Pressure Fluctuation Measurements in the NASA Langley 20-Inch Mach 6 Wind Tunnel

Rufer, Shann J.; Berridge, Dennis C.

doi:10.2514/6.2012-3262

Cited by 21 publications

(8 citation statements)

References 14 publications

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“…Finally, figure 14 relates the RMS pitot pressure fluctuations obtained in HEG and RWG to data obtained at comparable test conditions in the AEDC Tunnel 9 (Lafferty et al 70 ), the VKI H3 (Masutti et al 22 ), the BAM6QT at noisy conditions (Steen 71 ) and the NASA Langley 20-Inch Mach 6 Wind Tunnel (Rufer et al 26 ). It can be seen that the RWG Mach 6 data, ranging from 1.7 % to 2.6 %, show a good agreement with the RMS pitot fluctuations obtained in the VKI H3 facility at Mach 6.…”

Section: Resultsmentioning

confidence: 99%

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Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation

et al. 2018

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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. In combination with the experimental studies, direct numerical simulations of the leading-edge receptivity to fast and slow acoustic waves were performed for the slender 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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Section: Resultsmentioning

confidence: 99%

“…The RMS pitot pressure fluctuations measured in HEG scatter around 4 % and do not show a clear trend with increasing Reynolds number. 26 .…”

Section: Resultsmentioning

confidence: 99%

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation

et al. 2018

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“…Previous measurements in Ref. 12 indicate that the acoustic noise in the span of the tunnel should be fairly uniform. Figure 10 shows a discrepancy between the PCB sensor located on the centerline of the tunnel and the PCB sensor located at a location of 90.4 mm from the centerline.…”

Section: A Pressure Fluctuationsmentioning

confidence: 89%

“…13 Previous measurements of the acoustic noise in the facility by Rufer and Berridge show that the normalized noise level decreased with increasing Reynolds number and was approximately 1% of the pitot pressure at a freestream unit Reynolds number of 23 × 10 6 /m. 12 For the experiment presented in this paper, the total temperature was held relatively constantly at 493.4±2.3 K and the total pressure was varied to achieve nominal freestream unit Reynolds numbers ranging from 3.28 × 10 6 /m to 26.5 × 10 6 /m.…”

Section: A the 20-inch Mach 6 Tunnelmentioning

confidence: 99%

“…Previous measurements have been made to characterize the freestream fluctuations in the NASA Langley Research Center 20-Inch Mach-6 Tunnel. 2,4,[10][11][12] These measurements of freestream fluctuations are useful as a reference, but as changes to the facility flow path occur, these measurements should be repeated to ensure that the flow quality has not changed. Frequent calibration and characterization of the flow quality in hypersonic facilities may not be feasible due to cost and time constraints.…”

Section: Introductionmentioning

confidence: 99%

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Measurements of Freestream Fluctuations in the NASA Langley 20-Inch Mach 6 Tunnel

Chou

Leidy

Bathel

et al. 2018

2018 Fluid Dynamics Conference

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An experimental campaign was conducted to measure and to characterize the freestream disturbance levels in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel. A pitot rake was instrumented with fast pressure transducers, hot wires, and an atomic layer thermopile to quantify the fluctuation levels of pressure, mass flux, and heat flux, respectively. In conjunction with these probe-based measurements, focused laser differential interferometry was used to optically measure density fluctuations. Measurements were made at five nominal different unit Reynolds numbers ranging from 3.28-26.5×10 6 /m. The rake was positioned at two different streamwise locations and several different roll angles to measure flow uniformity within the test section. In general, noise levels were spatially consistent within the tested region. Pitot pressure fluctuation levels ranged from 0.84% at the highest Reynolds number tested to 1.89% at the lowest Reynolds number tested. Freestream mass-flux fluctuations remained relatively constant between 1.8-2.5% of the freestream. The pressure transducers were also used to determine the dominant disturbance speed and angle of propagation. The disturbances were estimated to travel at approximately 54-81% of the freestream speed at an angle of approximately 21-44°from the freestream direction, but these measurements had a significant amount of uncertainty. A comparison to previous measurements of pressure made in 2012 and of mass flux made in 1994 show almost no change in the RMS fluctuation of these flow quantities.

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Free-stream density perturbations in a reflected-shock tunnel

2014

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Pressure Fluctuation Measurements in the NASA Langley 20-Inch Mach 6 Wind Tunnel

Cited by 21 publications

References 14 publications

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation

Measurements of Freestream Fluctuations in the NASA Langley 20-Inch Mach 6 Tunnel

Free-stream density perturbations in a reflected-shock tunnel

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