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Journal articles on the topic 'Supersonic combustion ramjet engines'

Author: Grafiati
Published: 4 June 2021
Last updated: 6 September 2023

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1

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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2

Tunik, Yu V., and V. O. Mayorov. "Energy efficiency of detonation combustion in supersonic ramjet engines." Acta Astronautica 194 (May 2022): 488–95. http://dx.doi.org/10.1016/j.actaastro.2021.09.038.

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3

Kolosenok S.V., Kuranov A.L., Savarovskiy A.A., Bulat P.V., Galadzhun A.A., Levihin A.A., and Nikitenko A.B. "The application of supplementary fuels for the control of supersonic reacting air-fuel mix flows in the combustion chamber." Technical Physics Letters 48, no. 13 (2022): 40. http://dx.doi.org/10.21883/tpl.2022.13.53351.18764.

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Besides gas-dynamic methods, chemical ones are also suitable for the implementation of stable supersonic combustion of hydrocarbon fuels. Organoelemental compounds are known for their high reactivity, so attention was paid to organosilicon liquid during the research on the experimental model. The obtained estimates of the laminar flame speed in a mixture of vapors of this liquid with air were 0.72-0.8 m/s, which is higher than that of ethylene successfully used in supersonic combustion tests. The tested compound can be considered as a candidate for supplementary fuel to control the supersonic reactive flows in the combustion chambers of ramjet engines. Keywords: supersonic combustion, supplementary fuels, laminar flame speed, combustion efficiency
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4

Timoshenko, V. I., V. P. Halynskyi, and Yu V. Knyshenko. "Theoretical studies on rocket/space hardware aerogas dynamics." Technical mechanics 2021, no. 2 (June 29, 2021): 46–59. http://dx.doi.org/10.15407/itm2021.02.046.

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This paper presents the results of theoretical studies on rocket/space hardware aerogas dynamics obtained from 2016 to 2020 at the Department of Aerogas Dynamics and Technical Systems Dynamics of the Institute of Technical Mechanics of the National Academy of Sciences of Ukraine and the State Space Agency of Ukraine along the following lines: rocket aerodynamics, mathematical simulation of the aerogas thermodynamics of a supersonic ramjet vehicle, jet flows, and the hydraulic gas dynamics of low-thrust control jet engines. As to rocket aerodynamics, computational methods and programs (CMPs) were developed to calculate supersonic flow past finned rockets. The chief advantage of the CMPs developed is computational promptness and ease of adding wings and control and stabilization elements to rocket configurations. A mathematical simulation of the aerogas thermodynamics of a supersonic ramjet vehicle yielded new results, which made it possible to develop a prompt technique for a comprehensive calculation of ramjet duct flows and generalize it to 3D flow past a ramjet vehicle. Based on marching methods, CMPs were developed to simulate ramjet duct flows with account for flow past the airframe upstream of the air inlet, the effect of the combustion product jet on the airframe tail part, and its interaction with a disturbed incident flow. The CMPs developed were recommended for use at the preliminary stage of ramjet component shape selection. For jet flows, CMPs were developed for the marching calculation of turbulent jets of rocket engine combustion products with water injection into the jet body. This made it possible to elucidate the basic mechanisms of the effect of water injection, jet–air mixing, and high-temperature rocket engine jet afterburning in atmospheric oxygen on the flow pattern and the thermogas dynamic and thermalphysic jet parameters. CMPs were developed to simulate the operation of liquid-propellant low-thrust engine systems. They were used in supporting the development and ground firing tryout of Yuzhnoye State Design Office’s radically new system of control jet engines fed from the sustainer engine pipelines of the Cyclone-4M launch vehicle upper stage. The computed results made it possible to increase the informativity of firing test data in flight simulation. The CMPs developed were transferred to Yuzhnoye State Design Office for use in design calculations.
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5

Колосенок, С. В., А. Л. Куранов, А. А. Саваровский, П. В. Булат, А. А. Галаджун, А. А. Левихин, and А. Б. Никитенко. "Применение вспомогательных топлив для управления сверхзвуковыми потоками реагирующих топливно-воздушных смесей в канале камеры сгорания." Письма в журнал технической физики 47, no. 19 (2021): 19. http://dx.doi.org/10.21883/pjtf.2021.19.51507.18764.

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Besides gas-dynamic methods, chemical ones are also suitable for the implementation of stable supersonic combustion of hydrocarbon fuels. Organoelemental compounds are known for their high reactivity, so attention was paid to organosilicon liquid during the research on the experimental model. The obtained estimates of the laminar flame speed in a mixture of vapors of this liquid with air were 0.72-0.8 m/s, which is higher than that of ethylene successfully used in supersonic combustion tests. The tested compound can be considered as a candidate for supplementary fuel to control the supersonic reactive flows in the combustion chambers of ramjet engines.
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6

Кислов, Олег Владимирович, and Михаил Анатольевич Шевченко. "ОСОБЕННОСТИ РАСЧЕТА И РЕГУЛИРОВАНИЯ ДВУХКОНТУРНОГО ТУРБОРЕАКТИВНОГО ДВИГАТЕЛЯ С ФОРСАЖНОЙ КАМЕРОЙ СГОРАНИЯ В НАРУЖНОМ КОНТУРЕ НА ПРЯМОТОЧНЫХ РЕЖИМАХ РАБОТЫ." Aerospace technic and technology, no. 6 (November 27, 2020): 15–23. http://dx.doi.org/10.32620/aktt.2020.6.02.

