Статті в журналах: "Experimental Hypersonic flow"

1

GROENIG, Hans, and Herbert OLIVER. "Experimental Hypersonic Flow Research in Europe." JSME International Journal Series B 41, no. 2 (1998): 397–407. http://dx.doi.org/10.1299/jsmeb.41.397.

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2

Fan, Xiaoqiang, and Yuan Tao. "Investigation of Flow Control for the Hypersonic Inlets via Counter Flow." International Journal of Aerospace Engineering 2015 (2015): 1–8. http://dx.doi.org/10.1155/2015/956317.

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Experimental results show that there exist two flow fields in the hypersonic inlets when the forebody waves interact with the lip boundary, which is similar to the shock reflection ion hysteresis phenomenon. In order to improve the performance of the flow field, counterflow is applied to control the shock reflection configuration in the hypersonic inlets. For better understanding of the internal mechanism, inviscid numerical simulation is conducted. And the results demonstrate that it is feasible to realize the transition between the regular reflection configuration and the Mach reflection ion configuration in the hypersonic inlets. That is because the von Neumann criterion and detached criterion play a dominant role, respectively, in these transitions. In addition, the evolution process of Mach reflection ion in the hypersonic inlets can be divided into three stages: transmission of waves, emergence of Mach stem, and stabilization of flow field.

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3

Kalimuthu, R., R. C. Mehta, and E. Rathakrishnan. "Experimental investigation on spiked body in hypersonic flow." Aeronautical Journal 112, no. 1136 (October 2008): 593–98. http://dx.doi.org/10.1017/s0001924000002554.

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Abstract A spike attached to a hemispherical body drastically changes its flowfield and influences aerodynamic drag in a hypersonic flow. It is, therefore, a potential candidate for drag reduction of a future high-speed vehicle. The effect of the spike length, shape, spike nose configuration and angle-of-attack on the reduction of the drag is experimentally studied with use of hypersonic wind-tunnel at Mach 6. The effects of geometrical parameters of the spike and angle-of-attack on the aerodynamic coefficient are analysed using schlieren picture and measuring aerodynamic forces. These experiments show that the aerodisk is superior to the aerospike. The aerodisk of appropriate length, diameter and nose configuration may have the capability for the drag reduction. The inclusion of an aero disk at the leading edge of the spike has an advantage for the drag reduction mechanism if it is at an angle-of-attack, however consideration to be given for increased moment resulting from the spike is required.

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4

Papuccuoglu, Hakan. "Experimental investigation of hypersonic three-dimensional corner flow." AIAA Journal 31, no. 4 (April 1993): 652–56. http://dx.doi.org/10.2514/3.11599.

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5

Laitón, Sergio Nicolas Pachón, João Felipe de Araujo Martos, Israel da Silveira Rego, George Santos Marinho, and Paulo Gilberto de Paula Toro. "Experimental Study of Single Expansion Ramp Nozzle Performance Using Pitot Pressure and Static Pressure Measurements." International Journal of Aerospace Engineering 2019 (February 27, 2019): 1–11. http://dx.doi.org/10.1155/2019/7478129.

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In order to overcome the drag at hypersonic speed, hypersonic flight vehicles require a high level of integration between the airframe and the propulsion system. Propulsion system based on scramjet engine needs a close interaction between its aerodynamics and stability. Hypersonic vehicle nozzles which are responsible for generating most of the thrust generally are fused with the vehicle afterbody influencing the thrust efficiency and vehicle stability. Single expansion ramp nozzles (SERN) produce enough thrust necessary to hypersonic flight and are the subject of analysis of this work. Flow expansion within a nozzle is naturally 3D phenomena; however, the use of side walls controls the expansion approximating it to a 2D flow confined. An experimental study of nozzle performance traditionally uses the stagnation conditions and the area ratio of the diverging section of the tunnel for approaching the combustor exit conditions. In this work, a complete hypersonic vehicle based on scramjet propulsion is installed in the test section of a hypersonic shock tunnel. Therefore, the SERN inlet conditions are the real conditions from the combustor exit. The performance of a SERN is evaluated experimentally under real conditions obtained from the combustor exit. To quantify the SERN performance parameters such as thrust, axial thrust coefficient Cfx and lift L are investigated and evaluated. The generated thrust was determined from both static and pitot pressure measurements considering the installation of side walls to approximate 2D flow. Measurements obtained by a rake show that the flow at the nozzle exit is not symmetric. Pitot and pressure measurements inside the combustion chamber show nonuniform flow condition as expected due to side wall compression and boundary layer. The total axial thrust for the nozzle obtained with the side wall is slightly higher than without it. Static pressure measurements at the centerline of the nozzle show that the residence time of the flow in the expansion section is short enough and the flow of the central region of the nozzle is not altered by the lateral expansion when nozzle configuration does not include side walls.

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6

Juluru Sandeep and AVSS Kumara Swami Gupta. "Grid Adaptive Technique for Simulation of Scramjet Intake-Isolator at Hypersonic Speeds." Journal of Advanced Research in Fluid Mechanics and Thermal Sciences 101, no. 1 (January 18, 2023): 73–89. http://dx.doi.org/10.37934/arfmts.101.1.7389.

