Journal articles: 'Scramjet combustor'

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Xiong, Yuefei, Jiang Qin, Kunlin Cheng, Silong Zhang, and Yu Feng. "Quasi-One-Dimensional Model of Hydrocarbon-Fueled Scramjet Combustor Coupled with Regenerative Cooling." International Journal of Aerospace Engineering 2022 (August 8, 2022): 1–14. http://dx.doi.org/10.1155/2022/9931498.

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In order to rapidly predict the performance of hydrocarbon-fueled regeneratively cooled scramjet engine in system design, a quasi-one-dimensional model has been developed. The model consists of a supersonic combustor model with finite-rate chemistry and a cooling channel model with real gas working medium, which are governed by two sets of ordinary differential equations separately. Additional models for wall friction, heat transfer, sonic fuel injection, and mixing efficiency are also included. The two sets of ordinary differential equations are coupled and iteratively solved. The SUNDIALS code is used since the equations for supersonic combustion flow are stiff mathematically. The cooling channel model was verified by electric heating tube tests, and the supersonic combustor model was verified by experimental results for both hydrogen and hydrocarbon-fueled scramjet combustors. Three cases were comparatively studied: (1) scramjet combustor with an isothermal wall, (2) scramjet combustor with an adiabatic wall, and (3) scramjet combustor with regenerative cooling. Results showed that the model could predict the axial distributions of flow parameters in the supersonic combustor and cooling channel. Differences on ignition delay time and combustion efficiency for the three cases were observed.

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Li, Chaolong, Zhixun Xia, Likun Ma, Xiang Zhao, and Binbin Chen. "Numerical Study on the Solid Fuel Rocket Scramjet Combustor with Cavity." Energies 12, no. 7 (March 31, 2019): 1235. http://dx.doi.org/10.3390/en12071235.

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Scramjet based on solid propellant is a good supplement for the power device of future hypersonic vehicles. A new scramjet combustor configuration using solid fuel, namely, the solid fuel rocket scramjet (SFRSCRJ) combustor is proposed. The numerical study was conducted to simulate a flight environment of Mach 6 at a 25 km altitude. Three-dimensional Reynolds-averaged Navier–Stokes equations coupled with shear stress transport (SST) k − ω turbulence model are used to analyze the effects of the cavity and its position on the combustor. The feasibility of the SFRSCRJ combustor with cavity is demonstrated based on the validation of the numerical method. Results show that the scramjet combustor configuration with a backward-facing step can resist high pressure generated by the combustion in the supersonic combustor. The total combustion efficiency of the SFRSCRJ combustor mainly depends on the combustion of particles in the fuel-rich gas. A proper combustion organization can promote particle combustion and improve the total combustion efficiency. Among the four configurations considered, the combustion efficiency of the mid-cavity configuration is the highest, up to about 70%. Therefore, the cavity can effectively increase the combustion efficiency of the SFRSCRJ combustor.

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Athithan, A. Antony, S. Jeyakumar, Norbert Sczygiol, Mariusz Urbanski, and A. Hariharasudan. "The Combustion Characteristics of Double Ramps in a Strut-Based Scramjet Combustor." Energies 14, no. 4 (February 5, 2021): 831. http://dx.doi.org/10.3390/en14040831.

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This paper focuses on the influence of ramp locations upstream of a strut-based scramjet combustor under reacting flow conditions that are numerically investigated. In contrast, a computational study is adopted using Reynolds Averaged Navier Stokes (RANS) equations with the Shear Stress Transport (SST) k-ω turbulence model. The numerical results of the Deutsches Zentrum für Luft- und Raumfahrt or German Aerospace Centre (DLR) scramjet model are validated with the reported experimental values that show compliance within the range, indicating that the adopted simulation method can be extended for other investigations as well. The performance of the ramps in the strut-based scramjet combustor is analyzed based on parameters such as wall pressures, combustion efficiency and total pressure loss at various axial locations of the combustor. From the numerical shadowgraph, more shock interactions are observed upstream of the strut injection region for the ramp cases, which decelerates the flow downstream, and additional shock reflections with less intensity are also noticed when compared with the DLR scramjet model. The shock reflection due to the ramps enhances the hydrogen distribution in the spatial direction. The ignition delay is noticed for ramp combustors due to the deceleration of flow compared to the baseline strut only scramjet combustor. However, a higher flame temperature is observed with the ramp combustor. Because more shock interactions arise from the ramps, a marginal increase in the total pressure loss is observed for ramp combustors when compared to the baseline model.

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Ouyang, Hao, Weidong Liu, and Mingbo Sun. "Investigations on the Influence of the In-Stream Pylon and Strut on the Performance of a Scramjet Combustor." Scientific World Journal 2014 (2014): 1–10. http://dx.doi.org/10.1155/2014/309387.

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The influence of the in-stream pylon and strut on the performance of scramjet combustor was experimentally and numerically investigated. The experiments were conducted with a direct-connect supersonic model combustor equipped with multiple cavities. The entrance parameter of combustor corresponds to scramjet flight Mach number 4.0 with a total temperature of 947 K. The research results show that, compared with the scramjet combustor without pylon and strut, the wall pressure and the thrust of the scramjet increase due to the improvement of mixing and combustion effect due to the pylon and strut. The total pressure loss caused by the strut is considerable whereas pylon influence is slight.

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Yang, Pengnian, Zhixun Xia, Likun Ma, BinBin Chen, Yunchao Feng, Chaolong Li, and Libei Zhao. "Influence of the Multicavity Shape on the Solid Scramjet." International Journal of Aerospace Engineering 2021 (October 26, 2021): 1–14. http://dx.doi.org/10.1155/2021/9718537.

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In this paper, a modular solid scramjet combustor with multicavity was proposed. The influence of multicavity shape on the performance of solid scramjet was investigated by the direct-connected tests. The experiments simulated a flight Mach 5.5 at 25 km. The boron-based fuel-rich propellant was used. The microstructure of combustion products was analyzed by SEM. The experimental results show that the fuel-rich mixture produced by the gas generator would ignite rapidly in the solid scramjet combustor. The combustion process showed a typical characteristic of establishment-development-maintenance-attenuation. Compared to the flame-holding cavity, the other shapes of cavities, e.g., narrow and lobe, can improve mixing and combustion. In our experiment, the combustion efficiency increased from 0.41 to 0.48, and the total pressure recovery was 0.36. In summary, the proposed solid scramjet combustor can effectively solve the ignition delay problem of the fuel-rich mixture, and the narrow/lobe cavity shows the ability to improve the mixing and combustion of the fuel-rich mixture.

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Tahsini, AM. "Combustion efficiency and pressure loss balance for the supersonic combustor." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, no. 6 (December 18, 2019): 1149–56. http://dx.doi.org/10.1177/0954410019895885.