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A promising direction in aviation is the creation of anaircraft for supersonic cruise speeds (Mach 3...4). It is known that ramjet engines are more preferable for Mach numbers larger 3. However, they do not have starting thrust and uneconomical at subsonic flight speeds. At the same time, at subsonic flight speeds, turbofan engines are the most expedient. The combination of the positive properties of turbofan engines at subsonic speeds and a ramjet engines at supersonic speeds is possible by using duct-burning turbofan engine, which can operate at the ramjet mode with the blocked gas turbine duct at supersonic flight conditions. At this mode, duct-burning turbofan engine turns into ramjet engine, which, however, has special features due to the presence of fan in front of the combustion chamber, which operates in turbine mode or in zero power mode and also because of the outlet jet, which has annular shape, flows out from the duct causes the appearance of bottom drag. The presence of bottom drag requires both the development of a mathematical model for its calculation and taking into account its influence on the choice of the control law for the nozzle outlet area. The article presents a mathematical model of the working process of duct-burning turbofan engine at ramjet mode, taking into account the presence of fan in the flow path and bottom drug. Using the developed mathematical model, the regularities of changes in the internal and effective thrust, as well as the specific fuel consumption, depending on the relative fuel consumption and the critical section of the nozzle at a given altitude and flight speed are established. The critical section of the nozzle is the main regulating factor, and the relative fuel consumption is related to the main regulating factor - the fuel consumption. These patterns are useful for choosing a control program.There is such a combination of regulating factors whichprovides two extremes in the regularities of trust and specific fuel consumption changes: the mode of minimum specific fuel consumption and the mode of maximum thrust. In addition, the influence of gas underexpansion in the nozzle on the thrust-economic parameters of the engine and the required area of the nozzle outlet section were estimated. The obtained regularities are advisable to use when engine control program is chosen.
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7

Laube, Tomasz, and Janusz Piechna. "Analytical and Numerical Feasibility Analysis of a Contra-Rotary Ramjet Engine." Energies 13, no. 1 (December 30, 2019): 163. http://dx.doi.org/10.3390/en13010163.

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A new idea for a contra-rotary ramjet engine is presented. To define the theoretical limits of the non-typical, contra-rotary ramjet engine configuration, its analytical model was developed. The results obtained from that model and the analytical results were compared with those received from numerical simulations. The main weakness of existing rotary ramjet engine projects is the very high rotational speed of the rotor required for achieving supersonic inlet flow. In this paper, a new idea for a contra-rotary ramjet engine (CORRE) is presented and analyzed. This paper presents the results of analytical analysis and numerical simulations of a jet engine system with two rotors rotating in opposite directions. Contra-rotating rotors generate a supersonic air velocity at the inlet to the compressor at two times slower rotor’s speed. To determine the flow characteristics, combustion process, and engine efficiency of the double-rotor engine, a numerical solution of the average Navier-Stokes equations was used with the k-eps turbulence model and the non-premixed combustion model. The results of numerical simulations of flow and the combustion process inside the contra-rotary jet engine achieving a shockwave compression are shown and compared with similar data for a single-rotor engine design and analytical data. This paper presents only the calculation results of the flow processes and the combustion process, indicating the advantages of the proposed double-rotor design. The results of the numerical analysis were presented on the contours and diagrams of the pressure and flow velocity, temperature distribution, and mass fraction of the fuel.
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8

Козел, Дмитрий Викторович. "Выбор геометрических характеристик фронтового устройства и длины камеры сгорания прямоточного типа." Aerospace technic and technology, no. 4sup2 (August 27, 2021): 19–28. http://dx.doi.org/10.32620/aktt.2021.4sup2.03.

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A method has been developed for selecting the geometric characteristics of the front and the length of the direct-flow combustion chamber. Afterburner combustion chambers are of the ramjet type and are used for a short-term increase in the thrust of a gas turbine engine during takeoff, for overcoming the sound barrier by an aircraft and for flying at supersonic speed, and for making maneuvers. As part of ramjet engines, ramjet combustion chambers are used as the main combustion chambers in which the process of fuel combustion and heat supply to the working fluid is ensured. The developed method for selecting the geometric characteristics consists in optimizing the main operating characteristics of the combustion chamber. Mathematical models are proposed for describing the dependence of the total pressure loss, the combustion efficiency and the range of stable operation of the combustion chamber against the parameters of the flow at the inlet to the combustion chamber and the geometric characteristics of the front device and the length of the combustion chamber. The analysis of the dependences of the combustion chamber working characteristics on the geometric characteristics of the front-line device and its length is carried out. As a result of the analysis of mathematical models, a list of the main geometric characteristics of the front device was determined, on which the total pressure loss, the combustion efficiency and the range of stable operation of the combustion chamber depend. Optimization parameters, optimization criterion and limits for solving the optimization problem are determined. As an implementation of the optimization method, it is proposed to use a diagram of the combustion chamber performance in the coordinates of the optimization parameters. The developed method makes it possible to ensure the optimal basic operating characteristics of the combustion chamber - total pressure loss, combustion efficiency and combustion stability limits.
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9

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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10

IMAMURA, Osamu, Yuta ISHIKAWA, Shunsuke SUZUKI, Koshiro FUKUMOTO, Shunsuke NISHIDA, Yasushige UJIIE, and Mitsuhiro TSUE. "Combustion Characteristics of Liquid Normal Alkane Fuels in a Model Combustor of Supersonic Combustion Ramjet Engine." JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 58, no. 675 (2010): 116–22. http://dx.doi.org/10.2322/jjsass.58.116.

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11

WANG, JIANGFENG, CHEN LIU, and YIZHAO WU. "NUMERICAL SIMULATION OF SPRAY ATOMIZATION IN SUPERSONIC FLOWS." Modern Physics Letters B 24, no. 13 (May 30, 2010): 1299–302. http://dx.doi.org/10.1142/s0217984910023475.

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With the rapid development of the air-breathing hypersonic vehicle design, an accurate description of the combustion properties becomes more and more important, where one of the key techniques is the procedure of the liquid fuel mixing, atomizing and burning coupled with the supersonic crossflow in the combustion chamber. The movement and distribution of the liquid fuel droplets in the combustion chamber will influence greatly the combustion properties, as well as the propulsion performance of the ramjet/scramjet engine. In this paper, numerical simulation methods on unstructured hybrid meshes were carried out for liquid spray atomization in supersonic crossflows. The Kelvin-Helmholtz/Rayleigh-Taylor hybrid model was used to simulate the breakup process of the liquid spray in a supersonic crossflow with Mach number 1.94. Various spray properties, including spray penetration height, droplet size distribution, were quantitatively compared with experimental results. In addition, numerical results of the complex shock wave structure induced by the presence of liquid spray were illustrated and discussed.
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12

Скибина, Надежда Петровна. "Computational study of unsteady gas flow in the combustion chamber of a ramjet engine with heat transfer." Вычислительные технологии, no. 6 (January 19, 2021): 50–61. http://dx.doi.org/10.25743/ict.2020.25.6.003.

