be gray overestimates the wall temperature. Therefore, high-fidelity spectral models are recommended during the modelling of aero-engine combustion.

Mechanisms Leading to Stabilization and Incomplete Combustion in Lean CH4/H2 Swirling Wall-Impinging Flames

Abstract In gas turbines, confined highly turbulent flames unavoidably propagate in the vicinity of a relatively cool combustor liner, affecting both the local flame structure and global operation of the combustion system. In our recent work, we demonstrated, using simultaneous [OH] × [CH2O] planar laser-induced fluorescence (PLIF) and stereo-particle image velocimetry (stereo-PIV), that lean CH4/H2 flames at a high Karlovitz number can present a highly broken structure near wall, highlighted by a diffuse CH2O cloud, which suggests local quenching and incomplete oxidation. Such high Karlovitz numbers were achieved using an inclined plate, which substantially extended the lean flammability of the low swirl flames. Yet, how a cooled wall acting as a heat sink played a conducive role in stabilizing high Ka flames remains unanswered. In addition, the origin of the CH2O cloud is also unclear. Hence, in this work, we look to better understand the stabilization mechanisms for lean and ultralean flames on the same configuration, and how they may change with a parametric variation of plate incident angle, plate-nozzle distance, and bulk velocity up to the critical values that lead to flame blow off. The results show that the impinging swirling flow creates a low speed region that helps hold the flame, while the wall prevents mixing with ambient cold air. The production of diffuse CH2O, which indicates the occurrence of local quenching, is associated with a mean strain rate K beyond the extinction strain rate (ESR) Ke. For CH4 flames, most of the reaction zones reside within |K|/Ke&amp;lt;1; for 70% H2 flames at ϕ=0.4, the reaction zones are highly broken and scattered in a large area, where |K|/Ke&amp;lt;8, the interspace of which is fully filled by CH2O. In other words, high H2 fraction flames appear to be more robust to persistent strain rate, thus extending their stability envelope. However, these flames can subsist as highly broken flames featuring strong incomplete combustion.

Research on optimization of combustor liner structure based on arc-shaped slot hole

Research on optimization of combustor liner structure based on arc-shaped slot hole

IMPACT OF MEASUREMENTS TECHNIQUES ON HEAT TRANSFER CHARACTERISTICS IN AIR JET ARRAYS

Air impact processes have diverse applications in engineering, including backflow welding, textile drying, and gas turbine blade and combustion liner cooling. This research examines the influence of experimental methodologies and measurement tools on convective heat transfer in adjustable air jet assemblies. The experiment involves the use of heated targets made of thin stainless steel foil with constant heat flux boundary conditions. Thermography measures target surface temperatures by analyzing how internal passage cross-flow affects convective heat transfer via outflow adjustments. The experiments involve two arrays of jet nozzles: inline and staggering, each comprising 44 impingement jet nozzles arranged in 4 rows with 11 jet holes in each row. The study presents unsteady time average local and spatial Nusselt numbers as functions of jet Reynolds number (4630-14000) and explores their dependence on the jet nozzle diameter. Cross-flow levels significantly affect spatial and local Nusselt numbers in both local and span-wise averaged values, regardless of the Reynolds number. Strong cross-flow (single configuration) distributes flow causing turbulence and uneven heat distribution, reducing Nusselt numbers. In contrast, moderate cross-flow (double configuration) improves heat transfer and increases Nusselt numbers. The study emphasizes the crucial role of experimental techniques in heat transfer evaluation and demonstrates agreement with prior studies within a standard error below 5%.

Performance assessment and optimization of three-dimensional hybrid slot–effusion jet cooling configuration of an annular combustor liner

In this study, the film cooling performance of a three-dimensional (3D) hybrid slot–effusion configuration and its advantages over a baseline pure 3D slot cooling are investigated in detail using both experimental and numerical investigations. The numerical model is validated by in-house experiments conducted on a hybrid cooling configuration with a 3D slot and staggered effusion array. Infrared thermography and hot wire anemometry are used in this study to analyze the thermal and flow fields. The cooling performance of pure 3D slot and hybrid configurations is compared under identical coolant consumption. The coolant mass flow ratio (MFR) determines the cooling behavior in hybrid cooling. At low MFRs, hybrid cooling is less effective than pure 3D slot cooling. In addition, it is observed that there exists an optimal MFR where hybrid cooling outperforms 3D slot cooling. Subsequently, the effect of key design variables on the cooling performance of the hybrid configuration is studied, and they are optimized at typical combustor operating conditions using a genetic algorithm coupled to a Kriging surrogate model. The optimal configuration is experimentally tested at laboratory conditions and compared with the baseline 3D slot configuration, and it is noted that there is a significant improvement in the cooling effectiveness. Additionally, the film cooling performance of the optimized configuration at combustor representative operating conditions of maximum BR of 2 shows an enhancement of 16% in area-averaged adiabatic effectiveness and a 60% reduction in the standard deviation of local adiabatic effectiveness as compared to the baseline 3D slot cooling.

