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Статті в журналах з теми "Supersonic / hypersonic"

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Автор: Grafiati

Опубліковано: 7 липня 2024

Оновлено: 7 липня 2024

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1

Yang, Lung-Jieh, and Chao-Kang Feng. "A Unified Asymptotic Theory of Supersonic, Transonic, and Hypersonic Far Fields." Axioms 11, no. 11 (November 19, 2022): 656. http://dx.doi.org/10.3390/axioms11110656.

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The problems of steady, inviscid, isentropic, irrotational supersonic plane flow passing a body with a small thickness ratio was solved by the linearized theory, which is a first approximation at and near the surface but fails at far fields from the body. Such a problem with far fields was solved by W.D. Hayes’ “pseudo-transonic” nonlinear theory in 1954. This far field small disturbance theory is reexamined in this study first by using asymptotic expansion theory. A systematic approach is adopted to obtain the nonlinear Burgers’ equation for supersonic far fields. We also use the similarity method to solve this boundary value problem (BVP) of the inviscid Burgers’ equation and obtain the nonlinear flow patterns, including the jump condition for the shock wave. Secondly, the transonic and hypersonic far field equations were obtained from the supersonic Burgers’ equation by stretching the coordinate in the y direction and considering an expansion of the freestream Mach number in terms of the transonic and hypersonic similarity parameters. The mathematical structures of the far fields of the supersonic, transonic, and hypersonic flows are unified to be the same. The similar far field flow patterns including the shock positions of a parabolic airfoil for the supersonic, transonic, and hypersonic flow regimes are exemplified and discussed.

2

Li, Yong Hong, Xin Wu Tang, and Wei Qun Zhou. "Aerodynamic and Numerical Study on the Influence of Spike Shapes at Mach 1.5." Advanced Materials Research 1046 (October 2014): 177–81. http://dx.doi.org/10.4028/www.scientific.net/amr.1046.177.

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Taking into account the issue of configuration or aerodynamic heating, most supersonic and hypersonic flight vehicles have to use the blunt-nosed body. However, in supersonic especially in hypersonic flow the strong bow shock ahead of the blunt nose introduces a rather high shock drag that affects the aerodynamic performance of the vehicles seriously. A spike mounted on a blunt body during its flight pushes the strong bow shock away from the body surface and forms recirculation flow with low pressure ahead of the body surface, and then decreases the drag. The drag reduction effects of spikes in high supersonic and hypersonic flow had been validated through experimental and numerical methods. In order to analyze the influence of the spike on aerodynamic characteristics at low supersonic (M=1.5) flow past blunt-nosed bodies, numerical studies were carried out which included the influence of the spike shape, the analysis of the fluid flow structures and the effect on the aerodynamic characteristics of a blunt body.

3

de Araujo Martos, João Felipe, Israel da Silveira Rêgo, Sergio Nicholas Pachon Laiton, Bruno Coelho Lima, Felipe Jean Costa, and Paulo Gilberto de Paula Toro. "Experimental Investigation of Brazilian 14-X B Hypersonic Scramjet Aerospace Vehicle." International Journal of Aerospace Engineering 2017 (2017): 1–10. http://dx.doi.org/10.1155/2017/5496527.

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The Brazilian hypersonic scramjet aerospace vehicle 14-X B is a technological demonstrator of a hypersonic airbreathing propulsion system based on the supersonic combustion (scramjet) to be tested in flight into the Earth’s atmosphere at an altitude of 30 km and Mach number 7. The 14-X B has been designed at the Prof. Henry T. Nagamatsu Laboratory of Aerothermodynamics and Hypersonics, Institute for Advanced Studies (IEAv), Brazil. The IEAv T3 Hypersonic Shock Tunnel is a ground-test facility able to produce high Mach number and high enthalpy flows in the test section close to those encountered during the flight of the 14-X B into the Earth’s atmosphere at hypersonic flight speeds. A 1 m long stainless steel 14-X B model was experimentally investigated at T3 Hypersonic Shock Tunnel, for freestream Mach numbers ranging from 7 to 8. Static pressure measurements along the lower surface of the 14-X B, as well as high-speed Schlieren photographs taken from the 5.5° leading edge and the 14.5° deflection compression ramp, provided experimental data. Experimental data was compared to the analytical theoretical solutions and the computational fluid dynamics (CFD) simulations, showing good qualitative agreement and in consequence demonstrating the importance of these methods in the project of the 14-X B hypersonic scramjet aerospace vehicle.