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Hypersonic intake is one of the major components of Scramjet engine. It compresses the incoming hypersonic flow through a series of oblique shocks as the flow passes through intake-isolator section before entering the combustion chamber, which is essential for efficient combustion. The shocks generated inside in the intake interacts with boundary layer following shock boundary layer interaction and flow separation. The separated flow blocks the flow capture area such that engine expresses unstarting phenomenon. Understanding and mitigating such flow phenomenon is a challenging task. With respect to hypersonic speeds the experimental facilities are very limited. The only alternative to solve this problem is Computational Fluid Dynamics because of its capabilities. But validation of CFD results with analytical or experimental is the foremost prerequisite to chase computational analysis. Mostly at high speeds the precision of CFD results rest on the type of grid, number of elements and turbulence model used. So, in this paper, computational analysis of hypersonic intake is carried out through designed conditions to ensure the correct CFD process is used by varying number of elements in fluid domain by grid adaptive technique using ANSYS Fluent and satisfying Y+ parameter. The domain is analysed with various turbulence models and among them SST has predicted all the flow characteristics of scramjet intake-isolator at hypersonic speeds like separation bubble, shock reattachment, cowl shock etc similar to experimental results with the help of grid adaptive technique. So, grid adaptive technique is also proposed for simulation of scramjet intake at off-design conditions.

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7

Lanson, F., and J. L. Stollery. "Some hypersonic intake studies." Aeronautical Journal 110, no. 1105 (March 2006): 145–56. http://dx.doi.org/10.1017/s0001924000001123.

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Abstract A ‘two dimensional’ air intake comprising a wedge followed by an isentropic compression has been tested in the Cranfield Gun Tunnel at Mach 8·2. These tests were performed to investigate qualitatively the intake flow starting process. The effects of cowl position, Reynolds number, boundary-layer trip and introduction of a small restriction in the intake duct were investigated. Schlieren pictures of the flow on the compression surface and around the intake entrance were taken. Results showed that the intake would operate over the Reynolds number range tested. Tests with a laminar boundary layer demonstrated the principal influence of the Reynolds number on the boundary-layer growth and consequently on the flow structure in the intake entrance. In contrast boundary layer tripping produced little variation in flow pattern over the Reynolds number range tested. The cowl lip position appeared to have a strong effect on the intake performance. The only parameter which prevented the intake from starting was the introduction of a restriction in the intake duct. The experimental data obtained were in good qualitative agreement with the CFD predictions. Finally, these experimental results indicated a good intake flow starting process over multiple changes of parameters.

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8

V. Gromyko, Yuriy, Anatoliy A. Maslov, Andrey A. Sidorenko, Pavel A. Polivanov, and Ivan S. Tsyryulnikov. "Estimation of the Flow Parametrs in Hypersonic Wind Tunnels." Siberian Journal of Physics 6, no. 2 (July 1, 2011): 10–16. http://dx.doi.org/10.54362/1818-7919-2011-6-2-10-16.

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The paper describes the algorithm of the flow parameters calculation for hypersonic wind tunnels taking into account the real gas properties using air and carbon dioxide as a working gas. The results of the experimental measurements of the flow velocity at the contoured nozzle exit in the hypersonic wind tunnel IT-302M have been carried out for verification of the algorithm

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9

Creighton, S., and R. Hillier. "Experimental and computational study of unsteady hypersonic cavity flows." Aeronautical Journal 111, no. 1125 (November 2007): 673–88. http://dx.doi.org/10.1017/s0001924000004851.

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AbstractThis paper presents a combined experimental and computational study of annular cavities on a semi-angle cone in a Mach 8·9 flow. A range of cavity length-to-depth ratios has been considered, and a parameter has been determined that distinguishes between ‘weak oscillations’ and ‘strong oscillations’ of the cavity flow. Essentially the work identifies the transition from the case where the flow can be regarded as ‘pure cavity flow’ to that where the flow behaviour is tending towards that of a ‘spiked blunt body’. The CFD simulations also suggest that, for a certain range of cavity scale, the limiting cavity flow state depends upon the flow initialisation process; it may be weak or strongly oscillating.

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10

DONG, HAO, CHENG-PENG WANG, and KE-MING CHENG. "EXPERIMENTAL AND NUMERICAL INVESTIGATION OF HYPERSONIC JAWS INLET." Modern Physics Letters B 24, no. 13 (May 30, 2010): 1409–12. http://dx.doi.org/10.1142/s0217984910023748.

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In order to obtain the flow field characteristics and the influence of boundary layer, numerical simulations and wind tunnel tests are conducted for two streamline traced Jaws inlets at Mach number 7. The inlets are designed based on a flow field with 8-7 planar shock wave (the ramp in pitch plane is inclined at 8° to the free stream and in yaw plane is inclined at 7° to the free stream, yielding planar shocks). In the study, the static pressure distributions were measured and analyzed along the plane-symmetric centerline of the inlet with and without the boundary layer correction, respectively. Results show that boundary layer correction can obviously weaken the viscous influence to the inlet, increasing the mass flow coefficient and improving total pressure recovery.

