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Journal articles on the topic 'Hypersonic wind tunnels'

Author: Grafiati

Published: 4 June 2021

Last updated: 6 September 2023

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1

Stalker, R. J. "Modern developments in hypersonic wind tunnels." Aeronautical Journal 110, no. 1103 (January 2006): 21–39. http://dx.doi.org/10.1017/s0001924000004346.

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AbstractThe development of new methods of producing hypersonic wind-tunnel flows at increasing velocities during the last few decades is reviewed with attention to airbreathing propulsion, hypervelocity aerodynamics and superorbital aerodynamics. The role of chemical reactions in these flows leads to use of a binary scaling simulation parameter, which can be related to the Reynolds number, and which demands that smaller wind tunnels require higher reservoir pressure levels for simulation of flight phenomena. The use of combustion heated vitiated wind tunnels for propulsive research is discussed, as well as the use of reflected shock tunnels for the same purpose. A flight experiment validating shock-tunnel results is described, and relevant developments in shock tunnel instrumentation are outlined. The use of shock tunnels for hypervelocity testing is reviewed, noting the role of driver gas contamination in determining test time, and presenting examples of air dissociation effects on model flows. Extending the hypervelocity testing range into the superorbital regime with useful test times is seen to be possible by use of expansion tube/tunnels with a free piston driver.

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2

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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3

Griffith, Wayland C., William J. Yanta, and William C. Ragsdale. "Supercooling in hypersonic nitrogen wind tunnels." Journal of Fluid Mechanics 269 (June 25, 1994): 283–99. http://dx.doi.org/10.1017/s0022112094001564.

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Recent experimental observation of supercooling in large hypersonic wind tunnels using pure nitrogen identified a broad range of non-equilibrium metastable vapour states of the flow in the test cell. To investigate this phenomenon a number of real-gas effects are analysed and compared with predictions made using the ideal-gas equation of state and equilibrium thermodynamics. The observed limit on the extent of supercooling is found to be at 60% of the temperature difference from the sublimation line to Gibbs’ absolute limit on phase stability. The mass fraction then condensing is calculated to be 12–14%. Included in the study are virial effects, quantization of rotational and vibrational energy, and the possible role of vibrational relaxation and freezing in supercooling. Results suggest that use of the supercooled region to enlarge the Mach–Reynolds number test envelope may be practical. Data from model tests in supercooled flows support this possibility.

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4

Kurshin, Anatolyi Petrovich. "HYPERSONIC WIND TUNNELS BASED ON PRESSURE MULTIPLIERS PART I, PRACTICAL REQUIREMENTS: SCHEMES OF HYPERSONIC WIND TUNNELS." TsAGI Science Journal 49, no. 5 (2018): 527–41. http://dx.doi.org/10.1615/tsagiscij.2018029206.

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5

Tegler, Eric. "The Race To Hypersonic Supremacy." Aerospace Testing International 2022, no. 3 (September 2022): 18–24. http://dx.doi.org/10.12968/s1478-2774(23)50300-8.

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6

McDaniel, R. D., and H. A. Hassan. "Transition Mechanisms in Conventional Hypersonic Wind Tunnels." Journal of Spacecraft and Rockets 38, no. 2 (March 2001): 180–84. http://dx.doi.org/10.2514/2.3691.

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7

Kurshin, Anatolyi Petrovich. "HYPERSONIC WIND TUNNELS BASED ON PRESSURE MULTIPLIERS, PART II: CAPABILITIES OF HYPERSONIC WIND TUNNELS BASED ON VARIOUS SCHEMES." TsAGI Science Journal 49, no. 6 (2018): 625–43. http://dx.doi.org/10.1615/tsagiscij.2018029457.

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8

Nagai, Shinji, Shoichi Tsuda, Tadao Koyama, Noriaki Hirabayashi, Hideo Sekine, and Koichi Hozumi. "Uncertainty of Aerodynamic Coefficients at Hypersonic Wind Tunnels." JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 51, no. 591 (2003): 151–57. http://dx.doi.org/10.2322/jjsass.51.151.

