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Artykuły w czasopismach na temat „Scramjet engine”

Autor: Grafiati

Data publikacji: 6 września 2023

Data aktualizacji: 7 września 2023

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1

Jiang, Baohong. "Comprehensive Analysis of the Advanced Technologies for Scramjet". Highlights in Science, Engineering and Technology 43 (14.04.2023): 137–49. http://dx.doi.org/10.54097/hset.v43i.7413.

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Scramjet is a kind of aspirated engine, where oxygen in the atmosphere is used as oxidant to react with fuel in fuel bunker. Structural components are used in the scramjet to generate shock waves at high speed to compress the high-speed air flow, and realize the deceleration and pressurization of the air flow, which is different from engines where air compressors are used. Technologies related to the scramjet power/fuel are presented, and the features related to this kind of engines are highlighted in this paper. The development process of the scramjets in the application field both home and abroad is overviewed. The problems involved with scramjets in hypersonic vehicle application, combined cycle power system, design of thermal protection structures and high temperature materials are discussed. The critical technologies of scramjets, i.e., tail nozzle, combustion chamber, air inlet, fuel selection etc. are identified. The features of hydrocarbon fuel and its application in hypersonic vehicles are summarized. And the progress of research of the relevant technologies and personal prospects for scramjets are briefly described.

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2

Smart, M. "Scramjets". Aeronautical Journal 111, nr 1124 (październik 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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3

Jin, Liang, Xian Yu Wu, Jing Lei, Li Yan, Wei Huang i Jun Liu. "CFD Analysis of a Hypersonic Vehicle Powered by Triple-Module Scramjets". Applied Mechanics and Materials 390 (sierpień 2013): 71–75. http://dx.doi.org/10.4028/www.scientific.net/amm.390.71.

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A numerical investigation has been carried out to study the longitudinal performance of a hypersonic airbreathing vehicle with highly integrated triple-module scramjets. CFD-Fastran is used to evaluate the aerodynamic performance of the vehicle at inlet-open scramjet unpowered mode, and a chemical reacting code ChemTur3D has been built to evaluate the propulsion performance of the triple-module engines at scramjet powered mode. The flow conditions for the calculations include variations of angle of attack at Mach 5.85 test point. The wall pressure and surface friction are integrated to calculate drag, lift and pitching moment coefficients to predict the combined aeropropulsive force and moment characteristics during engine operation. Finally, numerical results is compared with available ground test data to assess solution accuracy, and a preflight aerodynamic database of the vehicle could be built for the hypersonic flight experiments.

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4

Daren, Y., C. Tao i B. Wen. "An idea of distributed parameter control for scramjet engines". Aeronautical Journal 111, nr 1126 (grudzień 2007): 787–96. http://dx.doi.org/10.1017/s0001924000001901.

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AbstractScramjet engines are used under extreme temperatures and with wide range of Mach numbers from 3 to 8 or higher and have shown different control properties from other airbreathing engines. New control problems involving distributed parameter control have been found concerning investigations of the control of scramjet engines whose physical states are spatially interacted and whose governing equations are partial differential equations. The work of this paper is based on the application of distributed parameter control conception to study the control problems of scramjet engines with the aim of achieving the desirable design properties and increasing control reliability. A new control idea based on shape control theory is put forward to realise the distributed parameter control of scramjet engines with the preconditions of proper space dimension and frequency-domain simplification. Simulation results and theoretic analysis for an axisymmetric, wall-injection scramjet engine show the feasibility and validity of the control idea.

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5

Cheng, Feng, Shuo Tang, Dong Zhang i Yi Li. "Quasi-One-Dimensional Modeling and Analysis of RBCC Dual-Mode Scramjet Engine". International Journal of Turbo & Jet-Engines 36, nr 2 (27.05.2019): 195–206. http://dx.doi.org/10.1515/tjj-2017-0055.

