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Статті в журналах з теми "Free-piston driven shock tunnel"

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Автор: Grafiati

Опубліковано: 6 вересня 2023

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1

Chang, Eric Won Keun, Wilson Y. K. Chan, Keill J. Hopkins, Timothy J. McIntyre, and Ananthanarayanan Veeraragavan. "Electrically-heated flat plate testing in a free-piston driven shock tunnel." Aerospace Science and Technology 103 (August 2020): 105856. http://dx.doi.org/10.1016/j.ast.2020.105856.

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2

Macrossan, M. N. "Hypervelocity flow of dissociating nitrogen downstream of a blunt nose." Journal of Fluid Mechanics 217 (August 1990): 167–202. http://dx.doi.org/10.1017/s0022112090000672.

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The nature of the non-equilibrium flow of strongly dissociating nitrogen has been investigated by a series of simulation calculations using non-equilibrium (finite rate) chemical reactions. These were made with the equilibrium flux method (EFM), and the results have been found to compare favourably with experimental results obtained with a free-piston driven shock-tube wind tunnel which was used to obtain interferograms of the flow of pure nitrogen over a blunt-nosed body, 65 mm long at three angles of incidence. No simple relation between the flow with non-equilibrium chemistry and those for frozen or equilibrium chemistry has been found. The problems of relating test flows produced in the shock tunnel to flight conditions are investigated by considering the test flows that might be produced by some ‘ideal equivalent wind tunnels’. It is shown that the degree of frozen dissociation in the test flow in a shock tunnel is not a serious matter, but that the large difference in Mach number between shock tunnel flows and flight conditions may be more important.

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3

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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4

ITOH, Katsuhiro, Tomoyuki KOMURO, Kazuo SATO, Shuichi UEDA, Hideyuki TANNO, and Masahiro TAKAHASHI. "Characteristics of Free-Piston Shock Tunnel HIEST. 1st Report. Tuned Operation of Free-Piston Driver." Transactions of the Japan Society of Mechanical Engineers Series B 68, no. 675 (2002): 2968–75. http://dx.doi.org/10.1299/kikaib.68.2968.

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5

Itoh, K., S. Ueda, T. Komuro, K. Sato, M. Takahashi, H. Miyajima, H. Tanno, and H. Muramoto. "Improvement of a free piston driver for a high-enthalpy shock tunnel." Shock Waves 8, no. 4 (August 1, 1998): 215–33. http://dx.doi.org/10.1007/s001930050115.

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6

Boyce, R. R., M. Takahashi, and R. J. Stalker. "Mass spectrometric measurements of driver gas arrival in the T4 free-piston shock-tunnel." Shock Waves 14, no. 5-6 (October 27, 2005): 371–78. http://dx.doi.org/10.1007/s00193-005-0276-3.

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7

Jacobs, P. A., and R. J. Stalker. "Mach 4 and Mach 8 axisymmetric nozzles for a high-enthalpy shock tunnel." Aeronautical Journal 95, no. 949 (November 1991): 324–34. http://dx.doi.org/10.1017/s0001924000024209.

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AbstractThis study examines the performance of two axisymmetric nozzles which were designed to produce uniform, parallel flow with nominal Mach numbers of 4 and 8. A free-piston-driven shock tube was used to supply the nozzle with high-temperature, high-pressure test gas. The inviscid design procedure treated the nozzle expansion in two stages. Close to the nozzle throat, the nozzle wall was specified as conical and the gas flow was treated as a quasi-one-dimensional chemically-reacting flow. At the end of the conical expansion, the gas was assumed to be calorically perfect and a contoured wall was designed (using Method-of-Characteristics) to convert the source flow into a uniform and parallel flow at the end of the nozzle. Performance was assessed by measuring Pitot pressures across the exit plane of the nozzles and, over the range of operating conditions examined, the nozzles produced satisfactory test flows. However, there were flow disturbances in the Mach 8 nozzle flow that persisted for significant times after flow initiation.

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8

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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9

Tani, K., K. Itoh, M. Takahashi, H. Tanno, T. Komuro, and H. Miyajima. "Numerical study of free-piston shock tunnel performance." Shock Waves 3, no. 4 (December 1994): 313–19. http://dx.doi.org/10.1007/bf01415829.

