Free stream noise and transition measurements in a Mach 3.5 pilot quiet tunnel

Beckwith, I. E.; Creel, T. R.; Chen, F.-J.; Kendall, J. M.

doi:10.2514/6.1983-42

Cited by 70 publications

(50 citation statements)

References 6 publications

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“…Malik (1984) investigated a supersonic boundary layer on a sharp cone (5 by Beckwith et al (1983) for Mach 3.5 (before shock). Malik (1984) showed that in these low free-stream disturbance experiments, transition could be predicted by the e N method with N ranging from 9 to 11.…”

Section: Flow Past a Circular Sharp Cone At Zero Angle Of Attackmentioning

confidence: 99%

Numerical Investigation of Laminar-Turbulent Transition in a Cone Boundary Layer at Mach 6

Sivasubramanian

2011

41st AIAA Fluid Dynamics Conference and Exhibit

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Section: Flow Past a Circular Sharp Cone At Zero Angle Of Attackmentioning

confidence: 99%

Numerical Investigation of Laminar-Turbulent Transition in a Cone Boundary Layer at Mach 6

Sivasubramanian

2011

41st AIAA Fluid Dynamics Conference and Exhibit

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“…By 80.5 cm from the throat, the RMS mass-flux reaches the 0.1% threshold that is commonly used to demarcate the end of the quiet-flow region. 11 With that definition, the quiet test core length for the present flow conditions is ∆x = 35.5 cm and the Reynolds number based on that length is Re ∆x = 3.48 × 10 6 . Beyond 80.5 cm, the RMS mass-flux increases rapidly due to radiated noise from the transitional nozzle wall boundary layer.…”

Section: Iiib Freestream Measurementsmentioning

confidence: 99%

“…Since boundarylayer transition is particularly sensitive to the disturbance environment, testing in a low-noise facility, such as the SLDT, is necessary to simulate transition in the low disturbance environment encountered during high-altitude flight. Details on the facility operation and performance characteristics are provided in Beckwith et al 11 and Chen, 12 so here we provide a brief overview of the facility.…”

Section: Introductionmentioning

confidence: 99%

High-Speed Boundary-Layer Transition Induced by an Isolated Roughness Element

Kegerise

Owens

King

2010

40th Fluid Dynamics Conference and Exhibit

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Progress on an experimental effort to quantify the instability mechanisms associated with roughness-induced transition in a high-speed boundary layer is reported in this paper. To simulate the low-disturbance environment encountered during high-altitude flight, the experimental study was performed in the NASA-Langley Mach 3.5 Supersonic LowDisturbance Tunnel. A flat plate trip sizing study was performed first to identify the roughness height required to force transition. That study, which included transition onset measurements under both quiet and noisy freestream conditions, confirmed the sensitivity of roughness-induced transition to freestream disturbance levels. Surveys of the laminar boundary layer on a 7• half-angle sharp-tipped cone were performed via hot-wire anemometry and pitot-pressure measurements. The measured mean mass-flux and Mach-number profiles agreed very well with computed mean-flow profiles. Finally, surveys of the boundary layer developing downstream of an isolated roughness element on the cone were performed. The measurements revealed an instability in the far wake of the roughness element that grows exponentially and has peak frequencies in the 150 to 250 kHz range.

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“…By 80.5 cm from the throat, the RMS mass-flux reaches the 0.1% threshold that is commonly used to demarcate the end of the quiet-flow region. 8 With that definition, the quiet test core length for the present flow conditions is ∆x = 35.5 cm and the Reynolds number based on that length is Re ∆x = 3.48 x 10 6 . Beyond 80.5 cm, the RMS mass-flux increases rapidly due to radiated noise from the transitional nozzle wall boundary layer.…”

Section: A Facilitymentioning

confidence: 99%

Off-Body Boundary-Layer Measurement Techniques Development for Supersonic Low-Disturbance Flows

Owens

Kegerise

Wilkinson

2011

49th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

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Investigations were performed to develop accurate boundary-layer measurement techniques in a Mach 3.5 laminar boundary layer on a 7° half-angle cone at 0° angle of attack. A discussion of the measurement challenges is presented as well as how each was addressed. A computational study was performed to minimize the probe aerodynamic interference effects resulting in improved pitot and hot-wire probe designs. Probe calibration and positioning processes were also developed with the goal of reducing the measurement uncertainties from 10% levels to less than 5% levels. Efforts were made to define the experimental boundary conditions for the cone flow so comparisons could be made with a set of companion computational simulations. The development status of the mean and dynamic boundary-layer flow measurements for a nominally sharp cone in a lowdisturbance supersonic flow is presented. = hot-wire temperature (K) t rad = radius of wedge probe body tip (mm) w = width of wedge probe body tip (mm) x = axial distance referenced to the cone apex (mm) ∆x = length of quiet test core (cm) y = wall-normal distance (mm) z = wall normal distance as defined for the probe redesign CFD only (mm) δ = boundary-layer thickness (mm) θ = cone azimuthal angle (deg) τ = hot-wire overheat ratio η = hot-wire recovery factor over bar = mean value of the quantity Nomenclature

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Free stream noise and transition measurements in a Mach 3.5 pilot quiet tunnel

Cited by 70 publications

References 6 publications

Numerical Investigation of Laminar-Turbulent Transition in a Cone Boundary Layer at Mach 6

Numerical Investigation of Laminar-Turbulent Transition in a Cone Boundary Layer at Mach 6

High-Speed Boundary-Layer Transition Induced by an Isolated Roughness Element

Off-Body Boundary-Layer Measurement Techniques Development for Supersonic Low-Disturbance Flows

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