Статті в журналах: "Aero-engine combustors"

1

Şöhret, Yasin, and T. Hikmet Karakoc. "Exergy indicators of a low-emission aero-engine combustor." Aircraft Engineering and Aerospace Technology 90, no. 2 (March 5, 2018): 344–50. http://dx.doi.org/10.1108/aeat-03-2016-0045.

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Purpose It is essential to develop more environment-friendly energy systems to prevent climate change and minimize environmental impact. Within this scope, many studies are performed on performance and environmental assessments of many types of energy systems. This paper, different from previous studies, aims to prove exergy performance of a low-emission combustor of an aero-engine. Design/methodology/approach It is a well-known fact that, with respect to previous exergy analysis, highest exergy destruction occurs in the combustor component of the engine. For this reason, it is required to evaluate a low-emission aero-engine combustor thermodynamically to understand the state of the art according to the authors’ best of knowledge. In this framework, combustor has been operated at numerous conditions (variable engine load) and evaluated. Findings As a conclusion of the study, the impact of emission reduction on performance improvement of the aero-engine combustors exergetically is presented. It is stated that exergy efficiency of the low-emission aero-engine combustor is found to be 64.69, 61.95 and 71.97 per cent under various operating conditions. Practical implications Results obtained in this paper may be beneficial for researchers who are interested in combustion and propulsion technology and thermal sciences. Originality/value Different from former studies, the impact of operating conditions on performance of a combustor is examined from the viewpoint of thermodynamics.

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Li, J., X. Sun, Y. Liu, and V. Sethi. "Preliminary aerodynamic design methodology for aero engine lean direct injection combustors." Aeronautical Journal 121, no. 1242 (June 21, 2017): 1087–108. http://dx.doi.org/10.1017/aer.2017.47.

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ABSTRACTThe Lean Direct Injection (LDI) combustor is one of the low-emissions combustors with great potential in aero-engine applications, especially those with high overall pressure ratio. A preliminary design tool providing basic combustor sizing information and qualitative assessment of performance and emission characteristics of the LDI combustor within a short period of time will be of great value to designers. In this research, the methodology of preliminary aerodynamic design for a second-generation LDI (LDI-2) combustor was explored. A computer code was developed based on this method covering the design of air distribution, combustor sizing, diffuser, dilution holes and swirlers. The NASA correlations for NOxemissions are also embedded in the program in order to estimate the NOx production of the designed LDI combustor. A case study was carried out through the design of an LDI-2 combustor named as CULDI2015 and the comparison with an existing rich-burn, quick-quench, lean-burn combustor operating at identical conditions. It is discovered that the LDI combustor could potentially achieve a reduction in liner length and NOxemissions by 18% and 67%, respectively. A sensitivity study on parameters such as equivalence ratio, dome and passage velocity and fuel staging is performed to investigate the effect of design uncertainties on both preliminary design results and NOxproduction. A summary on the variation of design parameters and their impact is presented. The developed tool is proved to be valuable to preliminarily evaluate the LDI combustor performance and NOxemission at the early design stage.

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3

Marudhappan, Raja, Chandrasekhar Udayagiri, and Koni Hemachandra Reddy. "Combustion chamber design and reaction modeling for aero turbo-shaft engine." Aircraft Engineering and Aerospace Technology 91, no. 1 (January 7, 2018): 94–111. http://dx.doi.org/10.1108/aeat-10-2017-0217.

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Purpose The purpose of this paper is to formulate a structured approach to design an annular diffusion flame combustion chamber for use in the development of a 1,400 kW range aero turbo shaft engine. The purpose is extended to perform numerical combustion modeling by solving transient Favre Averaged Navier Stokes equations using realizable two equation k-e turbulence model and Discrete Ordinate radiation model. The presumed shape β-Probability Density Function (β-PDF) is used for turbulence chemistry interaction. The experiments are conducted on the real engine to validate the combustion chamber performance. Design/methodology/approach The combustor geometry is designed using the reference area method and semi-empirical correlations. The three dimensional combustor model is made using a commercial software. The numerical modeling of the combustion process is performed by following Eulerian approach. The functional testing of combustor was conducted to evaluate the performance. Findings The results obtained by the numerical modeling provide a detailed understanding of the combustor internal flow dynamics. The transient flame structures and streamline plots are presented. The velocity profiles obtained at different locations along the combustor by numerical modeling mostly go in-line with the previously published research works. The combustor exit temperature obtained by numerical modeling and experiment are found to be within the acceptable limit. These results form the basis of understanding the design procedure and opens-up avenues for further developments. Research limitations/implications Internal flow and combustion dynamics obtained from numerical simulation are not experimented owing to non-availability of adequate research facilities. Practical implications This study contributes toward the understanding of basic procedures and firsthand experience in the design aspects of combustors for aero-engine applications. This work also highlights one of the efficient, faster and economical aero gas turbine annular diffusion flame combustion chamber design and development. Originality/value The main novelty in this work is the incorporation of scoops in the dilution zone of the numerical model of combustion chamber to augment the effectiveness of cooling of combustion products to obtain the desired combustor exit temperature. The use of polyhedral cells for computational domain discretization in combustion modeling for aero engine application helps in achieving faster convergence and reliable predictions. The methodology and procedures presented in this work provide a basic understanding of the design aspects to the beginners working in the gas turbine combustors particularly meant for turbo shaft engines applications.

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4

ZIEMANN, J. "Low-NOx combustors for hydrogen fueled aero engine." International Journal of Hydrogen Energy 23, no. 4 (April 1998): 281–88. http://dx.doi.org/10.1016/s0360-3199(97)00054-2.

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5

Bake, Friedrich, Ulf Michel, and Ingo Roehle. "Investigation of Entropy Noise in Aero-Engine Combustors." Journal of Engineering for Gas Turbines and Power 129, no. 2 (February 1, 2006): 370–76. http://dx.doi.org/10.1115/1.2364193.