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10

Terao, Kunio, and Tomonobu Furushoh. "Shock Tube Using Free Piston Driven by Detonation Waves." Japanese Journal of Applied Physics 33, Part 1, No. 5A (May 15, 1994): 2811–16. http://dx.doi.org/10.1143/jjap.33.2811.

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11

Jacobs, P. A. "Quasi-one-dimensional modeling of a free-piston shock tunnel." AIAA Journal 32, no. 1 (January 1994): 137–45. http://dx.doi.org/10.2514/3.11961.

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12

Labracherie, L., M. P. Dumitrescu, Y. Burtschell, and L. Houas. "On the compression process in a free-piston shock-tunnel." Shock Waves 3, no. 1 (March 1993): 19–23. http://dx.doi.org/10.1007/bf01414744.

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13

Tanno, H., T. Komuro, K. Sato, K. Fujita, and S. J. Laurence. "Free-flight measurement technique in the free-piston high-enthalpy shock tunnel." Review of Scientific Instruments 85, no. 4 (April 2014): 045112. http://dx.doi.org/10.1063/1.4870920.

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14

Stacey, C. H. B., and J. M. Simmons. "Measurement of shock-wave/boundary-layer interaction in a free-piston shock tunnel." AIAA Journal 30, no. 8 (August 1992): 2095–98. http://dx.doi.org/10.2514/3.11185.

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15

Mee, D. J., and C. P. Goyne. "Turbulent spots in boundary layers in a free-piston shock-tunnel flow." Shock Waves 6, no. 6 (December 1996): 337–43. http://dx.doi.org/10.1007/bf02511324.

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16

Altenhöfer, P., T. Sander, F. Koroll, and Ch Mundt. "LIGS measurements in the nozzle reservoir of a free-piston shock tunnel." Shock Waves 29, no. 2 (February 10, 2018): 307–20. http://dx.doi.org/10.1007/s00193-018-0808-2.

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17

Mee, D. J., and C. P. Goyne. "Turbulent spots in boundary layers in a free-piston shock-tunnel flow." Shock Waves 6, no. 6 (December 1, 1996): 337–1. http://dx.doi.org/10.1007/s001930050052.

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18

Krek, R. M., and R. J. Stalker. "Experiments on Space Shuttle Orbiter models in a free piston shock tunnel." Aeronautical Journal 96, no. 957 (September 1992): 249–59. http://dx.doi.org/10.1017/s0001924000050399.

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Анотація:

AbstractHeat transfer and pressure measurements were made on a model of the United States Space Shuttle Orbiter in the University of Queensland's T4 shock tunnel at three angles of attack, with stagnation enthalpies which varied by a factor of 11, from 2·1 MJ/kg to 23 MJ/kg and normal shock Reynolds numbers which varied by a factor of 38, from 2·1 x l04 to 8·l x 105. Leeward pressure results were obtained for comparison with flight data and equilibrium calculations, but the majority of the experiments were conducted to investigate the heat transfer distributions around the shuttle model. Both the windward and leeward heat transfer results exhibit the onset of transition to turbulent flow. The leeward results are compared with flight data as well as with conventional wind tunnel data, and are used to establish trends associated with variation in the major flow and geometry parameters. It was found that although high enthalpy effects could be important, Reynolds number effects played a dominant role in determining the flow.

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19

Abe, T., E. Ogura, S. Sato, and K. Funabiki. "Rupture-disk-less shock-tube with compression tube driven by free piston." Shock Waves 7, no. 4 (August 1, 1997): 205–9. http://dx.doi.org/10.1007/s001930050076.

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20

Boyce, R. R., M. Takahashi, and R. J. Stalker. "Mass spectrometric measurements of the freestream composition in the T4 free-piston shock-tunnel." Shock Waves 14, no. 5-6 (October 27, 2005): 359–70. http://dx.doi.org/10.1007/s00193-005-0275-4.

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21

Mundt, Christian, Russell Boyce, Peter Jacobs, and Klaus Hannemann. "Validation study of numerical simulations by comparison to measurements in piston-driven shock-tunnels." Aerospace Science and Technology 11, no. 2-3 (March 2007): 100–109. http://dx.doi.org/10.1016/j.ast.2006.12.002.

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22

McGilvray, M., A. G. Dann, and P. A. Jacobs. "Modelling the complete operation of a free-piston shock tunnel for a low enthalpy condition." Shock Waves 23, no. 4 (March 12, 2013): 399–406. http://dx.doi.org/10.1007/s00193-013-0437-8.