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11

LAURENCE, STUART J., R. DEITERDING, and G. HORNUNG. "Proximal bodies in hypersonic flow." Journal of Fluid Mechanics 590 (October 15, 2007): 209–37. http://dx.doi.org/10.1017/s0022112007007987.

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Hypersonic flows involving two or more bodies travelling in close proximity to one another are encountered in several important situations. The present work seeks to explore one aspect of the resulting flow problem by investigating the forces experienced by a secondary body when it is within the domain of influence of a primary body travelling at hypersonic speeds.An analytical methodology based on the blast wave analogy is developed and used to predict the secondary force coefficients for simple geometries in both two and three dimensions. When the secondary body is entirely inside the primary shocked region, the nature of the lateral force coefficient is found to depend strongly on the relative size of the two bodies. For two spheres, the methodology predicts that the secondary body will experience an exclusively attractive lateral force if the secondary diameter is larger than one-sixth of the primary diameter. The analytical results are compared with those from numerical simulations and reasonable agreement is observed if an appropriate normalization for the relative lateral displacement of the two bodies is used.Results from a series of experiments in the T5 hypervelocity shock tunnel are also presented and compared with perfect-gas numerical simulations, with good agreement. A new force-measurement technique for short-duration hypersonic facilities, enabling the experimental simulation of the proximal bodies problem, is described. This technique provides two independent means of measurement, and the agreement observed between the two gives a further degree of confidence in the results obtained.

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12

Kong, Wei Xuan, Rui Zhao, Jian Yu, and Chao Yan. "Numerical Investigation of Hypersonic Double-Cone Flow." Applied Mechanics and Materials 232 (November 2012): 240–45. http://dx.doi.org/10.4028/www.scientific.net/amm.232.240.

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Hypersonic flow of Mach number 8 past a 25°-50° double cone geometry is numerically simulated at ReD=4.8E5. Complicated flow structures, including Type V shock-shock interaction, shock-boundary layer interaction, separation and reattachment at the corner are presented and discussed. The surface pressure and heat transfer rate distributions are also calculated and compared with the experimental data. Results show that both the 2nd order MUSCL and 5th order WENO could accurately reproduce the shock structures, while the higher order scheme could predict a more accurate size of separation zone. Generally, the size of the separation zone is underestimated with an overvalued pressure distribution after reattachment employing the full turbulent models. On the other hand, transition induced by the reattachment shock has been calculated using transition model and the results of pressure peak and the size of separation zone show good agreement with the experimental measurements.

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13

Храмцов, П. П., В. А. Васецкий, В. М. Грищенко, М. В. Дорошко, М. Ю. Черник, А. И. Махнач та И. А. Ших. "Диагностика полей плотности фотометрическим теневым методом при гиперзвуковом обтекании конуса в легкогазовой баллистической установке". Журнал технической физики 89, № 10 (2019): 1506. http://dx.doi.org/10.21883/jtf.2019.10.48165.74-18.

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A new method of hypersonic flow generation is proposed and the results of an experimental study of hypersonic flow past cones with half-angles = 3° and = 12° are presented. The Mach numbers of the studied incident flows were = 18 ( = 3°) and = 14.4 ( = 12°). The use of a light-gas facility, where an accelerating channel was replaced with Laval nozzle, allows us to obtain a hypersonic outflow with optical density sufficiently high for flow visualization and diagnostics with the help of optical methods. The flow structure was visualized by means of the shadow method using the Foucault knife and the slit. Shadowgraphs were recorded by a high-speed camera with a frame rate of 300,000 fps and an exposure time of 1 µs. The Mach number for the incident flow was calculated from the inclination angle of the shock wave on shadowgraphs.

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14

Khalid, M., and K. A. Juhany. "Heat alleviation studies on hypersonic re-entry vehicles." Aeronautical Journal 122, no. 1257 (November 2018): 1673–96. http://dx.doi.org/10.1017/aer.2018.103.

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ABSTRACTA numerical simulation has been carried out to investigate the effects of leading edge blowing upon heat alleviation on the surface of hypersonic vehicles. The initial phase of this work deals with the ability of the present CFD-based techniques to solve hypersonic flow field past blunt-nosed vehicles at hypersonic speeds. Towards this end, the authors selected three re-entry vehicles with published flow field data against which the present computed results could be measured. With increasing confidence on the numerical simulation techniques to accurately resolve the hypersonic flow, the boundary condition at the solid blunt surface was then equipped with the ability to blow the flow out of the solid boundary at a rate of at least 0.01–0.1 times the free stream (ρ∞u∞) mass flow rate. The numerical iterative procedure was then progressed until the flow at the surface matched this new ‘inviscid like’ boundary condition. The actual matching of the flow field at the ejection control surface was achieved by iterating the flow on the adjacent cells until the flow conformed to the conditions prescribed at the control surface. The conditions at the surface could be submitted as a ρ∞u∞at the surface or could be equipped as a simple static pressure condition providing the desired flow rate. The comparison between the present engineering approach and the experimental data presented in this study demonstrate its ability to solve complex problems in hypersonic.