4

Milthorpe, J. F. "Simulation of supersonic and hypersonic flows." International Journal for Numerical Methods in Fluids 14, no. 3 (February 15, 1992): 267–88. http://dx.doi.org/10.1002/fld.1650140303.

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5

Xiao, Han-shan, Chao Ou, Hong-liang Ji, Zheng-chun He, Ning-yuan Liu, and Xian-xu Yuan. "Low-Cost and Aerodynamics-Aim Hypersonic Flight Experiment MF-1." MATEC Web of Conferences 316 (2020): 04006. http://dx.doi.org/10.1051/matecconf/202031604006.

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For increasing understanding of fundamental hypersonic phenomena, the flight test program, named MF-1, is to gather fundamental scientific and engineering data on the physics and technologies critical to future operational hypersonic flight with low-cost flight test platform, which is built on the retrofitted rockets. The MF-1 program is a hypersonic flight test program executed by China Aerodynamic Research and Development Center (CARDC). The MF-1 flight flew in December 2015. The flight focuses primarily on integration of instrumentation on the test vehicle, with application to boundary layer transition and shock interaction experiments. The MF-1 payload consists of a blunted 7°half angle cone, a cylinder and 33° flare configuration. The payload was boosted to Mach 5.32 utilizing a solid-rocket booster without control for the whole flight. The flight was fully successful, and measured transition under supersonic and hypersonic conditions. The heat flux data were given by the three-dimensional thermal identification method to discriminate transition zone. The preliminary analysis shows that the real-time flight data obtained by MF-1 are reliable and can be used to validate the transition predicting model and software. The results show that the existing model is able to predict the transition location of cone at a small angle-of-attack for supersonic or hypersonic flow. This paper describes the MF-1 mission and some general conclusions derived from the experiment.

6

Verhoff, A., and D. Stookesberry. "Prediction of inviscid supersonic/hypersonic aircraft flowfields." Journal of Aircraft 29, no. 4 (July 1992): 581–87. http://dx.doi.org/10.2514/3.46205.

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7

Huang, Wei, Jun-tao Chang, and Li Yan. "Mixing and combustion in supersonic/hypersonic flows." Journal of Zhejiang University-SCIENCE A 21, no. 8 (August 2020): 609–13. http://dx.doi.org/10.1631/jzus.a20mcsf1.

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8

Hu, Jiasen, and Arthur Rizzi. "Turbulent flow in supersonic and hypersonic nozzles." AIAA Journal 33, no. 9 (September 1995): 1634–40. http://dx.doi.org/10.2514/3.12861.

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9

James, Anthony. "Hot Property." Aerospace Testing International 2018, no. 3 (September 2018): 48–52. http://dx.doi.org/10.12968/s1478-2774(23)50116-2.

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With growing commercial interest in space, as well as in the development of supersonic and hypersonic aircraft, the unique capabilities of the Scirocco Plasma Wind Tunnel in Italy are increasingly in demand

10

Zhao, Lian Jin, Jia Lin, Jian Hua Wang, Jin Long Peng, De Jun Qu, and Lian Zhong Chen. "An Experimental Investigation on Transpiration Cooling for Supersonic Vehicle Nose Cone Using Porous Material." Applied Mechanics and Materials 541-542 (March 2014): 690–94. http://dx.doi.org/10.4028/www.scientific.net/amm.541-542.690.