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Strong evidence is presented that entropy noise is the major source of external noise in aero-engine combustion. Entropy noise is generated in the outlet nozzles of combustors. Low-frequency entropy noise, which was predicted earlier in theory and numerical simulations, was successfully detected in a generic aero-engine combustion chamber. It is shown that entropy noise dominates even in the case of thermo-acoustic resonances. In addition to this, a different noise generating mechanism was discovered that is presumably of even higher relevance to jet engines: There is strong evidence of broad band entropy noise at higher frequencies (1 to 3kHz in the reported tests). This unexpected effect can be explained by the interaction of small scale entropy perturbations (hot spots) with the strong pressure gradient in the outlet nozzle. The direct combustion noise of the flame zone seems to be of minor importance for the noise emission to the ambiance. The combustion experiments were supplemented by experiments with electrical heating. Two different methods for generating entropy waves were used, a pulse excitation and a sinusoidal excitation. In addition, high-frequency entropy noise was generated by steady electrical heating.

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6

Klose, G., R. Schmehl, R. Meier, G. Maier, R. Koch, S. Wittig, M. Hettel, W. Leuckel, and N. Zarzalis. "Evaluation of Advanced Two-Phase Flow and Combustion Models for Predicting Low Emission Combustors." Journal of Engineering for Gas Turbines and Power 123, no. 4 (October 1, 2000): 817–23. http://dx.doi.org/10.1115/1.1377010.

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The development of low-emission aero-engine combustors strongly depends on the availability of accurate and efficient numerical models. The prediction of the interaction between two-phase flow and chemical combustion is one of the major objectives of the simulation of combustor flows. In this paper, predictions of a swirl stabilized model combustor are compared to experimental data. The computational method is based on an Eulerian two-phase model in conjunction with an eddy dissipation (ED) and a presumed-shape-PDF (JPDF) combustion model. The combination of an Eulerian two-phase model with a JPDF combustion model is a novelty. It was found to give good agreement to the experimental data.

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7

Zhu, M., A. P. Dowling, and K. N. C. Bray. "Self-Excited Oscillations in Combustors With Spray Atomizers." Journal of Engineering for Gas Turbines and Power 123, no. 4 (October 1, 2000): 779–86. http://dx.doi.org/10.1115/1.1376717.

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Combustors with fuel-spray atomizers are susceptible to a low-frequency oscillation, particularly at idle and sub-idle conditions. For aero-engine combustors, the frequency of this oscillation is typically in the range 50–120 Hz and is commonly called “rumble.” In the current work, computational fluid dynamics (CFD) is used to simulate this self-excited oscillation. The combustion model uses Monte Carlo techniques to give simultaneous solutions of the Williams’ spray equation together with the equations of turbulent reactive flow. The unsteady combustion is calculated by the laminar flamelet presumed pdf method. A quasi-steady description of fuel atomizer behavior is used to couple the inlet flow in the combustor. A choking condition is employed at turbine inlet. The effects of the atomizer and the combustor geometry on the unsteady combustion are studied. The results show that, for some atomizers, with a strong dependence of mean droplet size on air velocity, the coupled system undergoes low-frequency oscillations. The numerical results are analyzed to provide insight into the rumble phenomena. Basically, pressure variations in the combustor alter the inlet air and fuel spray characteristics, thereby changing the rate of combustion. This in turn leads to local “hot spots,” which generate pressure fluctuations as they convect through the downstream nozzle.

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8

Corsini, A., F. Rispoli, and T. E. Tezduyar. "Stabilized finite element computation of NOx emission in aero-engine combustors." International Journal for Numerical Methods in Fluids 65, no. 1-3 (October 29, 2010): 254–70. http://dx.doi.org/10.1002/fld.2451.

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9

Tietz, S., and T. Behrendt. "Development and application of a pre-design tool for aero-engine combustors." CEAS Aeronautical Journal 2, no. 1-4 (September 13, 2011): 111–23. http://dx.doi.org/10.1007/s13272-011-0012-x.

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10

Hu, Bin, Yong Huang, Fang Wang, and Fa Xie. "CFD predictions of LBO limits for aero-engine combustors using fuel iterative approximation." Chinese Journal of Aeronautics 26, no. 1 (February 2013): 74–84. http://dx.doi.org/10.1016/j.cja.2012.12.014.

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11

TACHIBANA, Shigeru. "Combustion stability diagnostics by use of visualization in the development of aero-engine combustors." Journal of the Visualization Society of Japan 35, no. 138 (2015): 20–25. http://dx.doi.org/10.3154/jvs.35.138_20.

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12

Hu, Bin, Yong Huang, and Fang Wang. "FIA method for LBO limit predictions of aero-engine combustors based on FV model." Aerospace Science and Technology 28, no. 1 (July 2013): 435–46. http://dx.doi.org/10.1016/j.ast.2013.01.002.

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13

Agbadede, Roupa, and Biweri Kainga. "Effect of Water Injection into Aero-derivative Gas Turbine Combustors on NOx Reduction." European Journal of Engineering Research and Science 5, no. 11 (November 21, 2020): 1357–59. http://dx.doi.org/10.24018/ejers.2020.5.11.2180.

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Oxides of Nitrogen (NOx) generated from gas turbines causes enormous harm to human health and the environment. As a result, different methods have been employed to reduce NOx produced from gas turbine combustion process. One of such technique is the injection of water or steam into the combustion chamber to reduce the flame temperature. A twin shaft aero-derivative gas turbine was modelled and simulated using GASTURB simulation software. The engine was modelled after the GE LM2500 class of gas turbine engines. Water injection into the gas turbine combustor was simulated by implanting water-to-fuel ratios of 0 to 0.8, in an increasing order of 0.2. The results show that when water-to-fuel ratio was increased, the Nox severity index reduced. While heat rate and fuel flow increased with water-to-fuel ratio (injection flow rate).