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The purpose of this paper is to investigate the effects of intake’s compression process of the scramjet on its flight performance. The hydrogen injection to the supersonic cross-flow is considered as the problem configuration. The finite volume solver is developed to simulate the compressible reacting turbulent flow using the proper reaction mechanism as the finite rate chemistry. The combustion efficiency and the drag force are the most important parameters on the scramjet flight performance, and finding the design point to balance the higher combustion efficiency and the lower minimum drag, which depends on the total pressure loss, can be used to optimize the supersonic combustors. The performance of the supersonic intake is considered here using some oblique shock waves with equal flow-deflection angles to compute the combustor’s inlet condition. The variation of combustion efficiency and total pressure loss is presented for different combustor’s inlet conditions. The results are presented for the constant jet to inlet pressure ratios and also for the constant equivalence ratios, in which the last one is much appropriate and utilized to find the optimum design point of the intake and the combustor, for assumed flight condition.

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Lee, Jae-Hyuk, Eun-Sung Lee, Hyung-Seok Han, Min-Su Kim, and Jeong-Yeol Choi. "A Study on a Vitiated Air Heater for a Direct-Connect Scramjet Combustor and Preliminary Test on the Scramjet Combustor Ignition." Aerospace 10, no. 5 (April 28, 2023): 415. http://dx.doi.org/10.3390/aerospace10050415.

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Vitiation air heater (VAH) combustion characteristics for a direct-connect scramjet combustor (DCSC) were experimentally studied. The VAH consists of a head, modular chamber, and circular-to-rectangular shape transition (CRST) nozzle. The CRST nozzle transforms the circular cross-sectioned rocket-type VAH into a rectangular cross-sectioned scramjet combustor. The CRST nozzle exit Mach numbers at the top, middle, and bottom were measured using a tungsten wedge. The oblique shock formed by the wedge was captured using Schlieren visualization and recorded with a high-speed camera. The θ-β-M relation showed that the exit Mach number was 2.04 ± 0.04 with a chamber pressure of 1.685 ± 0.07 MPa. With the VAH design point verified, preliminary scramjet combustor ignition tests were conducted. As the fuel was not auto-ignited by the vitiated air, the forced ignition method, in which VAH ignition flame ignites the scramjet fuel, was used. The Schlieren images showed that a cavity shear layer combustion mode was formed and also showed that the forced ignition method could be used as a reference model for the ignitor-ignition method.

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Chen, Hao, Mingming Guo, Ye Tian, Jialing Le, Hua Zhang, and Fuyu Zhong. "Intelligent reconstruction of the flow field in a supersonic combustor based on deep learning." Physics of Fluids 34, no. 3 (March 2022): 035128. http://dx.doi.org/10.1063/5.0087247.

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The data-driven intelligent reconstruction of a flow field in a supersonic combustor aids the real-time monitoring of wave system evolution in a scramjet flow field structure, allowing the determination of the combustion state for active flow control. In this paper, a deep learning architecture based on a multi-branch fusion convolutional neural network (MBFCNN) is proposed to reconstruct the flow field in a supersonic combustor. Experiments on hydrogen-fueled scramjets with different equivalence ratios were carried out in a direct-connected supersonic pulse combustion wind tunnel with an inflow Mach number of 2.5 to establish a dataset for MBFCNN network training and testing. The trained model successfully reconstructed the flow field structure from measured wall pressure data. The flow field reconstruction model provided a rich information source for the evolution of the wave system structure under the self-ignition conditions of the hydrogen-fueled scramjet, greatly improving the detection accuracy. The proposed deep learning architecture method was compared with basic convolutional neural network and symmetric convolutional neural network methods. The three methods all accurately reconstructed the flow field of the supersonic combustor. However, the proposed MBFCNN provided the best reconstruction results, and its average linear correlation coefficient in the test set was 0.952. The proposed MBFCNN had a lower mean square error and higher peak signal-to-noise ratio than the other two methods, which verified that the proposed model is eminently able to reconstruct and predict the flow field of a supersonic combustor.

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Yang, Pengnian, Zhixun Xia, Likun Ma, Binbin Chen, Yunchao Feng, Chaolong Li, and Libei Zhao. "Direct-Connect Test of Solid Scramjet with Symmetrical Structure." Energies 14, no. 17 (September 6, 2021): 5589. http://dx.doi.org/10.3390/en14175589.

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The solid scramjet has become one of the most promising engine types. In this paper, we report the first direct-connect test of a solid scramjet with symmetrical structure, carried out using boron-based fuel-rich solid propellant as fuel. During the test, which simulated a flight environment at Mach 5.6 and 25 km, the performance of the solid scramjet was obtained by measuring the pressure, thrust, and mass flow. The results show that, due to the change in the combustion area of the propellant and the deposition of the throat in the gas generator during the test, the equivalence ratio gradually increased from 0.54 to 0.63. In a solid scramjet, it is possible to obtain a symmetrical distribution of the flow field within the combustor. Moreover, in a multi-cavity combustor, the combustion state expands from the cavity to the center of the flow channel. The performance of the solid scramjet increased during the test, reaching a combustion efficiency of about 42%, a total pressure recovery coefficient of 0.35, and a thrust gain specific impulse of about 418 s. The solid scramjet with symmetrical structure is feasible. The cavity configuration adopted in this paper can reduce the ignition delay time of fuel-rich gas and improve the combustion efficiency of gas-phase combustible components. The shock trains in the isolator are conducive to the recovery of the total pressure. The performance of the solid scramjet is limited by the low combustion efficiency of the particles.

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Zhao, Zhelong, and Xianyu Wu. "Control Oriented Model for Expander Cycle Scramjet." MATEC Web of Conferences 257 (2019): 01004. http://dx.doi.org/10.1051/matecconf/201925701004.

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As a efficient and simple design, expander cycle is widely applied in LRE engineering, but it is seldomly used on scramjet research. In order to establish a complete mathematical model for expander cycle scramjet, a control-oriented model for expander cycle scramjet is proposed in this paper. This model consists of four major parts: combustor, cooling channel, turbo pump and nozzle and gives the result of pressure, temperature, mach number and velocity distribution of combustor and cooling channel and is capable of simulate both pure supersonic combustion mode and supersonic shock wave mode of the combustor. Each part is given by specific mathematical description, which contains the calculation of airflow, combustion, heat transfer and thermal cracking of kerosene. By putting all these parts together, a complete model is formed. This model is proposed to calculate the performance and condition of the engine precisely, comprehensively, swiftly and can be directly used in further study.

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Ji, Zifei, Huiqiang Zhang, and Bing Wang. "Thrust control strategy based on the minimum combustor inlet Mach number to enhance the overall performance of a scramjet engine." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 13 (February 20, 2019): 4810–24. http://dx.doi.org/10.1177/0954410019830816.

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A lower combustor inlet Mach number is desirable in order to design a compact, lightweight combustor and boost the overall performance of the scramjet engine. In this study, a thrust control strategy is proposed for a hydrogen-fueled scramjet taking into account the operating limitations, which is called the minimum combustor inlet Mach number rule since the combustor inlet Mach number is used as the control variable. By scheduling the fuel supply and modifying the intake geometry, the combustor inlet Mach number can be minimized while ensuring a certain thrust output within the operation constraints. In this manner, the scramjet engine can be operated with high specific thrust and low fuel consumption throughout the flight envelope. The thrust control strategy is further applied to a hydrogen-fueled scramjet in the hypersonic flight regime. Because the combustor inlet Mach number varies with flight conditions, the thrust strategy can be applied in practice by monitoring the following aerothermodynamic parameters in different flight regimes instead: (1) combustor outlet Mach number, (2) combustor inlet static temperature, and (3) combustor outlet static temperature. Furthermore, the effects of the thrust output on the division of flight regime are investigated, and the overall performance of the hydrogen-fueled scramjet engine obtained from applying the thrust control strategy is discussed in detail.