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Abstract The quasi-one-dimensional method for the dual-mode scramjet (DMR) of the hypersonic RBCC powered vehicle was simplified in most of open researches. Furthermore, these simplified method can not fully capture the processes of wall heat transfer, changes in the boundary layer and the ratio of specific heat and the transonic flow in the reacting flow. Addressing this problem, we establish the models for processing core flow area, transonic flow and pre-combustion shock train (PCST) based on the governing equations for quasi-one-dimensional flow and certain assumptions. Thus the quasi-one-dimensional model of dual-mode scramjet engine that incorporates the changes in wall heat transfer and in the ratio of specific heat is built. Then, the reliability and accuracy of the model are assessed qualitatively and quantitatively by experiment and CFD numerical simulation. There is a high agreement between the theoretical calculations and the results of experimental data and CFD numerical simulation. This work expands the application scope and increases the reliability of quasi-one-dimensional model of dual-mode scramjet engine in RBCC engine. The results shed new light on the preliminary performance assessment and engineering application of dual-mode scramjet engine.

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6

Ji, Zifei, Huiqiang Zhang i 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, nr 13 (20.02.2019): 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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7

Relangi, Naresh, Lakshmi Narayana Phaneendra Peri, Caio Henrique Franco Levi Domingos, Amalia Fossella, Julia Meria Leite Henriques i Antonella Ingenito. "Design of Supersonic and Hybrid engine based Advanced Rocket (SHAR)". IOP Conference Series: Materials Science and Engineering 1226, nr 1 (1.02.2022): 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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8

Veeran, Sasha, Apostolos Pesyridis i Lionel Ganippa. "Ramjet Compression System for a Hypersonic Air Transportation Vehicle Combined Cycle Engine". Energies 11, nr 10 (25.09.2018): 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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9

Fan, Fa Qing, i Pei Yong Wang. "Investigation of the Non-Equilibrium Flow Phenomena in the Boundary Layer of the Scramjet Engine". Applied Mechanics and Materials 284-287 (styczeń 2013): 795–99. http://dx.doi.org/10.4028/www.scientific.net/amm.284-287.795.

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High-speed and high-temperature are the characteristics of the flow field in scramjet engine; the regular non-slip wall boundary condition requires zero speed at wall; in the same time, the material temperature limit does not allow high wall temperature; therefore the velocity gradient and temperature gradient in the engine boundary layer are huge. If these gradients are too large, the traditional assumption of the local thermal equilibrium in the fluid will fail, the Navier-Stokes equations are no longer valid in the boundary layer. For the first time, the non-equilibrium flow phenomena in Scramjet engine is studied here. Appropriate turbulence model and fine grid are used to analyze the turbulent boundary layer of the Hyshot scramjet engine with three different operating conditions. The result of the CFD simulation shows that the local Knudsen number in the engine boundary layer is greater than the critical value with the operating conditions 40Km/Ma8 and 30Km/Ma8; they are non-equilibrium flow and the Navier-Stokes equations fails. Special treatment of the boundary conditions are needed for these kinds of flow. With the operating condition of 20Km/Ma6, the local thermal equilibrium condition is observed and conventional CFD method is valid.

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10

Zhang, Fan, Huiqiang Zhang i Bing Wang. "Conceptual study of a dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 232, nr 5 (1.05.2017): 944–57. http://dx.doi.org/10.1177/0954410017703148.

