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Journal articles on the topic 'Air-breathing propulsion'

Author: Grafiati

Published: 4 June 2021

Last updated: 7 February 2022

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1

PATNAIK, SURYA N., THOMAS M. LAVELLE, and DALE A. HOPKINS. "Optimization of air-breathing propulsion engine concept." Communications in Numerical Methods in Engineering 13, no. 8 (August 1997): 635–41. http://dx.doi.org/10.1002/(sici)1099-0887(199708)13:8<635::aid-cnm89>3.0.co;2-x.

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2

Shi Lei, 石磊, 赵尚弘 Zhao Shanghonh, 周万银 Zhou Wanyin, 李勇军 Li Yongjun, 胥杰 Xu Jie, and 方绍强 Fang Shaoqiang. "Numerical Analysis of Air-Breathing Mode Laser Propulsion." Chinese Journal of Lasers 35, no. 1 (2008): 49–54. http://dx.doi.org/10.3788/cjl20083501.0049.

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3

Schoettle, U. M., H. Grallert, and F. A. Hewitt. "Advanced air-breathing propulsion concepts for winged launch vehicles." Acta Astronautica 20 (January 1989): 117–29. http://dx.doi.org/10.1016/0094-5765(89)90061-1.

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4

Marchioni, Francesco, and Mark A. Cappelli. "Extended channel Hall thruster for air-breathing electric propulsion." Journal of Applied Physics 130, no. 5 (August 7, 2021): 053306. http://dx.doi.org/10.1063/5.0048283.

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5

Zhang, Xi-bin, and Qun Zong. "Modeling and Analysis of an Air-Breathing Flexible Hypersonic Vehicle." Mathematical Problems in Engineering 2014 (2014): 1–9. http://dx.doi.org/10.1155/2014/264247.

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By using light-weighted material in hypersonic vehicle, the vehicle body can be easily deformed. The mutual couplings in aerodynamics, flexible structure, and propulsion system will bring great challenges for vehicle modeling. In this work, engineering estimated method is used to calculate the aerodynamic forces, moments, and flexible modes to get the physics-based model of an air-breathing flexible hypersonic vehicle. The model, which contains flexible effects and viscous effects, can capture the physical characteristics of high-speed flight. To overcome the analytical intractability of the model, a simplified control-oriented model of the hypersonic vehicle is presented with curve fitting approximations. The control-oriented model can not only reduce the complexity of the model, but also retain aero-flexible structure-propulsion interactions of the physics-based model and can be applied for nonlinear control.

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6

Göksel, B., and I. Ch Mashek. "First Breakthrough for Future Air-Breathing Magneto-Plasma Propulsion Systems." Journal of Physics: Conference Series 825 (April 12, 2017): 012005. http://dx.doi.org/10.1088/1742-6596/825/1/012005.

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7

Sundaresan, M. "Design aspects of launch vehicle sizing including air-breathing propulsion." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 220, no. 5 (May 2006): 487–98. http://dx.doi.org/10.1243/09544100jaero30.

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8

Bussard, Robert W., and Lorin W. Jameson. "Inertial-Electrostatic-Fusion Propulsion Spectrum: Air-Breathing to Interstellar Flight." Journal of Propulsion and Power 11, no. 2 (March 1995): 365–72. http://dx.doi.org/10.2514/3.51434.

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9

Urzay, Javier. "Supersonic Combustion in Air-Breathing Propulsion Systems for Hypersonic Flight." Annual Review of Fluid Mechanics 50, no. 1 (January 5, 2018): 593–627. http://dx.doi.org/10.1146/annurev-fluid-122316-045217.

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10

Mukundan, Vijith, Arnab Maity, Shashi Ranjan Kumar, and U. P. Rajeev. "Ascent Trajectory Optimization of Launch Vehicles with Air-Breathing Propulsion." IFAC-PapersOnLine 52, no. 12 (2019): 274–79. http://dx.doi.org/10.1016/j.ifacol.2019.11.255.

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11

Blankson, I. M. "Air-Breathing Hypersonic Cruise: Prospects for Mach 4–7 Waverider Aircraft." Journal of Engineering for Gas Turbines and Power 116, no. 1 (January 1, 1994): 104–15. http://dx.doi.org/10.1115/1.2906779.

