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Автор: Grafiati
Опубліковано: 6 вересня 2023

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Статті в журналах з теми "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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Анотація:
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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Анотація:
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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Анотація:
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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Анотація:
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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Дисертації з теми "Experimental Hypersonic flow"

1

Mohammed, Sohail. "Experimental investigation of shock wave and boundary layer interaction near convex corners in hypersonic flow." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1997. http://www.collectionscanada.ca/obj/s4/f2/dsk2/ftp01/MQ28817.pdf.

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2

Saad, Mohd Rashdan. "Experimental studies on shock boundary layer interactions using micro-ramps at Mach 5." Thesis, University of Manchester, 2013. https://www.research.manchester.ac.uk/portal/en/theses/experimental-studies-on-shock-boundary-layer-interactions-using-microramps-at-mach-5(71f1e11c-dbfd-443a-a9ee-e3fc160176f1).html.

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Анотація:
Shock boundary layer interactions (SBLI) is an undesirable event occurring in high-speed air-breathing propulsion system that stimulates boundary layer separation due to adverse pressure gradients and consequently lead to ow distortion and pressure loss in the intake section. Therefore it is essential to apply ow control mechanisms to prevent this phenomenon. This study involves a novel ow control device called micro-ramp, which is a part of the micro-vortex generator family that has shown great potential in solving the adverse phenomenon. The term micro refers to the height of the device, which is smaller than the boundary layer thickness, δ. It is important to highlight the two main novelties of this investigation. Firstly, it is the first micro-ramp study conducted in the hypersonic ow regime (Mach 5) since most of the previous micro-ramp studies were only performed in subsonic, transonic and supersonic flows. Another novelty is the various experimental techniques that were used in single study for example schlieren photography, oil-dot and oil- ow visualisation and conventional pressure transducers. In addition, advanced ow diagnostic tools such as infrared thermography, pressure sensitive paints (PSP) and particle image velocimetry (PIV) were also employed. T
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3

Boyd, Robert Raymond. "An Experimental and Computational Investigation on the Effect of Transonic Flow in Hypersonic Wind Tunnel Nozzles, Including Filtered Rayleigh Scattering Measurements /." The Ohio State University, 1996. http://rave.ohiolink.edu/etdc/view?acc_num=osu148793364864785.

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4

Chpoun, Amer. "Contribution a l'etude d'ecoulements hypersoniques (m=5) sur une rampe de compression en configuration 2-d et 3-d." Paris 6, 1988. http://www.theses.fr/1988PA06A005.

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Анотація:
Etude experimentale de l'influence de l'ecoulement transversal sur les distributions du flux thermique et de la pression parietale. Determination des grandeurs caracteristiques de la zone de decollement. Etude de l'apparition de la transition dans la zone du decollement en fonction du nombre de reynolds. Solution numerique pour la distribution de pression dans le cas de l'interaction laminaire
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5

Chpoun, Amer. "Contribution à l'étude d'écoulements hypersoniques (M=5) sur une rampe de compression en configuration 2-D et 3-D." Paris 6, 1988. http://www.theses.fr/1988PA066149.

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Анотація:
Etude expérimentale de l'influence de l'écoulement transversal sur les distributions du flux thermique et de la pression pariétale. Détermination des grandeurs caractéristiques de la zone de décollement. Etude de l'apparition de la transition dans la zone du décollement en fonction du nombre de Reynolds. Solution numérique pour la distribution de pression dans le cas de l'interaction laminaire
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6

Denman, Paul Ashley. "Experimental study of hypersonic boundary layers and base flows." Thesis, Imperial College London, 1996. http://hdl.handle.net/10044/1/45466.

