Fluctuating Disturbances in a Mach 5 Wind Tunnel

Anders, J. B.; Stainback, P. C.; Keefe, Laurence; Beckwith, I. E.

doi:10.2514/3.60763

Cited by 28 publications

(9 citation statements)

References 14 publications

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“…The experiments did not reveal any major difficulties. Heating the air by 30% did make the nozzle throat look cold, reducing the quiet-flow Reynolds number by about 15%, which appeared generally consistent with the earlier measurements of Demetriades [113], Anders et al [53], and Harvey et al [142].…”

Section: Four-inch Mach-4 Quiet Ludwieg Tubesupporting

confidence: 87%

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Development of Hypersonic Quiet Tunnels

Schneider

2008

Journal of Spacecraft and Rockets

218

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Since then, he has focused on high-speed boundary-layer transition, developing facilities and instrumentation for detailed measurements under quiet-flow conditions. A $1 million 9.5-in. Mach-6 quiet-flow Ludwieg tube was completed in 2001 and achieved high Reynolds number quiet flow in 2006. He was promoted to professor, School of Aeronautics and Astronautics, in 2004, and has written seven review articles on hypersonic transition. He is an Associate Fellow of AIAA.

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Section: Four-inch Mach-4 Quiet Ludwieg Tubesupporting

confidence: 87%

“…An axisymmetric Mach-5 nozzle was used in several early quiettunnel development studies [53]. Various modifications to the settling chamber and control valve included the installation of screens, steel wool, and Rigimesh® porous plates.…”

Section: Early Developmentsmentioning

confidence: 99%

Development of Hypersonic Quiet Tunnels

Schneider

2008

Journal of Spacecraft and Rockets

218

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“…Research found disturbances from the settling chamber upstream would travel through the nozzle and enter into the test section. The disturbances would increase the Reynold's number downstream and a turbulent boundary layer along the nozzle wall would develop sooner which would generate more noise (8). In theory, the suction port had great potential, however, in physical application, it was hindered by nonuniform suction as well as surface roughness.…”

Section: Suction Portmentioning

confidence: 99%

Supersonic Windtunnel Noise Attenuation

Wong¹

2021

Preprint

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The aerodynamic generated noise in the supersonic wind tunnel during operation at Ryerson University has exceeded the threshold of hearing damage. An acoustic silencer was to be designed and added to the wind tunnel to reduce the noise level. The main sources of noise generated from the wind tunnel with the silencer were identified to be located at the convergent divergent nozzle and the turbulent region downstream of the shock wave at the diffuser with the maximum acoustic power level of the entire wind tunnel at 161.09 dB. The designed silencer provided an overall sound pressure level reduction of 21.41 db which was considered as acceptable. Refinement to the mesh size and changes to the geometry of the mixing chamber was suggested for a more accurate result in noise output as well as flow conditions would match up to the physical flow. Additional acoustic treatment should be applied to the wind tunnel to further reduce sound pressure level since the noise level still exceeded the threshold of hearing loss.

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“…Previous work has concentrated on ways of maintaining laminar, nozzle-wall boundary layers, 8 or on ways of physically shielding the test model from the incoming noise. 7 An entirely new technique is presented herein for producing a relatively low-noise, supersonic test region at high unit Reynolds numbers where the nozzle-wall boundary layer is fully turbulent and where no physical shielding device is required in the flowfield.…”

Section: Contentsmentioning

confidence: 99%

New Technique for Reducing Test Section Noise in Supersonic Wind Tunnels

1980

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Nomenclature / = distance along nozzle centerline from upstream tip of test rhombus m -tunnel mass flow Ma, = freestream Mach number M w = inviscid Mach number at nozzle wall Pn = rms value of freestream static pressure Pn = mean value of freestream static pressure #00 = freestream unit Reynolds number r* = nozzle throat radius Ax = length of low-noise uniform flow test region y e = nozzle exit radius AbstractT HE major problem in the design of a low-noise supersonic wind tunnel is to reduce the noise radiated from the turbulent, nozzle-wall boundary layer. This broadband noise is generally the dominant disturbance mode in wind tunnels at Mach numbers greater than about 2.5 (Refs 1 and 2). These acoustic disturbances are propagated along Mach lines in supersonic flow; hence, their sources are located far upstream of the test section at so-called "acoustic origins" near the nozzle wall. It is shown that tailoring the nozzle contour to reduce the Mach number and boundary-layer thickness at the acoustic origins significantly reduces the noise levels in the upstream part of the test rhombus.

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Fluctuating Disturbances in a Mach 5 Wind Tunnel

Cited by 28 publications

References 14 publications

Development of Hypersonic Quiet Tunnels

Development of Hypersonic Quiet Tunnels

Supersonic Windtunnel Noise Attenuation

New Technique for Reducing Test Section Noise in Supersonic Wind Tunnels

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