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Gotowa bibliografia na temat „Supersonic compression structure”

Autor: Grafiati

Data publikacji: 6 września 2023

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Artykuły w czasopismach na temat "Supersonic compression structure"

Lin, E. P., Y. E. Kim i J. C. Hermanson. "Structure of Compression Waves on Supersonic Droplets". AIAA Journal 54, nr 2 (luty 2016): 777–81. http://dx.doi.org/10.2514/1.j054412.

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B Saheby, Eiman, Xing Shen i Anthony P. Hays. "Design and performance study of a parametric diverterless supersonic inlet". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, nr 2 (24.09.2019): 470–89. http://dx.doi.org/10.1177/0954410019875384.

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Diverterless supersonic inlet integration for a flight vehicle requires a three-dimensional compression surface (bump) design with an acceptable shock structure and boundary layer diversion; this results in a low drag induction system with acceptable propulsive efficiency. In this investigation, a computational fluid dynamics-based-generated bump is used to design an integrated diverterless supersonic inlet without any bleed mechanism on a forebody with a large wetted area. Numerical solution of the Navier–Stokes equations simulates the flow pattern of the configuration. The forebody design analysis includes simulating the effects of angle of attack and sideslip by dependent computational domains. Results demonstrate the ability of the bump surface to keep the shock structures in an operational mode even at high supersonic angles of attack. Analysis of shock structures and shock wave boundary layer interactions at supersonic maneuver conditions indicate that the aerodynamic efficiency of the diverterless supersonic inlet in conditions with a thick boundary layer and high angles of attack is sufficient to ensure operation throughout the supersonic flight envelope.

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Masud, J. "Flow field and performance analysis of an integrated diverterless supersonic inlet". Aeronautical Journal 115, nr 1170 (sierpień 2011): 471–80. http://dx.doi.org/10.1017/s0001924000006114.

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Abstract In this paper the computed flow and performance characteristics at low angle-of-attack (AOA) of an integrated diverterless supersonic inlet (DSI) are presented. The subsonic characteristics are evaluated at M∞ = 0·8 while the supersonic characteristics are evaluated at M∞= 1·7, which is near the design Mach number for the intake. In addition to the external flow features, the internal intake duct flow behaviour is also evaluated. The results of this study indicate effective boundary layer diversion due to the ‘bump’ compression surface in both subsonic and supersonic regimes. At M∞ = 1·7, the shockwave structure (oblique/normal shockwave) on the ‘bump’ compression surface and intake inlet is satisfactory at design (critical) mass flow rate. The intake duct flow behaviour at subsonic and supersonic conditions is generally consistent with ‘Y’ shaped intake duct of the present configuration. The secondary flow structure inside the duct has been effectively captured by present computations. The computed intake total pressure recovery at M∞ = 1·7 exhibits higher-than-conventional behaviour at low mass flow ratios, which is attributed to unique inlet design. Overall computed subsonic and supersonic total pressure recovery characteristics are satisfactory under the evaluated conditions and are also in agreement with wind tunnel test data.

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Knight, Doyle D., C. C. Horstman i Seymour Bogdonoff. "Structure of supersonic turbulent flow past a swept compression corner". AIAA Journal 30, nr 4 (kwiecień 1992): 890–96. http://dx.doi.org/10.2514/3.11006.

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Saheby, Eiman B., Xing Shen, Anthony P. Hays i Zhang Jun. "The inlet flow structure of a conceptual open-nose supersonic drone". Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 235, nr 12 (25.02.2021): 1687–705. http://dx.doi.org/10.1177/0954410020983043.

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This study describes the aerodynamic efficiency of a forebody–inlet configuration and computational investigation of a drone system, capable of sustainable supersonic cruising at Mach 1.60. Because the whole drone configuration is formed around the induction system and the design is highly interrelated to the flow structure of forebody and inlet efficiency, analysis of this section and understanding its flow pattern is necessary before any progress in design phases. The compression surface is designed analytically using oblique shock patterns, which results in a low drag forebody. To study the concept, two inlet–forebody geometries are considered for Computational Fluid Dynamic simulation using ANSYS Fluent code. The supersonic and subsonic performance, effects of angle of attack, sideslip, and duct geometries on the propulsive efficiency of the concept are studied by solving the three-dimensional Navier–Stokes equations in structured cell domains. Comparing the results with the available data from other sources indicates that the aerodynamic efficiency of the concept is acceptable at supersonic and transonic regimes.