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Анотація:
This experimental study documents the development and separation of a hypersonic boundary layer produced naturally on the cold surface of a sharp slender cone. At the base of the conical forebody, the equilibrium turbulent boundary layer was allowed to separate over an axisymmetric rearward facing step to form a compressible base flow. The investigation was conducted in the Imperial College No.2 gun tunnel at a freestream Mach number of 9 and unit Reynolds numbers of 15 and 55 million. The compressible boundary layer study was carried out at both of the available freestream unit Reynolds numbers and the measured data include distributions of wall static pressure and heat transfer rate, together with profiles of pitot pressure through the boundary layer. Using the chordwise distribution of surface heat flux as a means of transition detection, the cone transition Reynolds number was found to be 5.4x10^. This result, together with that obtained from flat plate studies conducted in the same test facility, provided a ratio of cone to flat plate transition Reynolds number of 0.8. Boundary layer integral quantities and shape factors are derived from velocity profiles and in most cases the measured data extended close enough to the wall to detect the peak values of the integrands. The separated flow region formed at the base of the cone was documented only at the higher unit Reynolds number, a condition under which the approaching turbulent boundary layer was found to be close to equilibrium. The data include pitot pressure profiles recorded normal to the surface downstream of reattachment, together with wall static pressure and heat transfer rate distributions measured throughout the base flow region. Reattachment occurred approximately two step heights downstream of separation and a surface flow visualisation study indicated the existence of Taylor-Goertler type vortices, emanating from the reattachment line in the downstream direction. A simple shear layer expansion model is developed and shown to provide a favourable prediction of the measured pitot pressure profiles recorded downstream of the reattachment line. The success of this second order model implies that the dynamics of the corner expansion process, except in the immediate vicinity of the wall, is governed largely by inviscid pressure mechanisms and that the supersonic region of the boundary layer expansion is essentially isentropic.
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7

Nieberding, Zachary J. "An Investigation of Acoustic Wave Propagation in Mach 2 Flow." University of Cincinnati / OhioLINK, 2014. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1406881591.

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8

O'Dowd, Devin Owen. "Aero-thermal performance of transonic high-pressure turbine blade tips." Thesis, University of Oxford, 2010. http://ora.ox.ac.uk/objects/uuid:e7b8e7d0-4973-4757-b4df-415723e7562f.

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9

Nagashetty, K. "Experimental Investigations on Hypersonic Waverider." Thesis, 2014. http://hdl.handle.net/2005/3195.

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In the flying field of space transportation domain, the increased efforts involving design and development of hypersonic flight for space missions is on toe to provide the optimum aerothermodynamic design data to satisfy mission requirements. Aerothermodynamics is the basis for designing and development of hypersonic space transportation flight vehicles such as X 51 a, and other programmes like planetary probes for Moon and Mars, and Earth re-entry vehicles such as SRE and space shuttle. It enables safe flying of aerospace vehicles, keeping other parameters optimum for structural and materials with thermal protection systems. In this context, the experimental investigations on hypersonic waverider are carried out at design Mach 6. The hypersonic waverider has high lift to drag ratio at design Mach number even at zero degree angle of incidence, and this seems to be one of the special characteristics for its shape at hypersonic flight regime. The heat transfer rates are measured using 30 thin film platinum gauges sputtered on a Macor material that are embedded on the test model. The waverider has 16 sensors on top surface and 14 on bottom surface of a model. The surface temperature history is directly converted to heat transfer rates. The heat transfer data are measured for design (Mach 6) and off-design Mach numbers (8) in the hypersonic shock tunnel, HST2. The results are obtained at stagnation enthalpy of ~ 2 MJ/kg, and Reynolds number range from 0.578 x 106 m-1 to 1.461 x 106 m-1. In addition, flow visualization is carried out by using Schlieren technique to obtain the shock structures and flow evolution around the Waverider. Some preliminary computational analyses are conducted using FLUENT 6.3 and HiFUN, which gave quantitative results. Experimentally measured surface heat flux data are compared with the computed one and both the data agree well. These detailed results are presented in the thesis.
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10

Kumar, Chintoo Sudhiesh. "Experimental Investigation Of Aerodynamic Interference Heating Due To Protuberances On Flat Plates And Cones Facing Hypersonic Flows." Thesis, 2013. http://etd.iisc.ernet.in/handle/2005/2621.

