Journal articles: 'Scramjet'

1

Jiang, Baohong. "Comprehensive Analysis of the Advanced Technologies for Scramjet." Highlights in Science, Engineering and Technology 43 (April 14, 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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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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Meng, Yu, Wenming Sun, Hongbin Gu, Fang Chen, and Ruixu Zhou. "Supersonic Combustion Mode Analysis of a Cavity Based Scramjet." Aerospace 9, no. 12 (December 15, 2022): 826. http://dx.doi.org/10.3390/aerospace9120826.

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Since flame stability is the key to the performance of scramjets, scramjet combustion mode and instability characteristics were investigated by using the POD method based on a cavity-stabilized scramjet. Experiments were developed on a directly connected scramjet model that had an inlet flow of Mach 2.5 with a cavity stabilizer. CH* chemiluminescence, schlieren, and a wall static pressure sensor were employed to observe flow and combustion behavior. Three typical combustion modes were classified by distinguishing averaged CH* chemiluminescence images of three ethylene fuel jet equivalence ratios. The formation reason was explained using schlieren images and pressure characteristics. POD modes (PDMs) were determined using the proper orthogonal decomposition (POD) of sequential flame CH* chemiluminescence images. The PSD (power spectral density) of the PDM spectra showed large peaks in a frequency range of 100–600 Hz for three typical stabilized combustion modes. The results provide oscillation characteristics of three scramjet combustion modes.

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Paull, A., R. J. Stalker, and D. J. Mee. "Scramjet thrust measurement in a shock tunnel." Aeronautical Journal 99, no. 984 (April 1995): 161–63. http://dx.doi.org/10.1017/s0001924000027147.

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This note reports tests in a shock tunnel in which a fully integrated scramjet configuration produced net thrust. The experiments not only showed that impulse facilities can be used for assessing thrust performance, but also were a demonstration of the application of a new technique(1) to the measurement of thrust on scramjet configurations in shock tunnels. These two developments are of significance because scramjets are expected to operate at speeds well in excess of 2 km/s, and shock tunnels offer a means of generating high Mach number flows at such speeds.

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Jin, Liang, Xian Yu Wu, Jing Lei, Li Yan, Wei Huang, and Jun Liu. "CFD Analysis of a Hypersonic Vehicle Powered by Triple-Module Scramjets." Applied Mechanics and Materials 390 (August 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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Stalker, R. J., N. K. Truong, R. G. Morgan, and A. Paull. "Effects of hydrogen–air non–equilibrium chemistry on the performance of a model scramjet thrust nozzle." Aeronautical Journal 108, no. 1089 (November 2004): 575–84. http://dx.doi.org/10.1017/s0001924000000403.

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AbstractTwo aspects of hydrogen-air non-equilibrium chemistry related to scramjets are nozzle freezing and a process called ‘kinetic afterburning’ which involves continuation of combustion after expansion in the nozzle. These effects were investigated numerically and experimentally with a model scramjet combustion chamber and thrust nozzle combination. The overall model length was 0·5m, while precombustion Mach numbers of 3·1±0·3 and precombustion temperatures ranging from 740K to 1,400K were involved. Nozzle freezing was investigated at precombustion pressures of 190kPa and higher, and it was found that the nozzle thrusts were within 6% of values obtained from finite rate numerical calculations, which were within 7% of equilibrium calculations. When precombustion pressures of 70kPa or less were used, kinetic afterburning was found to be partly responsible for thrust production, in both the numerical calculations and the experiments. Kinetic afterburning offers a means of extending the operating Mach number range of a fixed geometry scramjet.

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CHINZEI, Nobuo, and Goro MASUYA. "Scramjet Engines." Journal of the Society of Mechanical Engineers 94, no. 866 (1991): 75–80. http://dx.doi.org/10.1299/jsmemag.94.866_75.

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

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

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9

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

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

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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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11

Wu, Xianju, and Zhijun Wei. "Comparison of Dual-Combustion Ramjet and Scramjet Performances Considering Combustion Efficiency." Applied Sciences 13, no. 1 (December 29, 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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12

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

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

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13

Zhang, Jiang, Xiao Jun Pan, Jin Gang Dong, Yong Ming Qin, and Han Dong Ma. "Wind Tunnel Study on a Missile with Forward-Facing Cavity in Supersonic Flow." Advanced Materials Research 1016 (August 2014): 370–76. http://dx.doi.org/10.4028/www.scientific.net/amr.1016.370.

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A forward-facing cavity will be composed of the components of a scramjet from inlet to combustion chamber which has a uncovered inlet before the separation of the booster. Longitudinal oscillations are generated within the cavity under some certain flow conditions. Strong oscillations may damage the components of the scramjet, or induce bow-shock oscillations which may cause unsteady loads on the missile and affect the performance of the aerodynamical characteristics. An experimental study of missile model with a scramjet was conducted in a transonic wind tunnel. The characteristics of cavity flow were researched by both the dynamic force measurement and the fluctuation pressure measurement. In the experiments the oscillations within the cavity and the bow-shock in front of the inlet interacted. The oscillations of cavity flow and bow-shock affected the fluctuation pressure and the aerodynamical characteristics of missile remarkably. The amplitude of axial force was higher than the normal force's. The RMS of the fluctuation pressure of some measured place inside the scramjet reached a quarter of the total pressure, and the amplitude of the fluctuation reached half of the total pressure. Those might threaten the safety of the structure of the scramjet.