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There is currently a renewal of world-wide interest in hypersonic flight. Vehicle concepts being considered range from cruise missiles to SSTO and TSTO vehicles. The new characteristics of these vehicles are that they will be powered by air-breathing engines and have long residence times in the air-breathing corridor. In the Mach 4–7 regime, waverider aircraft are being considered as candidates for both long-range and short-range cruise missions, as hypersonic missiles, and as high-L/D highly maneuverable vehicles. This paper will discuss the potential for near-term and far-term application of air-breathing engines to the above-mentioned waverider vehicle concepts and missions. In particular, the cruise mission is discussed in detail and attempts are made to compare and contrast it with the accelerator mission. Past criticisms levied against waveriders alleging low volumetric efficiency, lack of engine/airframe integration studies, poor off-design performance, poor take-off and landing capability, have been shown by ongoing research to be unfounded. A discussion is presented of some of the technical challenges and ongoing research aimed at realizing such vehicles: from turboramjet and scramjet technology development, propulsion-airframe integration effects on vehicle performance, aeroservothermoelastic systems analysis, hypersonic stability and control with aeroservothermoelastic and propulsion effects, etc. A unique and very strong aspect of hypersonic vehicle design is the integration and interaction of the propulsion system, aerodynamics, aerodynamic heating, stability and control, and materials and structures. This first-order multidisciplinary situation demands the ability to integrate highly coupled and interacting elements in a fundamental and optimal fashion to achieve the desired performance. Some crucial technology needs are found in propulsion-airframe integration and its role in configuration definition, hypersonic boundary-layer transition and its impact on vehicle gross-weight and mission success, scramjet combustor mixing length and its impact on engine weight and, CFD (turbulence modeling, transition modeling, etc) as a principal tool for the design of hypersonic vehicles. Key technology implications in thermal management, structures, materials, and flight control systems will also be briefly discussed. It is concluded that most of the technology requirements in the Mach 4–7 regime are relatively conventional, making cited applications near-term, yet offering very significant advancements in aircraft technology.

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12

Ardema, M. D., J. V. Bowles, and T. Whittaker. "Near-optimal propulsion-system operation for an air-breathing launch vehicle." Journal of Spacecraft and Rockets 32, no. 6 (November 1995): 951–56. http://dx.doi.org/10.2514/3.26714.

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13

Patnaik, Surya N., Thomas M. Lavelle, Dale A. Hopkins, and Rula M. Coroneos. "Cascade Optimization Strategy for Aircraft and Air-Breathing Propulsion System Concepts." Journal of Aircraft 34, no. 1 (January 1997): 136–39. http://dx.doi.org/10.2514/2.2148.

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14

Mori, Koichi, Nobuyuki Harabe, Kimiya Komurasaki, and Yoshihiro Arakawa. "Blast-Wave Energy Conversion Processes in Air-Breathing RP Laser Propulsion." JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 51, no. 588 (2003): 23–30. http://dx.doi.org/10.2322/jjsass.51.23.

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15

Jing Chen, 陈静, 谭荣清 Rongqing Tan, 吴谨 Jin Wu, 卢远添 Yuantian Lu, 徐程 Cheng Xu, and 朱玉峰 Yufeng Zhu. "Air-breathing mode laser propulsion with a long-pulse TE CO2laser." Chinese Optics Letters 8, no. 8 (2010): 771–72. http://dx.doi.org/10.3788/col20100808.0771.

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16

Hong, Yanji, Junling Song, Cunyan Cui, and Qian Li. "Numerical study of energy conversion process in air-breathing laser propulsion." Applied Physics A 105, no. 1 (June 7, 2011): 189–96. http://dx.doi.org/10.1007/s00339-011-6479-9.

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17

Li, Yanwu, X. Chen, Danming Li, Yuhua Xiao, Peng Dai, and Chengshi Gong. "Design and analysis of vacuum air-intake device used in air-breathing electric propulsion." Vacuum 120 (October 2015): 89–95. http://dx.doi.org/10.1016/j.vacuum.2015.06.011.

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18

DOGRA, Bharat Ankur, Mehakveer SINGH, Tejinder Kumar JINDAL, and Subhash CHANDER. "Technological advancements in Pulse Detonation Engine Technology in the recent past: A Characterized Report." INCAS BULLETIN 11, no. 4 (December 8, 2019): 81–92. http://dx.doi.org/10.13111/2066-8201.2019.11.4.8.