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15

Xiao, Hong, Yuhe Shang, and Di Wu. "DSMC Simulation and Experimental Validation of Shock Interaction in Hypersonic Low Density Flow." Scientific World Journal 2014 (2014): 1–10. http://dx.doi.org/10.1155/2014/732765.

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Direct simulation Monte Carlo (DSMC) of shock interaction in hypersonic low density flow is developed. Three collision molecular models, including hard sphere (HS), variable hard sphere (VHS), and variable soft sphere (VSS), are employed in the DSMC study. The simulations of double-cone and Edney’s type IV hypersonic shock interactions in low density flow are performed. Comparisons between DSMC and experimental data are conducted. Investigation of the double-cone hypersonic flow shows that three collision molecular models can predict the trend of pressure coefficient and the Stanton number. HS model shows the best agreement between DSMC simulation and experiment among three collision molecular models. Also, it shows that the agreement between DSMC and experiment is generally good for HS and VHS models in Edney’s type IV shock interaction. However, it fails in the VSS model. Both double-cone and Edney’s type IV shock interaction simulations show that the DSMC errors depend on the Knudsen number and the models employed for intermolecular interaction. With the increase in the Knudsen number, the DSMC error is decreased. The error is the smallest in HS compared with those in the VHS and VSS models. When the Knudsen number is in the level of 10−4, the DSMC errors, for pressure coefficient, the Stanton number, and the scale of interaction region, are controlled within 10%.

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16

Ismagilov, D. R., and R. V. Sidelnikov. "Features of numerical simulation of hypersonic flow around simple bodies." Journal of «Almaz – Antey» Air and Space Defence Corporation, no. 2 (June 30, 2015): 49–54. http://dx.doi.org/10.38013/2542-0542-2015-2-49-54.

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The analysis of the possibility of using the numerical approximation schemes streams Roe FDS and AUSM + to meet the challenges of hypersonic aerodynamics and research trends in the perturbed region ahead streamlined blunt body to determine the laws of thermal and gas-dynamic processes and the establishment of the characteristics associated with the development of the necessary thermal protection of aircraft. Based on a comparison of the data with the experimental results revealed that the method of splitting the flow AUSM + is able to solve the problem of hypersonic flow around bodies with acceptable accuracy.

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17

Engblom, W. A., B. Yuceil, D. B. Goldstein, and D. S. Dolling. "Experimental and numerical study of hypersonic forward-facing cavity flow." Journal of Spacecraft and Rockets 33, no. 3 (May 1996): 353–59. http://dx.doi.org/10.2514/3.26767.

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18

Sun, Quan, Bangqin Cheng, Yinghong Li, Wei Cui, Yonggui Yu, and Junhun Jie. "Experimental Investigation of Hypersonic Flow and Plasma Aerodynamic Actuation Interaction." Plasma Science and Technology 15, no. 9 (September 2013): 908–14. http://dx.doi.org/10.1088/1009-0630/15/9/15.

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19

Simeonides, G., and W. Haase. "Experimental and computational investigations of hypersonic flow about compression ramps." Journal of Fluid Mechanics 283 (January 25, 1995): 17–42. http://dx.doi.org/10.1017/s0022112095002229.

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Comprehensive results of a joint experimental and computational study of the two-dimensional flow field over flat plate/compression ramp configurations at Mach 14 are presented. These geometries are aimed to simulate, in a simplified manner, the region around deflected control surfaces of hypersonic re-entry vehicles. The test cases considered cover a range of realistic flow conditions with Reynolds numbers to the hinge line varying between 4.5 × 105 and 2.6 × 106 (with a reference length taken as the distance between the leading edge and the hinge line) and a wall-to-total-temperature ratio of 0.12. The combination of flow and geometric parameters gives rise to fully laminar strong shock wave/boundary layer interactions with extensive separation, and transitional interactions with transition occurring near the reattachment point. A fully turbulent interaction is also considered which, however, was only approximately achieved in the experiments by means of excessive tripping of the oncoming hypersonic laminar boundary layer. Emphasis has been placed upon the quality and level of confidence of both experiments and computations, including a discussion on the laminar-turbulent transition process and the associated striation phenomenon. The favourable comparison between the experimental and computational results has profided the grounds for an enhanced understanding of the relevant flow processes and their modelling. Particularly in relation to transitional shock wave/boundary layer interactions, where laminar-turbulent transition is promoted by the adverse pressure gradient and flow concavity in the reattachment region, a method is proposed to compute extreme adverse effects in the interaction region avoiding such inhibiting requirements as transition modelling or turbulence modelling over separated regions.

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20

Hashimoto, Tokitada. "Experimental investigation of hypersonic flow induced separation over double wedges." Journal of Thermal Science 18, no. 3 (September 2009): 220–25. http://dx.doi.org/10.1007/s11630-009-0220-4.

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21

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.

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22

Huang, Guo, and Haiming Huang. "Numerical investigation of heat flux distribution in a deep gap based on chemical equilibrium." International Journal of Numerical Methods for Heat & Fluid Flow 27, no. 8 (August 7, 2017): 1662–74. http://dx.doi.org/10.1108/hff-03-2016-0119.