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Guimarães, Jefte Da Silva, Marco Antonio Sala Minucci, and Dermeval Carinhana Júnior. "ESTUDO DE UMA CÂMARA DE COMBUSTÃO SUPERSÔNICA USANDO UM TÚNEL DE CHOQUE." CIMATech 1, no. 7 (December 23, 2020): 126–36. http://dx.doi.org/10.37619/issn2447-5378.v7i1.297.126-136.

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Túneis de Choque são dispositivos capazes de gerar escoamentos de altas temperaturas e velocidades supersônicas, simulando as condições reais de escoamento na entrada de um combustor de motor Scramjet, viabilizando o estudo de combustão supersônica. Neste trabalho, são realizados ensaios aerotermodinâmicos e de combustão supersônica utilizando um Túnel de Choque, de modo a reproduzir as condições de escoamento na entrada de um combustor de motor Scramjet. O arranjo experimental consistiu na utilização do Túnel de Choque, um dispositivo com uma seção de alta pressão (Driver) e uma seção de baixa pressão (Driven), separadas entre si por um diafragma. Ao ser rompido o diafragma, ocorre a propagação do escoamento, o qual é acelerado a velocidades supersônicas através de um bocal bidimensional conectado à câmara de combustão supersônica. A metodologia adotada possibilitou a análise do escoamento no interior da câmara de combustão supersônica do motor Scramjet, gerando escoamentos de M=2.7.

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Cui, Tao, and Yang Ou. "Modeling of Scramjet Combustors Based on Model Migration and Process Similarity." Energies 12, no. 13 (June 30, 2019): 2516. http://dx.doi.org/10.3390/en12132516.

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Contributed by the low cost, the simulation method is considered an attractive option for the optimization and design of the supersonic combustor. Unfortunately, accurate and satisfactory modeling is time-consuming and cost-consuming because of the complex processes and various working conditions. To address this issue, a mathematical modeling for the combustor on the basis of the clustering algorithm, machine learning algorithm, and model migration strategy is developed in this paper. A general framework for the migration strategy of the combustor model is proposed among the similar combustors, and the base model, which is developed by training the machine learning model with data from the existing combustion processes, is amended to fit the unexampled combustor using the model migration strategy with a few data. The simulation results validate the effectiveness of the development strategy, and the migrated model is proved to be suitable for the new combustor in higher accuracy with less time and calculation.

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Huang, Juan-Chen, Yu-Hsuan Lai, Jeng-Shan Guo, and Jaw-Yen Yang. "Simulation of Two-Dimensional Scramjet Combustor Reacting Flow Field Using Reynolds Averaged Navier-Stokes WENO Solver." Communications in Computational Physics 18, no. 4 (October 2015): 1181–210. http://dx.doi.org/10.4208/cicp.190115.210715s.

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AbstractThe non-equilibrium chemical reacting combustion flows of a proposed long slender scramjet system were numerically studied by solving the turbulent Reynolds averaged Navier-Stokes (RANS) equations. The Spalart-Allmaras one equation turbulence model is used which produces better results for near wall and boundary layer flow field problems. The lower-upper symmetric Gauss-Seidel implicit scheme, which enables results converge efficiently under steady state condition, is combined with the weighted essentially non-oscillatory (WENO) scheme to yield an accurate simulation tool for scramjet combustion flow field analysis. Using the WENO schemes high-order accuracy and its non-oscillatory solution at flow discontinuities, better resolution of the hypersonic flow problems involving complex shock-shock/shock-boundary layer interactions inside the flow path, can be achieved. Two types of scramjet combustor with cavity-based and strut-based fuel injector were considered as the testing models. The flow characteristics with and without combustion reactions of the two types combustor model were studied with a transient hydrogen/oxygen combustion model. The detailed results of aerodynamic data are obtained and discussed, moreover, the combustion properties of varying the equivalent ratio of hydrogen, including the concentration of reacting species, hydrogen and oxygen, and the reacting products, water, are demonstrated to study the combustion process and performance of the combustor. The comparisons of flow field structures, pressure on wall and velocity profiles between the experimental data and the solutions of the present algorithms, showed qualitatively as well as the quantitatively in good agreement, and validated the adequacy of the present simulation tool for hypersonic scramjet reacting flow analysis.

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Li, Wei Qiang, and Fei Teng Luo. "Investigation of Performance of Scramjet Combustion Used Kerosene with Clean Air." Advanced Materials Research 601 (December 2012): 294–98. http://dx.doi.org/10.4028/www.scientific.net/amr.601.294.

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The aaviation kerosene is a practical candidate fuel for scramjet engine in the flight regimes of Mach number less than 8, in considerations of safety, economy and portable. In the present paper, the ignition characteristics of a scramjet combustor model fueled by aviation kerosene was experimentally investigated, using the resistance heating direct-connected supersonic combustion facility of Northwestern Polytechnical University. The inflow conditions at the direct-connected combustor entrance were specified as: the nominal Mach of 2.0, the total temperature of 870～930K and the total pressure of about 770kPa. Three combustor performance of difference ER were compared applied CFD simulated. The experimental result suggested that: the cavity design is a key problem for kerosene successful ignition is the combustor; That chooses appropriate gas-oil ratio could increased performance of combustor and kept isolator interfered by gas.

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Yang, Jun, Xian-yu Wu, and Zhen-guo Wang. "Parametric Study of Fuel Distribution Effects on a Kerosene-Based Scramjet Combustor." International Journal of Aerospace Engineering 2016 (2016): 1–14. http://dx.doi.org/10.1155/2016/7604279.

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Generally, the overall performance of scramjet combustor is greatly impacted by the fuel distribution scheme. The current paper mainly conducted a comprehensive parametric study of the impact of fuel distribution on the overall performance of a kerosene-based scramjet combustor. Herein, a 3D supersonic combustor with a recessed cavity and four injection orifices was taken into consideration. The combustor’s performance was analyzed by 3D RANS model. The fuel equivalence ratio for each injection port was taken as the design variables. And the combustion efficiency, the total pressure recovery coefficient, and the drag coefficient were chosen as the objective functions. Some novel data mining methods including DOE technique, Kriging approximation model, interaction analysis, and main effects analysis methods were employed to conduct the parametric study. The distributed fuel injection scheme was optimized by nondominated sorting genetic algorithm. The results show that three objective functions were remarkably affected by both of the total fuel equivalence ratio and the fuel distribution scheme. The objective functions cannot reach the optimal solution at the same time, and there must be a tradeoff among the objective functions.