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The liquid oxygen/methane staged cycle liquid-rocket engine is one of the most potential rocket engines in the future for its higher performance, higher fuel density and reusable capacity. Two working states of this liquid-rocket engine named as full-load state and half-load state are defined in this paper. Based on this liquid-rocket engine, a dual-rocket-based-combined-cycle propulsion system with liquid oxygen /air/methane as propellants is therefore proposed. The dual-rocket-based-combined-cycle system has then five working modes: the hybrid mode, pure ejector mode, ramjet mode, scramjet mode and pure rocket mode. In hybrid mode, the booster and ejector rockets driven by the full-load liquid-rocket engine work together with the purpose of reducing thrust demand on ejector rocket. In scramjet mode, the fuel-rich burned hot gas generated by the half-load liquid-rocket engine is used as fuel, which is helpful to reduce the technical difficulty of scramjet in hypersonic speed. The five working modes of dual-rocket-based-combined-cycle are highly integrated based on the full- or half-load state of the liquid oxygen/methane staged cycle liquid-rocket engine, and the unified single type fuel of liquid methane is adopted for the whole modes. Then a preliminary design of a horizontal takeoff two-stage-to-orbit launch vehicle is conducted based on the dual-rocket-based-combined-cycle propulsion system. Under an averaged baseline thrust and specific impulse, the launch trajectory to reach a low Earth orbit at 100 km is optimized via the pseudo-spectral method subject to maximizing the payload mass. It is shown that the two-stage-to-orbit vehicle based on the dual-rocket-based-combined-cycle can achieve the payload mass fraction of 0.0469 and 0.0576 for polar mission and equatorial mission, respectively. Conclusively, insights gained in this paper can be usefully applied to a more detailed design of the dual-rocket-based-combined-cycle powered two-stage-to-orbit launch vehicle.

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11

Liu, Xiaonan, i Yufei Ma. "Tunable Diode Laser Absorption Spectroscopy Based Temperature Measurement with a Single Diode Laser Near 1.4 μm". Sensors 22, nr 16 (15.08.2022): 6095. http://dx.doi.org/10.3390/s22166095.

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The rapidly changing and wide dynamic range of combustion temperature in scramjet engines presents a major challenge to existing test techniques. Tunable diode laser absorption spectroscopy (TDLAS) based temperature measurement has the advantages of high sensitivity, fast response, and compact structure. In this invited paper, a temperature measurement method based on the TDLAS technique with a single diode laser was demonstrated. A continuous-wave (CW), distributed feedback (DFB) diode laser with an emission wavelength near 1.4 μm was used for temperature measurement, which could cover two water vapor (H2O) absorption lines located at 7153.749 cm−1 and 7154.354 cm−1 simultaneously. The output wavelength of the diode laser was calibrated according to the two absorption peaks in the time domain. Using this strategy, the TDLAS system has the advantageous of immunization to laser wavelength shift, simple system structure, reduced cost, and increased system robustness. The line intensity of the two target absorption lines under room temperature was about one-thousandth of that under high temperature, which avoided the measuring error caused by H2O in the environment. The system was tested on a McKenna flat flame burner and a scramjet model engine, respectively. It was found that, compared to the results measured by CARS technique and theoretical calculation, this TDLAS system had less than 4% temperature error when the McKenna flat flame burner was used. When a scramjet model engine was adopted, the measured results showed that such TDLAS system had an excellent dynamic range and fast response. The TDLAS system reported here could be used in real engine in the future.

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12

Atashi-Abkenar, M. A. "Study on the Effect of Two Uncertainty Parameters on Scramjet Engine Using Monte Carlo Simulation". International Journal of Mathematical Models and Methods in Applied Sciences 16 (13.05.2022): 89–94. http://dx.doi.org/10.46300/9101.2022.16.16.

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Today, aerospace engines are developing rapidly, these engines are divided into two groups of gas turbines and without gas turbines. In this thesis, the thermodynamic performance of the scramjet engine is examined. This study is carried out with consideration of uncertainty parameters. Two parameters of the combustion chamber efficiency and heating value of fuel are considered as uncertainty parameters. Using Monte Carlo numerical simulation method, the functional curves of the scramjet engine were investigated, and Analysis is done. According to the use of uncertainty parameters, first, a brief explanation of the uncertainty illustrates according to calculate using their functions and the Monte Carlo method. Also, the uncertain effects on the functional charts are analyzed considering the variable taking into account each of the uncertainty parameters. According to the obtained results, it was determined that the uncertain effect of the combustion chamber is negligible compared to the heating value of the fuel, the number of different points of 100,200 and 300 is similar to each other, and according to the extracted functional charts With regard to the uncertainties, it was observed that the least compression efficiency and special fuel consumption would have the greatest effect from the uncertainties.