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9

Topchiyan, M. E., and A. M. Kharitonov. "Wind tunnels for hypersonic research (progress, problems, prospects)." Journal of Applied Mechanics and Technical Physics 35, no. 3 (May 1994): 383–95. http://dx.doi.org/10.1007/bf02369878.

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10

Duan, Lian, Meelan M. Choudhari, Amanda Chou, Federico Munoz, Rolf Radespiel, Thomas Schilden, Wolfgang Schröder, et al. "Characterization of Freestream Disturbances in Conventional Hypersonic Wind Tunnels." Journal of Spacecraft and Rockets 56, no. 2 (March 2019): 357–68. http://dx.doi.org/10.2514/1.a34290.

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11

Wagner, Alexander, Erich Schülein, René Petervari, Klaus Hannemann, Syed R. C. Ali, Adriano Cerminara, and Neil D. Sandham. "Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation." Journal of Fluid Mechanics 842 (March 13, 2018): 495–531. http://dx.doi.org/10.1017/jfm.2018.132.

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Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels were conducted by means of a slender wedge probe and direct numerical simulation. The study comprises comparative tunnel noise measurements at Mach 3, 6 and 7.4 in two Ludwieg tube facilities and a shock tunnel. Surface pressure fluctuations were measured over a wide range of frequencies and test conditions including harsh test environments not accessible to measurement techniques such as Pitot probes and hot-wire anemometry. A good agreement was found between normalized Pitot pressure fluctuations converted into normalized static pressure fluctuations and the wedge probe readings. Quantitative results of the tunnel noise are provided in frequency ranges relevant for hypersonic boundary-layer transition. Complementary numerical simulations of the leading-edge receptivity to fast and slow acoustic waves were performed for the applied wedge probe at conditions corresponding to the experimental free-stream conditions. The receptivity to fast acoustic waves was found to be characterized by an early amplification of the induced fast mode. For slow acoustic waves an initial decay was found close to the leading edge. At all Mach numbers, and for all considered frequencies, the leading-edge receptivity to fast acoustic waves was found to be higher than the receptivity to slow acoustic waves. Further, the effect of inclination angles of the acoustic wave with respect to the flow direction was investigated. An inclination angle was found to increase the response on the wave-facing surface of the probe and decrease the response on the opposite surface for fast acoustic waves. A frequency-dependent response was found for slow acoustic waves. The combined numerical and experimental approach in the present study confirmed the previous suggestion that the slow acoustic wave is the dominant acoustic mode in noisy hypersonic wind tunnels.

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12

Safronov, Oleksandr, Bohdan Semon, Oleksandr Nedilko, and Anna Horina. "Conditions for simulation of transonic flutter of aerodynamic control surfaces of supersonic aircraft in wind tunnels." Strength of Materials and Theory of Structures, no. 110 (June 26, 2023): 81–96. http://dx.doi.org/10.32347/2410-2547.2023.110.81-96.

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The transonic flutter of the aerodynamic control surfaces of supersonic and hypersonic aircraft refers to those phenomena of dynamic aeroelasticity, the assessment of whichin a flight experiment is dangerous. Since there is still no universally accepted mathematical model of the occurrence of this phenomenon, tests of dynamic-like models in wind tunnels can be classified as basic and safe methods for evaluating the characteristics of transonic flutter. Therefore, the substantiation of the conditions for modeling transonic flutter, which allow the transfer of the results of blowing of dynamic-like models in wind tunnels to full-scale aircraft designs, remains an actual scientific problem.The article proposes one of the possible approaches to justifying the conditions for modeling transonic flutter of aerodynamic control surfaces in wind tunnels, which is based on the analysis of a nonlinear mathematical model of the occurrence of this phenomenon.Based on the analysis of this mathematical model, it was determined that, in addition to the conditions for modeling transonic flutter in wind tunnels, which are due to the geometric similarity of the system "carrying aerodynamic surface – aerodynamic control surface", additional conditions for modeling this phenomenon should be the following dimensionless quantities of nature and model:- amplitudes of oscillations of nature and model;- numbers M at which transonic flutter occurs;- logarithmic decrements of oscillations of aerodynamic control surfaces;- Strouhal numbers;- ratio of gas density to material density of aerodynamic control surfaces;- the adiabatic index k=1.405, that is, the working body in the wind tunnel should be air.Blowing of dynamically similar models must be carried out in wind tunnels of a continuous type, subject to conditions (s-1).