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Проведено численное исследование нестационарного турбулентного сверхзвукового течения в камере сгорания прямоточного воздушно-реактивного двигателя. Описана методика экспериментального измерения температуры на стенке осесимметричного канала в камере сгорания двигателя. Математическое моделирование обтекания исследуемой модели двигателя проводилось для скоростей набегающего потока M = 5 ... 7. Начальные и граничные условия задачи соответствовали реальному аэродинамическому эксперименту. Проанализированы результаты численного расчета. Рассмотрено изменение распределения температуры вдоль стенки канала с течением времени. Проведена оценка согласованности полученных экспериментальных данных с результатами математического моделирования. Purpose. The aim of this study is a numerical simulation of unsteady supersonic gas flow in a working path of ramjet engine under conditions identical to aerodynamic tests. Free stream velocity corresponding to Mach numbers M=5 ... 7 are considered. Methodology. Presented study addresses the methods of physical and numerical simulation. The probing device for thermometric that allows to recording the temperature values along the wall of internal duct was proposed. To describe the motion of a viscous heat-conducting gas the unsteady Reynolds averaged Navier - Stokes equations are considered. The flow turbulence is accounted by the modified SST model. The problem was solved in ANSYS Fluent using finite-volume method. The initial and boundary conditions for unsteady calculation are set according to conditions of real aerodynamic tests. The coupled heat transfer for supersonic flow and elements of ramjet engine model are realized by setting of thermophysical properties of materials. The reliability testing of numerical simulation has been made to compare the results of calculations and the data of thermometric experimental tests. Findings. Numerical simulation of aerodynamic tests for ramjet engine was carried out. The agreement between the results of numerical calculations and experimental measurements for the velocity in the channel under consideration was obtained; the error was shown to be 2%. The temperature values were obtained in the area of contact of the supersonic flow with the surface of the measuring device for the external incident flow velocities for Mach numbers M = 5 ... 7. The process of heating the material in the channel that simulated the section of the engine combustion chamber was analyzed. The temperature distribution was studied depending on the position of the material layer under consideration relative to the contact zone with the flow. Value. In the course of the work, the fields of flow around the model of a ramjet engine were obtained, including the region of supersonic flow in the inner part of axisymmetric channel. The analysis of the temperature fields showed that to improve the quality of the results, it is necessary to take into account the depth of the calorimetric sensor. The obtained results will be used to estimate the time of interaction of the supersonic flow with the fuel surface required to reach the combustion temperature.
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13

Xu, Jing, Kunlin Cheng, Chaolei Dang, Yilin Wang, Zekuan Liu, Jiang Qin, and Xiaoyong Liu. "Performance comparison of liquid metal cooling system and regenerative cooling system in supersonic combustion ramjet engines." Energy 275 (July 2023): 127488. http://dx.doi.org/10.1016/j.energy.2023.127488.

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14

Zhang, Ji, Daoning Yang, Yi Wang, and Dongdong Zhang. "A Mixing Process Influenced by Wall Jet-Induced Shock Waves in Supersonic Flow." Applied Sciences 12, no. 16 (August 22, 2022): 8384. http://dx.doi.org/10.3390/app12168384.

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With the development of hypersonic air-breathing propulsion systems, such as the supersonic combustion ramjet (Scramjet) and rocket-based combined cycle (RBCC) engines, the mixing process of supersonic airstream with fuel in the engine combustor has been drawing more and more attention. Due to the compressibility effects, the mixing process in a supersonic condition is significantly inhibited. In the present paper, the novel strategy of wall-jet induced shock waves (WJISW) is put forward to realize mixing enhancement. The interaction process between WJISW and the supersonic mixing layer is researched and the enhanced-mixing mechanism is revealed, employing large eddy simulation (LES) methods. The fine vortex structures of the flow field are well captured and presented, utilizing the numerical schlieren technique. Detailed visualization results indicate that WJISW in a low frequency condition can result in the ‘region action mode’ (RAM) never reported before. The drastic dynamic behaviors including growth, deformation, and distortion in the interaction region can undoubtedly promote the mixing of upper and lower streams. The Reynolds stress distributions along the streamwise x-direction suggest that more intense fluctuations can be achieved with a low frequency WJISW. Moreover, a sharp increase in mixing layer thickness can be realized in the interaction region. The dynamic mode decomposition (DMD) analysis results show that the mixing layer evolution process is dominated by the mode induced by WJISW, which leads to the coexistence of both large- and small-scale structures in the flow field. The entrainment process corresponding to large-scale vortices and the nibbling process corresponding to small-scale vortices can obviously promote mixing enhancement. It is suggested that the present proposed strategy is a good candidate for enhanced-mixing with application to Scramjet and RBCC combustors.
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15

Srivastava, Srikant, P. N. Dwivedi, and D. M. Vinod Kumar. "3 Loop Structure based Fuel Flow Controller Design for Robust Operation of Ducted Ramjet Rocket." Defence Science Journal 71, no. 4 (July 1, 2021): 571–77. http://dx.doi.org/10.14429/dsj.71.16397.

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This paper is about the designing of 3-loop structure based fuel flow controller (FFC) for efficient operation of supersonic ramjet based propulsion system. The main objective of the control design is to vary appropriately the engine controllable parameters (throttle value area) such that commanded thrust is achieved by ramjet engine without endangering the engine stability and performance. Various factors such as combustion-intake interaction, atmospheric disturbance and other flight conditions have a significant impact on the air intake operation which lead to effect on engine performance. Due to above effect, intake un-start and buzzing phenomena can occur due to back pressure fluctuation and it disturb the intake pressure recovery and air mass flow rate. So, back pressure based extra loop introduce in 2-loop FFC design to have tight control on back pressure margin for smooth and efficient operation of air intake without need of extra hardware.
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16

Verma, Kumari Ambe, Sarken Kapayeva, Krishna Murari Pandey, and Kaushal Kumar Sharma. "The recent development of supersonic combustion ramjet engines for augmentation of the mixing performance and improvement in combustion Efficiency: A review." Materials Today: Proceedings 45 (2021): 7058–62. http://dx.doi.org/10.1016/j.matpr.2021.01.879.