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14

Agbadede, Roupa, and Isaiah Allison. "Effect of Water Injection into Aero-derivative Gas Turbine Combustors on NOx Reduction." European Journal of Engineering and Technology Research 5, no. 11 (November 21, 2020): 1357–59. http://dx.doi.org/10.24018/ejeng.2020.5.11.2180.

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Oxides of Nitrogen (NOx) generated from gas turbines causes enormous harm to human health and the environment. As a result, different methods have been employed to reduce NOx produced from gas turbine combustion process. One of such technique is the injection of water or steam into the combustion chamber to reduce the flame temperature. A twin shaft aero-derivative gas turbine was modelled and simulated using GASTURB simulation software. The engine was modelled after the GE LM2500 class of gas turbine engines. Water injection into the gas turbine combustor was simulated by implanting water-to-fuel ratios of 0 to 0.8, in an increasing order of 0.2. The results show that when water-to-fuel ratio was increased, the Nox severity index reduced. While heat rate and fuel flow increased with water-to-fuel ratio (injection flow rate).

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15

SUN, Lei, Yong HUANG, Xiwei WANG, Zekun ZHENG, Ruixiang WANG, and Xiang FENG. "Hybrid method based on flame volume concept for lean blowout limits prediction of aero engine combustors." Chinese Journal of Aeronautics 34, no. 5 (May 2021): 425–37. http://dx.doi.org/10.1016/j.cja.2020.12.033.

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Andreini, Antonio, Bruno Facchini, Andrea Giusti, and Fabio Turrini. "Assessment of Flame Transfer Function Formulations for the Thermoacoustic Analysis of Lean Burn Aero-engine Combustors." Energy Procedia 45 (2014): 1422–31. http://dx.doi.org/10.1016/j.egypro.2014.01.149.

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Pinelli, Lorenzo, Leonardo Lilli, Andrea Arnone, Paolo Gaetani, and Giacomo Persico. "Numerical Study of Entropy Wave Evolution within a HPT Stage." E3S Web of Conferences 197 (2020): 11011. http://dx.doi.org/10.1051/e3sconf/202019711011.

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Component reciprocal interaction and aero-thermal coupling are critical aspects in modern turbomachinery design. Combustors and highpressure turbine (HPT) interaction is extremely critical due to the compact and lightweight system design. In this context, computational and experimental analyses are thus necessary to study the interaction of the high temperature gas coming from combustor systems and entering the turbine in order to avoid engine mis-operations and to lower the indirect core noise generation. This paper presents a numerical study of pulsating temperature distortion (entropy wave) evolution within a high pressure turbine stage. Four different clocking positions between the 11 temperature spots and the 22 stators have been studied. The numerical results, obtained by URANS computations (TRAF code) and by a dedicated post-processing based on Fourier coefficients, have been compared with experimental measurements coming from the Laboratorio di Fluidodinamica delle Macchine (LFM) of the Politecnico di Milano (Italy) where the HP stage rig is located. The excellent agreement between numerical results and experimental acquisitions confirms the accuracy of the numerical approach. Such results also suggest recommendations for the thermal design of the rows and are the main prerequisite for the study of the indirect core noise generation.

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18

Zhao, Ziqiang, Xiaomin He, Guoyu Ding, Mingyu Li, Ping Jiang, and Weidong Huanga. "Effect of rotational direction of triple-swirler on cold flow characteristics of a model combustor." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 231, no. 5 (April 25, 2016): 918–30. http://dx.doi.org/10.1177/0954410016645126.

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Triple-swirler plays an important role for aero-engine combustors to achieve high temperature rise. In this paper, experimental investigations were carried out to explore the effect of triple-swirler rotational direction on swirling flow field in atmospheric condition. Two-dimensional-planar particle image velocimetry measurements show that the central toroidal recirculation zone (CTRZ) formation is significantly affected by the swirler rotational direction combinations: an obvious CTRZ can be formed by the triple-swirler with co-rotating intermediate swirler and outer swirler, while a much smaller CTRZ was obtained by the triple-swirler with a counter-rotating intermediate and outer swirler. Furthermore, the swirl level of the mixed flow is significantly affected by the rotational direction combination, and the integrated swirl numbers were calculated to help evaluating the swirl level generated by triple-swirlers. The rotational direction combination plays a key role on the tangential velocity distribution. The tangential velocity distribution is not only closely related to rotational direction, but also the swirl number combination and mass proportion of each swirler in a triple-swirler.

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Innocenti, A., A. Andreini, D. Bertini, B. Facchini, and M. Motta. "Turbulent flow-field effects in a hybrid CFD-CRN model for the prediction of NO and CO emissions in aero-engine combustors." Fuel 215 (March 2018): 853–64. http://dx.doi.org/10.1016/j.fuel.2017.11.097.

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20

Liu, Jian, Dingrui Zhang, Lingyun Hou, Jinhu Yang, and Gang Xu. "Laminar Burning Speed of Aviation Kerosene at Low Pressures." Energies 15, no. 6 (March 17, 2022): 2191. http://dx.doi.org/10.3390/en15062191.

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Aero-engine combustors may experience extreme low pressures in the case of an in-flight shutdown, which makes the study of aviation kerosene flame propagation characteristics at low pressures important. The present work examined flame propagation during the combustion of aviation kerosene over the pressure range from 25 to 100 kPa using a constant-volume bomb apparatus. The laminar burning speeds at different initial pressures, temperatures and equivalence ratios were measured and compared. In addition, numerical simulations were used to examine the reaction sensitivity of the laminar burning speed at low pressure. In trials at the lean flammability limit, the data indicated that it was more difficult to ignite the fuel under a lower pressure condition of 25 kPa and a lower temperature condition of 420 K. The experimental results of laminar burning speed were fitted to an equation providing the laminar burning speeds expected at different pressures (25–100 kPa), temperatures (400–480 K) and equivalence ratios (0.8–1.5). The temperature index (α=1.76) and pressure index (β=−0.15) of the fitting equation were obtained. Both hydrodynamic and diffusional thermal flame instabilities were found to be suppressed at low pressures. The negative effects of two specific reactions on laminar burning speed were greatly reduced at these same low pressures of 25 kPa.