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13

Lax, Philip A., and Sergey B. Leonov. "Semiempirical Model for Homogeneous Nitrogen Condensation in Hypersonic Wind Tunnels." AIAA Journal 58, no. 11 (November 2020): 4807–18. http://dx.doi.org/10.2514/1.j059519.

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14

Savino, R., R. Monti, and A. Esposito. "Behaviour of hypersonic wind tunnels diffusers at low Reynolds numbers." Aerospace Science and Technology 3, no. 1 (January 1999): 11–19. http://dx.doi.org/10.1016/s1270-9638(99)80018-2.

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15

Alferov, V. I., A. P. Labazkin, and A. P. Rudakova. "Hypersonic flow round models in wind tunnels of various classes." Fluid Dynamics 21, no. 2 (1986): 327–30. http://dx.doi.org/10.1007/bf01050193.

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16

Aulchenko, S. M., V. M. Galkin, V. I. Zvegintsev, and A. N. Shiplyuk. "Numerical design of multimodal axisymmetric hypersonic nozzles for wind tunnels." Journal of Applied Mechanics and Technical Physics 51, no. 2 (March 2010): 218–25. http://dx.doi.org/10.1007/s10808-010-0031-0.

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17

Buck, Gregory M. "Simultaneous luminescence pressure and temperature measurement system for hypersonic wind tunnels." Journal of Spacecraft and Rockets 32, no. 5 (September 1995): 791–94. http://dx.doi.org/10.2514/3.26685.

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18

Shimura, Takashi, Tohru Mitani, Noboru Sakuranaka, and Muneo Izumikawa. "Load Oscillations Caused by Unstart of Hypersonic Wind Tunnels and Engines." Journal of Propulsion and Power 14, no. 3 (May 1998): 348–53. http://dx.doi.org/10.2514/2.5287.

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19

Gutsche, Kevin, Ansgar Marwege, and Ali Gülhan. "Similarity and Key Parameters of Retropropulsion Assisted Deceleration in Hypersonic Wind Tunnels." Journal of Spacecraft and Rockets 58, no. 4 (July 2021): 984–96. http://dx.doi.org/10.2514/1.a34910.

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20

Drozdov, S. M., and A. S. Rtishcheva. "Comparative study of performance of heat exchangers used in hypersonic wind tunnels." Journal of Physics: Conference Series 1683 (December 2020): 022029. http://dx.doi.org/10.1088/1742-6596/1683/2/022029.

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21

Marineau, Eric C., Guillaume Grossir, Alexander Wagner, Madlen Leinemann, Rolf Radespiel, Hideyuki Tanno, Brandon C. Chynoweth, Steven P. Schneider, Ross M. Wagnild, and Katya M. Casper. "Analysis of Second-Mode Amplitudes on Sharp Cones in Hypersonic Wind Tunnels." Journal of Spacecraft and Rockets 56, no. 2 (March 2019): 307–18. http://dx.doi.org/10.2514/1.a34286.

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22

Borovoy, Volf Yakovlevich, Vladimir Nikolaevich Brazhko, Ivan Vladimirovich Egorov, Evgenii Gennadievich Zaitsev, and Arkadii Sergeyevich Skuratov. "DIAGNOSTICS AND NUMERICAL SIMULATION OF FLOWS IN HYPERSONIC SHORT-DURATION WIND TUNNELS." TsAGI Science Journal 44, no. 5 (2013): 609–26. http://dx.doi.org/10.1615/tsagiscij.2014010778.

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23

Kidd, Carl T., and John C. Adams. "Fast-Response Heat-Flux Sensor for Measurement Commonality in Hypersonic Wind Tunnels." Journal of Spacecraft and Rockets 38, no. 5 (September 2001): 719–29. http://dx.doi.org/10.2514/2.3738.