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13

Mitani, Tohru, Sadatake Tomioka, Takeshi Kanda, Koichiro Tani, Nobuo Chinzei i Toshinori Kouchi. "Scramjet Engine Performance Attained in RJTF Testing". JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 52, nr 600 (2004): 1–9. http://dx.doi.org/10.2322/jjsass.52.1.

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Mitnai, Tohru, Nobuo Chinzei i Nobuyuki Yatsuyanagi. "906 Scramjet Engine Testing in NAL-KPL". Proceedings of the Fluids engineering conference 2001 (2001): 118. http://dx.doi.org/10.1299/jsmefed.2001.118.

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15

Wang, Xiaodong, i Jialing Le. "Computations of inlet/isolator for SCRAMjet engine". Journal of Thermal Science 9, nr 4 (grudzień 2000): 334–38. http://dx.doi.org/10.1007/s11630-000-0073-3.

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16

Tsujikawa, Y., i M. Nagaoka. "Determination of Cycle Configuration of Gas Turbines and Aircraft Engines by an Optimization Procedure". Journal of Engineering for Gas Turbines and Power 113, nr 1 (1.01.1991): 100–105. http://dx.doi.org/10.1115/1.2906515.

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This paper is devoted to the analyses and optimization of simple and sophisticated cycles, particularly for various gas turbine engines and aero-engines (including the scramjet engine) to achieve maximum performance. The optimization of such criteria as thermal efficiency, specific output, and total performance for gas turbine engines, and overall efficiency, nondimensional thrust, and specific impulse for aero-engines has been performed by the optimization procedure with the multiplier method. Comparison of results with analytical solutions establishes the validity of the optimization procedure.

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17

Alhassani, Abdulla Khamis, Mohanad Tarek Mohamed, Mohammed Fares i Sharul Sham Dol. "Shock Waves Analysis of the Novel Intake Design System for a Scramjet Propulsion". WSEAS TRANSACTIONS ON SYSTEMS 20 (15.04.2021): 67–75. http://dx.doi.org/10.37394/23202.2021.20.9.

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The supersonic combustion scramjet in the inlet applies the shock waves compression mechanism tosubstitute the actual compressor from a gas turbine engine. The scramjet works with combustion of fuel throughthe air stream in supersonic condition at least with Mach 5. Novel design of a scramjet intake system was madewith variations in the angle of the fins and entrance width. The best combination of diameter and inclinationangle was 1.75 m and 15 degrees, respectively. The findings were able to increase the oblique shock waveinteractions and supplicate effective combustion and reduce pressure losses for the effective application ofscramjet system, which can be significant for aerospace industry.

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18

Kerrebrock, Jack L. "Some readily quantifiable aspects of scramjet engine performance". Journal of Propulsion and Power 8, nr 5 (wrzesień 1992): 1116–22. http://dx.doi.org/10.2514/3.23600.

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YATSUNAMI, Tomomi, Tohru MITANI, Kan KOBAYASHI i Goro MASUYA. "Effects of Residual Radicals on Scramjet Engine Testing." Journal of the Japan Society for Aeronautical and Space Sciences 48, nr 563 (2000): 411–17. http://dx.doi.org/10.2322/jjsass.48.411.

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20

Castrogiovanni, Anthony. "Review of "The Scramjet Engine, Processes and Characteristics"". AIAA Journal 48, nr 9 (wrzesień 2010): 2173–74. http://dx.doi.org/10.2514/1.50210.

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Chang, Juntao, Lei Wang, Wen Bao, Qinchun Yang i Jiang Qin. "Experimental Investigation of Hysteresis Phenomenon for Scramjet Engine". AIAA Journal 52, nr 2 (luty 2014): 447–51. http://dx.doi.org/10.2514/1.j052505.