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Purpose The purpose of this paper is to perform the simulation to explore the gap flow field under a hypersonic air flow. Thermal protection systems of hypersonic vehicles generally consist of thermal insulation tiles, and gaps between these tiles probably cause a severe local aerodynamic thermal effect. Design/methodology/approach The discretizations of convection flux term and temporal term in the governing equation with chemical equilibrium, respectively, take AUSM+-up flux-vector splitting scheme and the implicit lower-upper symmetric Gauss–Seidel method. Based on these, the flow field in a deep gap is simulated by means of the computer codes that the authors have written. Findings The numerical results show that the heat flux distribution in a gap has a good agreement with experimental results. Importantly, the distribution of heat flux is “U” shaped and the maximum of the heat flux occurs at the windward corner of a gap. Originality/value To explore the gap flow field under a hypersonic air flow, which is a chemically reacting, all speed and viscous flow, a novel model with an equivalent ratio of specific heats is presented. The investigation in this paper has a guide for the design of the thermal protection system in hypersonic vehicles.

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23

Gomes-Fernandes, R. "Flow features around double cones at hypersonic speed." Aeronautical Journal 117, no. 1193 (July 2013): 741–48. http://dx.doi.org/10.1017/s000192400000840x.

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AbstractAn experimental study of double cone geometries in hypersonic flow (at M∞of 8·2 and Redof 0·36 × 106) was performed to provide additional data to a computational simulation study. In this study, depending on how the flow was initialised, the numerical solution yielded a violent pulsation mode of instability or a steady flow. Finally, an analysis of the shock oscillation was made and discussed.

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24

Murray, N., R. Hillier, and S. Williams. "Experimental investigation of axisymmetric hypersonic shock-wave/turbulent-boundary-layer interactions." Journal of Fluid Mechanics 714 (January 2, 2013): 152–89. http://dx.doi.org/10.1017/jfm.2012.464.

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AbstractThis paper presents time-averaged data for high-Reynolds-number hypersonic shock-wave/boundary-layer interactions, using a body of revolution to achieve high standards of two-dimensionality. The data are collected at nominal Mach 8.9, but a calibration is included that permits weak flow gradients in the test section to be incorporated as part of the data interpretation or flow modelling. The axisymmetric turbulent test boundary layer is developed on a hollow cylinder, aligned axially with the flow. The shock-wave interaction with this boundary layer is then generated by two separate configurations. Firstly, an impinging shock-wave case, that uses a concentric cowl to radiate an axisymmetric shock system onto the test boundary layer: for this case both an attached flow and a separated flow interaction are formed. Secondly, use of a conical-flare afterbody to produce a separated flow interaction. Quantitative data are presented for surface pressures and heat transfer, supported by some schlieren visualization and surface oil flows. A restricted CFD programme is included to assist the interpretation of the experiments.

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25

Esch, T., and M. Giehrl. "Numerical Analysis of Nozzle and Afterbody Flow of Hypersonic Transport Systems." Journal of Engineering for Gas Turbines and Power 117, no. 3 (July 1, 1995): 389–93. http://dx.doi.org/10.1115/1.2814107.

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Using an implicit Finite-Volume Navier–Stokes code, the flow field in a Single Expansion Ramp Nozzle (SERN) for a hypersonic aircraft is studied. Comparisons between experimental data and CFD calculations for certain components of the integrated exhaust system (cold two-dimensional nozzle flow, high temperature reacting three-dimensional combustion chamber flow, and two-dimensional nozzle flow with external flow) are presented. To show the sensitivity of the considered components to off-design operating conditions, comprehensive numerical studies have been carried out. For the determination of nozzle performance a detailed two-dimensional analysis from transonic to hypersonic flight Mach numbers has been performed. A direct optimization method has been used to investigate the influence of the lower nozzle flap shape on the thrust vector.

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26

Kumar, D., and J. L. Stollery. "Hypersonic control flap effectiveness." Aeronautical Journal 100, no. 996 (July 1996): 197–208. http://dx.doi.org/10.1017/s0001924000067154.

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SummaryThe effects of flap deflection, configuration incidence and leading edge bluntness on boundary layer separation and transition have been studied experimentally. A quasi two-dimensional flat plate equipped with a full span trailing edge control flap has been employed for these tests.The studies have been carried out in a hypersonic gun tunnel at M∞ = 8·2 and Re∞/cm = 9·0 × 104. The flap deflection angles studied were in the range 0° ≤ β ≤ 30°. The incidence range was from zero to α = 10° (positive α is nose down). Leading edge bluntness effects were simulated by the introduction of a hemi-cylindrical leading edge.The flow structure was studied using high speed Schlieren photography as well as surface pressure and heat transfer measurements. Liquid crystals were employed to study the threedimensionality of the flow structure for selected configurations. Analytical theories have been developed to estimate the flap pressure and heat transfer levels for the sharp and blunt configurations. These are compared against experimental measurements.

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27

Sandham, N. D., E. Schülein, A. Wagner, S. Willems, and J. Steelant. "Transitional shock-wave/boundary-layer interactions in hypersonic flow." Journal of Fluid Mechanics 752 (July 4, 2014): 349–82. http://dx.doi.org/10.1017/jfm.2014.333.