Numerical investigation of cooling performance of 3D-printable cooling structure for aero-engine combustor

Thermal protection is an essential technology for gas turbine hot components to cope with the unprecedented thermal loads, and to ensure their enduring and efficient operation. A 3D-printable hybrid cooling cellular structure is suggested in this paper, the cooling property of which in a Can-type combustion chamber with chemical reaction considered is evaluated in detail. Reynolds Stress Turbulence Model together with the non-premixed combustion model are used to resolve the detailed flow structures and turbulence-chemistry interactions. Cases of cooling cellular with various film hole inclined angles are compared and analyzed. Results show that the cooling cellular can reduce the temperature of combustor liner noticeably, especially at the region near the second row of dilution holes. Meanwhile, the cooling performance is improved as the decrease of inclined angle, with an improvement of 29.1% as the inclined angle range from 90° to 30°, because the smaller incline angle is more resilient to coolant lift-off. A comparation with effusion cooling scheme is also made and concluded that the case of hybrid cooling exhibits a 673%-1409% higher porosity ratio, while the cooling effectiveness is reduced about 3.6%-8.3%. The lower performance of the hybrid cooling is explored further with a flat plate conjugate heat transfer numerical test, and ultimately attributed to the relative lower flow rate of coolant air induced by higher pressure loss of hybrid cooling cellular. Finally, the compound angle is further introduced into the hybrid cooling cellular. Results show that the impact of compound angle on cooling property depends highly on the angle between the film outflow and the swirling main-flow. The performance of compound angle cases will only increase when cooling cellular deflects in opposite directions of swirling flow, not vice versa. Overall, the hybrid cooling cellular shows a great potential in the cooling design of combustor liner. A smaller incline angle with compound angle deflected in the opposite direction of the swirling main-flow are also recommended for increased cooling performance.

Interrelationships of entropy production and turbulence kinetic energy associated with simple angle and compound angle full coverage film cooling

The present paper compares experimentally-measured distributions of entropy generation with numerically-predicted distributions of streamwise vorticity, turbulence kinetic energy, and production of turbulence kinetic energy. These comparisons are provided for two full-coverage film cooling environments which are designed to model the thermal management arrangements employed for combustor liners within gas turbine engines. One of these arrangements includes alternating compound angle values of +30° in one row of holes, followed by −30° in the next row of holes. The other arrangement includes simple angle holes with a zero degree compound angle value. The associated full-coverage film cooled boundary layers are investigated within a double-wall cooling test facility wherein the air for each full-coverage film cooling arrangement is provided by an impingement jet array. Experimentally-measured entropy generation values are obtained from film cooled boundary layer measurements of local total pressure variations, obtained in isothermal flow, relative to the freestream values outside of the boundary layer. The resulting entropy generation values quantify second law losses, which are associated with aerodynamic gains. To obtain the associated numerical-prediction results for a steady-state, three-dimensional flow field, employed is the ANSYS FLUENT Version 2022 R2 computational code with a SST k–ω turbulence closure model. The resulting comparisons of entropy generation, streamwise vorticity, turbulence kinetic energy, and production of turbulence kinetic energy are provided for a blowing ratio BR of 2.9 for simple angle film cooling with a main flow Reynolds number Rems of 138,000, and for compound angle film cooling with a main flow Reynolds number Rems of 142,000. Overall, evidence is provided of strong relationships between local turbulence kinetic energy and local entropy production since the two quantities are highly correlated with each other over s substantial range of experimental conditions and configurations. For example, ratio of turbulence kinetic energy to entropy generation data collapse for Z/de values from 0 to 3.5 for Y/de = 0.5, Y/de = 1.5, and both film cooling arrangements. Here, the value of the TKE/Sgen ratio is approximately constant for these conditions, which evidences direct and simple dependences of local turbulence kinetic energy upon local entropy generation, even for lower magnitudes of Z/de. The direct and simple dependences of local turbulence kinetic energy upon local entropy generation are further illustrated by the simplicity of linear correlation equations, which represent physical behavior for different conditions and configurations. These correlation equations are developed for three arrangements, which are associated with X/de = 54.54.