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21

Kirk, D. R., G. R. Guenette, S. P. Lukachko, and I. A. Waitz. "Gas Turbine Engine Durability Impacts of High Fuel-Air Ratio Combustors—Part II: Near-Wall Reaction Effects on Film-Cooled Heat Transfer." Journal of Engineering for Gas Turbines and Power 125, no. 3 (July 1, 2003): 751–59. http://dx.doi.org/10.1115/1.1606473.

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As commercial and military aircraft engines approach higher total temperatures and increasing overall fuel-to-air ratios, the potential for significant chemical reactions on a film-cooled surface is enhanced. Currently, there is little basis for understanding the effects on aero-performance and durability due to such secondary reactions. A shock tube experiment was employed to generate short duration, high temperature (1000–2800 K) and pressure (6 atm) flows over a film-cooled flat plate. The test plate contained two sets of 35 deg film cooling holes that could be supplied with different gases, one side using air and the other nitrogen. A mixture of ethylene and argon provided a fuel rich freestream that reacted with the air film resulting in near wall reactions. The relative increase in surface heat flux due to near wall reactions was investigated over a range of fuel levels, momentum blowing ratios (0.5–2.0), and Damko¨hler numbers (ratio of flow to chemical time scales) from near zero to 30. For high Damko¨hler numbers, reactions had sufficient time to occur and increased the surface heat flux by 30 percent over the inert cooling side. When these results are appropriately scaled, it is shown that in some situations of interest for gas turbine engine environments significant increases in surface heat flux can be produced due to chemical reactions in the film-cooling layer. It is also shown that the non-dimensional parameters Damko¨hler number (Da), blowing ratio (B), heat release potential (H*), and scaled heat flux Qs are the appropriate quantities to predict the augmentation in surface heat flux that arises due to secondary reactions.

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Mazzei, Lorenzo, Antonio Andreini, and Bruno Facchini. "Assessment of modelling strategies for film cooling." International Journal of Numerical Methods for Heat & Fluid Flow 27, no. 5 (May 2, 2017): 1118–27. http://dx.doi.org/10.1108/hff-03-2016-0086.

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Purpose Effusion cooling represents one the most innovative techniques for the thermal management of aero-engine combustors liners. The huge amount of micro-perforations implies a significant computational cost if cooling holes are included in computational fluid dynamics (CFD) simulations; therefore, many efforts are reported in literature to develop lower-order approaches aiming at limiting the number of mesh elements. This paper aims to report a numerical investigation for validating two approaches for modelling film cooling, distinguished according to the way coolant is injected (i.e. through either point or distributed mass sources). Design/methodology/approach The approaches are validated against experimental data in terms of adiabatic effectiveness and heat transfer coefficient distributions obtained for effusion cooled flat plates. Additional reynolds-averaged naver stokes (RANS) simulations were performed meshing also the perforation, so as to distinguish the contribution of injection modelling with respect to intrinsic limitations of turbulence model modelling. Findings Despite the simplified strategies for coolant injection, this work clearly shows the feasibility of obtaining a sufficiently accurate reproduction of coolant protection in conjunction with a significant saving in terms of computational cost. Practical/implications The proposed methodologies allow to take into account the presence of film cooling in simulations of devices characterized by a huge number of holes. Originality/value This activity represents the first thorough and quantitative comparison between two approaches for film cooling modelling, highlighting the advantages involved in their application.

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Li, Longbiao, Suyi Bi, and Youchao Sun. "Risk assessment method for aeroengine multiple failure risk using Monte Carlo simulation." Multidiscipline Modeling in Materials and Structures 12, no. 2 (August 8, 2016): 384–96. http://dx.doi.org/10.1108/mmms-06-2015-0028.

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Purpose – The purpose of this paper is to develop a method to predict the multi-failure risk of aero engine in service and to evaluate the effectiveness of different corrective actions. Design/methodology/approach – The classification of failure risk level, the determination of hazard ratio and the calculation of risk factor and the risk per flight have been proposed. The multi-failure risk assessment process of aero engine has been established to predict the occurrence of failure event and assess the failure risk level. According to the history aero engine failure data, the multi-failure risk, i.e., overheat, blade wounding, pump failure, blade crack, pipe crack and combustor crack, has been predicted considering with and without corrective action. Two corrective actions, i.e., reduce the maintenance interval and redesign the failure components, were adopted to analyze the decreasing of risk level. Findings – The multi-failure risk of aero engine with or without corrective action can be determined using the present method. The risk level of combustor crack decreases from high-risk level of 1.18×1e−9 without corrective action to acceptable risk level of 0.954×1e−9 by decreasing the maintenance interval from 1,000 to 800 h, or to 0.912×1e−9 using the redesign combustor. Research limitations/implications – It should be noted that probability of detection during maintenance actions has not been considered in the present analysis, which would affect the failure risk level of aero engine in service. Social implications – The method in the present analysis can be adapted to other types of failure modes which may cause significant safety or environment hazards, and used to determine the maintenance interval or choose appropriate corrective action to reduce the multi-failure risk level of aero engine. Originality/value – The maintenance interval or appropriate corrective action can be determined using the present method to reduce the multi-failure risk level of aero engine.

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Zhang, Junhong, Huwei Dai, Jiewei Lin, Yi Yuan, Zhiyuan Liu, Yubo Sun, and Kunying Ding. "Cracking analysis of an aero-engine combustor." Engineering Failure Analysis 115 (September 2020): 104640. http://dx.doi.org/10.1016/j.engfailanal.2020.104640.