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24

Joly, Veronique, and Claude Marmignon. "Effect of HO Vapor on the Vibrational Relaxation in Hypersonic Wind Tunnels." Journal of Thermophysics and Heat Transfer 11, no. 2 (April 1997): 266–72. http://dx.doi.org/10.2514/2.6233.

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25

Shneider, M. N., S. O. Macheret, and R. B. Miles. "Electric Charge Buildup in Hypersonic Wind Tunnels with Electron-Beam Energy Addition." AIAA Journal 45, no. 7 (July 2007): 1556–61. http://dx.doi.org/10.2514/1.16957.

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26

Freund, J. B., S. K. Lele, P. Moin, Veronique Joly, and Claude Marmignon. "Effect of H2O vapor on the vibrational relaxation in hypersonic wind tunnels." Journal of Thermophysics and Heat Transfer 11 (January 1997): 266–72. http://dx.doi.org/10.2514/3.889.

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27

Gregory, J. W., K. Asai, M. Kameda, T. Liu, and J. P. Sullivan. "A review of pressure-sensitive paint for high-speed and unsteady aerodynamics." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 222, no. 2 (February 1, 2008): 249–90. http://dx.doi.org/10.1243/09544100jaero243.

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The current paper describes the development of pressure-sensitive paint (PSP) technology as an advanced measurement technique for unsteady flow fields and short-duration wind tunnels. Newly developed paint formulations have step response times approaching 1 μs, making them suitable for a wide range of unsteady testing. Developments in binder technology are discussed, which have resulted in new binder formulations such as anodized aluminium, thin-layer chromatography plate, polymer/ceramic, and poly(TMSP) PSP. The current paper also details modeling work done to describe the gas diffusion properties within the paint binder and understand the limitations of the paint response characteristics. Various dynamic calibration techniques for PSP are discussed, along with summaries of typical response times. A review of unsteady and high-speed PSP applications is presented, including experiments with shock tubes, hypersonic tunnels, unsteady delta wing aerodynamics, fluidic oscillator flows, Hartmann tube oscillations, acoustics, and turbomachinery. Flowfields with fundamental frequencies as high as 21 kHz have been successfully measured with porous PSP formulations.

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28

Miller, Charles G. "Refinement of an 'alternate' method for measuring heating rates in hypersonic wind tunnels." AIAA Journal 23, no. 5 (May 1985): 810–12. http://dx.doi.org/10.2514/3.8989.

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29

Simeonides, G., J. P. Vermeulen, H. L. Boerrigter, and J. F. Wendt. "Quantitative heat transfer measurements in hypersonic wind tunnels by means of infrared thermography." IEEE Transactions on Aerospace and Electronic Systems 29, no. 3 (July 1993): 878–93. http://dx.doi.org/10.1109/7.220961.

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30

Zaccara, Mirko, Joshua B. Edelman, and Gennaro Cardone. "A general procedure for infrared thermography heat transfer measurements in hypersonic wind tunnels." International Journal of Heat and Mass Transfer 163 (December 2020): 120419. http://dx.doi.org/10.1016/j.ijheatmasstransfer.2020.120419.

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31

Gu, Sangdi, Jiaao Hao, and Chih-yung Wen. "Air thermochemistry in the converging section of de Laval nozzles on hypersonic wind tunnels." AIP Advances 12, no. 8 (August 1, 2022): 085320. http://dx.doi.org/10.1063/5.0106554.

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State-to-state simulations of nonequilibrium flow in nozzles are made for a range of reservoir conditions and geometries. The geometry of the converging section and throat has little influence on the thermochemistry of the flow. Higher reservoir pressure and temperature both drive the thermochemistry toward equilibrium. For reservoir temperatures of 1500, 4000, and 7000 K, the flow property that has the largest departure from equilibrium is the N2 vibrational temperature, the O mass fraction, and the N mass fraction, respectively. Even at the lowest reservoir pressure, these departures from equilibrium are only 14%, 8%, and 2% for the 1500, 4000, and 7000 K reservoirs, respectively. The differences in these flow properties at the throat between the nonequilibrium and equilibrium simulations are maintained throughout in the nonequilibrium simulations of the diverging section. Applying the simplification of equilibrium flow in the converging section and around the throat yields almost no observable errors in the vibrational population distributions in the diverging section. The simplification is recommended for most practical intents and purposes, and the current work provides important quantitative information to make informed judgments when applying it.