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Pulse Detonation Engine (PDE), is an emerging and promising propulsive technology all over the world in the past few decades. A pulse detonation engine (PDE) is a type of propulsion system that uses detonation waves to combust the fuel and oxidizer mixture. Theoretically, a PDE can be operate from subsonic to hypersonic flight speeds. Pulsed detonation engines offer many advantages over conventional air-breathing engines and are regarded as potential replacements for air-breathing and rocket propulsion systems, for platforms ranging from subsonic unmanned vehicles, long-range transportation, high-speed vehicles, space launchers to space vehicles. This article highlights the operating cycle of PDE, starting with the fuel-oxidizer mixture, combustion and Deflagration to detonation transition (DDT) followed by purging. PDE combustion process, a unique process, leads to consistent and repeatable detonation waves. This pulsed detonation combustion process causes rapid burning of the fuel-oxidizer mixture, which cannot be seen in any other combustion process as it is a thousand times faster than any other mode of combustion. PDE not only holds the capability of running effectively up to Mach 5 but it also changes the technicalities in space propulsion. The present paper is the extension of the previous study which is also a well characterized status report of PDE in different areas. The present study deals with the categorization of the design approach, computations & simulations, flow visualization, DDT & Thrust enhancement, PDRE’s, experimental detonation engines with some of the experience and research undertaken in Punjab Engineering College under the complete supervision and guidance of Prof. Tejinder Kumar Jindal followed by applications of PDE technology.

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19

Chinitz, Wallace, John I. Erdos, Oussama Rizkalla, G. Y. Anderson, and Dennis M. Bushnell. "Facility opportunities and associated stream chemistry considerations for hypersonic air-breathing propulsion." Journal of Propulsion and Power 10, no. 1 (January 1994): 6–17. http://dx.doi.org/10.2514/3.23705.

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20

Li Qian, 李倩, 洪延姬 Hong Yanji, 杨鹏涛 Yang Pengtao, and 柯发伟 Ke Fawei. "Numerical study on performance of airbreathing high frequency pulsed laser propulsion." High Power Laser and Particle Beams 23, no. 9 (2011): 2282–86. http://dx.doi.org/10.3788/hplpb20112309.2282.

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21

Xu, Jinyi, Zhiwen Wu, Pan Chen, Qimeng Xia, Kan Xie, and Xiangyang Liu. "Parametric Study of an Air-Breathing Electric Propulsion for Near-Space Vehicles." Journal of Propulsion and Power 34, no. 5 (September 2018): 1297–304. http://dx.doi.org/10.2514/1.b36131.

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22

Froning, H. D. "Investigation of antimatter air-breathing propulsion for single-stage-to-orbit ships." Acta Astronautica 17, no. 8 (August 1988): 853–61. http://dx.doi.org/10.1016/0094-5765(88)90168-3.

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23

Zhang, Dong, Shuo Tang, Lin Cao, Feng Cheng, and Fan Deng. "Research on control-oriented coupling modeling for air-breathing hypersonic propulsion systems." Aerospace Science and Technology 84 (January 2019): 143–57. http://dx.doi.org/10.1016/j.ast.2018.10.014.

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24

Hu, Peng, Yan Shen, Zhaopu Yao, Wei Mao, Yanlin Hu, and Xuhui Liu. "Study of multi-cusped plasma thruster applied to Air-Breathing Electric Propulsion." Vacuum 190 (August 2021): 110275. http://dx.doi.org/10.1016/j.vacuum.2021.110275.

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25

Rao, R. Bhaskara, and Harihar Singh. "Effect of Agglomeration on Combustion of Metallized Propellants for Air-Breathing Propulsion System." Defence Science Journal 46, no. 5 (January 1, 1996): 327–30. http://dx.doi.org/10.14429/dsj.46.4299.

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26

Dimri, Ankit, and Racheet Matai. "Improved Air Turbo Rocket for Space Applications Application to Orbital Vehicles and Reentry." Applied Mechanics and Materials 110-116 (October 2011): 2554–61. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.2554.

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An Air Turbo Rocket (ATR) is a propulsion system which combines a turbo jet with a rocket engine. Currently it is being touted as a propulsion system for future missile systems, as these engines have a higher thrust density when compared to other air breathing engines. This paper explores the possibility of modifying the ATR for use in space application as well as during spacecraft re-entry. Such modified ATR’s could be used to power space vehicles up to the Low Earth Orbit (LEO) to dock with the International Space Station (ISS). In addition, thrust reversal techniques on the ATR systems could be used to improve the accuracy of Ballistic Missiles and hypersonic space planes upon Re-entry. Challenges faced would be in this type of air breathing engine would be operating at different atmospheric conditions. This paper will explore an ATR design, which will operate at different modes namely conventional mode, which will be used during below absolute ceiling, and the mission mode, which will be employed during flight in vacuum. Lastly, the reentry mode, which can be used for lessening the entry velocity of a vehicle to reduce the risks associated with reentry. The paper will try to emphasize the advantages of ATR as an affordable launch system for space shuttles and satellites with high maneuverability.