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Zhang, Jin Chuan, Yun Wang, and Can Zhang. "Basic Research of the Intercooled Turbofan Aero-Engine." Advanced Materials Research 516-517 (May 2012): 544–47. http://dx.doi.org/10.4028/www.scientific.net/amr.516-517.544.

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In conventional turbofan aero-engine designs, the effective way of improvement of engine efficiency is through the increasing of overall pressure ratio and improving of combustor inlet gas temperature, but the further incresement of compressor overall pressor ratio is constricted by high pressure compressor outlet allowed temperature. The improvement of combustor outlet temperature is limited by turbine allowed inlet temperature during take-off and climbing. An intercooled core can be designed with a significantly higher overall pressure ratio also with reduced cooling air requirements, providing a higher thermal efficiency compared with a conventional core. Through the basic analysis of performance of intercooler aeroengines. It indicated that the intercooled aero-engines can decrese the feul consume clearly and have a further potential in future civil aircraft application.

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Xiao, Yinli, Zhibo Cao, and Changwu Wang. "The Effect of Dilution Air Jets on Aero-Engine Combustor Performance." International Journal of Turbo & Jet-Engines 36, no. 3 (August 27, 2019): 257–69. http://dx.doi.org/10.1515/tjj-2018-0045.

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Abstract In aero-engine combustor, the primary and secondary dilution air jets play a vital role to achieve efficient combustion and provide a satisfactory temperature profile at the combustor exit. In the current research, seven kinds of three-dome combustor sector associated with different dilution holes arrangements are simulated using Flamelet model. The influences of location, number and diameter of dilution air holes on flow field and performance of combustor are analysed in detail. The results demonstrate that the variation of airflow distribution, while keeping the total admission area constant, impact the combustion efficiency and pattern profile considerably, yet has insignificant effect on the total pressure coefficient. The maximum combustion efficiency and minimum pattern factor can be achieved simultaneously with deliberate dilution jet holes arrangement.

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Lv, Fengjun, Quan Li, and Guoru Fu. "Failure analysis of an aero-engine combustor liner." Engineering Failure Analysis 17, no. 5 (July 2010): 1094–101. http://dx.doi.org/10.1016/j.engfailanal.2010.01.003.

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Wang, Xiting, Ai He, and Zhongzhi Hu. "Transient Modeling and Performance Analysis of Hydrogen-Fueled Aero Engines." Processes 11, no. 2 (January 31, 2023): 423. http://dx.doi.org/10.3390/pr11020423.

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With the combustor burning hydrogen, as well as the strongly coupled fuel and cooling system, the configuration of a hydrogen-fueled aero engine is more complex than that of a conventional aero engine. The performance, and especially the dynamic behavior of a hydrogen-fueled aero engine, need to be fully understood for engine system design and optimization. In this paper, both the transient modeling and performance analysis of hydrogen-fueled engines are presented. Firstly, the models specific to the hydrogen-fueled engine components and systems, including the hydrogen-fueled combustor, the steam injection system, a simplified model for a quick NOx emission assessment, and the heat exchangers, are developed and then integrated to a conventional engine models. The simulations with both Simulink and Speedgoat-based hardware in the loop system are carried out. Secondly, the performance analysis is performed for a typical turbofan engine configuration, CF6, and for the two hydrogen-fueled engine configurations, ENABLEH2 and HySIITE, which are currently under research and development by the European Union and Pratt & Whitney, respectively. At last, the simulation results demonstrate that the developed transient models can effectively reflect the characteristics of hydrogen burning, heat exchanging, and NOx emission for hydrogen-fueled engines. In most cases, the hydrogen-fueled engines show lower specific fuel consumption, lower turbine entry temperature, and less NOx emissions compared with conventional engines. For example, at max thrust state, the advanced hydrogen-fueled engine can reduce the parameters mentioned above by about 68.5%, 3.7%, and 12.7%, respectively (a mean value of two configurations).

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Zhu, Lian Jun, Yu Cai Dong, Jian Guang Yuan, Liang Hai Yi, and Ge Hua Fan. "Prediction of the Total Sound Level of the Annular Combustor Noise of Aircraft Engine Based on SVR." Applied Mechanics and Materials 687-691 (November 2014): 33–36. http://dx.doi.org/10.4028/www.scientific.net/amm.687-691.33.

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Because of the high temperature and high pressure environment inside the annular combustor of an aircraft engine, the direct measurement of the burner noise is very difficult. This paper set up the model of the total sound level and the effect factor SVR though analyzing the relationship between the total sound level and noise parameters of the combustion chamber the annular combustor an aircraft engine and the influence factors,, and it is better than multiple regression mode and the projection pursuit regression model, and predict the predicting samples so it is important for aero-engine design and reliability analysis.

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30

Chen, Yi, Li Fei, Liming He, Lei Zhang, Chunchang Zhu, and Jun Deng. "The Influence of Dielectric Barrier Discharge Plasma on the Characteristics of Aero-Engine Combustion Chamber." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 37, no. 2 (April 2019): 369–77. http://dx.doi.org/10.1051/jnwpu/20193720369.

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A test platform was developed to investigate the performance of aero-engine combustor by the dielectric barrier discharge (DBD) plasma assisted combustion (PAC) in the simulated maximum condition. Conventional combustion experiments and plasma-assisted combustion conditions were conducted to study the effect of PAC on the performances including average outlet temperature, combustion efficiency and pattern factor under four different excessive air coefficients five different voltages. The comparative experiment shows that the combustion efficiency is improved after PAC compared with the normal conditions, the combustion efficiency of PAC increases 2.31% in the fuel-rich condition when Up-p is 40 kV. The uniformity of the outlet temperature field is also improved after PAC, the decrease of the pattern factor is more than 5% in the fuel-rich condition. These results offer certain reference value for the future application of PAC in aero-engine combustor and improving its performance.