A Computational Fluid Dynamics-Based Small-Scale Combustor Design Optimization Utilizing a Many-Objective Evolutionary Algorithm

Abstract This work demonstrates the capability of an open-source autonomous computational fluid dynamics (CFD) metamodeling environment (OpenACME) for optimizing small-scale combustor designs. OpenACME couples several object-oriented programing open-source codes for CFD-assisted design using a decomposition-based many-objective evolutionary algorithm. The CFD is based on steady, incompressible, three-dimensional simulations with k–ω SST RANS and flamelet/progress variable combustion model. There are five unconstrained design variables based on combustor liner dilution hole diameters. The CFD results are compared with existing experimental data in terms of combustion efficiency as a function of severity parameter. The comparison demonstrates that the CFD methods capture combustion efficiency trends. Next, more than 500 combustor designs are evaluated with OpenACME. A Pareto Frontier is generated in terms of combustion efficiency, pattern factor, and total pressure losses. Pseudo-weights are used to select a nondominated Pareto Frontier design point for future fabrication and experimental testing. OpenACME is demonstrated to be a viable tool for small-scale combustor design optimization.

A multiobjective optimization of 3D - slot jet configuration for enhancement of film cooling in an annular combustor liner

This study reports the results of a combined experimental and numerical investigation of the performance of a three-dimensional film cooling slot configuration and its optimization using an evolutionary-based genetic algorithm. The primary objective of the optimization is to maximize the area-averaged film cooling effectiveness (ηad.avg) and minimize its standard deviation (ση) at a fixed coolant mass flux for a single liner. The experimental study is conducted on flow and heat transfer characteristics to validate the numerical model under laboratory conditions (low temperature and pressure). A design space is created using the Latin hypercube sampling technique, and it is solved using steady-state RANS simulations with the validated numerical model to estimate the ηad.avg and ση under actual engine conditions (high temperature and pressure). A surrogate model is developed using the kriging technique to predict the objective functions with the geometrical parameters of the slot. Following this, the optimum configuration is identified using the genetic algorithm. The geometrical parameters are slot jet diameter (d), slot jet pitch (p), lip taper angle (α), and lip length (L). The numerical results show that the optimum configuration outperforms the two reference configurations of the baseline combustor. At a blowing ratio (BR) of 1, the optimum configuration enhances the ηad.avg by 8% and reduces the ση by 6.6% compared to reference -1 configuration. Similarly, compared to reference -2, the optimum configuration enhances the ηad.avg by 19.8% and reduces the ση by 5.2.%. Furthermore, under actual engine conditions, a numerical study is conducted on a subsequent row of liners to determine the optimum length of the cooling ring. The numerical results show that the optimum slot configuration cools an additional liner length (G/S) of 5.5 compared to the reference-1 slot and results in a reduction of one pair of cooling rings in the combustor, which contributes to a 16.6% less coolant mass flux.

Design and combustion characteristics analysis of a static shaft turbofan engine

Abstract A design scheme of a static shaft turbofan engine is proposed to meet the requirements of light weight and large thrust weight ratio of small aeroengine. As the core component of stable combustion, the thermal protection problem of the mid-mounted combustion chamber is particularly prominent. This paper designs a mid-mounted combustion chamber configuration that combines gas film cooling and central combustion. The influence of structural parameters on combustion characteristics is explored by numerical simulation, and the theoretical design and numerical simulation is verified based on combustion test results. The results show that the flame shape of the mid-mounted combustion chamber conforms to the characteristics of central combustion. The combustion effect of the nozzle with spray angle of 45° and flow rate of 1.87 kg/h meets the requirements of secondary combustion of the static shaft turbofan engine, and the air inlet of the combustion liner effectively increases the thickness of the cooling gas film. The experimental results are in good agreement with the numerical simulation results, and the temperature of the combustion liner wall can be reduced effectively. The above research provides a theoretical basis for the combustion chamber design of small static shaft turbofan engines and a reference for the thermal protection methods of small aeroengine combustion chambers.

Applications of weighted-sum-of-gray-gas model with soot to an experimental aero-engine combustion chamber

Due to the importance of radiation during combustion in the aero-engine, the weighted-sum-of-gray-gas (WSGG) model considering both gas and soot is applied to the combustion modelling for an experimental aero-engine in this work. Five gray gases are chosen to resolve the absorption coefficients of CO2-H2O mixture, with which the Planck-mean absorption coefficients of soot are used for the spectral consideration of gas-soot mixture. Together with the discrete ordinate method, the radiation effects using WSGG model on the combustion are presented. Comparisons are made between full-scale combustion simulations without radiation and those with radiation from gas only or gas-soot mixture. Results show that radiation from either gas or gas-soot mixture leads to no more than 3% differences in temperature and 5% differences in species concentrations. However, the temperature of combustor liner walls in the primary zone (i.e., the main combustion region) is significantly altered with an average of 100 K rise when considering radiation from gas only and a maximum of 500 K rise when considering radiation from gas-soot mixture. Results also show that the temperature of combustor liner walls with radiation from gas-soot mixture becomes more well-distributed than that either without radiation or with radiation from gas only.