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27

Chen, Hao, Lei Shi, Lihua Ma, and Yongan Chen. "Numerical simulation of air-breathing nanosecond laser propulsion considering subsonic inflow and multi-pulse." Optik 125, no. 14 (July 2014): 3444–48. http://dx.doi.org/10.1016/j.ijleo.2014.01.038.

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28

Vasil'ev, A. A. "The Principal Aspects of Application of Detonation in Propulsion Systems." Journal of Combustion 2013 (2013): 1–15. http://dx.doi.org/10.1155/2013/945161.

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The basic problems of application of detonation process in propulsion systems with impulse and continuous burning of combustible mixture are discussed. The results on propagation of detonation waves in supersonic flow are analyzed relatively to air-breathing engine. The experimental results are presented showing the basic possibility of creation of an engine with exterior detonation burning. The base results on optimization of initiation in impulse detonation engine are explained at the expense of spatial and temporal redistribution of an energy, entered into a mixture. The method and technique for construction of highly effective accelerators for deflagration to detonation transition are discussed also.

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29

Prisacariu, Vasile, Constantin Rotaru, Ionică Cîrciu, and Mihai Niculescu. "Numerical simulation and performances evaluation of the pulse detonation engine." MATEC Web of Conferences 234 (2018): 01001. http://dx.doi.org/10.1051/matecconf/201823401001.

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A pulse detonation engine (PDE) is a type of propulsion system that uses detonation waves to combust the fuel and oxidizer mixture. The engine is pulsed because the mixture must be renewed in the combustor between each detonation wave. Theoretically, a PDE can operate from subsonic up to hypersonic flight speed. Pulsed detonation engines offer many advantages over conventional propulsion systems and are regarded as potential replacements for air breathing and rocket propulsion systems, for platforms ranging from subsonic unmanned vehicles, long range transports, high-speed vehicles, space launchers to space vehicles. The article highlights elements of the current state of the art, but also theoretical and numerical aspects of these types of unconventional engines. This paper presents a numerical simulation of a PDE at h=10000 m with methane as working fluid for stoichiometric combustion, in order to find out the detonation conditions.

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30

Shi Lei, 石磊, 赵尚弘 Zhao Shanghong, 楚兴春 Chu Xingchun, 吴继礼 Wu Jili, 李晓亮 Li Xiaoliang, and 李国杰 Li Guojie. "Experimental and Numerical Study of Air-Breathing Mode Propulsion by Solid Nd:Glass Nanosecond Pulse Laser." Chinese Journal of Lasers 37, no. 1 (2010): 100–104. http://dx.doi.org/10.3788/cjl20103701.0100.

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31

Murray, James J., Abhijit Guha, and Alan Bond. "Overview of the development of heat exchangers for use in air-breathing propulsion pre-coolers." Acta Astronautica 41, no. 11 (December 1997): 723–29. http://dx.doi.org/10.1016/s0094-5765(97)00199-9.

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32

Erofeev, A. I., A. P. Nikiforov, G. A. Popov, M. O. Suvorov, S. A. Syrin, and S. A. Khartov. "Air-Breathing Ramjet Electric Propulsion for Controlling Low-Orbit Spacecraft Motion to Compensate for Aerodynamic Drag." Solar System Research 51, no. 7 (December 2017): 639–45. http://dx.doi.org/10.1134/s0038094617070048.

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33

Nirbito, Wahyu, Muhammad Arif Budiyanto, and Robby Muliadi. "Performance Analysis of Combined Cycle with Air Breathing Derivative Gas Turbine, Heat Recovery Steam Generator, and Steam Turbine as LNG Tanker Main Engine Propulsion System." Journal of Marine Science and Engineering 8, no. 9 (September 20, 2020): 726. http://dx.doi.org/10.3390/jmse8090726.

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This study explains the performance analysis of a propulsion system engine of an LNG tanker using a combined cycle whose components are gas turbine, steam turbine, and heat recovery steam generator. The researches are to determine the total resistance of an LNG tanker with a capacity of 125,000 m3 by using the Maxsurf Resistance 20 software, as well as to design the propulsion system to meet the required power from the resistance by using the Cycle-Tempo 5.0 software. The simulation results indicate a maximum power of the system of about 28,122.23 kW with a fuel consumption of about 1.173 kg/s and a system efficiency of about 48.49% in fully loaded conditions. The ship speed can reach up to 20.67 knots.