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During hypersonic flight or cruise in the near space, the aerodynamic heating causes a very high temperature on the leading edge of hypersonic vehicles. Transpiration cooling has been recognized the most effective cooling technology. This paper presents an experimental investigation on transpiration cooling using liquid water as coolant for a nose cone model of hypersonic vehicles. The nose cone model consists of sintered porous material. The experiments were carried out in the Supersonic Jet Arc-heated Facility (SJAF) of China Academy of Aerospace Aerodynamics (CAAA) in Beijing. The cooling effect in the different regions of the model was analyzed, and the shock wave was exhibited. The pressure variations of the coolant injection system were continuously recorded. The aim of this work is to provide a relatively useful reference for the designers of coolant driving system in practical hypersonic vehicles.

11

Weidner, J. P. "Hypersonic Propulsion-breaking the Thermal Barrier." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 207, no. 1 (January 1993): 47–59. http://dx.doi.org/10.1243/pime_proc_1993_207_246_02.

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The challenges of hypersonic propulsion impose unique features on the hypersonic vehicle—from large volume requirements to contain cryogenic fuel to airframe-integrated propulsion required to process sufficient quantities of air. Additional challenges exist in the design of the propulsion module that must be capable of efficiently processing air at very high enthalpies, adding and mixing fuel at supersonic speeds and expanding the exhaust products to generate thrust greater than drag. The paper explores the unique challenges of the integrated hypersonic propulsion system, addresses propulsion cycle selection to cope with the severe thermal environment and reviews the direction of propulsion research at hypervelocity speeds.

12

Rosato, Daniel A., Mason Thornton, Jonathan Sosa, Christian Bachman, Gabriel B. Goodwin, and Kareem A. Ahmed. "Stabilized detonation for hypersonic propulsion." Proceedings of the National Academy of Sciences 118, no. 20 (May 10, 2021): e2102244118. http://dx.doi.org/10.1073/pnas.2102244118.

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Future terrestrial and interplanetary travel will require high-speed flight and reentry in planetary atmospheres by way of robust, controllable means. This, in large part, hinges on having reliable propulsion systems for hypersonic and supersonic flight. Given the availability of fuels as propellants, we likely will rely on some form of chemical or nuclear propulsion, which means using various forms of exothermic reactions and therefore combustion waves. Such waves may be deflagrations, which are subsonic reaction waves, or detonations, which are ultrahigh-speed supersonic reaction waves. Detonations are an extremely efficient, highly energetic mode of reaction generally associated with intense blast explosions and supernovas. Detonation-based propulsion systems are now of considerable interest because of their potential use for greater propulsion power compared to deflagration-based systems. An understanding of the ignition, propagation, and stability of detonation waves is critical to harnessing their propulsive potential and depends on our ability to study them in a laboratory setting. Here we present a unique experimental configuration, a hypersonic high-enthalpy reaction facility that produces a detonation that is fixed in space, which is crucial for controlling and harnessing the reaction power. A standing oblique detonation wave, stabilized on a ramp, is created in a hypersonic flow of hydrogen and air. Flow diagnostics, such as high-speed shadowgraph and chemiluminescence imaging, show detonation initiation and stabilization and are corroborated through comparison to simulations. This breakthrough in experimental analysis allows for a possible pathway to develop and integrate ultra-high-speed detonation technology enabling hypersonic propulsion and advanced power systems.

13

Baysal, Oktay, and Wendy B. Hoffman. "Simulation of Three-Dimensional Shear Flow Around a Nozzle-Afterbody at High Speeds." Journal of Fluids Engineering 114, no. 2 (June 1, 1992): 178–85. http://dx.doi.org/10.1115/1.2910013.

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Turbulent shear flows at supersonic and hypersonic speeds around a nozzle-afterbody are simulated. The three-dimensional, Reynolds-averaged Navier-Stokes equations are solved by a finite-volume and implicit method. The convective and the pressure terms are differenced by an upwind-biased algorithm. The effect of turbulence is incorporated by a modified Baldwin-Lomax eddy viscosity model. The success of the standard Baldwin-Lomax model for this flow type is shown by comparing it to a laminar case. These modifications made to the model are also shown to improve flow prediction when compared to the standard Baldwin-Lomax model. These modifications to the model reflect the effects of high compressibility, multiple walls, vortices near walls, and turbulent memory effects in the shear layer. This numerically simulated complex flowfield includes a supersonic duct flow, a hypersonic flow over an external double corner, a flow through a non-axisymmetric, internal-external nozzle, and a three-dimensional shear layer. The specific application is for the flow around the nozzle-afterbody of a generic hypersonic vehicle powered by a scramjet engine. The computed pressure distributions compared favorably with the experimentally obtained surface and off-surface flow surveys.