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14

Daren, Y., C. Tao, and B. Wen. "An idea of distributed parameter control for scramjet engines." Aeronautical Journal 111, no. 1126 (December 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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15

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

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

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

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

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Alhassani, Abdulla Khamis, Mohanad Tarek Mohamed, Mohammed Fares, and Sharul Sham Dol. "Shock Waves Analysis of the Novel Intake Design System for a Scramjet Propulsion." WSEAS TRANSACTIONS ON SYSTEMS 20 (April 15, 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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Yi, Wan Shuang, Jin Jiang, Lin Lin Li, and Hong Gui Chen. "Research on the Model of Scramjet Fuel Supply System." Applied Mechanics and Materials 532 (February 2014): 382–87. http://dx.doi.org/10.4028/www.scientific.net/amm.532.382.

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In order to make a research on the Scramjet Fuel Supply System, I use the software named Flowmaster to make a numerical simulation. In the work, I first make some research on the steady-state response of the Scramjet Fuel Supply System, in which it mainly contains the change trend of the outlet flowrate and the temperature of the working fluid with the variation of the valve's opening, the heating transferred to the fluid and the pressure difference between the inlet and the outlet of the cooling channel. After the simulation of the steady-state characteristics, I make some analysis about the reason casing the Scramjet Fuel Supply System response as this. Then I conduct a calculation of transient process cased the sudden close of the control valve. After the calculation, I also analyze the tendency of pressure changes and the bubble volume generated by the abrupt close of the regulating valve. This may provide some useful theoretical basis and references in the study of characteristics of Scramjet Fuel Supply System.

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19

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

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

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

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

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

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

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

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

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

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

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

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

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26

Cheng, Feng, Shuo Tang, Dong Zhang, and Yi Li. "Quasi-One-Dimensional Modeling and Analysis of RBCC Dual-Mode Scramjet Engine." International Journal of Turbo & Jet-Engines 36, no. 2 (May 27, 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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27

Colket, Meredith B., and Louis J. Spadaccini. "Scramjet Fuels Autoignition Study." Journal of Propulsion and Power 17, no. 2 (March 2001): 315–23. http://dx.doi.org/10.2514/2.5744.

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28

Townend, L. H. "Domain of the Scramjet." Journal of Propulsion and Power 17, no. 6 (November 2001): 1205–13. http://dx.doi.org/10.2514/2.5865.

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29

CNINZEI, Nobuo, and Goro MASUYA. "Aerodynamic problems in scramjet." Journal of the Japan Society for Aeronautical and Space Sciences 38, no. 435 (1990): 187–93. http://dx.doi.org/10.2322/jjsass1969.38.187.

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30

SUNAMI, Tetsuji, Masatoshi KODERA, Katsuhiro ITOH, and Hideyuki TANNO. "Hypermixer Scramjet Flight Experiment." Proceedings of the Thermal Engineering Conference 2004 (2004): 45–46. http://dx.doi.org/10.1299/jsmeted.2004.45.

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31

Bushnell, D. M. "Hypervelocity Scramjet Mixing Enhancement." Journal of Propulsion and Power 11, no. 5 (September 1995): 1088–90. http://dx.doi.org/10.2514/3.51445.

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32

Jiang, Yuguang, Yu Feng, Silong Zhang, Jiang Qin, and Wen Bao. "Numerical heat transfer analysis of transcritical hydrocarbon fuel flow in a tube partially filled with porous media." Open Physics 14, no. 1 (January 1, 2016): 659–67. http://dx.doi.org/10.1515/phys-2016-0073.

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AbstractHydrocarbon fuel has been widely used in air-breathing scramjets and liquid rocket engines as coolant and propellant. However, possible heat transfer deterioration and threats from local high heat flux area in scramjet make heat transfer enhancement essential. In this work, 2-D steady numerical simulation was carried out to study different schemes of heat transfer enhancement based on a partially filled porous media in a tube. Both boundary and central layouts were analyzed and effects of gradient porous media were also compared. The results show that heat transfer in the transcritical area is enhanced at least 3 times with the current configuration compared to the clear tube. Besides, the proper use of gradient porous media also enhances the heat transfer compared to homogenous porous media, which could help to avoid possible over-temperature in the thermal protection.

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33

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

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

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34

Yao, Yu Feng. "Scramjet Flow and Intake SBLI: Technical Challenges and Case Study." Applied Mechanics and Materials 315 (April 2013): 344–48. http://dx.doi.org/10.4028/www.scientific.net/amm.315.344.