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34

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

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

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35

Mogrekar, Ashish, and Ramasami Sivakumar. "CFD Analysis of Micro-Ramps for Hypersonic Flows." Applied Mechanics and Materials 592-594 (July 2014): 1962–66. http://dx.doi.org/10.4028/www.scientific.net/amm.592-594.1962.

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Air intake is a crucial component for supersonic and hypersonic air breathing propulsion devices. The intake must provide the required mass flow rate of air with minimal loss of stagnation pressure. A major difficulty in the stable operation of an intake is associated with shock wave boundary layer interaction (SBLI). This causes boundary layer separation and adverse pressure gradients which lead to total pressure loss, flow unsteadiness and flow distortion in the intake system. Passive control devices such as micro-ramp, thick-vanes provide better boundary layer control and reduce parasitic drag. The proposed study aims to perform CFD analysis of micro-ramp for hypersonic flows and validate the results with the available experimental data. Two micro ramp models namely MR80 and MR40 are considered for this study. Results obtained show the presence of micro ramp successfully delayed the flow separation and helped to suppress SBLI.

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36

Moriondo, Andrea, Sylvain Mukenge, and Daniela Negrini. "Transmural pressure in rat initial subpleural lymphatics during spontaneous or mechanical ventilation." American Journal of Physiology-Heart and Circulatory Physiology 289, no. 1 (July 2005): H263—H269. http://dx.doi.org/10.1152/ajpheart.00060.2005.

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The role played by the mechanical tissue stress in supporting lymph formation and propulsion in thoracic tissues was studied in deeply anesthetized rats ( n = 13) during spontaneous breathing or mechanical ventilation. After arterial and venous catheterization and insertion of an intratracheal cannula, fluorescent dextrans were injected intrapleurally to serve as lymphatic markers. After 2 h, the fluorescent intercostal lymphatics were identified, and the hydraulic pressure in lymphatic vessels (Plymph) and adjacent interstitial space (Pint) was measured using micropuncture. During spontaneous breathing, end-expiratory Plymph and corresponding Pint were −2.5 ± 1.1 (SE) and 3.1 ± 0.7 mmHg ( P < 0.01), which dropped to −21.1 ± 1.3 and −12.2 ± 1.3 mmHg, respectively, at end inspiration. During mechanical ventilation with air at zero end-expiratory alveolar pressure, Plymph and Pint were essentially unchanged at end expiration, but, at variance with spontaneous breathing, they increased at end inspiration to 28.1 ± 7.9 and 28.2 ± 6.3 mmHg, respectively. The hydraulic transmural pressure gradient (ΔPtm = Plymph − Pint) was in favor of lymph formation throughout the whole respiratory cycle (ΔPtm = −6.8 ± 1.2 mmHg) during spontaneous breathing but not during mechanical ventilation (ΔPtm = −1.1 ± 1.8 mmHg). Therefore, data suggest that local tissue stress associated with the active contraction of respiratory muscles is required to support an efficient lymphatic drainage from the thoracic tissues.

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37

Mehta, U., J. Bowles, S. Pandya, J. Melton, L. Huynh, J. Kless, and V. Hawke. "Conceptual stage separation from widebody subsonic carrier aircraft for space access." Aeronautical Journal 118, no. 1209 (November 2014): 1279–309. http://dx.doi.org/10.1017/s0001924000009970.

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Abstract Stage separation is a critical technical issue for developing two-stage-to-orbit (TSTO) launch systems with widebody carrier aircraft that use air-breathing propulsion and launch vehicle stages that use rocket propulsion. During conceptual design phases, this issue can be addressed with a combination of engineering methods, computational fluid dynamics simulations, and trajectory analysis of the mated system and the launch vehicle after staging. The outcome of such analyses helps to establish the credibility of the proposed TSTO system and formulate a ground-based test programme for the preliminary design phase. This approach is demonstrated with an assessment of stage separation from the shuttle carrier aircraft. Flight conditions are determined for safe mated flight, safe stage separation, and for the launch vehicle as it commences ascending flight. Accurate assessment of aerodynamic forces and moments is critical during staging to account for interference effects from the proximities of the two large vehicles. Interference aerodynamics have a modest impact on the separation conditions and separated flight trajectories, but have a significant impact on the interaction forces.

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38

Schmid, Norbert R., Dirk C. Leinhos, and Leonhard Fottner. "Steady Performance Measurements of a Turbofan Engine With Inlet Distortions Containing Co- and Counterrotating Swirl From an Intake Diffuser for Hypersonic Flight." Journal of Turbomachinery 123, no. 2 (February 1, 2000): 379–85. http://dx.doi.org/10.1115/1.1343466.