14

Punniyakotti, S., and J. L. Stollery. "The estimation of aerodynamic forces on flat plate aerofoils at hypersonic and supersonic speed." Aeronautical Journal 116, no. 1185 (November 2012): 1207–15. http://dx.doi.org/10.1017/s0001924000007570.

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Abstract Existing theories are reviewed and compared. From the Tangent Wedge formula a simple expression for the normal force on thin wings is derived and compared with some existing experimental data, which cover the whole supersonic-hypersonic range of Mach numbers.

15

Nishio, Masatomi, Keiji Manabe, and Hiroaki Nakamura. "New Calculation Method of Supersonic/Hypersonic Flow: Application to Hypersonic MESUR Capsule Flowfield." Journal of Spacecraft and Rockets 43, no. 4 (July 2006): 916–18. http://dx.doi.org/10.2514/1.17817.

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16

Nishio, Masatomi, Keiji Manabe, Hiroaki Nakamura, and Shinji Sezaki. "A New Calculation Method of Supersonic/Hypersonic Flow." JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 51, no. 599 (2003): 683–89. http://dx.doi.org/10.2322/jjsass.51.683.

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17

Settles, Gary S., and Lori J. Dodson. "Supersonic and hypersonic shock/boundary-layer interaction database." AIAA Journal 32, no. 7 (July 1994): 1377–83. http://dx.doi.org/10.2514/3.12205.

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18

Neuenhahn, T., and H. Olivier. "Laminar incipient separation in supersonic and hypersonic flows." International Journal of Aerodynamics 2, no. 2/3/4 (2012): 114. http://dx.doi.org/10.1504/ijad.2012.049138.

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19

Spall, R. E., and M. R. Malik. "Goertler vortices in supersonic and hypersonic boundary layers." Physics of Fluids A: Fluid Dynamics 1, no. 11 (November 1989): 1822–35. http://dx.doi.org/10.1063/1.857508.

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20

El‐Hady, Nabil M. "Nonparallel instability of supersonic and hypersonic boundary layers." Physics of Fluids A: Fluid Dynamics 3, no. 9 (September 1991): 2164–78. http://dx.doi.org/10.1063/1.857898.

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21

Osiptsov, A. N., and M. A. Teverovskii. "Hypersonic flow past a supersonic two-phase source." Fluid Dynamics 33, no. 3 (May 1998): 407–18. http://dx.doi.org/10.1007/bf02698193.

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22

Cecere, D., A. Ingenito, E. Giacomazzi, L. Romagnosi, and C. Bruno. "Hydrogen/air supersonic combustion for future hypersonic vehicles." International Journal of Hydrogen Energy 36, no. 18 (September 2011): 11969–84. http://dx.doi.org/10.1016/j.ijhydene.2011.06.051.

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23

Masad, Jamal A., and Ridha Abid. "On transition in supersonic and hypersonic boundary layers." International Journal of Engineering Science 33, no. 13 (October 1995): 1893–919. http://dx.doi.org/10.1016/0020-7225(95)00046-z.

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24

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.

25

Gai, S. L. "Some features of steady separated flow from low speed to hypersonic." Aeronautical Journal 112, no. 1128 (February 2008): 109–13. http://dx.doi.org/10.1017/s0001924000002049.

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Steady non-vortex shedding base flow behind a bluff body is considered. Such a flow is characterised by the flow separation at the trailing edge of the body with an emerging shear layer which reattaches on the axis with strong recompression and recirculating flow bounded by the base, the shear layer, and the axis. Steady wake flows behind a bluff body at low speeds have been studied for more than a century (for example, Kirchhoff; Riabouchinsky). Recently, research on steady bluff body wake flow at low speeds has been reviewed and reinterpreted by Roshko. Roshko has also commented on some basic aspects of steady supersonic base flow following on from Chapman and Korst analyses. In the present paper, we examine the steady base flow features both at low speeds and supersonic speeds in the light of Roshko’s model and expand on some further aspects of base flows at supersonic and hypersonic speeds, not covered by Roshko.