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23

Mallinson, S. G., S. L. Gai, and N. R. Mudford. "Establishment of steady separated flow over a compression-corner in a free-piston shock tunnel." Shock Waves 7, no. 4 (August 1, 1997): 249–53. http://dx.doi.org/10.1007/s001930050080.

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24

Gildfind, David E., Chris M. James, Pierpaolo Toniato, and Richard G. Morgan. "Performance considerations for expansion tube operation with a shock-heated secondary driver." Journal of Fluid Mechanics 777 (July 20, 2015): 364–407. http://dx.doi.org/10.1017/jfm.2015.349.

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A shock-heated secondary driver is a modification typically applied to an expansion tube which involves placing a volume of helium between the primary diaphragm and the test gas. This modification is normally used to either increase the driven shock strength through the test gas for high-enthalpy conditions, or to prevent transmission of primary driver flow disturbances to the test gas for low-enthalpy conditions. In comparison to the basic expansion tube, a secondary driver provides an additional configuration parameter, adds mechanical and operational complexity, and its effect on downstream flow processes is not trivial. This paper reports on a study examining operation of a shock-heated secondary driver across the entire operating envelope of a free-piston-driven expansion tube, using air as the test gas. For high-enthalpy conditions it is confirmed that the secondary driver can provide a performance increase, and it is further shown how this device can be used to fine tune the flow condition even when the free-piston driver configuration is held constant. For low-enthalpy flow conditions, wave processes through the driven tube are too closely coupled, and the secondary driver no longer significantly influences the magnitude of the final test gas flow properties. It is found that these secondary driver operating characteristics depend principally on the initial density ratio between the secondary driver helium gas and the downstream test gas.

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25

Benhegouga, Islem, and Adel Boukehili. "A Design of a Mechanical Synthetic Jet Actuator." Applied Mechanics and Materials 152-154 (January 2012): 1516–21. http://dx.doi.org/10.4028/www.scientific.net/amm.152-154.1516.

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To show the influence of the synthetic jet on the wake of a cylinder and the aerodynamics performances, a piston-rod-crank actuator was designed to provide a synthetic jet with amplitude up to 20% of the free-stream velocity from the wind tunnel. The rod is driven in rotation by an electrical motor with variable frequency and maximum speed of 2800 rev/min corresponding to a frequency of 47 Hz.

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26

Hornung, Hans, Chihyung Wen, and Patrick Germain. "Hypervelocity Flow Simulation." Applied Mechanics Reviews 47, no. 6S (June 1, 1994): S14—S19. http://dx.doi.org/10.1115/1.3124393.

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Many of the flow problems associated with flight vehicles designed to reach or return from space can not be solved computationally. It is essential to address them by experiment, in particular, by ground simulation of the flow. The requirements and most successful simulation techniques are described, and their important limitations are discussed. Two selected examples are then presented from the free-piston reflected shock tunnel T5 at Caltech: Dissociating flow over spheres and transition from laminar to turbulent flow on a slender cone.

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27

Sriram, R., and G. Jagadeesh. "Shock-tunnel investigations on the evolution and morphology of shock-induced large separation bubbles." Aeronautical Journal 120, no. 1229 (June 7, 2016): 1123–52. http://dx.doi.org/10.1017/aer.2016.45.

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ABSTRACTShock-tunnel experiments are carried out to study the strong interaction between an impinging shock wave and boundary layer on a flat plate, accompanied bylargeseparation bubble with a length comparable to the distance of the location of shock impingement from the leading edge of the plate. For nominal freestream Mach numbers ranging from 6 to 8.5, moderate to high total enthalpies of 1.3MJ/kg to 6MJ/kg are simulated in the Indian Institute of Science's hypersonic shock tunnels HST-2 (a conventional Hypersonic Shock Tunnel) and Free Piston Shock Tunnel (FPST) with freestream Reynolds numbers ranging from 4 × 106/m to 0.3 × 106/m. The strong impinging shock is generated by a wedge (or shock generator) at an angle of 30.96° to the freestream. From the time-resolved Schlieren flow visualisations using a high-speed camera and surface pressure measurements on the flat plate using fast response sensors, a statistically steady flow field with a large separation bubble was established within the short test time of the shock tunnels (around 600µs in HST-2 and 300µs in FPST). The role of various parameters on the interaction – Mach number, location of shock impingement and flow total enthalpy – are investigated from the measured separation length and surface pressure distribution. For the nominal Mach number of 8.5, with shock impingement at 100mm from the leading edge, the separation length increased from 60mm to 70mm as the total enthalpy is increased from 1.6MJ/kg to 2.4MJ/kg; whereas it dropped drastically to 30-40mm at 6MJ/kg. This is due to the prominence of real gas effects at higher enthalpies.