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17

Ikawa, Hideo. "Rapid methodology for design and performance prediction of integrated supersonic combustion ramjet engine." Journal of Propulsion and Power 7, no. 3 (May 1991): 437–44. http://dx.doi.org/10.2514/3.23345.

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18

Smart, M. "Scramjets." Aeronautical Journal 111, no. 1124 (October 2007): 605–19. http://dx.doi.org/10.1017/s0001924000004796.

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Abstract The supersonic combustion ramjet, or scramjet, is the engine cycle most suitable for sustained hypersonic flight in the atmosphere. This article describes some of the challenges facing scramjet designers, and the methods currently used for the calculation of scramjet performance. It then reviews the HyShot 2 and Hyper-X flight programs as examples of how sub-scale flights are now being used as important steps towards the development of operational systems. Finally, it describes some recent advances in three-dimensional scramjets with application to hypersonic cruise and multi-stage access-to-space vehicles.
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19

Tam, Christopher K. W., and Fang Q. Hu. "The instability and acoustic wave modes of supersonic mixing layers inside a rectangular channel." Journal of Fluid Mechanics 203 (June 1989): 51–76. http://dx.doi.org/10.1017/s0022112089001370.

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At high supersonic convective Mach numbers the familiar Kelvin-Helmholtz instability of a thin unconfined two-dimensional shear layer becomes neutrally stable. In this paper, it is shown that when the same shear layer is put inside a rectangular channel the coupling between the motion of the shear layer and the acoustic modes of the channel produces new two-dimensional instability waves. The instability mechanism of these waves is examined. Extensive numerical computation of the properties of these new instability waves has been carried out. Based on these results two classes of these waves are identified. Some of the important characteristic features of these waves are reported in this paper. In addition to the unstable waves, a thorough analysis of the normal modes of a supersonic shear layer inside a rectangular channel reveals that there are basically two other families of neutral acoustic waves. Examples of some of the prominent characteristics of these neutral acoustic waves are also provided in this paper. The new instability waves are the dominant instabilities of a confined mixing layer at high supersonic convective Mach number. As such they are very relevant to the supersonic mixing and combustion processes inside a ramjet engine combustion chamber.
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20

Surya Narayanan, Subramanian, and Parammasivam K.M. "A review of computational studies on trapped vortex combustors for gas turbine applications." Aircraft Engineering and Aerospace Technology 95, no. 4 (October 20, 2022): 658–67. http://dx.doi.org/10.1108/aeat-12-2021-0366.

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Purpose The purpose of this paper is to comprehensively evaluate the progress in the development of trapped vortex combustors (TVCs) in the past three decades. The review aims to identify the needs, predict the scope and discuss the challenges of numerical simulations in TVCs applied to gas turbines. Design/methodology/approach TVC is an emerging combustion technology for achieving low emissions in gas turbine combustors. The overall operation of such TVCs can be on very lean mixture ratio and hence it helps in achieving high combustion efficiency and low overall emission levels. This review introduces the TVC concept and the evolution of this technology in the past three decades. Various geometries that were explored in TVC research are listed and their operating principles are explained. The review then categorically arranges the progress in computational studies applied to TVCs. Findings Analyzing extensive literature on TVCs the review discusses results of numerical simulations of various TVC geometries. Numerical simulations that were used to optimize TVC geometry and to enhance mixing are discussed. Reactive flow studies to comprehend flame stability and emission characteristics are then listed for different TVC geometries. Originality/value To the best of the authors’ knowledge, this review is the first of its kind to discuss extensively the computational progress in TVC development specific to gas turbine engines. Earlier review on TVC covers a wide variety of applications including land-based gas turbines, supersonic Ramjets, incinerators and hence compromise on the depth of analysis given to gas turbine engine applications. This review also comprehensively group the numerical studies based on geometry, flow and operating conditions.
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21

Billig, Frederick S. "Supersonic combustion ramjet missile." Journal of Propulsion and Power 11, no. 6 (November 1995): 1139–46. http://dx.doi.org/10.2514/3.23952.

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22

Veeran, Sasha, Apostolos Pesyridis, and Lionel Ganippa. "Ramjet Compression System for a Hypersonic Air Transportation Vehicle Combined Cycle Engine." Energies 11, no. 10 (September 25, 2018): 2558. http://dx.doi.org/10.3390/en11102558.

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This report assesses the performance characteristics of a ramjet compression system in the application of a hypersonic vehicle. The vehicle is required to be self-powered and perform a complete flight profile using a combination of turbojet, ramjet and scramjet propulsion systems. The ramjet has been designed to operate between Mach 2.5 to Mach 5 conditions, allowing for start-up of the scramjet engine. Multiple designs, including varying ramp configurations and turbo-ramjet combinations, were investigated to evaluate their merits and limitations. Challenges arose with attempting to maintain sufficient pressure recoveries and favourable flow characteristics into the ramjet combustor. The results provide an engine inlet design capable of propelling the vehicle between the turbojet and scramjet phase of flight, allowing for the completion of its mission profile. Compromises in the design, however, had to be made in order to allow for optimisation of other propulsion systems including the scramjet nozzle and aerodynamics of the vehicle; it was concluded that these compromises were justified as the vehicle uses the ramjet engine for a minority of the flight profile as it transitions between low supersonic to hypersonic conditions.
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Ma, Kaifu, Zijian Zhang, Yunfeng Liu, and Zonglin Jiang. "Aerodynamic principles of shock-induced combustion ramjet engines." Aerospace Science and Technology 103 (August 2020): 105901. http://dx.doi.org/10.1016/j.ast.2020.105901.