26

Niu, Yao Bin, and Zhong Wei Wang. "Flutter Analysis of Hypersonic Wings Subject to Thermal Load." Applied Mechanics and Materials 105-107 (September 2011): 466–69. http://dx.doi.org/10.4028/www.scientific.net/amm.105-107.466.

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Aeroelastic analysis plays a vital role in the design of hypersonic wings. The thermal flutter of hypersonic tail under different materials at different temperature is analyzed by using MSC.NASTRAN, and the hypersonic piston theory is used for the modeling of supersonic aerodynamic loads. The research results showed: nature frequency is reduced as temperature increased, elevated temperature tend to reduce model stiffness due to changes in material properties and development of thermal stresses. Increased temperature will decrease the flutter velocity, flutter speed of tail is different with different material of titanium, but the trend and the value of variety of flutter speed are same as temperature increase. When material is different, the value of variety of flutter speed is also different.

27

Mei, Chuh, K. Abdel-Motagaly, and R. Chen. "Review of Nonlinear Panel Flutter at Supersonic and Hypersonic Speeds." Applied Mechanics Reviews 52, no. 10 (October 1, 1999): 321–32. http://dx.doi.org/10.1115/1.3098919.

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A review of various analytical methods and experimental results of supersonic and hypersonic panel flutter is presented. The analytical methods are categorized into two main methods. The first category is the classical methods, which include Galerkin in conjunction with numerical integration, harmonic balance and perturbation methods. The second category is the finite element methods in either the frequency domain (eigensolution) or the time domain (numerical integration). A review of the experimental literature is given. The effects of different parameters on the flutter behavior are described. The parameters considered include inplane forces, thermal loading, flow direction, and initial curvature. Active control of composite panels at supersonic speeds and elevated temperatures is also considered. This review article cites 84 references.

28

Leite, Bernardo, Frederico Afonso, and Afzal Suleman. "Aerodynamic Shape Optimization of a Symmetric Airfoil from Subsonic to Hypersonic Flight Regimes." Fluids 7, no. 11 (November 15, 2022): 353. http://dx.doi.org/10.3390/fluids7110353.

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Hypersonic flight has been the subject of numerous research studies during the last eight decades. This work aims to optimize the aerodynamic performance of a two-dimensional baseline airfoil (NACA0012) at distinct flight regimes from subsonic to hypersonic speeds. A mission profile has been defined, where four points representing the subsonic, transonic, supersonic, and hypersonic flow conditions have been selected. A framework has been implemented based on high-fidelity RANS computational fluid dynamics simulations. Gradient-based optimizations have been conducted with the objective of minimizing the drag. The optimization results show an overall improvement in aerodynamic performance, including a decrease in the drag coefficient of up to 79.2% when compared to the baseline airfoil. In the end, a morphing strategy has been laid out based on the optimal shapes produced by the optimization.

29

Iannelli, G. S., and A. J. Baker. "Accuracy and Efficiency Assessments for a Weak Statement CFD Algorithm for High-Speed Aerodynamics." Journal of Engineering for Gas Turbines and Power 116, no. 3 (July 1, 1994): 468–73. http://dx.doi.org/10.1115/1.2906844.

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A bilinear finite element, implicit Runge-Kutta space-time discretization has been established for an aerodynamics weak statement CFD algorithm. The algorithm admits real-gas effect simulation, for reliable hypersonic flow characterization, via an equilibrium reacting air model. The terminal algebraic system is solved using an efficient block-tridiagonal quasi-Newton linear algebra procedure that employs tensor matrix product factorizations within a lexicographic mesh-sweeping protocol. A block solution-adaptive remeshing, for totally arbitrary convex elements, is also utilized to facilitate accurate shock and/or boundary layer flow resolution. Numerical validations are presented for representative benchmark supersonic-hypersonic aerodynamics problem statements.