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32

Sokolov, I. V. "Concept of low-entropy compression as applied to the development of chemically clean hypersonic wind tunnels." Technical Physics Letters 23, no. 10 (October 1997): 744–45. http://dx.doi.org/10.1134/1.1261785.

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33

Sun, Haoyu, Jiajie Wang, Yiping Han, Zhiwei Cui, Peng Sun, Xiaowei Shi, and Wenjuan Zhao. "Backward Scattering Characteristics of a Reentry Vehicle Enveloped by a Hypersonic Flow Field." International Journal of Antennas and Propagation 2018 (August 19, 2018): 1–14. http://dx.doi.org/10.1155/2018/5478580.

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The hypersonic flow field around a reentry vehicle has a significant influence on the ground-vehicle communication as well as on the detection and recognition of the reentry vehicle. Backward scattering characteristics of a reentry vehicle enveloped by a hypersonic flow field are analyzed using a high-order auxiliary differential equation finite difference time-domain (ADE-FDTD) method in this paper. Flow field parameters, including electron density, neutral particle density, and temperature, are obtained by solving the Navier-Stokes (NS) equations numerically. According to the flow field parameters, distributions of the plasma frequency and the collision frequency are then derived. Based on a validity of the physical model and the high-order ADE-FDTD method, backward radar cross sections (RCSs) of a perfect electrical conductor (PEC) sphere enveloped by a hypersonic flow field under different Mach numbers, heights, and incident angles of the electromagnetic (EM) wave are then investigated. Numerical results show that the incident angle of the EM wave exerts noticeable effects on the backward RCS, which is due to an inhomogeneous distribution of the plasma. The flight height and Mach number have significant influences on the distribution of the plasma that they play an important role in the variation of the RCSs. The results presented in this paper provide useful reference data for practical tests in ballistic range or in the high-frequency plasma wind tunnels, where a sphere target is usually used due to its simple shape.

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34

Brazier, Jean-Philippe, Jan Martinez Schramm, Sébastien Paris, Thomas Gawehn, and Bodo Reimann. "An overview of HyFIE Technical Research Project: cross-testing in main European hypersonic wind tunnels on EXPERT body." CEAS Space Journal 8, no. 3 (April 23, 2016): 167–76. http://dx.doi.org/10.1007/s12567-016-0117-5.

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35

Grossir, G., B. Dias, O. Chazot, and T. E. Magin. "High temperature and thermal non-equilibrium effects on the determination of free-stream flow properties in hypersonic wind tunnels." Physics of Fluids 30, no. 12 (December 2018): 126102. http://dx.doi.org/10.1063/1.5058098.

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36

Mena, Jesús Garicano, Raffaele Pepe, Andrea Lani, and Herman Deconinck. "Assessment of Heat Flux Prediction Capabilities of Residual Distribution Method: Application to Atmospheric Entry Problems." Communications in Computational Physics 17, no. 3 (March 2015): 682–702. http://dx.doi.org/10.4208/cicp.070414.211114a.

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AbstractIn the present contribution we evaluate the heat flux prediction capabilities of second-order accurate Residual Distribution (RD) methods in the context of atmospheric (re-)entry problems around blunt bodies. Our departing point is the computation of subsonic air flows (with air modeled either as an inert ideal gas or as chemically reacting and possibly out of thermal equilibrium gas mixture) around probe-like geometries, as those typically employed into high enthalpy wind tunnels. We confirm the agreement between the solutions obtained with the RD method and the solutions computed with other Finite Volume (FV) based codes.However, a straightforward application of the same numerical technique to hypersonic cases involving strong shocks exhibits severe deficiencies even on a geometry as simple as a 2D cylinder. In an attempt to mitigate this problem, we derive new variants of RD schemes. A comparison of these alternative strategies against established ones allows us to derive a diagnose for the shortcomings observed in the traditional RD schemes.