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Clauss, W., R. Sontgen, A. Feinauer, R. Guerra, and W. Waidmann. "CARS TEMPERATURE MEASUREMENTS IN A SUPERSONIC RAMJET COMBUSTION CHAMBER." International Journal of Energetic Materials and Chemical Propulsion 3, no. 1-6 (1994): 132–44. http://dx.doi.org/10.1615/intjenergeticmaterialschemprop.v3.i1-6.140.

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Cohen-Zur, Abraham, and Benveniste Natan. "Experimental Investigation of a Supersonic Combustion Solid Fuel Ramjet." Journal of Propulsion and Power 14, no. 6 (November 1998): 880–89. http://dx.doi.org/10.2514/2.5379.

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26

Brummund, U., and F. Scheel. "IMAGING OF MIXING AND COMBUSTION PROCESSES IN A SUPERSONIC COMBUSTION RAMJET CHAMBER." International Journal of Energetic Materials and Chemical Propulsion 5, no. 1-6 (2002): 762–72. http://dx.doi.org/10.1615/intjenergeticmaterialschemprop.v5.i1-6.790.

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27

Paull, A., R. J. Stalker, and D. J. Mee. "Experiments on supersonic combustion ramjet propulsion in a shock tunnel." Journal of Fluid Mechanics 296 (August 10, 1995): 159–83. http://dx.doi.org/10.1017/s0022112095002096.

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Measurements have been made of the propulsive effect of supersonic combustion ramjets incorporated into a simple axisymmetric model in a free piston shock tunnel. The nominal Mach number was 6, and the stagnation enthalpy varied from 2.8 to 8.5 MJ kg−1. A mixture of 13% silane and 87% hydrogen was used as fuel, and experiments were conducted at equivalence ratios up to approximately 0.8. The measurements involved the axial force on the model, and were made using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels. A net thrust was experienced up to a stagnation enthalpy of 3.7 MJ kg−1, but as the stagnation enthalpy increased, an increasing net drag was recorded. Pilot and static pressure measurements showed that the combustion was supersonic.The results were found to compare satisfactorily with predictions based on established theoretical models, used with some simplifying approximations. The rapid reduction of net thrust with increasing stagnation enthalpy was seen to arise from increasing precombustion temperature, showing the need to control this variable if thrust performance was to be maintained over a range of stagnation enthalpies. Both the inviscid and viscous drag were seen to be relatively insensitive to stagnation enthalpy, with the combustion chambers making a particularly significant contribution to drag. The maximum fuel specific impulse achieved in the experiments was only 175 s, but the theory indicates that there is considerable scope for improvement on this through aerodynamic design.
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Savino, Raffaele, and Giuseppe Pezzella. "Numerical analysis of supersonic combustion ramjet with upstream fuel injection." International Journal for Numerical Methods in Fluids 43, no. 2 (2003): 165–81. http://dx.doi.org/10.1002/fld.602.

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29

Crump, James E., Klaus C. Schadow, Vigor Yang, and Fred E. C. Culick. "Longitudinal combustion instabilities in ramjet engines Identification of acoustic modes." Journal of Propulsion and Power 2, no. 2 (March 1986): 105–9. http://dx.doi.org/10.2514/3.22852.

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30

Timoshenko, V. I., and V. P. Halynskyi. "Methods and programs for comprehensive calculations of supersonic flow about ramjet flying vehicles." Technical mechanics 2022, no. 2 (June 30, 2022): 3–16. http://dx.doi.org/10.15407/itm2022.02.003.

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This paper discusses the use of the authors’ fast methods and programs for the calculation of 3D supersonic flow about a flying vehicle and thermogas dynamic processes in the components of an airframe-integrated ramjet. To conduct fast comprehensive calculations, use is made of marching methods, which are two to three orders of magnitude faster than pseudoviscosity methods. 3D supersonic flows about the airframe, in the inlet section of the air intake, and in the exhaust jet are calculated using a “viscous layer” model or Godunov’s scheme for the inviscid approximation. Subsonic flows in the outlet section of the air intake and in the combustion chamber are calculated using a “narrow channel” or a quasi-one-dimensional model. The elements of the presented methods and programs that complement a previously proposed fast comprehensive model are described in more detail. A method for assigning the spatial shape of the flying vehicle surface and the ramjet duct walls is described. A simplified approach to determining the critical area of the exit nozzle in the one-dimensional approximation is proposed. The paper substantiates the advantages of marching methods over pseudoviscosity ones in the predesigning of ramjets with direct account for flow choking, which may occur in the combustion chamber or the exit nozzle. The calculated 3D flows in the individual components and the full assembly of a stylized-shape flying vehicle are presented. The main advantages of the proposed methods and programs are their comprehensiveness and fast computation speed. Their use in the calculation of 3D supersonic flow about a ramjet flying vehicle shortens the ramjet component predesigning time.
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Tunik, Yu V. "Actual schemes of supersonic ramjet on detonation combustion. Unsteady detonation waves." Physical-Chemical Kinetics in Gas Dynamics 21, no. 1 (2020): 1–12. http://dx.doi.org/10.33257/phchgd.21.1.871.

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32

Nair, Prasanth P., Amsha S, Abhilash Suryan, and Sandro Nizetic. "Investigation of flow characteristics in supersonic combustion ramjet combustor toward improvement of combustion efficiency." International Journal of Energy Research 45, no. 1 (March 6, 2020): 231–53. http://dx.doi.org/10.1002/er.5257.

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33

Levi, Reuben VR, and Amrutha Rajamani. "Study on Performance of Ramjet Intake by Changing the Cowl Angle." IOP Conference Series: Materials Science and Engineering 1258, no. 1 (October 1, 2022): 012042. http://dx.doi.org/10.1088/1757-899x/1258/1/012042.