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31

Wang, Tianyu, Jinlu Yu, Bingbing Zhao, Weida Cheng, Lei Zhang, and Yongkun Sun. "Study on plasma combustion process in aero engine combustor." Journal of Physics: Conference Series 2228, no. 1 (March 1, 2022): 012034. http://dx.doi.org/10.1088/1742-6596/2228/1/012034.

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Abstract In order to explore the effect of plasma combustion on the performance of aeroengine combustion chamber, a simplified 9-Step plasma reaction mechanism was added to the 16 component 25 step reaction mechanism of aviation kerosene, and a kerosene combustion chemical reaction process model considering the excited state relaxation reaction of plasma electron collision reaction and the chemical reaction involved in excited state was established, The numerical calculation of combustion process in combustion chamber was carried out, and the numerical calculation results of combustion process with and without plasma combustion support were compared and analyzed. The results showed that after adding plasma in the combustion chamber, the active particles produced by plasma reaction made kerosene burn more fully in the main combustion zone. Under the appropriate chemical ratio, the average temperature of the outlet section of the combustion chamber increases from 2208.5K to 2543.5K, and the combustion efficiency increases by 11% to 95.6% on the basis of 84.6%; The outlet temperature field was more uniform, and the temperature distribution coefficient (OTDF) was reduced from 0.131 to 0.111, which improved the performance of the combustion chamber.

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32

Antoshkiv, O., Th Poojitganont, L. Jehring, and C. Berkholz. "Main aspects of kerosene and gaseous fuel ignition in aero-engine." Aeronautical Journal 121, no. 1246 (December 2017): 1779–94. http://dx.doi.org/10.1017/aer.2017.113.

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ABSTRACTVarious liquid and gaseous alternative fuels have been proposed to replace the kerosene as aircraft fuel. Furthermore, new combustion technologies were developed to reduce the emissions of aero-engine. A staged fuel injection arrangement for a lean burn combustion system was applied to improve the operability of an aero-engine by achieving high flame stability at reduced combustion emissions. Originally, both circuits (pilot and main) are fuelled by kerosene; moreover, the pilot injector is operating at low power (engine idle and approach) and the pilot flame is anchored in an airflow recirculation zone. In the case of the performed research, the pilot injector was modified to allow the use of gaseous fuels. Thus, the burner model allows a flexible balancing of the mass flows for gaseous and liquid fuel. The present paper describes the investigation of ignitability for the proposed staged combustor model fuelled by gaseous and liquid fuels. A short overview on physical properties of used fuels is given. To investigate atomisation and ignition, different measurements systems were used. The effectiveness of two ignitor types (spark plug and laser ignitor) was analysed. The ignition performance of the combustor operating on various fuels was compared and discussed in detail.

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33

Khandelwal, B., A. Karakurt, V. Sethi, R. Singh, and Z. Quan. "Preliminary design and performance analysis of a low emission aero-derived gas turbine combustor." Aeronautical Journal 117, no. 1198 (December 2013): 1249–71. http://dx.doi.org/10.1017/s0001924000008848.

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Abstract Modern gas turbine combustor design is a complex task which includes both experimental and empirical knowledge. Numerous parameters have to be considered for combustor designs which include combustor size, combustion efficiency, emissions and so on. Several empirical correlations and experienced approaches have been developed and summarised in literature for designing conventional combustors. A large number of advanced technologies have been successfully employed to reduce emissions significantly in the last few decades. There is no literature in the public domain for providing detailed design methodologies of triple annular combustors. The objective of this study is to provide a detailed method designing a triple annular dry low emission industrial combustor and evaluate its performance, based on the operating conditions of an industrial engine. The design methodology employs semi-empirical and empirical models for designing different components of gas turbine combustors. Meanwhile, advanced DLE methods such as lean fuel combustion, premixed methods, staged combustion, triple annular, multi-passage diffusers, machined cooling rings, DACRS and heat shields are employed to cut down emissions. The design process is shown step by step for design and performance evaluation of the combustor. The performance of this combustor is predicted, it shows that NO x emissions could be reduced by 60%-90% as compared with conventional single annular combustors.

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34

Xiao, Yinli, Changwu Wang, Zhibo Cao, and Wenyan Song. "Laser holography measurement and theoretical analysis of a pressure-swirl nozzle spray." Advances in Mechanical Engineering 10, no. 12 (December 2018): 168781401881325. http://dx.doi.org/10.1177/1687814018813253.

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The characteristic of the spray within combustion chamber is one of the determining factors that affect the performance and exhaust gas emissions of an aero-engine. Recently, the holography technique has been successfully applied to spray atomization measurement due to its significant advances. In this article, an atmospheric test rig of pressure-swirl nozzle is built. The kerosene spray generated at the atmospheric condition and in an aero-engine combustor is measured. The Sauter mean diameter of the spray droplets is obtained. In addition, the theoretical analysis of film formation and sheet breakup processes are conducted. Comparison of theoretical analysis and experimental results on the spray atomization of a pressure-swirl nozzle is presented.

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35

Mishra, R. K., and Sunil Chandel. "Soot Formation and Its Effect in an Aero Gas Turbine Combustor." International Journal of Turbo & Jet-Engines 36, no. 1 (March 26, 2019): 61–73. http://dx.doi.org/10.1515/tjj-2016-0062.

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Abstract Soot formation and the effect of soot deposit on the performance and integrity on an aero gas turbine combustor has been studied. Defective atomizer or blockage of air passages creates a fuel rich mixture which promotes soot formation in combustor primary zone. The temperature field and soot concentration inside the liner has been analyzed at high equivalence ratio using computational model in CFX. The peak temperature in primary zone increases till equivalence ratio reaches ϕ=1.1. But at high equivalence ratio, i. e., ϕ≥1.2, the peak temperature in primary zone decreases and that in dilution zone increases. Soot concentration increases at liner front end as well as in dilution zone when equivalence ratio increases from 1.25 to 3.0. Erosion and distortion of atomizer flow passages cause higher spray cone angle which again increases the soot concentration. Soot deposit inside liner has detrimental effect on the life and performance of the combustor as well as of the aero engine.