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37

Alferov, V. I., A. S. Bushmin, and I. V. Egorov. "Experimental investigation of flow past simple model bodies in hypersonic wind tunnels at similar values of the Mach and Reynolds numbers but at different physical flow velocities." Fluid Dynamics 50, no. 1 (January 2015): 109–17. http://dx.doi.org/10.1134/s0015462815010123.

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38

Corke, Thomas, Alexander Arndt, Eric Matlis, and Michael Semper. "Control of stationary cross-flow modes in a Mach 6 boundary layer using patterned roughness." Journal of Fluid Mechanics 856 (October 11, 2018): 822–49. http://dx.doi.org/10.1017/jfm.2018.636.

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Experiments were performed to investigate passive discrete roughness for transition control on a sharp right-circular cone at an angle of attack at Mach 6.0. A cone angle of attack of $6^{\circ }$ was set to produce a mean cross-flow velocity component in the boundary layer over the cone by which the cross-flow instability was the dominant mechanism of turbulent transition. The approach to transition control is based on exciting less-amplified (subcritical) stationary cross-flow modes through the addition of discrete roughness that suppresses the growth of the more-amplified (critical) cross-flow modes, and thereby delays transition. The passive roughness consisted of indentations (dimples) that were evenly spaced around the cone at an axial location that was just upstream of the first linear stability neutral growth branch for cross-flow modes. The experiments were performed in the air force academy (AFA) Mach 6.0 Ludwieg Tube Facility. The cone model was equipped with a motorized three-dimensional traversing mechanism that mounted on the support sting. The traversing mechanism held a closely spaced pair of fast-response total pressure Pitot probes. The measurements consisted of surface oil flow visualization and off-wall azimuthal profiles of mean and fluctuating total pressure at different axial locations. These documented an 25 % increase in the transition Reynolds number with the subcritical roughness. In addition, the experiments revealed evidence of a nonlinear, sum and difference interaction between stationary and travelling cross-flow modes that might indicate a mechanism of early transition in conventional (noisy) hypersonic wind tunnels.

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39

Miles, Richard B., Garry L. Brown, Walter R. Lempert, Richard Yetter, George J. Williams, Seymour M. Bogdonoff, Douglas Natelson, and Jeffrey R. Guest. "Radiatively driven hypersonic wind tunnel." AIAA Journal 33, no. 8 (August 1995): 1463–70. http://dx.doi.org/10.2514/3.12568.

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40

Li, Shichao, Zihao Liu, Fan Zhao, and Hongli Gao. "A New Hypersonic Wind Tunnel Force Measurement System to Reduce Additional Bending Moment and Avoid Time-Varying Stiffness." Sensors 22, no. 7 (March 27, 2022): 2572. http://dx.doi.org/10.3390/s22072572.

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In order to improve traditional hypersonic wind tunnel airframe/propulsion integrated aerodynamic testing technology for hypersonic vehicles, a new force measurement system called the aerodynamic force measuring support (AFMS) was designed. The AFMS effectively overcomes the defect that the traditional internal box-balance occupies a large amount of internal space in the aircraft test model, which makes the airframe/propulsion integrated aerodynamic test more difficult. The AFMS also alleviates the interference of the additional bending moment caused by the non-coincidence between the calibration center of traditional external box-balance and the gravity center of the aircraft test model, innovatively designing a convex structure in the joint part of the force measuring system. Furthermore, the AFMS effectively overcomes the time-varying stiffness of joints caused by test model vibration in hypersonic wind tunnel testing, which eventually leads to test errors. Compared with the traditional box-balance, the AFMS proposed in this study has sufficient design space. This ensures more thorough aerodynamic decomposition of the AFMS and less interference between channels, whilst also having the advantages of the large support stiffness of traditional box-balance. Thus, the AFMS provides a new technical path for airframe/propulsion integrated aerodynamic testing of air-breathing hypersonic vehicles in a hypersonic wind tunnel.