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Kanda, Takeshi, Tetsuo Hiraiwa, Tohru Mitani, Sadatake Tomioka i Nobuo Chinzei. "Mach 6 Testing of a Scramjet Engine Model". Journal of Propulsion and Power 13, nr 4 (lipiec 1997): 543–51. http://dx.doi.org/10.2514/2.5201.

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Kanda, Takeshi, Tetsuji Sunami, Sadatake Tomioka, Kouichiro Tani i Tohru Mitani. "Mach 8 Testing of a Scramjet Engine Model". Journal of Propulsion and Power 17, nr 1 (styczeń 2001): 132–38. http://dx.doi.org/10.2514/2.5718.

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SIMONE, Domenico, i Claudio BRUNO. "Modeling LiH Combustion in Solid Fuelled Scramjet Engine". TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 8, ists27 (2010): Pa_47—Pa_56. http://dx.doi.org/10.2322/tastj.8.pa_47.

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Kim, Jae Won, i Oh Joon Kwon. "Modeling of incomplete combustion in a scramjet engine". Aerospace Science and Technology 78 (lipiec 2018): 397–402. http://dx.doi.org/10.1016/j.ast.2018.04.044.

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Verma, Ansh. "Ameliorative Study of a Scramjet Engine by Regenerative Cooing using Finite Element". International Journal of Engineering and Technology 2, nr 6 (2010): 592–97. http://dx.doi.org/10.7763/ijet.2010.v2.187.

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Hu, Jichao, Juntao Chang, Lei Wang, Shibin Cao i Wen Bao. "Unstart Coupling Mechanism Analysis of Multiple-Modules Hypersonic Inlet". Scientific World Journal 2013 (2013): 1–10. http://dx.doi.org/10.1155/2013/254376.

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The combination of multiplemodules in parallel manner is an important way to achieve the much higher thrust of scramjet engine. For the multiple-modules scramjet engine, when inlet unstarted oscillatory flow appears in a single-module engine due to high backpressure, how to interact with each module by massflow spillage, and whether inlet unstart occurs in other modules are important issues. The unstarted flowfield and coupling characteristic for a three-module hypersonic inlet caused by center module II and side module III were, conducted respectively. The results indicate that the other two hypersonic inlets are forced into unstarted flow when unstarted phenomenon appears on a single-module hypersonic inlet due to high backpressure, and the reversed flow in the isolator dominates the formation, expansion, shrinkage, and disappearance of the vortexes, and thus, it is the major factor of unstart coupling of multiple-modules hypersonic inlet. The coupling effect among multiple modules makes hypersonic inlet be more likely unstarted.

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Zhang, Linqing, Juntao Chang, Wenxiang Cai, Hui Sun i Yingkun Li. "A Preliminary Research on Combustion Characteristics of a Novel-Type Scramjet Combustor". International Journal of Aerospace Engineering 2022 (30.12.2022): 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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Wang, Shirui. "Analysis and Possible Improvements of Scramjet Engines: The Effective Thrust and the Combustion Stability Problems". Theoretical and Natural Science 5, nr 1 (25.05.2023): 204–9. http://dx.doi.org/10.54254/2753-8818/5/20230403.

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The scramjet is the representative of the future jet engine, and it has irreplaceable advantages in terms of working performance and service life. Although it has been developed more than fifty years, there are many problems in the driven system of the scramjet engine. This paper mainly focuses on the problems of the effective thrust and the combustion stability. The research shows that how the pressure ratio affects the positive thrust and the combustion stability. Furthermore, this paper also shows that the elements which can affect the pressure ratio. The lower inlet air temperature is beneficial to enhance the thrust. According to the relevant research which apply the CJ detonation theorem to find the possible reason for the combustion instability. For the possible improvements, the research found that the engine thrust at different speed and the allowable maximum speed. The type of fuel and the reactivity of the fuel can influence the thrust.