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36

Jakubowski, Robert. "Study of Bypass Ratio Increasing Possibility for Turbofan Engine and Turbofan with Inter Turbine Burner." Journal of KONES 26, no. 2 (June 1, 2019): 61–68. http://dx.doi.org/10.2478/kones-2019-0033.

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Abstract Current trends in the high bypass ratio turbofan engines development are discussed in the beginning of the paper. Based on this, the state of the art in the contemporary turbofan engines is presented and their change in the last decade is briefly summarized. The main scope of the work is the bypass ratio growth analysis. It is discussed for classical turbofan engine scheme. The next step is presentation of reach this goal by application of an additional combustor located between high and low pressure turbines. The numerical model for fast analysis of bypass ratio grows for both engine kinds are presented. Based on it, the numerical simulation of bypass engine increasing is studied. The assumption to carry out this study is a common core engine. For classical turbofan engine bypass ratio grow is compensated by fan pressure ratio reduction. For inter turbine burner turbofan, bypass grown is compensated by additional energy input into the additional combustor. Presented results are plotted and discussed. The main conclusion is drawing that energy input in to the turbofan aero engine should grow when bypass ratio is growing otherwise the energy should be saved by other engine elements (here fan pressure ratio is decreasing). Presented solution of additional energy input in inter turbine burner allow to eliminate this problem. In studied aspect, this solution not allows to improve engine performance. Specific thrust of such engine grows with bypass ratio rise – this is positive, but specific fuel consumption rise too. Classical turbofan reaches lower specific thrust for higher bypass ratio but its specific fuel consumption is lower too. Specific fuel consumption decreasing is one of the goal set for future aero-engines improvements.

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37

Li, L., X. F. Peng, and T. Liu. "Combustion and cooling performance in an aero-engine annular combustor." Applied Thermal Engineering 26, no. 16 (November 2006): 1771–79. http://dx.doi.org/10.1016/j.applthermaleng.2005.11.023.

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38

Zheng, Min, Fan Shen, and Pei Luo. "Vibration Fatigue Analysis of the Structure under Thermal Loading." Advanced Materials Research 853 (December 2013): 559–64. http://dx.doi.org/10.4028/www.scientific.net/amr.853.559.

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The fatigue problem of structures under concurrent thermal and vibration loading has not been thoroughly studied even though it is common in applications of aero-engine combustor liners. Here we attempt to explore such a problem using a simplified combustor liner model that is implemented by the commercial finite element software ANSYS Workbench. The modal parameters at various temperatures are calculated and the fatigue behavior under stochastic base excitation and thermal environment are analyzed. The results show that thermal loading not only has an effect on dynamic characteristics but also reduces the vibration fatigue life of the structure.

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39

Mishra, R. K., and Sunil Chandel. "Numerical Analysis of Exhaust Emission from an Aero Gas Turbine Combustor under Fuel-Rich Condition." International Journal of Turbo & Jet-Engines 36, no. 4 (November 18, 2019): 411–24. http://dx.doi.org/10.1515/tjj-2016-0079.

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Abstract Emission characteristic of an aero gas turbine combustor has been studied under fuel-rich condition. The combustor has been designed to operate at rich fuel-air mixture and may experience very high mixture up to equivalence ratio of 3.0 in the primary zone in the event of atomizer deterioration and flow passage distortion or during some maneuver. The concentration of major engine emission/pollutant constituents at combustor exhaust such as CO, NOx, soot, unburned hydrocarbons and atmospheric pollutants such as CO2 and water vapor has been analyzed at fuel-rich high equivalence ratio using computational model in CFX. The pollutant concentrations are found to be increased with increase in equivalence ratio. Atomization quality in terms of fuel droplet size is found to have a significant contribution in the concentration of various species at combustor exit.

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40

Ehsan, Kianpour, Nor Azwadi Che Sidik, and Mohsen Agha Seyyed Mirza Bozorg. "Thermodynamic Analysis of Flow Field at the End of Combustor Simulator." Applied Mechanics and Materials 225 (November 2012): 261–66. http://dx.doi.org/10.4028/www.scientific.net/amm.225.261.

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This study was carried out to investigate the effects of different cooling holes configurations on the thermal field characteristics inside a combustor simulator. In this research, a three-dimensional presentation of a true Pratt and Whitney aero-engine was simulated and analyzed. This combustor simulator combined the interaction of two rows of dilution jets, which were staggered in the stream wise direction and aligned in the span wise direction. The findings of the study indicate that the thickness of the film-cooling layer was thicker for the greater penetration depth. Furthermore, for the combustor simulator with more cooling holes, the temperature near the wall and between the jets was slightly increased. Also at the leading edge of the jet, the gradients of temperature were quite high at the jet-mainstream interface.

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41

Zeng, Wen, Hai-xia Li, Bao-dong Chen, and Hong-an Ma. "Kinetic Simulation of Combustion Process in the Combustor of the Aero-Engine." Combustion Science and Technology 186, no. 8 (June 26, 2014): 1097–114. http://dx.doi.org/10.1080/00102202.2014.902815.

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42

Huang, Shengfang, Zhibo Zhang, Huimin Song, Yun Wu, and Yinghong Li. "A Novel Way to Enhance the Spark Plasma-Assisted Ignition for an Aero-Engine Under Low Pressure." Applied Sciences 8, no. 9 (September 1, 2018): 1533. http://dx.doi.org/10.3390/app8091533.