30

Yang, Tian-Peng, Jiang-Feng Wang, Fa-Ming Zhao, Xiao-Feng Fan, and Yu-Han Wang. "Numerical analysis of exhaust jet secondary combustion in hypersonic flow field." Modern Physics Letters B 32, no. 12n13 (May 10, 2018): 1840045. http://dx.doi.org/10.1142/s0217984918400456.

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The interaction effect between jet and control surface in supersonic and hypersonic flow is one of the key problems for advanced flight control system. The flow properties of exhaust jet secondary combustion in a hypersonic compression ramp flow field were studied numerically by solving the Navier–Stokes equations with multi-species and combustion reaction effects. The analysis was focused on the flow field structure and the force amplification factor under different jet conditions. Numerical results show that a series of different secondary combustion makes the flow field structure change regularly, and the temperature increases rapidly near the jet exit.

31

Ito, M. "International collaboration in super/hypersonic propulsion system research project (HYPR)." Aeronautical Journal 104, no. 1040 (October 2000): 445–51. http://dx.doi.org/10.1017/s0001924000091934.

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Abstract A ten-year engineering research project on a super/hypersonic transport propulsion system (HYPR) initiated by AIST (Agency of Industrial Science and Technology) in 1989 has been successfully completed. The HYPR project was aimed at establishing the technology base for a propulsion system of supersonic/hypersonic transport (SST/HST) aircraft that will fly as fast as Mach 5. The project proved the feasibility of a combined cycle engine (CCE) for SST/HST by developing and running the world's first CCE , a combination of a variable cycle turbofan engine (VCE) and a methane-fuelled ramjet. The development of the CCE began with a variety of component tests, followed by the development of the high temperature core engine (HTCE) and VCE. This step-by-step approach led to steady and solid success. Based on this success, a new project was started in 1999, titled "Research and development of environmentally compatible propulsion system for next generation supersonic transport (ESPR)." This paper introduces the various achievements during the ten years of the HYPR project, and the prospects of the upcoming ESPR project.

32

Meaburn, J., M. Bryce, D. J. Harman, and J. A. López. "Ubiquitous High Speed Ejecta in Pne - MyCn 18." Symposium - International Astronomical Union 209 (2003): 515. http://dx.doi.org/10.1017/s0074180900209509.

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Supersonic and even hypersonic outflows have been found in a wide variety of PNe using the Manchester Echelle spectrometer (MES; Meaburn et al. 1984). Dramatic examples of these extreme motions found over many years are highlighted here for comparison with their most recent discovery in the bipolar PN, MyCn 18.

33

Mateescu, Dan. "Explicit Exact and Third-Order-Accurate Pressure-Deflection Solutions for Oblique Shock and Expansion Waves." Open Aerospace Engineering Journal 3, no. 1 (February 18, 2010): 1–8. http://dx.doi.org/10.2174/1874146001003010001.

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This paper presents explicit analytical solutions of the pressure coefficient and the pressure ratio across the oblique shock and expansion waves in function of the flow deflection angle. These new explicit pressure-deflection solutions can be efficiently used in solving applied aerodynamic problems in supersonic flows, such as the aerodynamics of airfoils and wings in supersonic-hypersonic flows and the shock and expansion waves interactions, and can be also used to increase the computational efficiency of the numerical methods based on the Riemann problem solution requiring the pressure-deflection solution of the oblique shock and expansion waves, such as the Godunov method.

34

Mortensen, Clifton H. "Toward an understanding of supersonic modes in boundary-layer transition for hypersonic flow over blunt cones." Journal of Fluid Mechanics 846 (May 10, 2018): 789–814. http://dx.doi.org/10.1017/jfm.2018.246.