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AbstractStrong interactions of shock waves with boundary layers lead to flow separations and enhanced heat transfer rates. When the approaching boundary layer is hypersonic and transitional the problem is particularly challenging and more reliable data is required in order to assess changes in the flow and the surface heat transfer, and to develop simplified models. The present contribution compares results for transitional interactions on a flat plate at Mach 6 from three different experimental facilities using the same instrumented plate insert. The facilities consist of a Ludwieg tube (RWG), an open-jet wind tunnel (H2K) and a high-enthalpy free-piston-driven reflected shock tunnel (HEG). The experimental measurements include shadowgraph and infrared thermography as well as heat transfer and pressure sensors. Direct numerical simulations (DNS) are carried out to compare with selected experimental flow conditions. The combined approach allows an assessment of the effects of unit Reynolds number, disturbance amplitude, shock impingement location and wall cooling. Measures of intermittency are proposed based on wall heat flux, allowing the peak Stanton number in the reattachment regime to be mapped over a range of intermittency states of the approaching boundary layer, with higher overshoots found for transitional interactions compared with fully turbulent interactions. The transition process is found to develop from second (Mack) mode instabilities superimposed on streamwise streaks.

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28

He, Miao, Zijie Li, and Hao Wang. "Experiment on gas flow field of large-caliber hypersonic balance gun in half-space." AIP Advances 13, no. 2 (February 1, 2023): 025020. http://dx.doi.org/10.1063/5.0134738.

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As one of the main methods to study the characteristics of an object, experiment is the main way to verify the correctness and reliability of numerical results. The hypersonic projectile near the ground launch involves many complex physical phenomena, so the ground surface has an important impact on the flight stability of the projectile and the structural characteristics of the gas flow field. In this paper, we carried out a whole process experiment of a hypersonic projectile from the bottom of the tube to the muzzle based on the 300 mm balance gun. The muzzle pressure and the initial velocity of the muzzle were monitored by the data acquisition system. We obtained the influence of the ground on the gas flow field of the hypersonic projectile through the overpressure data on the ground near the muzzle. Meanwhile, the dynamic characteristics of the launch process are recorded by high-speed photography. The experimental results show that the muzzle shock wave meets the ground, and the reflection phenomenon occurs, forming a wavefront with opposite directions of propagation. It makes the pressure in the lower half of the domain significantly higher than that in the upper half. By analyzing the experimental phenomena and experimental data, this paper shows the development characteristics of the projectile’s flow field, as well as the influence of the ground on the flow field and the motion stability of the projectile.

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29

Zhang, Lin, Junli Yang, Tiecheng Duan, Jie Wang, Xiuyi Li, and Kunyuan Zhang. "Numerical and Experimental Investigation on Nosebleed Air Jet Control for Hypersonic Vehicle." Aerospace 10, no. 6 (June 9, 2023): 552. http://dx.doi.org/10.3390/aerospace10060552.

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A new idea of nosebleed air jets with strong coupled internal and external flow is put forward using the lateral jet control principle to improve the maneuverability and fast reaction capabilities of hypersonic vehicles. The hypersonic vehicle’s nose stagnant high-pressure and high-temperature gas is utilized as the drive source for long-term jet control. The significant coupled jet interaction of the internal and external flow changes the aerodynamic characteristics. As a result, the structure is basic and does not rely on any external source to achieve flight attitude control. The complicated flow characteristics of the nosebleed jet in supersonic crossflow surrounding the vehicle were numerically and experimentally investigated. The jet interaction characteristics and the aerodynamic characteristic changes generated by the nosebleed air jet are verified by comparing the flow field with and without the jet. Results indicate that the nosebleed air jet alters the center-of-pressure coefficient, which is subsequently coupled with the interference aerodynamic force. This results in a variation in pitch moment. The jet decreases the pitching moment coefficient when compared with the case without a jet. It is probable that combining nosebleed air jets with model centroid adjustment yields an optimal trim angle of attack.

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30

Park, Seong-Hyeon, Junemo Kim, Ilsung Choi, and Gisu Park. "Experimental study of separation behavior of two bodies in hypersonic flow." Acta Astronautica 181 (April 2021): 414–26. http://dx.doi.org/10.1016/j.actaastro.2021.01.037.

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31

Fu Jia, Yi Shi-He, Wang Xiao-Hu, Zhang Qing-Hu, and He Lin. "Experimental study on flow visualization of hypersonic flat plate boundary layer." Acta Physica Sinica 64, no. 1 (2015): 014704. http://dx.doi.org/10.7498/aps.64.014704.

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32

ZHONG, Ce, Kojiro SUZUKI, and Yasumasa WATANABE. "Experimental and Numerical Study of Hypersonic Flow over Backward-Facing Step." TRANSACTIONS OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES, AEROSPACE TECHNOLOGY JAPAN 19, no. 5 (2021): 735–43. http://dx.doi.org/10.2322/tastj.19.735.

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33

Menezes, Viren, S. Saravanan, G. Jagadeesh, and K. P. J. Reddy. "Experimental Investigations of Hypersonic Flow over Highly Blunted Cones with Aerospikes." AIAA Journal 41, no. 10 (October 2003): 1955–66. http://dx.doi.org/10.2514/2.1885.