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Tian, Ye, Shunhua Yang, and Jialing Le. "Study on the Effect of Air Throttling on Flame Stabilization of an Ethylene Fueled Scramjet Combustor." International Journal of Aerospace Engineering 2015 (2015): 1–10. http://dx.doi.org/10.1155/2015/504684.

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The effect of air throttling on flame stabilization of an ethylene fueled scramjet combustor was investigated by numerical simulation and experiments in this paper. The results were obtained under the inflow condition with Mach number of 2.0, total temperature of 900 K, total pressure of 0.8 MPa, and total equivalence ratio of 0.5. The shock train generated by air throttling had a big effect on the flow structure of the scramjet combustor. Compared with the combustor without air throttling, the flow field with air throttling had a lower velocity and higher pressure, temperature, and vortices intensity. Air throttling was an effective way to achieve flame stabilization; the combustion in the combustor without air throttling was nearly blowout. In the experiment, the combustion was nearly blowout with air throttling location of 745 mm, and the fuel/air mixture in the combustor with air throttling location of 875 mm was burned intensively. It was important to choose the location and time sequence of air throttling for fuel ignition and flame stabilization. The numerical simulation results agreed well with experimental measurements.

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Kim, Min-Su, In-Hoi Koo, Keon-Hyeong Lee, Eun-Sung Lee, Hyung-Seok Han, Seung-Min Jeong, Holak Kim, and Jeong-Yeol Choi. "Experimental Study on the Ignition Characteristics of Scramjet Combustor with Tandem Cavities Using Micro-Pulse Detonation Engine." Aerospace 10, no. 8 (August 11, 2023): 706. http://dx.doi.org/10.3390/aerospace10080706.

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This experimental investigation focused on the ignition and combustion characteristics of a tandem cavity-based scramjet combustor with side-by-side identical cavities. This study utilized the Pusan National University-direct connect scramjet combustor (PNU-DCSC), which was capable of simulating flight conditions at Mach number 4.0–5.0 and altitudes of 20–25 km using the vitiated air heater (VAH). The combustion tests were conducted under off-design point conditions corresponding to low inlet enthalpy. It is a condition in which self-ignition does not occur, and a micro pulse detonation engine (μPDE) ignitor is used. The results revealed that as the injection pressure of the gaseous hydrogen fuel (GH2) and the corresponding equivalence ratio increased, the combustion mode transitioned from the cavity-shear layer flame to the jet-wake flame. Furthermore, the measured wall static pressure profiles along the isolator and scramjet combustor indicated that the region of elevated pressure distribution caused by the shock train expanded upstream with higher equivalence ratios. When ignited from the secondary cavity, the combustion area did not extend to the primary cavity at lower equivalence ratios, while it expanded upstream faster with higher equivalence ratios. Therefore, the combustion characteristics of the tandem cavity were found to vary based on the overall equivalence ratio of the main fuel (GH2) and ignition position.

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Zhang, Linqing, Juntao Chang, Wenxiang Cai, Hui Sun, and Yingkun Li. "A Preliminary Research on Combustion Characteristics of a Novel-Type Scramjet Combustor." International Journal of Aerospace Engineering 2022 (December 30, 2022): 1–18. http://dx.doi.org/10.1155/2022/3930440.

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In this work, a new configuration of strut-based scramjet is proposed, and a series of simulations are conducted to investigate its possibility of practical application. The simulation results are verified via the classical DLR ramjet and an experiment conducted on the connected pipe facility. The inlet area ( A in ) and air intake height ( H ) of the combustor are varied independently to investigate their performance. The results indicate that the flow field and shock wave structure of such engine reveal similar characteristics as the classical DLR engine, and the variation in engine geometry can significantly affect its combustion characteristics. Moreover, the combustion efficiency could be enhanced by 2% as the A in varied from 900π mm2 to 1100π mm2; increasing the air intake path ( H ) to 12 mm can increase the combustion efficiency by 25%. In general, the present work proposes a new geometry of the scramjet combustor; this combustor has possibility of practical application, but a further and detailed investigation is still needed.

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Rouzbar, R., and S. Eyi. "Reacting flow analysis of a cavity-based scramjet combustor using a Jacobian-free Newton–Krylov method." Aeronautical Journal 122, no. 1258 (December 2018): 1884–915. http://dx.doi.org/10.1017/aer.2018.110.

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ABSTRACTThe scramjet is a rather a new technology and there are many issues related to their operation, especially when it comes to the combustion processes. Combustion in high-speed flows causes various problems such as flame instability and poor fuel–air mixing efficiency. One of the methods used to overcome these problems is to recess a cavity in the combustor wall where a secondary flow is generated. In this study, a computational fluid dynamics (CFD) code is developed to analyse the reacting flow passing through the cavity-based scramjet combustor. The developed code is based on three-dimensional coupled Navier–Stokes and finite rate chemistry equations. An ethylene-air reduced chemical reaction model is used as a fuel–air combination. The Spalart–Allmaras model is utilised for turbulence closure. The non-dimensional form of the flow and chemical reaction equations are discretised using a finite volume method. The Jacobian-Free Newton–Krylov (JFNK) method is used to solve the coupled system of non-linear equations. The JFNK is a matrix-free solution method which improves the computational cost of Newton’s method. The parameters that affect the performance of the JFNK method are studied in the analysis of a scramjet combustor. The influence of the forcing term on the convergence of the JFNK method is studied in the analysis of scramjet combustor. Different upwind flux vector splitting methods are utilised. Various flux limiter techniques are employed for the calculations of higher order flux vectors. The effects of flux vector splitting and flux limiter methods on the convergence and accuracy of the JFNK method are evaluated. Moreover, the variations of the mixing efficiency with fuel injection angles are studied.

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Ingle, Rahul, and Debasis Chakraborty. "Numerical Simulation of Dual-Mode Scramjet Combustor with Significant Upstream Interaction." International Journal of Manufacturing, Materials, and Mechanical Engineering 2, no. 3 (July 2012): 60–74. http://dx.doi.org/10.4018/ijmmme.2012070105.

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This paper is concerned with a numerical study corresponding to experimental investigation of Chinzei and co-workers on hydrogen fueled dual-mode scramjet engine essentially to understand the key features of upstream interaction, mixing and combustion. Three dimensional Navier Stokes equations along with a K-? turbulence model and infinitely fast kinetics are solved using commercial CFD software. Reasonable agreement has been obtained between the computed surface pressure with experimental values and the results of other numerical simulations. Insights into the flow features inside the combustor are obtained through analysis of various thermochemical parameters. The comparison of surface pressure with experimental results and other numerical results demonstrated that simple kinetics and turbulence – chemistry interaction model may be adequate to address the overall flow features in the combustor. A principal conclusion is that the boundary layer at the combustor entry has a pronounced effect on the flow development in the dual-mode scramjet combustor and causes significant upstream interaction.

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Roga, Sukanta, and Krishna Murari Pandey. "Computational Analysis of Hydrogen-Fueled Scramjet Combustor Using Cavities in Tandem Flame Holder." Applied Mechanics and Materials 772 (July 2015): 130–35. http://dx.doi.org/10.4028/www.scientific.net/amm.772.130.