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Realistic flight vehicles for reentry into the Earth’s atmosphere are commonly similar to blunted cones. The main reason for blunting a cone is to mitigate high heat loads at the nose. Another reason for blunting the cone is to delay boundary-layer transition. It is commonly understood that the second mode is damped in flow over a cone as the nose radius is increased. This is thought to lead to the delay in transition. Here, a blunted cone at a realistic reentry trajectory point with significant real-gas effects is studied. It is shown, using linear stability theory and direct numerical simulation, that there exist multiple unstable modal instabilities in the boundary layer. One of these modal instabilities is called the supersonic mode, as its phase velocity is supersonic relative to the flow velocity at the edge of the boundary layer. Its growth rate is found to increase with increasing nose radius until a certain nose radius is reached. After this radius, any further increase in nose radius decreases its growth rate. There is adequate agreement between theory and direct numerical simulation for the growth rate, phase velocity and eigenfunction of the supersonic mode. At the reentry conditions tested, the supersonic mode is more likely the cause of boundary-layer transition than the second mode for blunted cones with a small wall-temperature ratio. Initial parametric studies confirm that a decrease in wall temperature amplifies the supersonic mode. Also, the supersonic mode’s growth rate is shown to be a maximum when its phase velocity is aligned with the flow velocity.

35

Zhang, Silong, Jiang Qin, Wen Bao, and Long Zhang. "Numerical Analysis of Supersonic Film Cooling in Supersonic Flow in Hypersonic Inlet with Isolator." Advances in Mechanical Engineering 6 (January 2014): 468790. http://dx.doi.org/10.1155/2014/468790.

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36

Zhao, X. H., S. H. Yi, Q. Mi, Y. F. Hu, and H. L. Ding. "Skin Friction Reduction of Hypersonic Body by Supersonic Layer." Fluid Dynamics 57, no. 5 (September 21, 2022): 686–96. http://dx.doi.org/10.1134/s0015462822050123.

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37

Bhutta, Bilal A., and Clark H. Lewis. "Supersonic/hypersonic flowfield predictions over typical finned missile configurations." Journal of Spacecraft and Rockets 30, no. 6 (November 1993): 674–81. http://dx.doi.org/10.2514/3.26372.

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Zhi, Chen, YI Shihe, Zhu Yangzhu, Zhang Qinghu, and Wu Yu. "Application of Nano Technique in Measuring Supersonic/Hypersonic Flow." Modeling and Numerical Simulation of Material Science 03, no. 01 (2013): 1–3. http://dx.doi.org/10.4236/mnsms.2013.31b001.

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39

Lobbia, Marcus A. "Rapid Supersonic/Hypersonic Aerodynamics Analysis Model for Arbitrary Geometries." Journal of Spacecraft and Rockets 54, no. 1 (January 2017): 315–22. http://dx.doi.org/10.2514/1.a33514.

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40

NUSCA, MICHAEL J. "AEROTHERMODYNAMIC ANALYSIS FOR AXISYMMETRIC PROJECTILES AT SUPERSONIC/HYPERSONIC SPEEDS." Engineering Computations 10, no. 5 (May 1993): 423–45. http://dx.doi.org/10.1108/eb023918.

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41

Hui, W. H., and J. J. Hu. "Space-marching gridless computation of steady supersonic/hypersonic flow." International Journal of Computational Fluid Dynamics 20, no. 1 (January 2006): 55–59. http://dx.doi.org/10.1080/10618560600578476.

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42

Huang, Wei, Zhao-bo Du, Li Yan, and Zhi-xun Xia. "Supersonic mixing in airbreathing propulsion systems for hypersonic flights." Progress in Aerospace Sciences 109 (August 2019): 100545. http://dx.doi.org/10.1016/j.paerosci.2019.05.005.

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43

Bonnet, J. P., D. Grésillon, and J. P. Taran. "NONINTRUSIVE MEASUREMENTS FOR HIGH-SPEED, SUPERSONIC, AND HYPERSONIC FLOWS." Annual Review of Fluid Mechanics 30, no. 1 (January 1998): 231–73. http://dx.doi.org/10.1146/annurev.fluid.30.1.231.

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44

Neill, Stephen, and Apostolos Pesyridis. "Modeling of Supersonic Combustion Systems for Sustained Hypersonic Flight." Energies 10, no. 11 (November 18, 2017): 1900. http://dx.doi.org/10.3390/en10111900.

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45

Abouali, Omid, and Goodarz Ahmadi. "A model for supersonic and hypersonic impactors for nanoparticles." Journal of Nanoparticle Research 7, no. 1 (February 2005): 75–88. http://dx.doi.org/10.1007/s11051-004-7910-3.