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Most gas turbine engines require the Mach range on the engine face to be at a mild subsonic pace (around Mach 04). Therefore, for a supersonic plane with a gas turbine engine, the characteristic of the air inlet is to slow the supersonic unfastened flow to a subsonic pace and offer a matched air mass float charge to the engine. The gas turbine engine calls for the delivery of uniform excessive general stress restoration air for proper overall performance and operation because the quality of the airflow on the engine face substantially influences the overall performance of the engine, in particular, the total stress loss, which influences the engine thrust and therefore the gasoline consumption. A 2-D supersonic intake is designed for Mach No. 2.4 at an altitude of 11,000 m. The design of the mixed compression supersonic intake is done numerically, and the design with the k- turbulent module is done using CFD. Optimization of the intake is done by changing the cowl deflection in order to get maximum total pressure recovery. Results are compared to knowing the optimal design. The main advantage of mixed compression is that it gives the maximum total pressure recovery at any Mach number. The computation is done with the ANSYS software.
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Wang, Tai-Yu, Zun Cai, Bin An, Jiao-Ru Wang, Ming-Bo Sun, Chang-Hai Liang, and Zhen-Guo Wang. "Behaviors of the reacting flowfield during the spontaneous formation of ramjet mode under a supersonic inflow." Physics of Fluids 35, no. 2 (February 2023): 024105. http://dx.doi.org/10.1063/5.0135294.

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This work experimentally studied the formation process of the ramjet mode occurring in a cavity-based combustor operating at a high-enthalpy supersonic flow. The ramjet mode is featured by the phenomenon that the incoming supersonic inflow is decelerated to be subsonic before it enters the combustor, which is caused by the strong heat release under a high equivalence ratio. In the experiments, the ignition is performed after a steady fuel mass flow rate has been achieved. According to the flame behavior and the flowfield structure, the formation process of the ramjet mode can be divided into three stages, among which stage 1 (from ignition to the cavity shear-layer mode) is shortest, while stage 3 (from the lifted shear-layer mode to the ramjet mode) consumes the longest time. In stage 2, flashback occurs and shock–shock interactions are found to be strongly coupled with the local combustion which have an influence on the propagation velocity of the backpressure. A thickening boundary layer upstream of the separation shock is observed when the separation shock has interwoven with the jet-induced bow-shock. The thickening process could be extremely short (in 100 μs) before the thickened boundary layer separates, during which the propagation velocity of the backpressure can be apparently decelerated. The same phenomena shown in the supplementary experiments confirm that the thickening boundary layer and its deceleration effect on the propagation of the backpressure are not accidental but more likely to be inherent to the flashback occurring under a supersonic flow.
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35

Li, Chunlei, Yingkun Li, Weixuan Li, Liang Zhu, Xiong Chen, Shuifeng Yang, and Yan Wu. "Numerical investigation on fluid characteristics of supersonic mixing layers with splitter plate in a confined space." Journal of Physics: Conference Series 2235, no. 1 (May 1, 2022): 012063. http://dx.doi.org/10.1088/1742-6596/2235/1/012063.

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Abstract High-speed airbreathing engines has been going on for more than a century, since French engineer Reina Lauren first developed the concept of the ramjet. Kelvin-Helmholtz (K-H) instability and compressibility effects on the evolution of supersonic mixing layer growing rate have been widely investigated in experimental and computational ways. The present study proposed a quasi-DNS solver which is written in Fortran 90, the mixing process of the supersonic planar mixing layers is analyzed in detail, especially, the process of the vortex rolling up, stretching, pairing and merging are also reported in present study. It is noteworthy that the mechanisms of K-H instability are analysed in detail, and the baroclinic term has a great influence on the growth of the supersonic mixing layer in confined space.
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36

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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37

Berglund, M., E. Fedina, C. Fureby, J. Tegnér, and V. Sabel'nikov. "Finite Rate Chemistry Large-Eddy Simulation of Self-Ignition in Supersonic Combustion Ramjet." AIAA Journal 48, no. 3 (March 2010): 540–50. http://dx.doi.org/10.2514/1.43746.

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38

Chandra Bose, G., S. Thanigaiarasu, S. Elangovan, and E. Rathakrishnan. "Experimental Investigation of Shape Transition Effects on Isolator Performance." International Journal of Turbo & Jet-Engines 35, no. 4 (December 19, 2018): 331–38. http://dx.doi.org/10.1515/tjj-2015-0022.

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Abstract Isolator is a critical component in supersonic air breathing engine and it is usually situated between the inlet and the combustor of a dual-mode ramjet/scramjet engine. In the present study, shape transition effects on isolator performance have been studied by carrying out experimental investigations on square, square to circular and square to elliptical transition ducts. The length of the isolator chosen in this study is 180 mm and the cross-sectional area of 900 mm2 is maintained constant along the length for all the ducts. Experiments were carried out for isolator inlet Mach 2, using a contoured nozzle. Varying the pressure of the settling chamber varied the expansion level at the nozzle exit, which run the nozzle. The wall static pressure along the length of the isolator and the Pitot pressure at the exit plane of the isolator were measured for all the configurations. Shadowgraph technique was employed for visualizing the shock-train in the isolator. The square to circular transition isolator is found to be more efficient in achieving the static pressure rise across the isolator than the square and square to elliptical transition ducts.
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39

Zhang, Xiao Yuan, Li Zi Qin, and Yu Liu. "Effects of Chemical Reaction Models on Simulations of Scramjet Nozzle Flow." Applied Mechanics and Materials 490-491 (January 2014): 931–35. http://dx.doi.org/10.4028/www.scientific.net/amm.490-491.931.

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The chemical non-equilibrium flow of supersonic combustion ramjet (scramjet) nozzle is numerical simulated with different chemical kinetic models to research the effects on numerical results of the nozzle performance. The numerical results show that total temperature is increased due to the recombination of dissociation compositions and the combustion of the residual fuel. The effect of the combustion of the residual fuel is more obvious in this paper, and the effect to the performance of the nozzle is noticeable. The species of the compositions in the models influence the quantity of heat sending out when it get equilibrium, so the 9-species chemical kinetic models are more suitable in the simulation of the scramjet nozzle chemical non-equilibrium flows.
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40

Tretyakov, P. K., A. V. Tupikin, and V. N. Zudov. "Kerosene Combustion in a Pseudoshock with Varied Conditions at the Supersonic Ramjet Combustor Model." Combustion, Explosion, and Shock Waves 57, no. 6 (November 2021): 635–39. http://dx.doi.org/10.1134/s0010508221060010.