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Анотація:
With the age of hypersonic flight imminent just beyond the horizon, researchers are working hard at designing work-arounds for all the major problems as well as the minor quirks associated with it. One such issue, seemingly innocuous but one that could be potentially deadly, is the problem of interference heating due to surface protuberances. Although an ideal design of the external surfaces of a high-speed aircraft dictates complete smoothness to reduce drag, this is not always possible in reality. Control surfaces, sheet joints, cable protection pads etc. generate surface discontinuities of varying geometries, in the form of both protrusions as well as cavities. These discontinuities are most often small in dimension, comparable to the local boundary layer thickness at that location. Such protuberances always experience high rates of heat transfer, and therefore should be appropriately shielded. However, thermal shielding of the protrusions alone is not a full solution to the problem at hand. The interference caused to the boundary layer by the flow causes the generation of local hot spots in the vicinity of the protuberances, which should be properly mapped and adequately addressed. The work presented in this thesis aims at locating and measuring the heat flux values at these hot spots near the protrusions, and possibly formulating empirical correlations to predict the hot spot heat flux for a given set of flow conditions and protrusion geometry. Experimental investigations were conducted on a flat plate model and a cone model, with interchangeable sharp and blunt nose tips, with attached 3D protuberances. Platinum thin-film sensors were placed around the protrusion so that the heat fluxes could be measured in its vicinity and the hottest spot located. These experiments were carried out at five different hypersonic free stream flow conditions generated using two shock tunnels, one of the conventional type, and the other of the free-piston driven type. The geometry of the protrusions, i.e., the height and the deflection angle, was also parametrically varied to study its effect on the hot spot heat flux. The results thus obtained for the flat plate case were compared to existing correlations in the open literature from a similar previous study at a much higher Reynolds number range. Since a mismatch was observed between the results of the current experiments and the existing correlations, a new empirical correlation has been developed to predict the hot spot heat flux, that is valid within the range of flow conditions studied here. A similar attempt was made for the case of the cone model, for which no previous correlations exist in the open literature. However, a global correlation covering the entire range of flow conditions used here could not be formed. A correlation that is valid for just one out of the five flow conditions used here is presented for the cones with sharp and blunt nose tips separately. Schlieren flow visualization was carried out to obtain a better understanding of the shock structures near the protuberances on both models. For most cases, where the protrusion height and deflection angle were large enough to cause flow separation immediately upstream of the protuberance, a separation shock was manifested which deflected some part of the boundary layer above the protuberance, while the rest of the fluid in the boundary layer entered a recirculating region in the separated zone before escaping to the side. Some preliminary computational analysis was conducted which confirmed this qualitatively. However, the quantitative match of surface heat flux between the simulations and experiments were not encouraging. Schlieren visualization revealed that for the flat plate case, the foot of the separation shock was located at a distance of 10.5 to 12 times the protrusion height ahead of it, whereas in the case of the sharp cone, it was at a distance of 9 to 10.5 times the protrusion height. The unsteady nature of the separation shock was also captured and addressed. Some preliminary experiments on boundary layer tripping were also conducted, the results of which have been presented here. From this analysis, it has become evident that a single global correlation cannot be formed which could be used for a wide range of flow conditions to predict the hot spot heat flux in interference interactions. The entire range of conditions that may be encountered during hypersonic flight has to be broken down into sections, and the interference heating pattern should be studied in each of these sections individually. By doing so, a series of different correlations can be formed at the varying flow conditions which will then be available for high-speed aircraft designers.
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Книги з теми "Experimental Hypersonic flow"

1

R, Keener Earl, Hui Frank C. L, and Ames Research Center, eds. Experimental results for a hypersonic nozzle/afterbody flow field. Moffett Field, Calif: National Aeronautics and Space Administration, Ames Research Center, 1995.

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2

Hughson, M. C. Hypersonic prediction comparisons with experimental data for a cone-cylinder at Mach 6.86. Eglin Air Force Base, Fl: Air Force Armament Laboratory, Eglin Air Force Base, 1989.

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3

W, Paulson John, Weilmuenster K. James, and United States. National Aeronautics and Space Administration., eds. Experimental and computational analysis of shuttle orbiter hypersonic trim anomaly. Washington, DC: American Institute of Aeronautics and Astronautics, 1995.

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4

Nagamatsu, Henry T. Computational, theoretical, and experimental investigation of flow over a sharp flat plate, M1 = 10 - 25. Washington, D. C: American Institute of Aeronautics and Astronautics, 1994.

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5

Center, Langley Research, ed. Mach 10 experimental database of a three-dimensional scramjet inlet flow field. Hampton, Va: National Aeronautics and Space Administration, Langley Research Center, 1995.