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GAO, B., i Z. N. WU. "A study of the flow structure for Mach reflection in steady supersonic flow". Journal of Fluid Mechanics 656 (21.05.2010): 29–50. http://dx.doi.org/10.1017/s0022112010001011.

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In this paper we study the waves generated over the slipline and their interactions with other waves for Mach reflection in steady two-dimensional supersonic flow. We find that a series of expansion and compression waves exist over the slip line, even in the region immediately behind the leading part of the reflected shock wave, previously regarded as a uniform flow. These waves make the leading part of the slipline, previously regarded as straight, deviate nonlinearly towards the reflecting surface. When the transmitted expansion waves from the upper corner first intersect the slipline, an inflexion point is produced. Downstream of this inflexion point, compression waves are produced over the slipline. By considering the interaction between the various expansion or compression waves, we obtain a Mach stem height, the shape and position of the slipline and reflected shock wave, compared well to computational fluid dynamics (CFD) results. We also briefly consider the case with a subsonic portion behind the reflected shock wave. The global flow pattern is obtained through CFD and the starting point of the sonic line is identified through a simple analysis. The sonic line appears to coincide with the first Mach wave from the upper corner expansion fan after transmitted from the reflected shock wave.

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Cassel, K. W., A. I. Ruban i J. D. A. Walker. "An instability in supersonic boundary-layer flow over a compression ramp". Journal of Fluid Mechanics 300 (10.10.1995): 265–85. http://dx.doi.org/10.1017/s0022112095003685.

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Separation of a supersonic boundary layer (or equivalently a hypersonic boundary layer in a region of weak global interaction) near a compression ramp is considered for moderate wall temperatures. For small ramp angles, the flow in the vicinity of the ramp is described by the classical supersonic triple-deck structure governing a local viscous-inviscid interaction. The boundary layer is known to exhibit recirculating flow near the corner once the ramp angle exceeds a certain critical value. Here it is shown that above a second and larger critical ramp angle, the boundary-layer flow develops an instability. The instability appears to be associated with the occurrence of inflection points in the streamwise velocity profiles within the recirculation region and develops as a wave packet which remains stationary near the corner and grows in amplitude with time.

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Liu, Yi, Zhi Guo Dou i Li Wei Duan. "Numerical Investigation of Cavity Flow Field Characteristics in Supersonic Flow". Applied Mechanics and Materials 789-790 (wrzesień 2015): 368–72. http://dx.doi.org/10.4028/www.scientific.net/amm.789-790.368.

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The cold flow field in a two dimensional cavity of supersonic combustor has been simulated numerically by using the compressible flow Navier-Stokes equation with theκ-ωSST turbulence model. The flow field structure of different cavity aft wall slope angle (16°,30° and 90°) , different fore aft wall height ratio (1 and 2) and different length depth ratio (3 and 5) are analyzed. The conclusions are as follows: As cavity aft wall slope angle decreases, the compression wave formed at cavity leading separation corner shifts into expansion wave, the shear layer moves into cavity gradually; As cavity fore aft wall height ratio increases from one to two, the expansion wave formed at cavity leading separation corner strengthens and there is no compression wave formed at;As cavity length depth ratio increases from three to five, the compression or expansion wave formed at cavity leading separation corner weakens, cavity bottom wall pressure tends to be constant and aft wall pressure rises.

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Niu, Keishiro. "Implosion motion and fuel compression in direct or indirect driven target". Laser and Particle Beams 7, nr 3 (sierpień 1989): 505–9. http://dx.doi.org/10.1017/s0263034600007473.

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When a one-shell three-layer cryogenic target is irradiated by a driver beam of total energy 10 MJ and pulse width 30 ns, the pusher pressure increases to 1013 Pa, accelerating fuel toward target center, and the fuel implosion velocity reaches 3 × 105 m/s. A spherical hollow target plays the role of a supersonic converging nozzle, and the fuel is compressed to 269 times the solid density in the supersonic region and to 3·51 × 104 times in subsonic region. Nonuniform beam-energy-deposition in pusher layer causes nonuniform pusher pressure and hence nonuniform implosion, which reduces fuel compression significantly. The smoothing of pusher pressure by radiative energy transfer, or gas-filled target instead of cryogenic hollow target can be used to reduce the defect of nonuniform implosion. At last, the structure of an indirect driven target is proposed to smooth out pusher pressure in spite of nonuniform beam irradiation.