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This work presents the computational analysis of scramjet combustor using cavities in tandem flame holder by means of 3D. The fuel used by scramjet combustor with cavities in tandem flame holder is hydrogen, the fluid flow and the work is based on the species transport combustion with standard k-ε viscous model. The Mach number at inlet is 2.47 and stagnation temperature and static pressure for vitiated air are 1000K and 100kPa respectively. These computational analysis is mainly aimed to study the flow structure and combustion efficiency. The computational results are compared qualitatively and quantitatively with experimental results and these are agreed as well. Due to the combustion, the recirculation region behind the cavity injector becomes larger as compared to mixing case which acts as a flame holder. From the analysis, the maximum Mach number of 2.33 is observed in the recirculation areas.

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Jeong, Seung-Min, and Jeong-Yeol Choi. "Combined Diagnostic Analysis of Dynamic Combustion Characteristics in a Scramjet Engine." Energies 13, no. 15 (August 4, 2020): 4029. http://dx.doi.org/10.3390/en13154029.

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In this work, the dynamic combustion characteristics in a scramjet engine were investigated using three diagnostic data analysis methods: DMD (Dynamic Mode Decomposition), STFT (Short-Time Fourier Transform), and CEMA (Chemical Explosive Mode Analysis). The data for the analyses were obtained through a 2D numerical experiment using a DDES (Delayed Detached Eddy Simulation) turbulence model, the UCSD (University of California at San Diego) hydrogen/oxygen chemical reaction mechanism, and high-resolution schemes. The STFT was able to detect that oscillations above 50 kHz identified as dominant in FFT results were not the dominant frequencies in a channel-type combustor. In the analysis using DMD, it was confirmed that the critical point that induced a complete change of mixing characteristics existed between an injection pressure of 0.75 MPa and 1.0 MPa. A combined diagnostic analysis that included a CEMA was performed to investigate the dynamic combustion characteristics. The differences in the reaction steps forming the flame structure under each combustor condition were identified, and, through this, it was confirmed that the pressure distribution upstream of the combustor dominated the dynamic combustion characteristics of this scramjet engine. From these processes, it was confirmed that the combined analysis method used in this paper is an effective approach to diagnose the combustion characteristics of a supersonic combustor.

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Relangi, Naresh, Lakshmi Narayana Phaneendra Peri, Caio Henrique Franco Levi Domingos, Amalia Fossella, Julia Meria Leite Henriques, and Antonella Ingenito. "Design of Supersonic and Hybrid engine based Advanced Rocket (SHAR)." IOP Conference Series: Materials Science and Engineering 1226, no. 1 (February 1, 2022): 012031. http://dx.doi.org/10.1088/1757-899x/1226/1/012031.

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Abstract The paper deals with the design of a two-stage to orbit rocket launcher loaded with a solid rocket booster, scramjet, and hybrid rocket for delivering a 100kg payload in 200 km circular orbit. The possibility of implementing a cavity-based axisymmetric circular combustor in a scramjet is proposed. Computational analysis on various injector locations in a circular combustor and their validation with the test bench results were performed. The utilisation of a hybrid rocket in the final stage of the launcher to deliver the payload is discussed and the performance characteristics of the circular scramjet combustor and the hybrid rocket are shown. The overall mission proposed based on the sustainable and reusable characteristics.

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Fureby, Christer, Guillaume Sahut, Alessandro Ercole, and Thommie Nilsson. "Large Eddy Simulation of Combustion for High-Speed Airbreathing Engines." Aerospace 9, no. 12 (December 1, 2022): 785. http://dx.doi.org/10.3390/aerospace9120785.

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Large Eddy Simulation (LES) has rapidly developed into a powerful computational methodology for fluid dynamic studies, between Reynolds-Averaged Navier–Stokes (RANS) and Direct Numerical Simulation (DNS) in both accuracy and cost. High-speed combustion applications, such as ramjets, scramjets, dual-mode ramjets, and rotating detonation engines, are promising propulsion systems, but also challenging to analyze and develop. In this paper, the building blocks needed to perform LES of high-speed combustion are reviewed. Modelling of the unresolved, subgrid terms in the filtered LES equations is highlighted. The main families of combustion models are presented, focusing on finite-rate chemistry models. The density-based finite volume method and the reaction mechanisms commonly employed in LES of high-speed H2-air combustion are briefly reviewed. Three high-speed combustor applications are presented: an experiment of supersonic flame stabilization behind a bluff body, a direct connect facility experiment as a transition case from ramjet to scramjet operation mode, and the STRATOFLY MR3 Small-Scale Flight Experiment. Several combinations of turbulence and combustion models are compared. Comparisons with experiments are also provided when available. Overall, the results show good agreement with experimental data (e.g., shock train, mixing, wall heat flux, transition from ramjet to scramjet operation mode).

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Relangi, Naresh, Antonella Ingenito, and Suppandipillai Jeyakumar. "The Implication of Injection Locations in an Axisymmetric Cavity-Based Scramjet Combustor." Energies 14, no. 9 (May 4, 2021): 2626. http://dx.doi.org/10.3390/en14092626.

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This paper presents the effect of cavity-based injection in an axisymmetric supersonic combustor using numerical investigation. An axisymmetric cavity-based angled and transverse injections in a circular scramjet combustor are studied. A three-dimensional Reynolds-averaged Navier–Stokes (RANS) equation along with the k-ω shear-stress transport (SST) turbulence model and species transport equations are considered for the reacting flow studies. The numerical results of the non-reacting flow studies are validated with the available experimental data and are in good agreement with it. The performance of the injection system is analyzed based on the parameters like wall pressures, combustion efficiency, and total pressure loss of the scramjet combustor. The transverse injection upstream of the cavity and at the bottom wall of the cavity in a supersonic flow field creates a strong shock train in the cavity region that enhances complete combustion of hydrogen-air in the cavity region compared to the cavity fore wall injection schemes. Eventually, the shock train in the flow field enhances the total pressure loss across the combustor. However, a marginal variation in the total pressure loss is observed between the injection schemes.

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Zhang, Junlong, Guangjun Feng, Guowei Luan, Hongchao Qiu, and Wen Bao. "Research on combustion performance improvement by strut/wall combined injection in scramjet combustor." Thermal Science, no. 00 (2023): 92. http://dx.doi.org/10.2298/tsci220917092z.

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Strut/wall combined fuel injection scheme was adopted to improve mixing and combustion efficiency in a scramjet combustor fueled with liquid kerosene in the condition of Mach 6. Injectors were placed on the front of the strut and the side wall of the combustor. A series of numerical simulations and experiments were carried out to improve the combustor performance under conditions of different incoming flow velocity, injection methods and fuel distribution ratios. The value of pressure was obtained by pressure sensor and the flame images were captured by the high-speed camera in experiment. By processing and analyzing the basic data, characteristics of fuel mixing and combustion performance were discussed in this paper. Then, the influence mechanism of the strut/wall combined injection on the performance of the combustor was explained based on the performance with influence factors. Results indicated that the mixing and combustion efficiency was related to condition, injection method and nozzle arrangement. The strut/wall combined injection dispersed the heat release, which could reduce the pressure rise and total temperature. The fuel distribution ratio between the strut injection and wall injection is also a key factor affecting the performance of the combustor. These results in this paper are valuable for the combustion organization in the supersonic combustor and the improvement of the combustor performance.