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46

Davis, Dominic A. R., and Frank T. Smith. "On subsonic, supersonic and hypersonic inflectional-wave/vortex interaction." Journal of Engineering Mathematics 30, no. 6 (November 1996): 611–45. http://dx.doi.org/10.1007/bf00042785.

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47

d’Humières, G., and J. L. Stollery. "Drag reduction on a spiked body at hypersonic speed." Aeronautical Journal 114, no. 1152 (February 2010): 113–19. http://dx.doi.org/10.1017/s0001924000003584.

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AbstractFitting a spike on a blunt body provides a drag reduction at supersonic and hypersonic speeds. In this study, the laminar flow over a spiked, conical body terminated by a spherical cap, inspired by the Apollo re-entry capsule design, was investigated using a hypersonic wind tunnel. Schlieren pictures revealed the absence of flow unsteadiness for the range of spike lengths tested, and force measurements showed a maximum reduction of 77% of the unspiked body drag.A simple theoretical model based on the pressure drag generated by a solid cone showed good agreement with the experimental data. The measured shock stand-off distance agreed well with predictions.

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Bruno, Claudio, and Antonella Ingenito. "Some Key Issues in Hypersonic Propulsion." Energies 14, no. 12 (June 21, 2021): 3690. http://dx.doi.org/10.3390/en14123690.

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This paper summarizes and discusses some critical aspects of flying hypersonically. The first is the L/D (lift over drag) ratio determining thrust and that in turn depends on the slenderness Küchemann’s τ parameter. This second parameter is found to depend on the relative importance of wave versus friction drag. Ultimately, all engineering drag is argued to depend on vorticity formed at the expense of the vehicle kinetic energy, thus requiring work by thrust. Different mixing strategies are discussed and shown to depend also on mechanisms forming vorticity when the regime is compressible. Supersonic combustion is briefly analyzed and found, at sufficiently high combustor Mach, to take place locally at constant volume, unlike conventional Brayton cycles.

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Liu, D. D., P. C. Chen, Z. X. Yao, and D. Sarhaddi. "Recent advances in lifting surface methods." Aeronautical Journal 100, no. 998 (October 1996): 327–40. http://dx.doi.org/10.1017/s0001924000067038.

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AbstractRecent advances in two lifting surface methods are reported. The present subsonic constant pressure method improves upon the program robustness of the doublet lattice method. The present unified supersonic-hypersonic method extends the range of applicability of lifting surface theory to the Newtonian limit. Based on a uniformly-valid series formulation, the method accurately approximates the non-linear thickness effect, whereas this effect is neglected by the linear theory and overestimated by piston theory. In fact, the present results generally agree well with the trends of several Euler solutions. Among other findings, cases computed by the unified method confirm that the effect of thickness is to reduce the supersonic flutter speed.

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Sinha, Jayanta, Sanjay Singh, Om Prakash, and Dhruv Panchal. "Passive flow modification over the supersonic and the hypersonic air-intake system using bleed." FME Transactions 51, no. 3 (2023): 329–37. http://dx.doi.org/10.5937/fme2303329s.

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The air intake should be operated at design conditions to achieve a high total pressure recovery and optimum mass capture ratio. The current research focuses on the numerical simulation of the supersonic and hypersonic air inlet and its starting and unstarting characteristics. 2D RANS equation for supersonic and hypersonic intake has been solved using the k-oSST turbulence model. The in-house code and the algorithm based on the RANS equation have also been validated in due process and used for subsequent simulations. The sudden drop in mass capture ratio indicates the unstart condition of the intake. The presence of a bleed section has a commendable effect on the performance parameter of the air intake. A separation bubble was observed at the intake's entrance during the off-design conditions, resulting in performance losses. Four different bleed sections ranging in size from 1.6mm to 8.6mm were used, and simulations with bleed were run for different Mach numbers ranging from 3 to 8. The optimum bleed size of 3mm has been found quite effective in modifying Total pressure recovery within the optimum mass flow rate over the wide range of Mach numbers.