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The influence of distorted inlet flow on the steady and unsteady performance of a turbofan engine, which is a component of an air-breathing combined propulsion system for a hypersonic transport aircraft, is reported in this paper. The performance and stability of this propulsion system depend on the behavior of the turbofan engine. The complex shape of the intake duct causes inhomogeneous flow at the engine inlet plane, where total pressure and swirl distortions are present. The S-bend intakes are installed axisymmetrically left and right into the hypersonic aircraft, generating axisymmetric mirror-inverted flow patterns. Since all turbo engines of the propulsion system have the same direction of rotation, one distortion corresponds to a corotating swirl at the low pressure compressor (LPC) inlet while the mirror-inverted image counterpart represents a counterrotating swirl. Therefore the influence of the distortions on the performance and stability of the ‘CO’ and ‘COUNTER’ rotating turbo engine are different. The distortions were generated separately by an appropriate simulator at the inlet plane of a LARZAC 04 engine. The results of low-frequency measurements at different engine planes yield the relative variations of thrust and specific fuel consumption and hence the steady engine performance. High-frequency measurements were used to investigate the different influence of CO and COUNTER inlet distortions on the development of LPC instabilities.

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39

Michalski, Quentin, Bastien Boust, and Marc Bellenoue. "Experimental Investigation of Ignition Stability in a Cyclic Constant-Volume Combustion Chamber Featuring Relevant Conditions for Air-Breathing Propulsion." Flow, Turbulence and Combustion 102, no. 2 (February 2019): 279–98. http://dx.doi.org/10.1007/s10494-019-00015-1.

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40

Dunn, M. G., C. Padova, and R. M. Adams. "Response of an Operational Turbofan Engine to a Simulated Nuclear Blast." Journal of Fluids Engineering 109, no. 2 (June 1, 1987): 121–29. http://dx.doi.org/10.1115/1.3242631.

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This paper describes the results of a measurement program designed to determine the transient response of an air-breathing propulsion system to simulated nuclear blast waves. A Ludwieg-tube facility, incorporating a driver technique consisting of an activating chamber and a nonfrangible diaphragm, was used to create the required shock waves. Detailed measurements were performed at incident shock overpressures of approximately 6.9, 10.3, 13.8, and 17.2 kPa (1.0, 1.5, 2.0, and 2.5 psi). For each of these overpressures, data were obtained for engine speeds of 0, 80, 90, and 100 percent of maximum speed. Typical results are presented for distortion patterns at the fan face for both an extended bellmouth and a S-shaped inlet at either 0 or 20 deg yaw angle.

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41

Dutta, T., K. P. Sinhamahapatra, and S. S. Bandyopadhyay. "CFD Analysis of Energy Separation in Ranque-Hilsch Vortex Tube at Cryogenic Temperature." Journal of Fluids 2013 (November 14, 2013): 1–14. http://dx.doi.org/10.1155/2013/562027.

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Study of the energy separation phenomenon in vortex tube (VT) at cryogenic temperature (temperature range below 123 K) has become important because of the potential application of VT as in-flight air separator in air breathing propulsion. In the present study, a CFD model is used to simulate the energy separation phenomenon in VT with gaseous air at cryogenic temperature as working fluid. Energy separation at cryogenic temperature is found to be considerably less than that obtained at normal atmospheric temperature due to lower values of inlet enthalpy and velocity. Transfer of tangential shear work from inner to outer fluid layers is found to be the cause of energy separation. A parametric sensitivity analysis is carried out in order to optimize the energy separation at cryogenic temperature. Also, rates of energy transfer in the form of sensible heat and shear work in radial and axial directions are calculated to investigate the possible explanation of the variation of the hot and cold outlet temperatures with respect to various geometric and physical input parameters.

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Vinogradov, Viacheslav A., Yurii M. Shikhman, and Corin Segal. "A Review of Fuel Pre-injection in Supersonic, Chemically Reacting Flows." Applied Mechanics Reviews 60, no. 4 (July 1, 2007): 139–48. http://dx.doi.org/10.1115/1.2750346.