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Osaka, Jun, Yoshitaka Uriuda, Osamu Imamura, Kiyotaka Yamashita, Shuhei Takahashi, Mitsuhiro Tsue, and Michikata Kono. "Combustion Characteristics of Kerosene in a Scramjet Combustor." JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 55, no. 637 (2007): 98–103. http://dx.doi.org/10.2322/jjsass.55.98.

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Kay, I. W., W. T. Peschke, and R. N. Guile. "Hydrocarbon-fueled scramjet combustor investigation." Journal of Propulsion and Power 8, no. 2 (March 1992): 507–12. http://dx.doi.org/10.2514/3.23505.

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Deepu, M., S. S. Gokhale, and S. Jayaraj. "Numerical Modelling of Scramjet Combustor." Defence Science Journal 57, no. 4 (July 20, 2007): 367–79. http://dx.doi.org/10.14429/dsj.57.1784.

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Li, Zhen, and Hongbin Gu. "Investigation for Effects of Jet Scale on Flame Stabilization in Scramjet Combustor." Energies 15, no. 10 (May 21, 2022): 3790. http://dx.doi.org/10.3390/en15103790.

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Jet scale affects the mixing and combustion of fuel and inflow. With the increase in the scale of scramjet combustors, the study of large-scale jets is particularly significant. The effects of jet scale on flame stability in scramjet combustors were studied by direct-connect combustion experiments. In this paper, the flame distribution characteristics of different jet scales were compared by using similar jet/inflow momentum ratios. The inflow Mach numbers were 2.4 and 3.0, and the total temperature was 1265 K and 1600 K, respectively. The results show that, when the equivalence ratio increases, the combustion intensity increases. Under the condition of same momentum ratio, the increase of jet scale is conducive to fuel injection into the core mainstream, increasing heat release, and the flame stabilization mode will change from cavity stabilization mode to jet-wake stabilization mode. Increasing the distance between jet orifices is not beneficial to combustion, and may even lead to blowoff.

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Wei, Bao Xi, Qiang Gang, Yan Zhang, Rong Jian Liu, Liang Tian, and Xu Xu. "A Study on Performance Comparison of Integrated Aerodynamic-Ramp-Injector/ Gas-Portfire Flame Holder with Cavity." Applied Mechanics and Materials 390 (August 2013): 8–11. http://dx.doi.org/10.4028/www.scientific.net/amm.390.8.

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Experimental study of an integrated aerodynamic-ramp-injector /gas-portfire (aero-ramp/G-P) has been conducted in a hydrogen-fueled scramjet combustor. The aero-ramp injectors consisted of four flush-walled holes, arranged to induce vorticular motion and enhance mixing. For comparison, a recessed cavity with four low downstream-angled circular injector holes was also examined. The combustor models were investigated experimentally using the scramjet direct connected test facility at the Beihang University. The facility can deliver a continuous supersonic flow of Mach number 2 with a total temperature of 1200K. The hot experimental results showed that the combustion efficiency and air specific impulse of aero-ramp/G-P are 85% and 35s while the corresponding values of cavity are 92% and 34s. These results justify the feasibility of aero-ramp/G-P flame holder.

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Manna, P., and D. Chakraborty. "Numerical investigation of transverse sonic injection in a non-reacting supersonic combustor." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 219, no. 3 (March 1, 2005): 205–15. http://dx.doi.org/10.1243/095441005x30261.

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Efficient combustion and heat release in scramjet flows depend on effective mixing of the fuel in supersonic streams. Usually, transverse sonic injection in-stages are employed as one of the suitable means for efficient supersonic combustor design. Numerical simulations are carried out to study the mixing characteristics of staged sonic air injections in supersonic stream ( M = 2.07) behind a backward-facing step in scramjet combustor by solving three-dimensional Navier-Stokes equations along with K-ε turbulence model with a commercial CFD software CFX-TASCFlow. Computed results of the jet penetration and spreading show very good agreement with the experimental values and the results of other computations. A good overall match has been obtained between the experimental values and the computation for various flow profiles at various axial locations in the combustor. However, the values differ in the near-field region at the injection plane. The assumed uniformity of the flow-field properties at the injection orifice and/or the inadequacy of the turbulence model considered in this study is conjectured to be the cause of the difference.

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Fureby, C. "Large eddy simulation modelling of combustion for propulsion applications." Philosophical Transactions of the Royal Society A: Mathematical, Physical and Engineering Sciences 367, no. 1899 (July 28, 2009): 2957–69. http://dx.doi.org/10.1098/rsta.2008.0271.

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Predictive modelling of turbulent combustion is important for the development of air-breathing engines, internal combustion engines, furnaces and for power generation. Significant advances in modelling non-reactive turbulent flows are now possible with the development of large eddy simulation (LES), in which the large energetic scales of the flow are resolved on the grid while modelling the effects of the small scales. Here, we discuss the use of combustion LES in predictive modelling of propulsion applications such as gas turbine, ramjet and scramjet engines. The LES models used are described in some detail and are validated against laboratory data—of which results from two cases are presented. These validated LES models are then applied to an annular multi-burner gas turbine combustor and a simplified scramjet combustor, for which some additional experimental data are available. For these cases, good agreement with the available reference data is obtained, and the LES predictions are used to elucidate the flow physics in such devices to further enhance our knowledge of these propulsion systems. Particular attention is focused on the influence of the combustion chemistry, turbulence–chemistry interaction, self-ignition, flame holding burner-to-burner interactions and combustion oscillations.

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Cheng, Wen Qiang, Jing Yuan Liu, and Rakesh Shrestha. "Insight in the Performance of Scramjet Combustor Based on Orthogonal Experimental Design." Advanced Materials Research 705 (June 2013): 463–67. http://dx.doi.org/10.4028/www.scientific.net/amr.705.463.

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A numerical insight was accomplished to optimize the scramjet combustor configuration based on orthogonal experimental design. Parametric modeling of combustor configurations was performed by the orthogonal array with 13 factors at 3 levels. Numerical simulations were proceeded by k-ε standard turbulence model and eddy-dissipation model in the combustion process. The performance indexes of combustion efficiency, total pressure recovery coefficient and thrust gain coefficient were evaluated. Detailed comparison with the effect of the factors on the performance was also carried out to demonstrate the main factors and determine the optimal configuration. The analysis of the extreme differences of the factors indicates that the main factors affecting combustion efficiency were the length of the wedge, the length depth ratio of the cavity, the depth of the cavity, and the length of the expanding section; The main factors affecting total pressure recovery coefficient are the angle of the primary combustor, the length of the expanding section, and the thickness of the strut; The main factors affecting thrust gain coefficient are the thickness of the strut, the length of the expanding section, and the angle of the secondary combustor. Validation of the optimal configuration is then confirmed that its performance is higher than the rest of the configurations, with the combustion efficiency of 0.915 and the total pressure recovery coefficient of 0.486, which are 31.5% and 65.9% higher than the experimental results, respectively.