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34

Meng, B. Q., G. L. Han, C. K. Yuan, C. Wang, and Z. L. Jiang. "Experimental and Numerical Study on Hypersonic Flow over Double-Wedge Configuration." AIAA Journal 55, no. 9 (September 2017): 3227–30. http://dx.doi.org/10.2514/1.j055546.

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35

Chanetz, B., R. Benay, J. M. Bousquet, R. Bur, T. Pot, F. Grasso, and J. Moss. "Experimental and numerical study of the laminar separation in hypersonic flow." Aerospace Science and Technology 2, no. 3 (March 1998): 205–18. http://dx.doi.org/10.1016/s1270-9638(98)80054-0.

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36

Vetlutskii, V. N., A. A. Maslov, S. G. Mironov, T. V. Poplavskaya, and A. N. Shiplyuk. "Hypersonic flow on a flat plate. Experimental results and numerical modeling." Journal of Applied Mechanics and Technical Physics 36, no. 6 (November 1995): 848–54. http://dx.doi.org/10.1007/bf02369381.

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37

Prakash, R., L. M. Le Page, L. P. McQuellin, S. L. Gai, and S. O’Byrne. "Direct simulation Monte Carlo computations and experiments on leading-edge separation in rarefied hypersonic flow." Journal of Fluid Mechanics 879 (October 2, 2019): 633–81. http://dx.doi.org/10.1017/jfm.2019.692.

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A comprehensive study of the fundamental characteristics of leading-edge separation in rarefied hypersonic flows is undertaken and its salient features are elucidated. Separation of a boundary layer undergoing strong expansion is typical in many practical hypersonic applications such as base flows of re-entry vehicles and flows over deflected control surfaces. Boundary layer growth under such conditions is influenced by effects of rarefaction and thermal non-equilibrium, thereby differing significantly from the conventional no-slip Blasius type. A leading-edge separation configuration presents a fundamental case for studying the characteristics of such a flow separation but with minimal influence from a pre-existing boundary layer. In this work, direct simulation Monte Carlo computations have been performed to investigate flow separation and reattachment in a low-density hypersonic flow over such a configuration. Distinct features of leading-edge flow, limited boundary layer growth, separation, shear layer, flow structure in the recirculation region and reattachment are all explained in detail. The fully numerical shear layer profile after separation is compared against a semi-theoretical profile, which is obtained using the numerical separation profile as the initial condition on existing theoretical concepts of shear layer analysis based on continuum flow separation. Experimental studies have been carried out to determine the surface heat flux using thin-film gauges and computations showed good agreement with the experimental data. Flow visualisation experiments using the non-intrusive planar laser-induced fluorescence technique have been performed to image the fluorescence of nitric oxide, from which velocity and rotational temperature distributions of the separated flow region are determined.

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38

Han, Junhao, Lin He, Xiwang Xu, and Zhengbang Wu. "Experimental Investigation of a Roughness Element Wake on a Hypersonic Flat Plate." Aerospace 9, no. 10 (October 2, 2022): 574. http://dx.doi.org/10.3390/aerospace9100574.

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An experimental investigation was performed on the wake flow field of an isolated roughness element of a flat plate at Mach 6 by employing the nanoparticle-based planar laser scattering (NPLS) approach. The three-dimensional features and causes of the flow field structure were scrutinized by transient flow field images of roughness elements on various planes. The time-resolved NPLS technique was implemented to examine the time evolution characteristics of the wake flow field of roughness elements. In the following, the process of dynamic evolution of large-scale vortex structures in the wake flow field was methodically assessed. Additionally, the influences of roughness element heights on the wake vortex structure were evaluated and the obtained results were compared.

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39

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.

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40

McIntyre, T. J., H. Kleine, and A. F. P. Houwing. "Optical imaging techniques for hypersonic impulse facilities." Aeronautical Journal 111, no. 1115 (January 2007): 1–16. http://dx.doi.org/10.1017/s0001924000001718.

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Abstract The application of optical imaging techniques to hypersonic facilities is discussed and examples of experimental measurements are provided. Traditional Schlieren and shadowgraph techniques still remain as inexpensive and easy to use flow visualisation techniques. With the advent of faster cameras, these methods are becoming increasingly important for time-resolved high-speed imaging. Interferometry’s quantitative nature is regularly used to obtain density information about hypersonic flows. Recent developments have seen an extension of the types of flows that can be imaged and the measurement of other flow parameters such as ionisation level. Planar laser induced fluorescence has been used to visualise complex flows and to measure such quantities as temperature and velocity. Future directions for optical imaging are discussed.

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41

JACKSON, A. P., R. HILLIER, and S. SOLTANI. "Experimental and computational study of laminar cavity flows at hypersonic speeds." Journal of Fluid Mechanics 427 (January 25, 2001): 329–58. http://dx.doi.org/10.1017/s0022112000002433.