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Watanabe, Yasumasa, Alec Houpt i Sergey Leonov. "Plasma-Assisted Control of Supersonic Flow over a Compression Ramp". Aerospace 6, nr 3 (12.03.2019): 35. http://dx.doi.org/10.3390/aerospace6030035.

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This study considers the effect of an electric discharge on the flow structure near a 19.4° compression ramp in Mach-2 supersonic flow. The experiments were conducted in the supersonic wind tunnel SBR-50 at the University of Notre Dame. The stagnation temperature and pressure were varied in a range of 294–600 K and 1–3 bar, respectively, to attain various Reynolds numbers ranging from 5.3 × 105 to 3.4 × 106 based on the distance between the exit of the Mach-2 nozzle and the leading edge of the ramp. Surface pressure measurements, schlieren visualization, discharge voltage and current measurements, and plasma imaging with a high-speed camera were used to evaluate the plasma control authority on the ramp pressure distribution. The plasma being generated in front of the compression ramp shifted the shock position from the ramp corner to the electrode location, forming a flow separation zone ahead of the ramp. It was found that the pressure on the compression surface reduced almost linearly with the plasma power. The ratio of pressure change to flow stagnation pressure was also an increasing function of the ratio of plasma power to enthalpy flux, indicating that the task-related plasma control effectiveness ranged from 17.5 to 25.

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Rozprawy doktorskie na temat "Supersonic compression structure"

Kempf, Severin Gabriel. "Numerical Study of the Stability of Embedded Supersonic Compressor Stages". Thesis, Virginia Tech, 2003. http://hdl.handle.net/10919/34506.

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A numerical case study of a multistage compressor with relative supersonic rotors is presented. The purpose of the investigation was to determine the flow instability mechanism of the UEET compressor and its relation to the rotor shock structure in the relative velocity reference frame. The computational study was conducted with the NASA code ADPAC , utilizing the mixing-plane assumption for the boundary condition between adjacent, relatively-rotating blade rows. A steady, five-blade-row, numerical simulation using the Baldwin-Lomax turbulence model was performed, creating several constant speed lines. The results are presented, highlighting the role shock structure plays in the stability of the compressor. The shock structure in the downstream rotor isolates the upstream rotor from the exit conditions until the shock detaches from the leading edge. At this point the shock structure in the upstream rotor moves, changing the conditions for the downstream rotor. This continues with increasing pressure at the exit until the shock in the upstream rotor detaches from the leading edge. This event causes an instantaneous drop in the mass flow rate, initiating positive incident separation on the suction side of stator-two.
Master of Science

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Książki na temat "Supersonic compression structure"

Rostafiński, Wojciech. Analysis of fully stalled compressor. [Cleveland, Ohio: National Aeronautics and Space Administration, Lewis Research Center, 1986.

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Streszczenia konferencji na temat "Supersonic compression structure"

Lin, Eric P., i James C. Hermanson. "Compression Wave Structure on Droplets under Supersonic Conditions". W 50th AIAA/ASME/SAE/ASEE Joint Propulsion Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2014. http://dx.doi.org/10.2514/6.2014-3946.

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Cheng, Yao, Zhansheng Liu, Ruixian Ma i Guanghui Zhang. "Numerical Vibration Analysis of Supersonic Mixed-Compression Intake". W ASME Turbo Expo 2013: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/gt2013-95444.

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Vibration analysis of a supersonic mixed compression intake is carried out under periodic back pressure by using one-way fluid structure interaction method. Steady flow and transient flow in the intake are simulated with finite volume method. The total pressure recovery and the exit flow distortion of the steady flow solution are very close to Anderson and Wong’s experiment results. The frequency components of pressure transferring to the intake cowl wall are analyzed using FFT method for the unsteady flow, and are compared with the natural frequencies of the intake structure to avoid the resonance of the intake structure. The dynamic response of the intake structure is simulated with finite element method considering geometric nonlinearity. The displacement history of structure is obtained, and the region with large amplitude of vibration is identified. It indicates that the forced vibration occurs in the intake under the excitation of the oscillating back pressure. The static displacement is two times of the cowl wall thickness, while the peak value of the dynamic displacement is 10% of the cowl wall thickness.

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Zapryagaev, V. I., I. N. Kavun i L. P. Trubitsyna. "Structure of a supersonic separated flow over a compression corner with side walls". W INTERNATIONAL CONFERENCE ON THE METHODS OF AEROPHYSICAL RESEARCH (ICMAR 2020). AIP Publishing, 2021. http://dx.doi.org/10.1063/5.0052018.