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Wang, Z. P., H. B. Gu, L. W. Cheng, F. Q. Zhong, and X. Y. Zhang. "CH* Luminance Distribution Application and a One-Dimensional Model of the Supersonic Combustor Heat Release Quantization." International Journal of Turbo & Jet-Engines 36, no. 1 (March 26, 2019): 45–50. http://dx.doi.org/10.1515/tjj-2016-0064.

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Abstract One-dimensional model is an important way to evaluate the performance and flow characteristics of dual-mode scramjet combustor. Current work is based on a modified one-dimensional model assisted by measurements acquired on a direct-connected scramjet facility. CH* images and gas-sampling facility have been employed to quantify heat release for optimizing one-dimensional model. The results show that modified one-dimensional model gives a better evaluation of axis parameters distribution, especially for Mach number, which is the standard parameter to evaluate combustion mode. The ram/scram mode derived by the analytical results has been investigated. Intensive heat release is beneficial to obtain more stable pre-combustion shock and subsonic flow in the recirculation zone.

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Suppandipillai, Jeyakumar, Jayaraman Kandasamy, R. Sivakumar, Mehmet Karaca, and Karthik K. "Numerical investigations on the hydrogen jet pressure variations in a strut based scramjet combustor." Aircraft Engineering and Aerospace Technology 93, no. 4 (April 5, 2021): 566–78. http://dx.doi.org/10.1108/aeat-08-2020-0162.

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Purpose This paper aims to study the influences of hydrogen jet pressure on flow features of a strut-based injector in a scramjet combustor under-reacting cases are numerically investigated in this study. Design/methodology/approach The numerical analysis is carried out using Reynolds Averaged Navier Stokes (RANS) equations with the Shear Stress Transport k-ω turbulence model in contention to comprehend the flow physics during scramjet combustion. The three major parameters such as the shock wave pattern, wall pressures and static temperature across the combustor are validated with the reported experiments. The results comply with the range, indicating the adopted simulation method can be extended for other investigations as well. The supersonic flow characteristics are determined based on the flow properties, combustion efficiency and total pressure loss. Findings The results revealed that the augmentation of hydrogen jet pressure via variation in flame features increases the static pressure in the vicinity of the strut and destabilize the normal shock wave position. Indeed, the pressure of the mainstream flow drives the shock wave toward the upstream direction. The study perceived that once the hydrogen jet pressure is reached 4 bar, the incoming flow attains a subsonic state due to the movement of normal shock wave ahead of the strut. It is noticed that the increase in hydrogen jet pressure in the supersonic flow field improves the jet penetration rate in the lateral direction of the flow and also increases the total pressure loss as compared with the baseline injection pressure condition. Practical implications The outcome of this research provides the influence of fuel injection pressure variations in the supersonic combustion phenomenon of hypersonic vehicles. Originality/value This paper substantiates the effect of increasing hydrogen jet pressure in the reacting supersonic airstream on the performance of a scramjet combustor.

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Kanda, Takeshi, Nobuo Chinzei, Kenji Kudo, Atsuo Murakami, and Tetsuo Hiraiwa. "Autoignited Combustion Testing in a Water-Cooled Scramjet Combustor." Journal of Propulsion and Power 20, no. 4 (July 2004): 657–64. http://dx.doi.org/10.2514/1.11381.

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Dharavath, Malsur, P. Manna, and Debasis Chakraborty. "Thermochemical exploration of hydrogen combustion in generic scramjet combustor." Aerospace Science and Technology 24, no. 1 (January 2013): 264–74. http://dx.doi.org/10.1016/j.ast.2011.11.014.

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Li, Xiang, Qingchun Lei, Xiaocun Zhao, Wei Fan, Shuang Chen, Li Chen, Ye Tian, and Quan Zhou. "Combustion Characteristics of a Supersonic Combustor with a Large Cavity Length-to-Depth Ratio." Aerospace 9, no. 4 (April 14, 2022): 214. http://dx.doi.org/10.3390/aerospace9040214.

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The combustion characteristics of a hydrogen-fueled supersonic combustor featuring a large cavity length-to-depth ratio (i.e., 11) were examined by performing experimental trials while varying the fuel injector positions and equivalence ratios. During these trials, flame chemiluminescence images were acquired simultaneously from the side and bottom of the combustor under Mach 2.0 inflow conditions. The flame was observed to stabilize inside the cavity under all conditions. Proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) analyses of sequential flame chemiluminescence images demonstrated the important effects of oblique shocks induced by fuel injection and heat release on flame stabilization. Because fluctuations in the locations of the flame and of the intense heat release zone were not observed and no dominant frequency was identified in POD and DMD analyses, the present configuration was evidently able to suppress combustion instability. The present research provides preliminary guidance for exploring the feasibility of using cavity combustors with large length-to-depth ratios in scramjet engines.

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Shi, Deyong, Wenyan Song, Jingfeng Ye, Bo Tao, Yanhua Wang, and Qiang Fu. "Experimental Investigation of Reacting Flow Characteristics in a Dual-Mode Scramjet Combustor." International Journal of Turbo & Jet-Engines 35, no. 4 (December 19, 2018): 321–30. http://dx.doi.org/10.1515/tjj-2015-0014.

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Abstract In this work, a hydrogen fueled dual-mode scramjet combustor was investigated experimentally. Clean and dry air was supplied to the combustor through a Mach 2 Nozzle with a total temperature of 800 K and a total pressure of 800 kPa. The high enthalpy air was provided by an electricity resistance heater. Room temperature hydrogen was injected with sonic speed from injector orifices vertically, and downstream the injector a tandem cavity flame holder was mounted. Except wall pressure profiles, velocity and temperature profiles in and at exit of the combustor were also measured using hydroxyl tagging velocimetry (HTV) and tunable diode laser absorption spectroscopy (TDLAS) respectively. Results showed that combustion occurred mainly at the bottom side of the combustor. And there were also an extreme disparity of the velocity and temperature profiles along the Y direction, i.e. the transverse direction.

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Bordoloi, Namrata, Krishna Murari Pandey, and Kaushal Kumar Sharma. "Numerical Investigation on the Effect of Inflow Mach Numbers on the Combustion Characteristics of a Typical Cavity-Based Supersonic Combustor." Mathematical Problems in Engineering 2021 (September 8, 2021): 1–14. http://dx.doi.org/10.1155/2021/3526454.