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Developing an efficient, supersonic combustion-based, air breathing propulsion cycle operating above Mach 3.5, especially when conventional hydrocarbon fuels are sought and particularly when liquid fuels are preferred to increase density, requires mostly effective mechanisms to improve mixing efficiency. One way to extend the time available for mixing is to inject part of the fuel upstream of the vehicle’s combustion chamber. Injection from the wall remains one of the most challenging problems in supersonic aerodynamics, including the requirement to minimize impulse losses, improve fuel-air mixing, reduce inlet∕combustor interactions, and promote flame stability. This article presents a review of studies involving liquid and, in selected cases, gaseous fuel injected in supersonic inlets or in combustor’s insulators. In all these studies, the fuel was injected from a wall in a wake of thin swept pylons at low dynamic pressure ratios (qjet∕qair=0.6–1.5), including individual pylon∕injector geometries and combinations in the inlet and combustor’s isolator, a variety of injection conditions, different injectants, and evaluated their effects on fuel plume spray, impulse losses, and mixing efficiency. This review article cites 47 references.

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Zhai, Jian, Chen-An Zhang, Fa-Min Wang, and Wei-Wei Zhang. "Alleviation of lateral spillage of two-dimensional hypersonic inlet using waverider-configuration chines." International Journal of Modern Physics B 34, no. 14n16 (June 4, 2020): 2040074. http://dx.doi.org/10.1142/s0217979220400743.

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Hypersonic inlet is an important part of the propulsion system of hypersonic air-breathing vehicles. However, the performance of the two-dimensional hypersonic inlet, a major type of hypersonic inlets, is considerably deteriorated for lateral spillage. In this study, waverider-configuration chines mounted on the lateral sides of a two-dimensional three-staged external-compression hypersonic inlet for a Mach number of 6.0 are investigated to determine their ability to alleviate the lateral spillage. The chines are built by using a waverider design method. The numerical results suggest that a severe flow spillage induced by three-dimensional effect shows up near the lateral edge of the inlet without chines, which degrades the mass-flow ratio and flow uniformity. In contrast, the waverider-configuration chines effectively alleviate the lateral spillage. Consequently, the mass-flow ratio and the flow uniformity are both improved significantly.

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Llobet, J. R., K. D. Basore, R. J. Gollan, and I. H. Jahn. "Experimental and numerical heat transfer from vortex-injection interaction in scramjet flowfields." Aeronautical Journal 124, no. 1280 (May 11, 2020): 1545–67. http://dx.doi.org/10.1017/aer.2020.39.

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ABSTRACTAir-breathing propulsion has the potential to decrease the cost per kilogram for access-to-space, while increasing the flexibility of available low earth orbits. However, to meet the performance requirements, fuel-air mixing inside of scramjet engines and thermal management still need to be improved.An option to address these issues is to use intrinsically generated vortices from scramjet inlets to enhance fuel-air mixing further downstream, leading to shorter, less internal drag generating, and thus more efficient engines. Previous works have studied this vortex-injection interaction numerically, but validation was impractical due to lack of published experimental data. This paper extends upon these previous works by providing experimental data for a canonical geometry, obtained in the T4 Stalker Tube at Mach 8 flight conditions, and assesses the accuracy of numerical methodologies such as RANS CFD to predict the vortex-injection interaction.Focus is placed on understanding the ability of the numerical methodology to replicate the most important aspects of the vortex-injection interaction. Results show overall good agreement between the numerical and experimental results, as all major features are captured. However, limitations are encountered, especially due to a localised region of over predicted heat flux.

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Mehta, U., J. Bowles, J. Melton, L. Huynh, and P. Hagseth. "Water injection pre-compressor cooling assist space access." Aeronautical Journal 119, no. 1212 (February 2015): 145–71. http://dx.doi.org/10.1017/s0001924000010319.

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AbstractAdvances in space activity are linked to reductions in launch cost. Air-breathing propulsion-assisted flight systems offer the potential for revolutionary change of the space operations paradigm. Horizontal launch of a space-access system provides mission flexibility, responsiveness, and affordability. One way to reduce launch cost is to increase the Mach number at which a launch vehicle is staged from a carrier aircraft. Without exceeding the engine and airframe design limits, the pre-compressor cooling technology allows an operational aircraft to operate at Mach numbers and altitudes beyond its basic operational limits. This is an essential, near-term technology for reducing launch cost to place small-weight payloads in low Earth orbit. The advantage of this technology is assessed with a modified McDonnell Douglas QF-4C aircraft. Payloads are unachievable or marginal with an unmodified QF-4C. However, payloads weighing around 150 pounds are plausible with this aircraft when incorporating the water injection pre-compressor cooling (WIPCC) technology.

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Wicks, Frank. "Nuclear Navy." Mechanical Engineering 126, no. 01 (January 1, 2004): 30–36. http://dx.doi.org/10.1115/1.2004-jan-2.