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41

HUANG, Ju, Yongneng YANG, Qi LIU, Haibin YANG, and Wei ZHANG. "Developing and applying Mach 4.5 nozzle in hypersonic wind tunnel." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 39, no. 5 (October 2021): 1064–69. http://dx.doi.org/10.1051/jnwpu/20213951064.

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Mach 4.5 tests in a conventional trans-supersonic wind tunnel are often accompanied by the air liquefaction phenomenon, resulting in the low reliability of test data. The Mach 4.5 nozzle developed in a hypersonic wind tunnel is able to heat airflow and provide more accurate test data. At present, China does not have the capability to test the Mach 4.5 nozzle in the 0.5-meter hypersonic wind tunnel. This gap may be filled by developing the Mach 4.5 nozzle in the hypersonic wind tunnel. The axisymmetric nozzle profile was calculated by the inviscid flow calculation method, and the boundary layer was modified by the Sivells-Payne method. Then, the numerical simulation was carried out, and the simulation results prove that the nozzle profile thus calculated meets the design requirements of the Mach number. For its structural design, a three-section design method is adopted to ensure the continuity and smoothness of the inner surface so as to better calibrate the flow field. Standard model tests were also carried out. The test results show that the velocity field of the Mach 4.5 nozzle we developed meets technical requirements. The standard model test data provide data reliable support for the development of aircraft.

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42

Wang, Xiaoguang, Miaojiao Peng, Zhenghong Hu, Yueshi Chen, and Qi Lin. "Feasibility investigation of large-scale model suspended by cable-driven parallel robot in hypersonic wind tunnel test." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 231, no. 13 (August 8, 2016): 2375–83. http://dx.doi.org/10.1177/0954410016662067.

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Cable-driven parallel robot is a special kind of robot, which is actuated by cables. It is already applied in the low speed wind tunnel to get aerodynamic measurement of aircraft model, and the aircraft pose could be adjusted by changing the cable length. Whether it can be used in hypersonic wind tunnel still needs further discussion. This paper presents the dynamics and aerodynamics analysis of a large-scale model supported by 6-DOF cable-driven parallel robot to investigate the feasibility of this special kind of suspension system in hypersonic wind tunnel. The description of this setup with a X-51A-like model is given, and then based on the system dynamic equations, aerodynamic force and stiffness matrix are derived. In the simulation, properties of dynamics and aerodynamics are mainly concerned. A typical shock tunnel with flow duration of about 100 milliseconds is taken as an example, and results show that the system is stable enough to meet the fundamental static wind tunnel test. From the cable tension variation under impact load and the sensitivity analysis, it is likely accessible to derive the aerodynamic forces. Compared with the sting suspension method, cable-driven parallel robot has the priority of higher inherent frequency and more flexible degrees. The interference to the flow field induced by cables is also preliminarily proved to be small by the CFD simulation, which can be acceptable and corrected. Researches conducted show the feasibility of cable-driven parallel robot’s application in hypersonic wind tunnel.

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Dong, Hao, Shicheng Liu, Xi Geng, and Keming Cheng. "Study on Oil-Film Interferometry Measurement Technique of Hypersonic Boundary Layer Transition." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 36, no. 6 (December 2018): 1156–61. http://dx.doi.org/10.1051/jnwpu/20183661156.

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Prediction of boundary layer transition is important for the design of hypersonic aircrafts. The study of boundary layer transition of hypersonic flow around a flat plate using oil-film interferometry was investigated at Φ500mm traditional hypersonic wind tunnel. In order to measure the skin friction fast and precisely on the hypersonic wind tunnel, the traditional oil-film interferometry technique is improved. A high-speed camera is used to capture the images of fringes and the viscosity of the silicon oil is modified according to the wall temperature measured by thermocouples during the test. The skin frictions of smooth surface and the surface with single square roughness element were measured. For the smooth surface, the boundary layer is laminar. However, the boundary layer transition is promoted by wake vortices induced by the roughness element. Both the results of skin friction with and without the roughness element are in good agreement with the simulation results correspondingly, indicating high accuracy of the oil film interferometry technique.