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28

Paull, A., R. J. Stalker, and D. J. Mee. "Experiments on supersonic combustion ramjet propulsion in a shock tunnel." Journal of Fluid Mechanics 296 (August 10, 1995): 159–83. http://dx.doi.org/10.1017/s0022112095002096.

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Measurements have been made of the propulsive effect of supersonic combustion ramjets incorporated into a simple axisymmetric model in a free piston shock tunnel. The nominal Mach number was 6, and the stagnation enthalpy varied from 2.8 to 8.5 MJ kg−1. A mixture of 13% silane and 87% hydrogen was used as fuel, and experiments were conducted at equivalence ratios up to approximately 0.8. The measurements involved the axial force on the model, and were made using a stress wave force balance, which is a recently developed technique for measuring forces in shock tunnels. A net thrust was experienced up to a stagnation enthalpy of 3.7 MJ kg−1, but as the stagnation enthalpy increased, an increasing net drag was recorded. Pilot and static pressure measurements showed that the combustion was supersonic.The results were found to compare satisfactorily with predictions based on established theoretical models, used with some simplifying approximations. The rapid reduction of net thrust with increasing stagnation enthalpy was seen to arise from increasing precombustion temperature, showing the need to control this variable if thrust performance was to be maintained over a range of stagnation enthalpies. Both the inviscid and viscous drag were seen to be relatively insensitive to stagnation enthalpy, with the combustion chambers making a particularly significant contribution to drag. The maximum fuel specific impulse achieved in the experiments was only 175 s, but the theory indicates that there is considerable scope for improvement on this through aerodynamic design.

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29

Jayaram, Vishakantaiah, Singh Preetam, and K. P. J. Reddy. "Experimental Investigation of Nano Ceramic Material Interaction with High Enthalpy Argon under Shock Dynamic Loading." Applied Mechanics and Materials 83 (July 2011): 66–72. http://dx.doi.org/10.4028/www.scientific.net/amm.83.66.

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Indigenously designed and fabricated free piston driven shock tube (FPST) was used to generate strong shock heated test gases for the study of aero-thermodynamic reactions on ceramic materials. The reflected shock wave at the end of the shock tube generates high pressure and temperature test gas (Argon, Ar) for short duration. Interaction of materials with shock heated Ar gas leads to formation of a new solid or stabilization of a material in new crystallographic phase. In this shock tube, the generated shock waves was utilized to heat Ar to a very high temperature (11760 K) at about 40-55 bar for 2-4 ms. We confirmed the phase transformation and electronic structure of the material after exposure to shock by XRD and XPS studies. This high enthalpy gas generated in the shock-tube was utilized to synthesize cubic perovskite CeCrO3from fluorite Ce0.5Cr0.5O2+δoxide. We were able to demonstrate that this ceramic materials undergoes phase transformations with the interaction of high enthalpy gas under shock dynamic loading.

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30

He, Y., and R. G. Morgan. "Transition of compressible high enthalpy boundary layer flow over a flat plate." Aeronautical Journal 98, no. 972 (February 1994): 25–34. http://dx.doi.org/10.1017/s0001924000050181.

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AbstractThis paper presents the results of an experimental investigation into the characteristics of boundary layer transition to turbulence in hypervelocity air flows. A series of experiments was conducted using a flat plate model, equipped with static pressure and thin film heat transfer transducers, in a free piston shock tunnel. Transition was observed in the stagnation enthalpy range of 2·35 to 19·2 MJ/kg. The transition Reynolds number correlates well with the unit Reynolds number through a simple empirical relation. The influences of Mach number, pressure and wall cooling are examined. The measured heat transfer rates in laminar and turbulent regions are compared with empirical predictions. Freestream disturbances of the test flow were also measured and analysed.

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31

Neely, A. J., and R. G. Morgan. "The Superorbital Expansion Tube concept, experiment and analysis." Aeronautical Journal 98, no. 973 (March 1994): 97–105. http://dx.doi.org/10.1017/s0001924000050107.