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This article discusses how atomic weapons had been the vision and accomplishment of physicists. The very limited space available would require a high power-density reactor and would present unprecedented heat transfer challenges. Diesel electric submarines had to surface to run air breathing engines, which needed to charge their batteries. Range and submerged time of a submarine with a reliable nuclear propulsion system would be limited only by food supply and endurance of the crew. The rumor among the first nuclear submariners was that they would surface every 4 years to reenlist. The reason that the submarine was the first application of controlled nuclear power is a tribute to the vision of another man. Back in 1953, while the Nautilus was being constructed, Rickover’s naval career appeared to be over. The Navy’s promotion process was not favorable for an engineering duty officer. Nuclear power put Rickover even farther out of the promotion mainstream.

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

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

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Ferrero, Andrea. "Control of a Supersonic Inlet in Off-Design Conditions with Plasma Actuators and Bleed." Aerospace 7, no. 3 (March 19, 2020): 32. http://dx.doi.org/10.3390/aerospace7030032.

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Supersonic inlets are a key component of present and future air-breathing propulsion systems for high-speed flight. The inlet design is challenging because of several phenomena that must be taken under control: shock waves, boundary layer separation and unsteadiness. Furthermore, the intensity of these phenomena is strongly influenced by the working conditions and so active control systems can be particularly useful in off-design conditions. In this work, a mixed compression supersonic inlet with a double wedge ramp is considered. The flow field was numerically investigated at different values of Mach number. The simulations show that large separations appear at the higher Mach numbers on both the upper and lower walls of the duct. In order to improve the performances of the inlet two different control strategies were investigated: plasma actuators and bleed. Different locations of the plasma actuator are considered in order to also apply this technology to configurations with a diverter which prevents boundary layer ingestion. The potential of the proposed solutions is investigated in terms of total pressure recovery, flow distortion and power consumption.

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FREUND, JONATHAN B., PARVIZ MOIN, and SANJIVA K. LELE. "Compressibility effects in a turbulent annular mixing layer. Part 2. Mixing of a passive scalar." Journal of Fluid Mechanics 421 (October 25, 2000): 269–92. http://dx.doi.org/10.1017/s0022112000001634.

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The mixing of fuel and oxidizer in a mixing layer between high-speed streams is important in many applications, especially air-breathing propulsion systems. The details of this process in a turbulent annular mixing layer are studied with direct numerical simulation. Convective Mach numbers of the simulations range from Mc = 0.1 to Mc = 1.8. Visualizations of the scalar field show that at low Mach numbers large intrusions of nearly pure ambient or core fluid span the mixing region, whereas at higher Mach numbers these intrusions are suppressed. Increasing the Mach number is found to change the mixture fraction probability density function from non-marching to marching and the mixing efficiency from 0.5 at Mc = 0.1 to 0.67 at Mc = 1.5. Scalar concentration fluctuations and the axial velocity fluctuations become highly correlated as the Mach number increases and a suppressed role of pressure in the axial momentum equation is found to be responsible for this. Anisotropy of scalar flux increases with Mc, and is explained via the suppression of transverse turbulence lengthscale.

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Deepu, M., M. P. Dhrishit, and S. Shyji. "Numerical simulation of high speed reacting shear layers using AUSM+- up scheme-based unstructured finite volume method solver." International Journal of Modeling, Simulation, and Scientific Computing 08, no. 03 (September 2017): 1750020. http://dx.doi.org/10.1142/s1793962317500209.

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Development of an Advection Upstream Splitting Method (AUSM[Formula: see text]-up) scheme-based Unstructured Finite Volume (UFVM) solver for the simulation of two-dimensional axisymmetric/planar high speed compressible turbulent reacting shear layers is presented. The inviscid numerical flux is evaluated using AUSM[Formula: see text]-up upwind scheme. An eight-step hydrogen–oxygen finite rate chemistry model is used to model the development of chemical species in a supersonic reacting flow field. The chemical species terms are alone solved implicitly in this explicit flow solver by rescaling the equation in time. The turbulence modeling has been done using RNG-based [Formula: see text]–[Formula: see text] model. Three-stage Runge–Kutta method has been used for explicit time integration. The nonreacting two-dimensional Cartesian version of the same solver has been successfully validated against experimental and numerical results reported for the wall static pressure data in sonic slot injection to supersonic stream. Detailed validation studies for reacting flow solver has been done using experimental results reported for a coaxial supersonic combustor, in which species profile at various axial locations has been compared. Present numerical solver is useful in simulating combustors of high speed air-breathing propulsion devices.

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