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This paper presents a combined experimental/computational study of a surface cavity in a low Reynolds number Mach 9 flow. The geometry is based on a body of revolution, which produces highly two-dimensional time-averaged flow for all experimental test cases. A range of cavity length-to-depth ratios, up to a maximum of 8, is investigated. These correspond to ‘closed’ cavity flows, with the free shear layer bridging the entire cavity. For most cases the free shear layer is laminar. However, there is evidence of three-dimensional unsteadiness which is believed to be the consequence of Taylor–Görtler-type vortex formation. The effect of this is first detected on the cavity floor but progressively spreads as the cavity length is increased. For the longest cavities the flow is also influenced by the early stages of laminar–turbulent transition in the free shear layer.

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42

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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43

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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44

Hoffman, Eugene N. A., Elijah J. LaLonde, Angelina Andrade, Ivana Chen, Hayden A. Bilbo, and Christopher S. Combs. "Flow Characterization of the UTSA Hypersonic Ludwieg Tube." Aerospace 10, no. 5 (May 16, 2023): 463. http://dx.doi.org/10.3390/aerospace10050463.

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The characterization of a hypersonic impulse facility is performed using a variety of methods including Pitot probe scans, particle image velocimetry, and schlieren imaging to verify properties such as the velocity, Mach number, wall boundary layer thickness, and freestream turbulence intensity levels. The experimental results are compared to the numerical simulations of the facility performed with Ansys Fluent to compare the design and operational conditions. The presentation of results in this manuscript is prefaced by a description of the facility and its capabilities. The UTSA Ludwieg tube facility can produce a hypersonic freestream flow with a Mach number of 7.2 ± 0.2 and unit Reynolds numbers of up to 200 × 106 m−1. The Pitot probe profiles of the 203-mm-square test section indicate a 152 ± 10 mm square freestream core with turbulence intensity values ranging from 1% to 2%. Schlieren imaging of the oblique shockwaves on a 15° wedge model provided an alternate means of verifying the Mach number. Particle image velocimetry and previous molecular tagging velocimetry results showed a good agreement with the Pitot probe data and numerical simulations in the key parameters including freestream velocity, wall boundary layer velocity profiles, and wall boundary layer thickness.

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45

Liu, Meikuan, Guilai Han, Zongxian Li, and Zonglin Jiang. "Experimental study on the effects of the cone nose-tip bluntness." Physics of Fluids 34, no. 10 (October 2022): 101703. http://dx.doi.org/10.1063/5.0110928.

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In this Letter, hypersonic boundary-layer transition was investigated on a large-scale cone with a height of 3 m and a half-cone angle of 7° at a zero angle of attack in the JF-12 hypersonic flight duplicate shock tunnel. For the same freestream unit Reynolds number, with the increase in the bluntness Reynolds number, the transition Reynolds number has a trend of first increasing and then decreasing, showing a “transition reversal” phenomenon. As the bluntness increased, the high/low-frequency instability waves in the boundary-layer were modulated, which caused the boundary-layer transition to be delayed and then advanced.

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46

Maslov, Eugene, Irina Zharova, Valery Faraponov, Eugene Kozlov, and Vladislav Matskevich. "Study of heat transfer processes in the flowing part of hypersonic air-ramjet engine." MATEC Web of Conferences 194 (2018): 01037. http://dx.doi.org/10.1051/matecconf/201819401037.

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The technique and results of the experimental-theoretical study of gas dynamics, heat transfer and the structure of gas flow in the flowing channel of a model hypersonic air-ramjet engine are presented for Mach numbers M = (5; 6).

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47

Saad, Mohd Rashdan, Azam Che Idris, and Konstantinos Kontis. "Flow Diagnostics in Shock Wave-Boundary Layer Interaction Experiments in Hypersonic Flow." Applied Mechanics and Materials 660 (October 2014): 669–73. http://dx.doi.org/10.4028/www.scientific.net/amm.660.669.

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Shock Wave-Boundary Layer Interaction (SBLI) is a phenomenon occurring in high-speed propulsion systems that is highly undesirable. Numerous methods have been tested to manipulate and control SBLI which includes both active and passive flow control techniques. To determine the improvements brought by the flow control techniques, advanced and state-of the-art flow diagnostics and experimental techniques are required, especially when it involves high-speed flows. In this study, a number of advanced flow diagnostics were employed to investigate the effect of micro-vortex generators in controlling SBLI in Mach 5 such as Pressure Sensitive Paints (PSP), Particle Image Velocimetry (PIV), schlieren photography and oil-flow visualization. The flow diagnostics successfully visualized the boundary layer separation and also the improvements brought by the micro-vortex generators.

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48

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

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

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49

Matsumoto, Kazuki, Mutsuo Kotake, Hajime Itoh, and Masatomi Nishio. "An Experimental Investigation of Shock-Wave Boundary-Layer Interaction in Hypersonic Flow." Journal of the Visualization Society of Japan 19, Supplement2 (1999): 267–70. http://dx.doi.org/10.3154/jvs.19.supplement2_267.

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50

Tsuboi, Nobuyuki, and Yoichiro Matsumoto. "Experimental and Numerical Study of Hypersonic Rarefied Gas Flow over Flat Plates." AIAA Journal 43, no. 6 (June 2005): 1243–55. http://dx.doi.org/10.2514/1.10950.

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