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Xiong, Yuefei, Jiang Qin, Kunlin Cheng, Silong Zhang i Yu Feng. "Quasi-One-Dimensional Model of Hydrocarbon-Fueled Scramjet Combustor Coupled with Regenerative Cooling". International Journal of Aerospace Engineering 2022 (8.08.2022): 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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Yang, Pengnian, Zhixun Xia, Likun Ma, Binbin Chen, Yunchao Feng, Chaolong Li i Libei Zhao. "Direct-Connect Test of Solid Scramjet with Symmetrical Structure". Energies 14, nr 17 (6.09.2021): 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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Yang, Qingchun, Juntao Chang i Wen Bao. "Richtmyer-Meshkov Instability Induced Mixing Enhancement in the Scramjet Combustor with a Central Strut". Advances in Mechanical Engineering 6 (1.01.2014): 614189. http://dx.doi.org/10.1155/2014/614189.

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Experimental and numerical study of Richtmyer-Meshkov instability (RMI) induced mixing enhancement has been conducted in a liquid-fueled scramjet engine with a central strut. To generate the RMI in the scramjet engine, transverse high temperature jets are employed downstream the strut injector. Compared to the transverse ordinary temperature jet, the jet penetration into the supersonic airstream of high temperature jet increases by 60%. The numerical results indicate that the RMI phenomenon markedly enhances the mixing efficiency (up to 43%), which is necessary to initiate the chemical reactions. Ground experiments were carried out in the combustor, which verify the numerical method from the perspective of wall pressures of the combustor. In particular, the experiment results indicate that the RMI can benefit flame-holding due to the mixing enhancement.

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Sarosh, Ali, Dong Yun Feng i Muhammad Adnan. "An Aerothermodynamic Design Approach for Scramjet Combustors and Comparative Performance of Low-Efficiency Systems". Applied Mechanics and Materials 110-116 (październik 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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Wu, Xianju, i Zhijun Wei. "Comparison of Dual-Combustion Ramjet and Scramjet Performances Considering Combustion Efficiency". Applied Sciences 13, nr 1 (29.12.2022): 480. http://dx.doi.org/10.3390/app13010480.

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The performances of a dual-combustion ramjet (DCR) and a scramjet were compared via computational fluid dynamics numerical simulation to provide theoretical guidance for engine selection for a hypersonic vehicle. Kerosene, C12H23, with an equivalence ratio of 0.8, was employed as the fuel, and the reactive flow was modeled using six-species and four-step chemistry. The results show that the DCR has a central combustion mode, which has a smaller temperature gradient and more uniform heat release, resulting in higher combustion efficiency, compared to the near-wall combustion mode of the scramjet. The total pressure recovery coefficient of scramjet is 0.9% lower than that of DCR under the Ma6 condition, but 5.6% higher than that of DCR under the Ma7 condition. The combustion efficiency of DCR is 35.6% and 25.4% higher than that of the scramjet under Ma6 and Ma7 conditions, respectively. The decrease in the combustion efficiency of the DCR is caused by the increase in the dissociation rate of CO2 into CO with the increase in temperature. The performance of DCR is better than that of scramjet under both conditions. However, the performance advantage of DCR decreases as the Mach number increases. Specifically, under the conditions of Ma6 and Ma7, the specific impulse or specific thrust of DCR was 2.67 times and 1.51 times that of scramjet, respectively.

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Chang, Juntao, Lei Wang i Wen Bao. "Mathematical modeling and characteristic analysis of scramjet buzz". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 228, nr 13 (29.01.2014): 2542–52. http://dx.doi.org/10.1177/0954410014521055.

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Buzz is an important issue for a scramjet engine. A mathematical model of buzz oscillations is necessary for control system design. Control-oriented models of hypersonic vehicle propulsion systems require a reduced-order model that is accurate to some extent but requires less than a few seconds of computational time. To achieve this goal, a reduced-order model of buzz oscillations for a scramjet engine is built by introducing the modeling idea of Moore–Greitzed model for compressors. The introduction of characteristic lines avoids the complex interactions in hypersonic inlet, such as shock–shock interactions and shock–boundary layer interaction. And the inlet characteristics are obtained from the pressure signal of combustor. Based on the established buzz model, we can predict the inlet performance, characterize the stability margin of inlet, reflect the oscillatory characteristics of inlet buzz including the dominant amplitude and frequency and describe the transition process of inlet buzz.