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6

Holland, Scott D. Mach 10 experimental database of a three-dimensional scramjet inlet flow field. Hampton, Va: National Aeronautics and Space Administration, Langley Research Center, 1995.

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7

United States. National Aeronautics and Space Administration., ed. Experimental aerothermodynamic research of hypersonic aircraft: Technical progress report (substitution) for the period October 1, 1988 - March 31, 1989. [Washington, DC: National Aeronautics and Space Administration, 1989.

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8

United States. National Aeronautics and Space Administration., ed. Experimental aerothermodynamic research of hypersonic aircraft: Technical progress report (substitution) for the period March 1, 1987 - September 30, 1987. [Washington, DC: National Aeronautics and Space Administration, 1987.

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9

Avery, Don E. Experimental aerodynamic heating to simulated space shuttle tiles in laminar and turbulent boundary layers with variable flow angles at a nominal Mach number of 7. Washington, D.C: National Aeronautics and Space Administration, Scientific and Technical Information Branch, 1985.

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10

Wells, William L. Surface flow and heating distributions on a cylinder in near wake of Aeroassist Flight Experiment (AFE) configuration at incidence in Mach 10 air. Hampton, Va: Langley Research Center, 1990.

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Частини книг з теми "Experimental Hypersonic flow"

1

Henckels, A., and F. Maurer. "Hypersonic Corner Flow Detailed Experimental Study." In Hypersonic Flows for Reentry Problems, 441–54. Berlin, Heidelberg: Springer Berlin Heidelberg, 1992. http://dx.doi.org/10.1007/978-3-642-77922-0_41.

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2

Vetter, M., H. Olivier, and H. Grönig. "Flow over double ellipsoid and sphere — Experimental results." In Hypersonic Flows for Reentry Problems, 489–500. Berlin, Heidelberg: Springer Berlin Heidelberg, 1992. http://dx.doi.org/10.1007/978-3-642-77922-0_45.

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3

Chazot, Olivier. "Aerospace Flight Modeling and Experimental Testing." In Uncertainty in Engineering, 131–47. Cham: Springer International Publishing, 2021. http://dx.doi.org/10.1007/978-3-030-83640-5_9.

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AbstractValidation processes for aerospace flight modeling require to articulate uncertainty quantification methods with the experimental approach. On this note, the specific strategies for the reproduction of re-entry flow conditions in ground-based facilities are reviewed. It shows how it combines high-speed flow physics with the hypersonic wind tunnel capabilities.
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Aupoix, B. "Experimental Validation of Hypersonic Viscous Flow Models." In New Trends in Instrumentation for Hypersonic Research, 41–50. Dordrecht: Springer Netherlands, 1993. http://dx.doi.org/10.1007/978-94-011-1828-6_4.

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5

Aymer, D., T. Alziary, L. De Luca, and G. Carlomagno. "Experimental Study of the Flow Around a Double Ellipsoid Configuration." In Hypersonic Flows for Reentry Problems, 335–57. Berlin, Heidelberg: Springer Berlin Heidelberg, 1991. http://dx.doi.org/10.1007/978-3-642-76527-8_27.

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6

Linde, Magnus. "Leeside Flow over Delta Wing at M = 7.15 Experimental results for Test Case 7.1.2." In Hypersonic Flows for Reentry Problems, 927–32. Berlin, Heidelberg: Springer Berlin Heidelberg, 1991. http://dx.doi.org/10.1007/978-3-642-76527-8_60.

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7

Havermann, M., and F. Seiler. "Boundary Layer Influence on Supersonic Jet/Cross-Flow Interaction in Hypersonic Flow." In New Results in Numerical and Experimental Fluid Mechanics V, 281–88. Berlin, Heidelberg: Springer Berlin Heidelberg, 2006. http://dx.doi.org/10.1007/978-3-540-33287-9_35.

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8

Henckels, A., and F. Maurer. "Experimental Study of the Longitudinal Hypersonic Corner Flow Field Hermes-R&D Research Program, Problem No. 5." In Hypersonic Flows for Reentry Problems, 315–31. Berlin, Heidelberg: Springer Berlin Heidelberg, 1991. http://dx.doi.org/10.1007/978-3-642-76527-8_26.