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Menaa, Mohamed. "On the Structure of Turbulence in a Supersonic Compression Surface Using Reynolds Stress Model". W 14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2006. http://dx.doi.org/10.2514/6.2006-7942.

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TROTSYUK, A. V. "DETONATION STRUCTURES IN A SUPERSONIC ANNULAR RAMJET CHAMBER". W International Colloquia on Pulsed and Continuous Detonations. TORUS PRESS, 2021. http://dx.doi.org/10.30826/icpcd12b09.

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The results of a systematic study of the structures and flow regimes with an oblique detonation wave (ODW) in an annular ramjet straight-flow detonation chamber (DC) of a new type have been presented. In the combustor with a new design, detonation burning of the reacting mixture is organized by using a compression body shaped as a continuous monofilar spiral with a constant pitch angle. Numerical simulations are performed for a supersonic flow of a stoichiometric hydrogen–air mixture with Mach number M0=3 and 5 at the combustor entrance. A mathematical model of the reacting flow in the combustor is developed in a two-dimensional (2D) unsteady formulation. The flow dynamics for the two different operation start types of DC and the final structure of the steady flow in the combustor are numerically studied. Various geometric parameters of the DC are considered (length, radius, and spiral pitch angle). A bifurcation of the steady flow structures with respect to the initial conditions of combustor start is detected for some combinations of the DC geometric parameters.

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Yang, Ling, Jingjun Zhong i Ji-ang Han. "Numerical Research of the Ram-Rotor With Different Geometric Parameters". W ASME 2011 Turbo Expo: Turbine Technical Conference and Exposition. ASMEDC, 2011. http://dx.doi.org/10.1115/gt2011-46051.

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With the design methods of typical supersonic aircraft intakes and the advantages of shock wave compression, ram-rotors have become a new attractive compression system. Lots of research work has been carried out on rampressors, but the influence of the geometric parameters on the shock wave structure and compression performance of the ram-rotor has not been studied systematically. Therefore, a thorough study on ram-rotor with different geometric parameters is required. In this paper, a steady three-dimensional Navier-Stokes equation adopted in the “Fluent” software package is carried out on a large parallel computer. Six factors which may influence the ram-rotor performance are investigated numerically. These geometric parameters are strake section shape, throat length-height ratio, strake stagger angle, compression ramp angle, subsonic divergent angle and throat contraction ratio. The study is composed of two parts. The aim of the first part is to understand the influence of the geometric parameters listed above on the shock wave structure and compression performance of the ram-rotor by comparison and analysis of the relative Mach number and static pressure in the flow-path. The aim of the second part is to obtain the optimal geometric structure of the ram-rotor by comparison and analysis of the structure of the flow fields, the compression performance and the ram-rotor properties. First of all, the numerical method is validated by comparing the numerical results of the flow field of a supersonic intake with experimental results in this paper. Secondly, the flow field structures in the ram-rotor, especially the number and position of shock waves and the separation zone, are studied. Thirdly, the influence of the geometric parameters on the rotor performance is studied. Some parameter distributions, such as the flow angle, adiabatic efficiency, total pressure ratio, total pressure recovery coefficient, are compared and analyzed. The rules of the ram-rotor performance variation with different geometric parameters are also presented. Finally, some advice for improving the overall performance of the ram-rotor is given according to the flow field analysis.

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Alhussan, Khaled. "Study the Structure of Three Dimensional Oblique Shock Waves Over Conical Rotor-Vane Surfaces". W ASME 2005 Fluids Engineering Division Summer Meeting. ASMEDC, 2005. http://dx.doi.org/10.1115/fedsm2005-77440.

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This paper will explain the numerical analysis and the mapping of the flow in a confined region. In this paper some characteristics of non-steady, compressible, flow are explored, including compression and expansion wave interactions and creation. The results will show a promising achievement, first, to understand the flow structure inside a supersonic confined region, second, to use this knowledge to interpolate the numerical results in order to achieve a design methodology that will benefit the industrial applications for example in turbomachinery. Results including contour plots of static pressure, total pressure, and Mach number will show the structure of oblique shock waves in a complex three-dimensional conical surface. A CFD analysis enables one to understand the complex flow structure inside this confined region. Through this computational analysis, a better interpretation of the physical phenomenon of the three dimensional rotting oblique shock waves can be achieved. It is essential to evaluate the ability of numerical technique that can solve problems in which compression and expansion waves occur. In particular it is necessary to understand the details of developing a mesh that will allow resolution of some discontinuities in similar flow.