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The air-breathing engines, commonly known as Supersonic Combustor Ramjet (SCRAMJET) engines, are one of the most prominent technologies among researchers due to their high thrust-to-weight ratio. The researchers are constantly making efforts for improved performance of the combustor under the required boundary conditions. The present working computational model studies a hydrogen-fueled parallel cavity scramjet combustor to recognize the complex flow field characteristics and performance of the combustor in Ansys 15.0. The computational model developed is a replica of an experiment conducted in China which slightly modified the boundary conditions. The standard two-equation K- ε turbulence model and Reynolds averaged Navier Stokes (RANS) equation with finite-rate/eddy dissipation species reaction model are used to simulate the problem. The validation of the present model is achieved by comparing the results with already available experimental data in conformity with the literature. The results of the simulations are in satisfactory accord with the experimental data and images. Furthermore, to achieve the stated objective, different incoming Mach numbers, namely, 2.25, 2.52, and 2.75, are considered for a more clear understanding of variables that affects the characteristics of the flow field. The temperature, Mach number, density pressure, and H2O mass fraction contours were studied to facilitate proper understanding. The maximum temperature rise observed is 2711.467 K for M = 2.25. Additionally, the performance parameters, namely, combustion and mixing efficiencies, are also studied. The maximum combustion and mixing efficiencies are 87.47% and 98.15% for M = 2.25 and 2.75, respectively.

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Wang, Guangyu, Shijie Liu, Haoyang Peng, and Weidong Liu. "Experimental Investigation of a Cylindrical Air-Breathing Continuous Rotating Detonation Engine with Different Nozzle Throat Diameters." Aerospace 9, no. 5 (May 16, 2022): 267. http://dx.doi.org/10.3390/aerospace9050267.

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A continuous detonation engine with various exhaust nozzles, analogous to typical scramjet cavity combustors with variable rear-wall heights, was adopted to perform a succession of cylindrical air-breathing continuous rotating detonation experiments fueled by a non-premixed ethylene/air mixture. The results show that the detonation combustion was observed to self-sustain in the combustor through simultaneous high-speed imaging covering the combustor and isolator. A long test, lasting more than three seconds, was performed in this unique configuration, indicating that the cylindrical isolator–combustor engine exhibits potential for practical applications. Three distinct combustion modes were revealed with varied equivalent ratios (hybrid mode, sawtooth wave mode, and deflagration mode). The diameter of the nozzle throat was critical in the formation of rotating detonation waves. When the nozzle throat diameter was larger than the specific value, the detonation wave could not form and self-sustain. The upstream boundary of the shock train was supposed to be close to the isolator entrance in conditions of a high equivalence ratio and small nozzle throat diameter. In addition, it was verified that periodic high-frequency pressure oscillation could cause substantial impacts on the incoming flow as compared with the steady deflagration with the same combustor pressure.

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Tao, C., Y. Daren, and B. Wen. "Distributed parameter control arithmetic for an axisymmetrical dual-mode scramjet." Aeronautical Journal 112, no. 1135 (September 2008): 557–65. http://dx.doi.org/10.1017/s0001924000002517.

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AbstractDual-mode scramjet is one of the candidates for hypersonic flight propulsion system which will be used in wide range of flight Mach numbers from 4 to 12 or higher, wherein dual-mode scramjet should be well designed to be suitable for subsonic/supersonic combustion operation according to the flight conditions. Therefore this system is required to operate in a finite number of operational modes that necessitate robust, stable, and smooth transitions between them by which selective operability of supersonic/subsonic combustion modes and efficient combustor operation in these modes may be realised. A key issue in making mode transition efficient and stable is mode transition control. The major problem in mode transition control is the handling of the various flow and combustion coupling effects of dual-mode scramjet whose physical states are spatially coupled and whose governing equations are partial differential equations. Involving these distributed parameter issues, our basic idea is using the shape control theory to study the control problems of mode transition for dual-mode scramjet with the aim of achieving the desirable design properties and increasing control reliabilities. This specific approach is motivated by the promise of novel techniques in control theory developed in recent years. Concrete control arithmetic of this approach, such as shape control model, sensitivity analysis and gradient-based optimisation procedure, are given in this paper. Simulation results for an axisymmetric, wall-injection dual-mode scramjet show the feasibility and validity of the method.

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Sarosh, Ali, Dong Yun Feng, and Muhammad Adnan. "An Aerothermodynamic Design Approach for Scramjet Combustors and Comparative Performance of Low-Efficiency Systems." Applied Mechanics and Materials 110-116 (October 2011): 4652–60. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.4652.

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This paper is aimed at development of an integrated approach based on analytical and computational aerothermodynamics for the special case of design of a 75% (low process-efficiency), hydrogen-fuelled, constant area combustor of a hypersonic airbreathing propulsion (HAP) system thereafter undertaking study of two types of HAP systems. The results of configurational aerothermodynamics implied that the most appropriate constant area configuration had a 30 degrees downstream wall-mounted fuel injector with a single acoustically stable cavity placed downstream of the fuel injection point. Moreover for identical flow inlet parameters and system configurations at lower levels of thermodynamic process efficiencies, the constant combustor-area (i.e. Scramjet 1) engine is superior in its performance to the constant combustor-pressure (i.e. Scramjet 2) engine for all values of fuel-air ratios.

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Zhang, Qi, Weibing Zhu, Dongchao Yang, and Hong Chen. "Numerical Investigation on Flame Stabilization of Cavity-Based Scramjet Combustor Using Compressible Modified FGM Model." International Journal of Chemical Engineering 2022 (February 8, 2022): 1–21. http://dx.doi.org/10.1155/2022/8388827.

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The technology of flame stabilization on cavity-based scramjet combustor has great significance in the field of future spacecraft. In this paper, a compressible modified FGM model was established based on the idea of the flamelet model, which was adopted to simulate the unsteady combustion process of the hydrogen transverse jet in the upper cavity of the scramjet. The results show that the compressible modified FGM model can accurately reflect the flow field and the propagation process of the flame in the supersonic cavity, and can capture the fine shock structure in the flow field. The coupling effect of shock waves and shear layer cause the shear layer to quickly destabilize, resulting in the turbulence effect, which promotes the mixing of air and fuel. The boundary layer separation at the upper wall of the combustion chamber will reduce the stability of the shear layer.

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Tian, Ye, Wen Shi, Mingming Guo, Yuan Liu, Chenlin Zhang, and Jialing Le. "Investigation of combustion characteristics in a hydrogen-fueled scramjet combustor." Acta Astronautica 186 (September 2021): 486–95. http://dx.doi.org/10.1016/j.actaastro.2021.06.021.

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TAKAHASHI, Shuhei, Kazunori WAKAI, Sadatake TOMIOKA, Mitsuhiro TSUE, and Michikata KONO. "Interaction between Combustion and Flowfield in a Rectangular Scramjet Combustor." Journal of the Japan Society for Aeronautical and Space Sciences 46, no. 538 (1998): 633–39. http://dx.doi.org/10.2322/jjsass1969.46.633.

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Takahashi, Shuhei, Kazunori Wakai, Sadatake Tomioka, Mitsuhiro Tsue, and Michikata Kono. "Effects of combustion on flowfield in a model scramjet combustor." Symposium (International) on Combustion 27, no. 2 (January 1998): 2143–50. http://dx.doi.org/10.1016/s0082-0784(98)80062-3.

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Nordin-Bates, K., C. Fureby, S. Karl, and K. Hannemann. "Understanding scramjet combustion using LES of the HyShot II combustor." Proceedings of the Combustion Institute 36, no. 2 (2017): 2893–900. http://dx.doi.org/10.1016/j.proci.2016.07.118.

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Journal articles on the topic 'Scramjet combustor'

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