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41

Bogomolov, Iaroslav, and Vladimir Malinin. "DETERMINATION OF BREAKING CHARACTERISTICS IN THE PRE-COMBUSTION CHAMBER OF A COMBINED RAMJET ENGINE ON A POWDERED ALUMINUM FUEL TAKING INTO ACCOUNT AIR BRAKING IN AIR INTAKE DEVICE." Perm National Research Polytechnic University Aerospace Engineering Bulletin, no. 65 (2021): 105–11. http://dx.doi.org/10.15593/2224-9982/2021.65.11.

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The paper considers a combined ramjet engine powered by powdered aluminum fuel . The prototype was a combined solid-fuel ramjet engine. The advantages of the engine under consideration are given. An example of the combustion of finely dispersed aluminum powder is considered, from which it follows that the initial temperature of the aluminum powder will affect the stall characteristics in the pre-chamber. The following characteristics of the PAF ramjet were determined: the temperature of air stagnation in the air intake device, the temperature of the mixture of aluminum powder and the stalled air flow, the excess air ratio, and the stall characteristics in the pre-chamber taking into account the air stagnation temperatures. All parameters are de-termined for engine operating altitudes equal to 0.5, 10 and 18 km. A comparison is made of the limiting flame propagation ve-locities under standard conditions and with an incident air flow. Based on the obtained values of the characteristics, the interval of values of the coefficient of air sampling from the inlet to the pre-chamber of the ramjet engine on the PAF was determined, corresponding to the maximum possible areas of the engine operating parameters. Using the obtained values of characteristics, the combustion process of powdered aluminum in the pre-chamber of a ramjet engine takes place without flame blowout, which excludes unstable operation of the propulsion system as a whole. The use of powdered aluminum makes it possible to regulate the thrust in a wide range of values, and a high initial temperature of the air entering the pre-chamber for repeated switching on and off of the engine. Based on the available data, the type of engines under consideration is suitable for combat missiles of various classes, but the most suitable for aircraft-based missiles.
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42

Du, Zhao-bo, Wei Huang, and Li Yan. "Investigation on gaseous jet in forebody/inlet for shock-induced combustion ramjet (shcramjet) engines." Acta Astronautica 152 (November 2018): 262–74. http://dx.doi.org/10.1016/j.actaastro.2018.08.030.

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43

Langston, Lee S. "Detonation Gas Turbines." Mechanical Engineering 135, no. 12 (December 1, 2013): 50–54. http://dx.doi.org/10.1115/1.2013-dec-4.

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This article focuses on various technical and functional aspects of detonation gas turbines. Detonation combustion involves a supersonic flow, with the chemical reaction front accelerating, driving a shock wave system in its advancement. In the 1990s, detonation-based power concepts began with pulse detonation engines (PDEs), and have now moved into the continuous detonation mode, termed rotating detonation engines (RDEs). Modern gas turbine combustors are compact, robust, tolerant of a wide variety of fuels, and provide the highest combustion intensities. The single-spool RDE gas turbine is represented by a detonation cycle, which accounts for the supersonic features of the heat addition, starting at station 2.5′. Continued research and development by the RDE technical community is needed to see if the promise of improved performance and downsized turbomachinery for a detonation cycle is real.
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44

Pismennyi, V. L. "Hyper Afterburner Jet Engines." Proceedings of Higher Educational Institutions. Маchine Building, no. 01 (718) (January 2020): 51–62. http://dx.doi.org/10.18698/0536-1044-2020-1-51-62.

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This paper introduces a thrust augmentation method for super- and hypersonic jet engines by means of applying water at the engine intake. This method expands the use of jet engines with subsonic combustion, allowing velocities up to Mach 8 and altitude up to 45 km. At velocities higher than 3–4 Mach, stagnation temperature of the air is getting higher than the critical temperature of water, which makes the existence of water at the gas turbine engine intake impossible. Water vapour as a working medium of a jet engine creates the so-called inner thermodynamic circle. This phenomenon defines the physics of the thrust augmentation method proposed. The author discusses three variants of hyper afterburner application: hyper afterburner turbojet, hyper afterburner ramjet, and hyper afterburner turbo ejecting engine. The presented basic specifications of the hyper afterburner engines qualitatively differ from those of their prototypes (engines without the hyper afterburner thrust augmentation function). The proposed thrust augmentation method of jet engines is of a special interest for the aerospace field, particularly, for creating air launch systems. It is shown that the application of hyper afterburner in turbo ejecting engines can increase velocity and altitude of the launch aircraft up to Mach 7 and 40 km respectively, thus opening new avenues in space exploration.
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45

Sokolova, E. I., E. S. Studennikov, and O. G. Chelebyan. "Estimating the applicability of kinetic schemes in hydrogen combustion simulation in combustion chambers of aircraft engines." Journal of Physics: Conference Series 2057, no. 1 (October 1, 2021): 012068. http://dx.doi.org/10.1088/1742-6596/2057/1/012068.

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Abstract This paper presents the results of numerical simulation of hydrogen combustion in a supersonic flow of an oxidizing medium in a model combustion chamber using various models of chemical kinetics. The best scheme, which most accurately describes the combustion processes, is revealed. Comparison of the calculated distribution of molar fractions with experimental data is carried out, and relative deviations for the piloted mode of operation of the chamber are obtained.
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46

da Silva Guimarães, Jefte, Valéria Serrano Faillace Oliveira Leite, Marco Antonio Sala Minucci, and Dermeval Carinhana. "Case study of the additive manufacturing application in the supersonic flow researches." Rapid Prototyping Journal 27, no. 8 (August 23, 2021): 1480–88. http://dx.doi.org/10.1108/rpj-05-2020-0105.