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Xu, Hongyang, Yonghua Fan, Xi Tong i Jie Yan. "Designing Control System of Hypersonic Vehicle with Dynamic Pressure Constraints Considered". Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 37, nr 1 (luty 2019): 41–47. http://dx.doi.org/10.1051/jnwpu/20193710041.

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An airbreathing hypersonic vehicle(AHV) generally adopts a scramjet engine as its propulsion, which needs strict conditions of flight dynamic pressure. A new dynamic pressure control system is presented, in which the AHV autopilot directly uses dynamic pressure to track the dynamic pressure command. Firstly, the dynamic pressure model of the AHV is established and the state equations of dynamic pressure control are given. Then, the dynamic pressure control system is implemented using the LQR optimal control theory. The dynamic pressure error is augmented in order to add an integral control to track the dynamic pressure with zero steady error. The structure of the dynamic pressure control system is also obtained. The simulation results show that the dynamic pressure control system has a good performance for guaranteeing the dynamic pressure of the scramjet engine with the disturbance of drag and thrust considered.

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Han, Seoeum, Sangyoon Lee i Bok Jik Lee. "Numerical Analysis of Thermochemical Nonequilibrium Flows in a Model Scramjet Engine". Energies 13, nr 3 (31.01.2020): 606. http://dx.doi.org/10.3390/en13030606.

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This numerical study was conducted to investigate the flow properties in a model scramjet configuration of the experiment in the T4 shock tunnel. In most numerical simulations of flows in shock tunnels, the inflow conditions in the test section are determined by assuming the thermal equilibrium of the gas. To define the inflow conditions in the test section, the numerical simulation of the nozzle flow with the given nozzle reservoir conditions from the experiment is conducted by a thermochemical nonequilibrium computational fluid dynamics (CFD) solver. Both two-dimensional (2D) and three-dimensional (3D) numerical simulations of the flow in a model scramjet were conducted without fuel injection. Simulations were performed for two types of inflow conditions: one for thermochemical nonequilibrium states obtained from the present nozzle simulation and the other for the data available using the thermal equilibrium and chemical nonequilibrium assumptions. The four results demonstrate the significance of the modelling approach for choosing between 2D or 3D, and thermal equilibrium or nonequilibrium.

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Gao, Jin, Ziyi Kang, Weiheng Sun, Youyin Wang, Junlong Zhang i Wen Bao. "Feasibility and Performance Analysis of High-Energy-Density Hydrocarbon-Fueled Turboexpander Engine". Aerospace 10, nr 9 (25.08.2023): 753. http://dx.doi.org/10.3390/aerospace10090753.

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With the in-depth research on hypersonic aerodynamics and hypersonic propulsion technology, humans are growing closer to space travel. Recent studies have shown that the pre-cooled air-turborocket (ATR) or turboexpander engines are some of the potential propulsion methods for reusable space vehicles and single stage-to-orbit (SSTO) missions because they have a high specific impulse at low Mach numbers, which can overcome the problem of the “thrust gap” in turbine-based combined-cycle (TBCC) engines. The ATR engine needs an additional oxidizing agent and the turboexpander engine usually uses hydrogen as fuel, which has low energy density and poor safety. To address this problem, this paper proposed a high-energy-density (HED) hydrocarbon-fueled turboexpander engine, and its feasibility has been proven through a simplified thermodynamic model. Through detailed thermodynamic analysis based on the energy and pressure balance, this paper analyzed the performance characteristics of the engine to evaluate its capacity to work in a wide speed range at low Mach numbers. The results show that the endothermic hydrocarbon-fueled turboexpander engine has good specific impulse in Mach 0∼4 at an equivalence ratio of 0.7∼1.3, and the turboexpander engine can be combined with the dual-mode scramjet and become an efficient acceleration method for SSTO missions and the reusable spacecraft.