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Finding a new ignition strategy for ignition enhancement in a lean-burn combustor has always been the biggest challenge for high-altitude, long-endurance unmanned aerial vehicles (UAVs). It is of great importance for the development of high-altitude, long-endurance aircraft to improve the secondary ignition ability of the aero-engine at high altitude where the ignition capability of the aero-engine igniter rapidly declines. An innovative ignition mode is therefore urgently needed. A novel plasma-assisted ignition method based on a multichannel discharge jet-enhanced spark (MDJS) was proposed in this study. Compared to the conventional spark igniter (SI), the arc discharge energy of the MDJS was increased by 13.6% at 0.12 bar and by 14.7% at 0.26 bar. Furthermore, the spark plasma penetration depth of the MDJS was increased by 49% and 103% at 0.12 bar and 0.26 bar, respectively. The CH* radicals showed that the MDJS obtained a larger initial spark kernel and reached a higher spark plasma penetration depth, which helped accelerate the burning velocity. Ignition tests in a model swirl combustor showed that the lean ignition limit was extended 24% from 0.034 to 0.026 at 25 m/s with 20 °C kerosene and 17% from 0.075 to 0.062 at 12 m/s with −30 °C kerosene maximally. The MDJS was a unique plasma-assisted ignition method, activated by the custom ignition power supply instead of a special power supply with an extra gas source. The objective of this study was to provide a novel multichannel discharge jet-enhanced spark ignition strategy which would help to increase the arc discharge energy, the spark plasma penetration depth and the activated area without changing the power supply system and to improve the safety and performance of aero-engines.

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43

Li, Le, Jianqin Suo, Han Yu, and Longxi Zhang. "Optimal Design and Application of Gas Analysis System." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 38, no. 1 (February 2020): 104–13. http://dx.doi.org/10.1051/jnwpu/20203810104.

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To accurately measure the performance parameters of an aero-engine combustor under high-temperature and high-pressure environment, such as outlet temperature, combustion efficiency and pollutant emission, the optimal design of the existing gas sampling and analysis system was carried out, the data post-processing method was corrected, and the full-component enthalpy conservation method suitable for calculating combustor outlet temperature was established. A lean direct injection and low-pollution combustor was used to evaluate the performance and application of its gas analysis system and its data post-processing method. The investigation results indicate that the optimized gas analysis system meets the International Civil Aviation Organization's pollutant emission measurement requirements and that it has a fast response and a high data quality. The comparison of the measurement results of the gas analysis system with those of the traditional measurement methods shows that the gas analysis system has a more accurate measurement, a wider condition range and a better stability, thus accurately evaluating the outlet temperature, combustion efficiency and emission characteristics of a gas turbine combustor.

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44

Zhao, Jiazi, Yasong Sun, Yifan Li, and Changhao Liu. "Investigation of coupled radiation-conduction heat transfer in cylindrical systems by discontinuous spectral element method." Journal of the Global Power and Propulsion Society 6 (December 30, 2022): 354–66. http://dx.doi.org/10.33737/jgpps/156350.

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Nowadays, in order to achieve higher efficiency in aero-engines, the increase of turbine inlet temperature in aero-engine is in urgent need. At present, the turbine inlet temperature is around 2,000 K, which means the radiation and coupled radiation-conduction heat transfer play more and more important roles in hot section of aero-engines. As we all konw, considering the cylindrical symmetry of aero-engines. It is convenient to adopt the cylindrical coordinate to simplify the description of these systems, such as annular combustor, exhaust nozzle, etc. In this paper, Discontinuous Spectral Element Method (DSEM) is extended to solve the radiation and coupled radiation-coduction heat transfer in cylindrical coordinate system. Both the spatial and angular computational domains of radiative transfer equation (RTE) are discretized and solved by DSEM. For coupled radiation-conduction heat transfer problem, Discontinuous Spectral Element Method-Spectral Element Method (DSEM-SEM) scheme is used to avoid using two sets of grid which would cause the increase of computational cost and the decrease of accuracy. Then, the effects of various geometric and thermal physical parameters are comprehensively investigated. Finally, these methods are further extended to 2D cylindrical system.

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45

Choi, Myeung Hwan, Dongsoo Shin, Youngbin Yoon, and Jaye Koo. "Swirl Number of Radial Swirler Design for Combustor in Aero Gas Turbine Engine." Journal of the Korean Society for Aeronautical & Space Sciences 47, no. 12 (December 31, 2019): 848–55. http://dx.doi.org/10.5139/jksas.2019.47.12.848.

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46

Dai, Huwei, Junhong Zhang, Yanyan Ren, Nuohao Liu, Bin Wu, Kunying Ding, and Jiewei Lin. "Failure mechanism of thermal barrier coatings of an ex-service aero-engine combustor." Surface and Coatings Technology 380 (December 2019): 125030. http://dx.doi.org/10.1016/j.surfcoat.2019.125030.

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47

Eberle, Christian, Peter Gerlinger, Klaus Peter Geigle, and Manfred Aigner. "Numerical Investigation of Transient Soot Evolution Processes in an Aero-Engine Model Combustor." Combustion Science and Technology 187, no. 12 (July 6, 2015): 1841–66. http://dx.doi.org/10.1080/00102202.2015.1065254.

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48

Kim, Daesik, Seungchai Jung, and Heeho Park. "Acoustic damping characterization of a double-perforated liner in an aero-engine combustor." Journal of Mechanical Science and Technology 33, no. 6 (June 2019): 2957–65. http://dx.doi.org/10.1007/s12206-019-0544-2.

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49

Chen, Y. D., Q. Feng, Y. R. Zheng, and X. F. Ding. "Formation of hole-edge cracks in a combustor liner of an aero engine." Engineering Failure Analysis 55 (September 2015): 148–56. http://dx.doi.org/10.1016/j.engfailanal.2015.05.018.

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50

Chen, Yi, Li-Ming He, Li Fei, Jun Deng, Jian-Ping Lei, and Han Yu. "Experimental study of dielectric barrier discharge plasma-assisted combustion in an aero-engine combustor." Aerospace Science and Technology 99 (April 2020): 105765. http://dx.doi.org/10.1016/j.ast.2020.105765.

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