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Abstract In response to the need for ground testing facilities for super orbital re-entry research, a small scale facility has been set up at the University of Queensland to demonstrate the Superorbital Expansion Tube concept. This unique device is a free piston driven, triple diaphragm, impulse shock facility which uses the enthalpy multiplication mechanism of the unsteady expansion process and the addition of a secondary shock driver to further heat the driver gas. The pilot facility has been operated to produce quasi-steady test flows in air with shock velocities in excess of 13 km/s and with a usable test flow duration of the order of 15 μs. An experimental condition produced in the facility with total enthalpy of 108 MJ/kg and a total pressure of 335 MPa is reported. A simple analytical flow model which accounts for non-ideal rupture of the light tertiary diaphragm and the resulting entropy increase in the test gas is discussed. It is shown that equilibrium calculations more accurately model the unsteady expansion process than calculations assuming frozen chemistry. This is because the high enthalpy flows produced in the facility can only be achieved if the chemical energy stored in the test flow during shock heating of the test gas is partially returned to the flow during the process of unsteady expansion. Measurements of heat transfer rates to a flat plate demonstrate the usability of the test flow for aerothermodynamic testing and comparison of these rates with empirical calculations confirms the usable accuracy of the flow model.

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32

MALLINSON, S. G., S. L. GAI, and N. R. MUDFORD. "The interaction of a shock wave with a laminar boundary layer at a compression corner in high-enthalpy flows including real gas effects." Journal of Fluid Mechanics 342 (July 10, 1997): 1–35. http://dx.doi.org/10.1017/s0022112097005673.

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Анотація:

The high-enthalpy, hypersonic flow over a compression corner has been examined experimentally and theoretically. Surface static pressure and heat transfer distributions, along with some flow visualization data, were obtained in a free-piston shock tunnel operating at enthalpies ranging from 3 MJ kg−1 to 19 MJ kg−1, with the Mach number varying from 7.5 to 9.0 and the Reynolds number based on upstream fetch from 2.7×104 to 2.7×105. The flow was laminar throughout. The experimental data compared well with theories valid for perfect gas flow and with other relevant low-to-moderate enthalpy data, suggesting that for the current experimental conditions, the real gas effects on shock wave/boundary layer interaction are negligible. The flat-plate similarity theory has been extended to include equilibrium real gas effects. While this theory is not applicable to the current experimental conditions, it has been employed here to determine the potential maximum effect of real gas behaviour. For the flat plate, only small differences between perfect gas and equilibrium gas flows are predicted, consistent with experimental observations. For the compression corner, a more rapid rise to the maximum pressure and heat transfer on the ramp face is predicted in the real gas flows, with the pressure lying slightly below, and the heat transfer slightly above, the perfect gas prediction. The increase in peak heat transfer is attributed to the reduction in boundary layer displacement thickness due to real gas effects.

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33

SENGUPTA, T. K., T. T. LIM, SHARANAPPA V. SAJJAN, S. GANESH, and J. SORIA. "Accelerated flow past a symmetric aerofoil: experiments and computations." Journal of Fluid Mechanics 591 (October 30, 2007): 255–88. http://dx.doi.org/10.1017/s0022112007008464.

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Accelerated flow past a NACA 0015 aerofoil is investigated experimentally and computationally for Reynolds number Re = 7968 at an angle of attack α = 30°. Experiments are conducted in a specially designed piston-driven water tunnel capable of producing free-stream velocity with different ramp-type accelerations, and the DPIV technique is used to measure the resulting flow field past the aerofoil. Computations are also performed for other published data on flow past an NACA 0015 aerofoil in the range 5200 ≤ Re ≤ 35000, at different angles of attack. One of the motivations is to see if the salient features of the flow captured experimentally can be reproduced numerically. These computations to solve the incompressible Navier–Stokes equation are performed using high-accuracy compact schemes. Load and moment coefficient variations with time are obtained by solving the Poisson equation for the total pressure in the flow field. Results have also been analysed using the proper orthogonal decomposition technique to understand better the evolving vorticity field and its dependence on Reynolds number and angle of attack. An energy-based stability analysis is performed to understand unsteady flow separation.

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34

Collen, Peter, Luke J. Doherty, Suria D. Subiah, Tamara Sopek, Ingo Jahn, David Gildfind, Rowland Penty Geraets, et al. "Development and commissioning of the T6 Stalker Tunnel." Experiments in Fluids 62, no. 11 (October 10, 2021). http://dx.doi.org/10.1007/s00348-021-03298-1.