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Schrijer, F. F. J., R. Caljouw, F. Scarano, and B. W. van Oudheusden. "Three dimensional experimental investigation of a hypersonic double-ramp flow." In Shock Waves, 719–24. Berlin, Heidelberg: Springer Berlin Heidelberg, 2009. http://dx.doi.org/10.1007/978-3-540-85168-4_116.

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Rolim, T. C., and F. K. Lu. "Experimental Investigation of a Hypersonic Inlet with Variable Sidewall for Flow Control." In 29th International Symposium on Shock Waves 2, 1003–8. Cham: Springer International Publishing, 2015. http://dx.doi.org/10.1007/978-3-319-16838-8_33.

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Тези доповідей конференцій з теми "Experimental Hypersonic flow"

1

Hutchins, Kelley, Maruthi Akella, Noel Clemens, and Jeffrey Donbar. "Detection and Transient Dynamics Modeling of Experimental Hypersonic Inlet Unstart." In 6th AIAA Flow Control Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2012. http://dx.doi.org/10.2514/6.2012-2808.

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2

Li, Suxun, Yongkang Chen, Yulin Li, Suxun Li, Yongkang Chen, and Yulin Li. "Hypersonic flow over double-ellipsoid - Experimental investigation." In 15th Applied Aerodynamics Conference. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1997. http://dx.doi.org/10.2514/6.1997-2287.

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3

Bez, Jean-Pierre, Jean-Jacques Chattot, and Francois Noel. "Some Validations by Experimental Results in Hypersonic Flow Computations." In Aerospace Vehicle Conference. 400 Commonwealth Drive, Warrendale, PA, United States: SAE International, 1988. http://dx.doi.org/10.4271/880925.

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4

de Luca, Luigi, and Gennaro Cardone. "Experimental analysis of Goertler vortices in hypersonic wedge flow." In Aerospace Sensing, edited by Jan K. Eklund. SPIE, 1992. http://dx.doi.org/10.1117/12.58544.

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SPAID, FRANK, and EARL KEENER. "Experimental results for a hypersonic nozzle/afterbody flow field." In 28th Joint Propulsion Conference and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1992. http://dx.doi.org/10.2514/6.1992-3915.

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6

Singh, Amarjit, and J. Stollery. "Experimental investigation of hypersonic flow over a wing-body combination." In International Aerospace Planes and Hypersonics Technologies. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1995. http://dx.doi.org/10.2514/6.1995-6083.

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Engblom, W., B. Yuceil, D. Goldstein, and D. Dolling. "Hypersonic forward-facing cavity flow - An experimental and numerical study." In 33rd Aerospace Sciences Meeting and Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 1995. http://dx.doi.org/10.2514/6.1995-293.

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Kaga, T., T. Matsu-Ura, and T. Fujiwara. "Experimental study of ablation gas jet opposing to hypersonic flow." In 40th AIAA Aerospace Sciences Meeting & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2002. http://dx.doi.org/10.2514/6.2002-297.

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Karpuzcu, Irmak Taylan, Deborah A. Levin, Derek Mamrol, Lauren N. Wagner, Mark E. Noftz, Joseph S. Jewell, and Sonya T. Smith. "Jet Flow - Shockwave Interactions in a Hypersonic Flow using Experimental and Kinetic Methods." In AIAA SCITECH 2022 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2022. http://dx.doi.org/10.2514/6.2022-1577.

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Zhongjie, Shao, Xing Chen, Dan Wang, and Wang Yuan. "Experimental study of hypersonic boundary layer flow control based on isolated roughness." In THE 6th NTERNATIONAL CONFERENCE ON FLUID FLOW, HEAT AND MASS TRANSFER. Avestia Publishing, 2019. http://dx.doi.org/10.11159/ffhmt19.153.

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Звіти організацій з теми "Experimental Hypersonic flow"

1

Schneider, Steven P. Completion of the 9.5-Inch Mach-6 Ludwieg Tube: Enabling Low-Cost Hypersonic Quiet-Flow Experiments. Fort Belvoir, VA: Defense Technical Information Center, February 2001. http://dx.doi.org/10.21236/ada387557.

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