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Nagai, Shinji, Shoichi Tsuda, Tadao Koyama, Noriaki Hirabayashi, and Hideo Sekine. "Humidity Management for a Hypersonic Wind Tunnel." JOURNAL OF THE JAPAN SOCIETY FOR AERONAUTICAL AND SPACE SCIENCES 55, no. 645 (2007): 483–89. http://dx.doi.org/10.2322/jjsass.55.483.

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NOMURA, Shigeaki, Seizou SAKAKIBARA, Kouichi HOZUMI, and Kunio SOGA. "On NAL's New Large Hypersonic Wind Tunnel." Journal of the Japan Society for Aeronautical and Space Sciences 42, no. 480 (1994): 32–38. http://dx.doi.org/10.2322/jjsass1969.42.32.

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NOMURA, Ryo, Hideki KAWAMOTO, Hidenori YOSHIDA, Takeshi YONEDA, and Shigeru AOKI. "Geometry Optimizatin of Hypersonic Wind Tunnel Nozzle." Transactions of the Japan Society of Mechanical Engineers Series B 70, no. 700 (2004): 3051–57. http://dx.doi.org/10.1299/kikaib.70.3051.

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

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

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Zhu, Xiao-Jun, and Feng Li. "Exploration on application of dredging thermal protection in the leading edge of the wing." International Journal of Modern Physics B 34, no. 14n16 (April 20, 2020): 2040105. http://dx.doi.org/10.1142/s0217979220401050.

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Aiming at the severe aerodynamic heating problem in the leading edge of the hypersonic vehicle, in order to ensure the sharp shape of the leading edge of the wing, a dredging thermal protection structure is proposed, and the built-in high-temperature heat pipe structure is used to provide thermal protection for the leading edge of the wing. By means of numerical simulation and arc wind tunnel test, the dredging thermal protection structure of the leading edge of the wing is analyzed, and the thermal protection effect of the built-in high-temperature heat pipe is obtained. The numerical results show that under certain thermal conditions, the temperature at the leading edge of the wing decreases by 304 K, and the minimum temperature of the tail increases by 130 K. The heat flow is dredged from the high-temperature zone to the low-temperature zone, and the thermal load of the leading edge of the wing is weakened. The same result can be obtained by the arc wind tunnel test, which verifies the accuracy of the numerical method and the feasibility of the dredging thermal protection structure with high-temperature heat pipe embedded in the leading edge of the wing.

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Kawashima, Takashi, Hidenori Yoshida, Hiromitsu Kiyose, and Hideki Kawamoto. "A Numerical Simulation of Hypersonic Wind Tunnel Flow." Transactions of the Japan Society of Mechanical Engineers Series B 60, no. 580 (1994): 4082–88. http://dx.doi.org/10.1299/kikaib.60.4082.

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Huang, Min, Zhong-wei Wang, Xing-Bao Yang, Zhen-yun Guo, and Yao-bin Niu. "Preliminary Validation of the Wind Tunnel Based Flight Control System Evaluation Method." MATEC Web of Conferences 179 (2018): 03021. http://dx.doi.org/10.1051/matecconf/201817903021.

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As mathematical models of aircraft aerodynamics and rudder loadings always have to be built in the hardware-in-the-loop simulation, wind tunnel based flight control system (FCS) evaluation methods were proposed in order to test and evaluate the flight control systems under real aerodynamic and rudder loading environment. To validate the evaluation method, a wind tunnel based flight control system test was performed in a hypersonic wind tunnel facility. As the aircraft support rig in the wind tunnel is static, the aircraft angle of attack cannot be changed in this test. During the test, the elevator response, the lift force and the pitching moment were measured. By analysing the measured data, the elevator control performance of the pitch control system was determined, and the pitch angle was successfully predicted, but the open-loop pitch control performance was not determined. These results validate the feasibility of evaluating the elevator control performance and predicting the pitch angle of a FCS by the wind tunnel based FCS evaluation method.

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