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Yang, Inyoung, Yang-Ji Lee, Young-Moon Kim i Kyung-Jae Lee. "Combustion Test of a Mach 5 Scramjet Engine Model". Journal of the Korean Society of Propulsion Engineers 17, nr 3 (1.06.2013): 9–14. http://dx.doi.org/10.6108/kspe.2013.17.3.009.

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Ha, Jeong Ho, Rajarshi Das, Foluso Ladeinde, Tae Ho Kim i Heuy Dong Kim. "Numerical Study on Mode Transition in a Scramjet Engine". Journal of the Korean Society of Propulsion Engineers 21, nr 6 (7.12.2017): 21–31. http://dx.doi.org/10.6108/kspe.2017.21.6.021.

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O'Neill, Mary Kae Lockwood, i Mark J. Lewis. "Design tradeoffs on scramjet engine integrated hypersonic waverider vehicles". Journal of Aircraft 30, nr 6 (listopad 1993): 943–52. http://dx.doi.org/10.2514/3.46438.

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Ravi Teja, Gonnabathula S. B., Kumar Pakki Bharani Chandra i Gopal Jee. "Development of a Control Oriented Scramjet Engine Inlet Model". IFAC-PapersOnLine 55, nr 1 (2022): 198–203. http://dx.doi.org/10.1016/j.ifacol.2022.04.033.

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HARADA, Nobuhiro. "Application of MHD Power Generation Technology to Scramjet Engine". Journal of The Institute of Electrical Engineers of Japan 128, nr 1 (2008): 32–35. http://dx.doi.org/10.1541/ieejjournal.128.32.

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Kaminaga, Susumu, Sadatake Tomioka i Hiroyuki Yamasaki. "Performance of Scramjet Engine with MHD Energy Bypass System". JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 53, nr 623 (2005): 554–61. http://dx.doi.org/10.2322/jjsass.53.554.

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Shimura, Takashi, Noboru Sakuranaka, Tetsuji Sunami i Kouichiro Tani. "Thrust, Lift, and Pitching Moment of a Scramjet Engine". Journal of Propulsion and Power 17, nr 3 (maj 2001): 617–21. http://dx.doi.org/10.2514/2.5786.

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Sadatake TOMIOKA, By, Ryohei KOBAYASHI, Atsuo MURAKAMI, Shuichi UEDA, Tomoyuki KOMURO i and Katsuhiro ITOH. "Combustion Enhancement in Scramjet-Operation of a RBCC Engine". TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 10, ists28 (2012): Pa_55—Pa_61. http://dx.doi.org/10.2322/tastj.10.pa_55.

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Berglund, M., i C. Fureby. "LES of supersonic combustion in a scramjet engine model". Proceedings of the Combustion Institute 31, nr 2 (styczeń 2007): 2497–504. http://dx.doi.org/10.1016/j.proci.2006.07.074.

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TSUJIKAWA, Y. "Effects of hydrogen active cooling on scramjet engine performance". International Journal of Hydrogen Energy 21, nr 4 (kwiecień 1996): 299–304. http://dx.doi.org/10.1016/0360-3199(95)00077-1.

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Yang, Qingchun, Youhai Zong i Wen Bao. "Constant static-temperature heating for hydrogen fueled scramjet engine". International Journal of Hydrogen Energy 41, nr 3 (styczeń 2016): 2002–10. http://dx.doi.org/10.1016/j.ijhydene.2015.11.014.

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Zhang, Shikong, Jiang Li, Fei Qin, Zhiwei Huang i Rui Xue. "Numerical investigation of combustion field of hypervelocity scramjet engine". Acta Astronautica 129 (grudzień 2016): 357–66. http://dx.doi.org/10.1016/j.actaastro.2016.09.028.

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