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Yang, Ling, Jingjun Zhong i Ji-ang Han. "Numerical Research on the Flow Field and Performance of a Ram-Rotor and a Scrampressor". W ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-69358.

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The design methods of typical supersonic aircraft intakes and shock wave compression technology have been applied to ram-rotor, a new attractive compression system. A ram-rotor is a typical structure including the compression ramp, the throat and the subsonic diffuser; a scrampressor is similar to ram-rotor, the only different is that scrampressor has no subsonic diffuser. Base on the preparatory work, it has been found that these two structures have different advantages respectively. So, in this paper, the three dimensional Reynolds-averaged Navier-Stokes equations and the Spalart-Allmaras turbulent model are used to simulate numerically the flow field of the ram-rotor and the scrampressor at the design and at the off-design conditions. The back pressure and rotational speed are mainly considered which may affect the flow field and the total performance. It has been found that back pressure can not have influence on the flow field before the throat outlet obviously. With increasing of the back pressure, the position of the flow separation zone and shock train move forwards to the inlet. The rotational speed changes the shock wave structure of the ram-rotor and scrampressor evidently. With the rotational speed increasing, each shock wave moves to the outlet and the shock wave number decreases. The ram-rotor and scrampressor structure is similar, except the ram-rotor flow structure has a large flow separation zone after the throat outlet. The compression capability of the ram-rotor is higher than that of the scrampressor. The total performance of the scrampressor is better than the ram-rotor.

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Naziar, Javaid, Rich Couch i Milt Davis. "An Approach for the Development of an Aerodynamic-Structural Interaction Numerical Simulation for Aeropropulsion Systems". W ASME 1996 International Gas Turbine and Aeroengine Congress and Exhibition. American Society of Mechanical Engineers, 1996. http://dx.doi.org/10.1115/96-gt-480.

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Traditionally, aeropropulsion structural performance and aerodynamic performance have been designed separately and later mated together via flight testing. In today’s atmosphere of declining resources, it is imperative that more productive ways of designing and verifying aeropropulsion performance and structural interaction be made available to the aerospace industry. One method of obtaining a more productive design and evaluation capability is through the use of numerical simulations. Currently, Lawrence Livermore National Laboratory has developed a generalized fluid/structural interaction code known as ALE3D. This code is capable of characterizing fluid and structural interaction for components such as the combustor, fan/stators, inlet and/or nozzles. This code solves the 3D Euler equations and has been applied to several aeropropulsion applications such as a supersonic inlet and a combustor rupture simulation. To characterize aerodynamic-structural interaction for rotating components such as the compressor, appropriate turbomachinery simulations would need to be implemented within the ALE3D structure. The Arnold Engineering Development Center is currently developing a three-dimensional compression system code known as TEACC (Turbine Engine Analysis Compressor Code). TEACC also solves the 3D Euler equations and is intended to simulate dynamic behavior such as inlet distortion, surge or rotating stall. The technology being developed within the TEACC effort provides the necessary turbomachinery simulation for implementation into ALE3D. This paper describes a methodology to combine three-dimensional aerodynamic turbomachinery technology into the existing aerodynamic-structural interaction simulation, ALE3D to obtain the desired aerodynamic and structural integrated simulation for an aeropropulsion system.

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Alhussan, Khaled. "Oblique Shock Waves Interaction in a Non-Steady Three Dimensional Rotating Flow". W ASME 2005 Fluids Engineering Division Summer Meeting. ASMEDC, 2005. http://dx.doi.org/10.1115/fedsm2005-77442.

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This paper will explain the numerical analysis and the mapping of the flow in a confined region. In this paper some characteristics of non-steady, compressible, flow are explored, including compression and expansion wave interactions and creation. The results will show a promising achievement, first, to understand the flow structure inside a supersonic confined region, second, to use this knowledge to interpolate the numerical results in order to achieve a design methodology that will benefit the industrial applications for example in turbomachinery. Results including contour plots of static pressure, total pressure, Mach number, temperature and velocity vectors will show the structure of rotating oblique shock waves in a complex three-dimensional conical surface. A CFD analysis enables one to understand the complex flow structure inside this confined region. Through this computational analysis, a better interpretation of the physical phenomenon of the three dimensional rotting oblique shock waves can be achieved. It is essential to evaluate the ability of numerical technique that can solve problems in which compression and expansion waves occur. In particular it is necessary to understand the details of developing a mesh that will allow resolution of some discontinuities in similar flow.

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