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Abstract The T6 Stalker Tunnel is a multi-mode, high-enthalpy, transient ground test facility. It is the first of its type in the UK. The facility combines the original free-piston driver from the T3 Shock Tunnel with modified barrels from the Oxford Gun Tunnel. Depending on test requirements, it can operate as a shock tube, reflected shock tunnel or expansion tube. Commissioning tests of the free-piston driver are discussed, including the development of four baseline driver conditions using piston masses of either 36 kg or 89 kg. Experimental data are presented for each operating mode, with comparison made to numerical simulations. In general, high-quality test flows are observed. The calculated enthalpy range of the experimental conditions achieved varies from $$2.7\hbox { MJ kg}^{-1}$$ 2.7 MJ kg - 1 to $$115.0\hbox { MJ kg}^{-1}$$ 115.0 MJ kg - 1 . Graphical abstract

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Chan, W. Y. K., R. W. Whitside, M. K. Smart, D. E. Gildfind, P. A. Jacobs, and T. Sopek. "Nitrogen driver for low-enthalpy testing in free-piston-driven shock tunnels." Shock Waves, May 15, 2021. http://dx.doi.org/10.1007/s00193-021-01002-0.

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36

Deep, Sneh, and Gopalan Jagadeesh. "Gas-Surface Energy Exchange Characterization Around a Cone in the Free-Piston-Driven Shock Tunnel." Journal of Thermophysics and Heat Transfer, February 21, 2021, 1–16. http://dx.doi.org/10.2514/1.t6016.

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37

Kumar, C. S., and K. P. J. Reddy. "Experimental Investigation of Heat Fluxes in the Vicinity of Protuberances on a Flat Plate at Hypersonic Speeds." Journal of Heat Transfer 135, no. 12 (September 27, 2013). http://dx.doi.org/10.1115/1.4024667.

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Heat transfer rates measured in front and to the side of a protrusion on an aluminum flat plate subjected to hypersonic flow at zero angle of attack are presented for two flow enthalpies of approximately 2 MJ/kg and 4.5 MJ/kg. Experiments were conducted in the hypersonic shock tunnel (HST2) and free piston driven HST3 at a freestream Mach number of 8. Heat transfer data was obtained for different geometries of the protrusion of a height of 4 mm, which is approximately the local boundary layer thickness. Comparatively high rates of heat transfer were obtained at regions of flow circulation in the separated region, with the hottest spot generally appearing in front of the protuberance. Experimental values showed moderate agreement with existing empirical correlations at higher enthalpy but not at all for the lower enthalpy condition, although the correlations were coined at enthalpy values nearer to the lower value. Schlieren visualization was also done to investigate the flow structures qualitatively.

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Sander, Tobias, Jens Weber, and Christian Mundt. "Simultaneous thermometry and velocimetry for a shock tunnel using homodyne and heterodyne detection." Applied Physics B 128, no. 8 (July 23, 2022). http://dx.doi.org/10.1007/s00340-022-07850-7.

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AbstractAt our institute a piston-driven shock tunnel is operated to investigate structures of space transportation systems under reentry and propelled flight conditions. For temperature measurements in the nozzle reservoir under single-shot conditions, laser-induced thermal grating spectroscopy is used to date to measure the speed of sound of the test gas. The temperature then can be calculated from this data. The existing experimental setup has already been successfully used to measure flows up to an enthalpy of 2.1 MJ/kg. Since conducting the experiments is extremely time-consuming, it is desirable to extract as much data as possible from the test runs. To additionally measure the velocity of the test gas, the test setup was extended. Besides, extensive improvements have been implemented to increase the signal-to-noise ratio. As the experiments can be conducted much faster at the double-diaphragm shock tube of the institute without any restrictions on the informative value, the development of the heterodyne detection technique is carried out at this test facility. A series of 36 single-shot temperature and velocity measurements is presented for enthalpies of up to 1.0 MJ/kg. The averaged deviation between the measured values and the values calculated from the shock equations of all measurements related to the average of the calculated values is 2.0% for the Mach number, 0.9% for the velocity after the incident shock and 4.8% for the temperature after the incident shock.