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This paper reviews some basic research areas associated with Scramjet-powered hypersonic flying vehicle, particularly the forebody boundary-layer transition and intake shock-wave boundary-layer interactions (SBLI). Some technical and physical challenges in aerodynamics, aero-thermodynamics, aero-design are visited with focuses being placed on hypersonic boundary-layer transition process and its underlying physical mechanics, feasible physics-based engineering transition prediction methods, and physics-based modelling of shock-shock, shock-wave/boundary-layer interactions of Scramjet flows. Experimental, analytical and numerical studies of previously relevant studies have also been summarized with a total of twelve transition/intake configurations that can be used as benchmarks for validating physical model development and numerical simulation tools. A case study of Scramjet intake SBLI has been carried out by using computational fluid dynamics approach to understand shock induced flow separation and its consequent influences on combustion performance, along with research perspectives discussed accordingly.

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35

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

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

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36

Zeng, Yao Yuan, Wen Tao Zhao, and Zheng Hua Wang. "Parallel Simulation of Scramjet with Multilevel Hypergraph Partitioning." Advanced Materials Research 706-708 (June 2013): 1479–82. http://dx.doi.org/10.4028/www.scientific.net/amr.706-708.1479.

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As one of the most considerable methods to study hypersonic flight vehicle, the numerical simulation of supersonic combustion ramjet (scramjet) has drawn an ever increasing attention at present. Nevertheless, the traditional serial simulation is ungratified for current research requirements because of high calculation precision, avaricious memory overhead and overlong computation time. Meanwhile, the efficiency of parallel simulation using the domain decomposition method is not very satisfactory. In this paper, we study on a general algorithm for scramjet design, and subdivide the computing domain by using a multilevel hypergraph partitioning algorithm. In order to reduce computation while enhancing the degree of parallelism, overlapping communication with computation and non blocking communication is adopted to decrease the communication time when dealing with global communication. Finally, experimental results testing on a China-made supercomputer show the smallest value of parallel efficiency is more than 48% when the number of processors is 256. In conclusion, the result indicates that our parallel algorithm is simple, effective and practical in scramjet simulation.

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37

Roux, J. A. "Parametric Ideal Scramjet Cycle Analysis." Journal of Thermophysics and Heat Transfer 25, no. 4 (October 2011): 581–85. http://dx.doi.org/10.2514/1.t3758.

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38

Schütte, Gerrit, and Stephan Staudacher. "Probabilistic Aspects of Scramjet Design." Journal of Propulsion and Power 25, no. 2 (March 2009): 281–88. http://dx.doi.org/10.2514/1.38195.

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39

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

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40

Goldfeld, Marat A., Alexey V. Starov, and Viacheslav V. Vinogradov. "Experimental Study of Scramjet Module." Journal of Propulsion and Power 17, no. 6 (November 2001): 1222–26. http://dx.doi.org/10.2514/2.5867.

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41

Burakhanov, B. M., A. P. Likhachev, S. A. Medin, V. A. Novikov, V. I. Okunev, V. Yu Rickman, and V. A. Zeigarnik. "Advancement of Scramjet Magnetohydrodynamic Concept." Journal of Propulsion and Power 17, no. 6 (November 2001): 1247–52. http://dx.doi.org/10.2514/2.5871.

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42

Nelson, H. F. "Radiative Heating in Scramjet Combustors." Journal of Thermophysics and Heat Transfer 11, no. 1 (January 1997): 59–64. http://dx.doi.org/10.2514/2.6201.

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43

Schetz, Joseph A., Frederick S. Billig, and Stanley Favin. "Modular analysis of scramjet flowfields." Journal of Propulsion and Power 5, no. 2 (March 1989): 172–80. http://dx.doi.org/10.2514/3.23133.

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44

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

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45

Seleznev, R. K. "History of scramjet propulsion development." Journal of Physics: Conference Series 1009 (April 2018): 012028. http://dx.doi.org/10.1088/1742-6596/1009/1/012028.

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46

Evrard, Thomas, Richard Butler, Steven W. Hughes, J. Ranjan Banerjee, and H. F. Nelson. "Radiative heating in scramjet combustors." Journal of Thermophysics and Heat Transfer 11 (January 1997): 59–64. http://dx.doi.org/10.2514/3.858.

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47

Townend, Leo H. "The domain of the scramjet." Philosophical Transactions of the Royal Society of London. Series A: Mathematical, Physical and Engineering Sciences 357, no. 1759 (August 1999): 2317–34. http://dx.doi.org/10.1098/rsta.1999.0433.

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48

Guizzo, E. "Hypersonic flight [scramjet aircraft propulsion]." IEEE Spectrum 41, no. 1 (January 2004): 66. http://dx.doi.org/10.1109/mspec.2004.1317885.

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49

Mitani, Tohru. "Ignition problems in scramjet testing." Combustion and Flame 101, no. 3 (May 1995): 347–59. http://dx.doi.org/10.1016/0010-2180(94)00218-h.

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50

Chuang, Chien-Hsiung, and Jason L. Speyer. "Periodic Optimal Hypersonic Scramjet Cruise." Optimal Control Applications and Methods 8, no. 3 (October 29, 2007): 231–42. http://dx.doi.org/10.1002/oca.4660080305.

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Journal articles on the topic 'Scramjet'

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