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Purpose The purpose of this paper is to demonstrate the aerodynamic behavior of a supersonic combustion test bench (SCTB) components, as the transition piece and the combustor of a scramjet (supersonic combustion ramjet), manufactured by 3D printing or additive manufacturing (AM). Design/methodology/approach For the dimensional and structural analysis of the manufactured models, a portable 3D scanner was used to generate the mesh of its dimensions, and to compare them before and after the experiments, a roughness measuring system was also used to verify the roughness inside the models before and after the tests, as roughness is an important parameter because it directly affects the boundary layer. For the visualization of the flow, the non-intrusive schlieren optical technique was used. Findings The experiments were carried out on the SCBT for Mach 2 flows, using the manufactured prototypes and showed that there was no structural and dimensional change of the model after the test batteries. It was found that the roughness presented by the material did not affect the quality of the flow generated. This shows that the investigated material can also be applied in experiments with supersonic flow. Originality/value This paper presents that it is possible to use in ground test facilities, for the studies of supersonic flow (in cold condition), pieces and models manufactured by 3D printing without affecting the quality of the flow generated during the experiments. This study presents a new perspective to approach AM applied in the studies of supersonic flows.
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47

Nayal, Sanchayata, Sampada Lamb, Devabrata Sahoo, N. V. Raghavendra, and S. A. I. Bellary. "Computational analysis of novel cavity-based flameholder designs for supersonic combustion engines." IOP Conference Series: Materials Science and Engineering 814 (June 19, 2020): 012018. http://dx.doi.org/10.1088/1757-899x/814/1/012018.

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48

Wolański, P. "RDE research and development in Poland." Shock Waves 31, no. 7 (October 2021): 623–36. http://dx.doi.org/10.1007/s00193-021-01038-2.

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AbstractA very short survey of research conducted in Poland on the development of the rotating detonation engine (RDE) is presented. Initial studies conducted in cooperation with Japanese partners lead to development of a joint patent on RDE. Then, an intensive basic and applied research was started at the Institute of Heat Engineering of the Warsaw University of Technology. One of the first achievements was the demonstration of performance of the rocket engine with an aerospike nozzle utilizing continuously rotating detonation (CRD), and research was directed into development of a small turbofan engine utilizing such a combustion regime. These activities promoted international cooperation and stimulated RDE development not only in Poland but also in other countries. A research directed to measure and calculate flow parameters as well as to analyze the use of liquid fuels was conducted. In the Institute of Aviation in Warsaw, research on the application of the CRD to turbine engines as well as rocket, ramjet, and combined cycle engines was carried out. In the paper, a special emphasis is given to international cooperation in this area with partners from many countries engaged in the development of the pressure gain combustion to propulsion systems.
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Jia, Bingyue, Yining Zhang, Hao Meng, Fanxiao Meng, Hu Pan, and Yanji Hong. "Experimental Study on the Propagation Characteristics of Rotating Detonation Wave with Liquid Hydrocarbon/High-Enthalpy Air Mixture." Aerospace 10, no. 8 (July 31, 2023): 682. http://dx.doi.org/10.3390/aerospace10080682.

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Rotating detonation engines (RDEs) are a promising propulsion technology featuring high thermal efficiency and a simple structure. To adapt the practical engineering applications of ramjet RDEs, rotating detonation combustion using a liquid hydrocarbon and pure air mixture will be required. This paper presents an experimental study on the propagation characteristics of rotating detonation waves with a liquid hydrocarbon and high-enthalpy air mixture in a hollow cylindrical chamber. The parameters, such as the equivalence ratio and inlet mass flux, are considered in this experiment. The frequency and the propagation velocity of rotating detonation combustion are analyzed under typical operations. The experimental results show that the peak pressure and propagation velocity of the rotating detonation wave are close to the C-J theoretical values under the inlet mass flux of 400 kg/(m2s). Both the propagation velocity and peak pressure of the rotating detonation wave decrease as the mass flux and equivalence ratio are reduced while the number of detonation wavefronts increases. Detonation wave instability tends to occur when the inlet mass flux decreases. There is a transition progress from thermo-acoustic combustion to rotating detonation combustion in the experiment under the condition of mass flux 350 kg/(m2s) and the equivalent ratio 0.8. The static pressure in the chamber is higher during detonation combustion than during thermo-acoustic combustion. These experimental results provide evidence that rotating detonation waves have the potential to significantly improve propulsion performance. The findings can serve as a valuable reference for the practical engineering application of rotating detonation engines.
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Li, Zhijie, Jie Pan, Wei Li, Xiangting Wang, Haiqiao Wei, and Jiaying Pan. "New Insights into Abnormal Combustion Phenomena Induced by Diesel Spray-Wall Impingement under Engine-Relevant Conditions." Energies 15, no. 8 (April 17, 2022): 2941. http://dx.doi.org/10.3390/en15082941.

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High altitude and low temperature is the common extreme environment for internal combustion engines. Under such operating conditions, heavy-duty diesel engines often suffer from serious abnormal combustion, such as knocking combustion, which results in piston crown breakdown and cylinder head erosion. Spray-wall impingement and pool fires are considered potential causes; however, the detailed mechanism remains poorly understood owing to the lack of research data. In this study, for the first time, the destructive abnormal combustion induced by diesel spray-wall impingement was identified using an optical rapid compression machine under engine-relevant conditions at high altitudes. Combining instantaneous pressure and temperature measurements with simultaneously recorded high-speed photography gives useful insights into understanding the detailed combustion processes. The experimental results show that depending on the extent of diesel spray-wall impingement, supersonic detonation-like reaction fronts featuring bright luminosity can be observed. The propagation of these reaction fronts in-cylinder results in severe pressure oscillations with an amplitude approaching hundreds of atmospheres, which is like the super-knock events in boosted direct-injection spark-ignition engines. Further parametric analysis indicates that the interplay between the diffusion combustion controlled by diesel spray and the premixed combustion dominated by attached film evaporation results in the formation of abnormal combustion. Destructive reaction fronts tend to occur at a prolonged ignition delay time, which facilitates the mixing between diesel evaporation and hot air.
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