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Lynch, K. P., T. Grasser, R. Spillers, C. Downing, K. A. Daniel, E. R. Jans, S. Kearney, B. J. Morreale, R. Wagnild, and J. L. Wagner. "Design and characterization of the Sandia free-piston reflected shock tunnel." Shock Waves, May 23, 2023. http://dx.doi.org/10.1007/s00193-023-01127-4.

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Sudarshan, B., H. A. Pranav, and A. V. Sanjay. "Hypersonic flow study in a pneumatically operated academic shock tunnel." Review of Scientific Instruments 94, no. 5 (May 1, 2023). http://dx.doi.org/10.1063/5.0142147.

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A hypersonic shock tunnel is a primary tool used for basic experimental research and may be used in engineering and university courses to study compressible flows involving shock waves. In the present study, a pneumatically operated shock tunnel is demonstrated for hypersonic flow studies. The high-pressure nitrogen gas is used to drive a pneumatic cylinder, which is used to burst the thin metal diaphragms. Tunnel-free stream conditions are quantified using the measured pressure values and by applying shock tube relations. The free-stream Mach number of 5.5–7.2 is achieved by varying the bursting pressure and test gas pressure from 2.1 to 4.5 bars and 0.2 to 0.5 bar, respectively. The simulation is performed and the shock standoff distance quantified, and the stagnation pressures measured. The results demonstrate that the pneumatically operated tunnel enhanced operation capacity compared to the manually operated tunnel and well suits the academic hypersonic research and developmental activities.

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Tanno, Marie, and Hideyuki Tanno. "Aerodynamic characteristics of a free-flight scramjet vehicle in shock tunnel." Experiments in Fluids 62, no. 7 (July 2021). http://dx.doi.org/10.1007/s00348-021-03229-0.

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Abstract A multi-component aerodynamic test for an airframe-engine integrated scramjet vehicle model was conducted in the free-piston shock tunnel HIEST. A free-flight force measurement technique was applied to the scramjet vehicle model named MoDKI. A new method using multiple piezoelectric accelerometers was developed based on overdetermined system analysis. Its unique features are the following: (1) The accelerometer’s mounting location can be more flexible. (2) The measurement precision is predicted to be improved by increasing the number of accelerometers. (3) The angular acceleration can be obtained with single-axis translational accelerometers instead of gyroscopes. (4) Through the averaging process of the multiple accelerometers, model natural vibration is expected to be mitigated. With eight model-onboard single-axis accelerometers, the three-component aerodynamic coefficients (Drag, Lift, and Pitching moment) of MoDKI were successfully measured at the angle of attack from 0.7 to 3.4 degrees under a Mach 8 free-stream test flow condition. A linear regression fitting revealed a 95% prediction interval as the measurement precision of each aerodynamic coefficient. Graphical abstract

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Selcan, C., T. Sander, and Ch Mundt. "In situ nozzle reservoir thermometry by laser-induced grating spectroscopy in the HELM free-piston reflected shock tunnel." Shock Waves, January 9, 2021. http://dx.doi.org/10.1007/s00193-020-00982-9.

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AbstractExperimental determination of test gas caloric quantities in high-enthalpy ground testing is impeded by excessive pressure and temperature levels as well as minimum test timescales of short-duration facilities. Yet, accurate knowledge of test gas conditions and stagnation enthalpy prior to nozzle expansion is crucial for a valid comparison of experimental data with numerical results. To contribute to a more accurate quantification of nozzle inlet conditions, an experimental study on non-intrusive in situ measurements of the post-reflected shock wave stagnation temperature in a large-scale free-piston reflected shock tunnel is carried out. A series of 20 single-shot temperature measurements by resonant homodyne laser-induced grating spectroscopy (LIGS) is presented for three low-/medium-enthalpy conditions (1.2–2.1 MJ/kg) at stagnation temperatures 1100–1900 K behind the reflected shock wave. Prior limiting factors resulting from impulse facility recoil and restricted optical access to the high-pressure nozzle reservoir are solved, and advancement of the optical set-up is detailed. Measurements in air agree with theoretical calculations to within 1–15%, by trend reflecting greater temperatures than full thermo-chemical equilibrium and lesser temperatures than predicted by ideal gas shock jump relations. For stagnation pressures in the range 9–22 MPa, limited influence due to finite-rate vibrational excitation is conceivable. LIGS is demonstrated to facilitate in situ measurements of stagnation temperature within full-range ground test facilities by superior robustness under high-pressure conditions and to be a useful complement of established optical diagnostics for hypersonic flows.

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