Fluctuating wall pressures measured beneath a supersonic turbulent boundary layer

Beresh, Steven J.; Henfling, John F.; Spillers, Russell Wayne; Pruett, Brian Owen Matthew

doi:10.1063/1.3609271

Cited by 97 publications

(82 citation statements)

References 36 publications

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“…The wall-pressure spectrum for both Mach numbers varies rather weakly with frequency as ω → 0. The absence of the more rapid and incompressible ω 2 scaling at low frequencies in the wall-pressure spectrum is consistent with the measurements by Beresh et al 19 and the DNS by Bernardini & Pirozzoli 42 at supersonic Mach numbers. At high frequencies, the wall spectrum for the Mach 6 case has significantly higher energy than the Mach 2.5 case, and the spectrum for both cases exhibits a slightly more rapid decay than the ω −5 scaling predicted theoretically by Blake.…”

Section: Frequency Spectrasupporting

confidence: 80%

“…[17][18][19] As pointed out by several authors, 19,20 there are few (if any) reliable measurements of the variance of the wall pressure fluctuations and its frequency spectra at high speeds, due to either a poor spatial resolution of the pressure transducers and/or the limited frequency response of these sensors. Laufer's measurements of the acoustic fluctuations in the freestream region 1 are also subject to analogous sources of experimental error.…”

Section: Introductionmentioning

confidence: 99%

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Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

Duan

Choudhari

2013

51st AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

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Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by a Mach 6 turbulent boundary layer with nominal freestream Mach number of 6 and Reynolds number of Reτ ≈ 464. The emphasis is on comparing the primarily vortical pressure signal at the wall with the acoustic freestream signal under higher Mach number conditions. Moreover, the Mach-number dependence of pressure signals is demonstrated by comparing the current results with those of a supersonic boundary layer at Mach 2.5 and Reτ ≈ 510. It is found that the freestream pressure intensity exhibits a strong Mach number dependence, irrespective of whether it is normalized by the mean wall shear stress or by the mean pressure, with the normalized fluctuation amplitude being significantly larger for the Mach 6 case. Spectral analysis shows that both the wall and freestream pressure fluctuations of the Mach 6 boundary layer have enhanced energy content at high frequencies, with the peak of the premultiplied frequency spectrum of freestream pressure fluctuations being at a frequency of ωδ/U∞ ≈ 3.1, which is more than twice the corresponding frequency in the Mach 2.5 case. The space-time correlations indicate that the pressure-carrying eddies for the higher Mach number case are of smaller size, less elongated in the spanwise direction, and convect with higher convection speeds relative to the Mach 2.5 case. The demonstrated Mach-number dependence of the pressure field, including radiation intensity, directionality, and convection speed, is consistent with the trend exhibited in experimental data and can be qualitatively explained by the notion of 'eddy Mach wave' radiation. NomenclatureC p heat capacity at constant pressure, J/(K·kg) C pp Space-time correlation coefficient of the pressure field, dimensionless C v heat capacity at constant volume, J/(K·kg)

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Section: Frequency Spectrasupporting

confidence: 80%

Section: Introductionmentioning

confidence: 99%

Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

Duan

Choudhari

2013

51st AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

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“…The general character of the surface pressure spectra is similar to that of low-speed boundary layers, which have been studied extensively over the years (see, for instance, the review by Bull 8 and the monograph by Blake 29 ). In particular, the spectrum compares well with the experimental measurements in a low-speed boundary layer by Farabee & Casarella 30 and in a Mach 2 supersonic boundary layer by Beresh et al, 11 as well as with the DNS results by Bernardini & Pirozzoli.…”

Section: B Frequency Spectrasupporting

confidence: 74%

“…The recent data acquired by Beresh et al showed similar large scatter when compared with a broad compilation of high-speed measurements. As pointed out by several authors, 11,12 there are few (if any) reliable measurements of the variance of the wall pressure fluctuations and its frequency spectra, due to the poor spatial resolution of pressure transducers or limitations in the frequency response of pressure sensors. As far as the acoustic fluctuations in the freestream region are concerned, the only detailed measurements have been reported by Laufer 2 and these measurements are subject to similar sources of experimental error as the wall-pressure measurements.…”

Section: Introductionmentioning

confidence: 99%

Numerical Study of Pressure Fluctuations Due to High-Speed Turbulent Boundary Layers

Duan

Choudhari

2012

42nd AIAA Fluid Dynamics Conference and Exhibit

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Direct numerical simulations (DNS) are used to examine the pressure fluctuations generated by fully developed turbulence in supersonic turbulent boundary layers with an emphasis on both pressure fluctuations at the wall and the acoustic fluctuations radiated into the freestream. The wall and freestream pressure fields are first analyzed for a zero pressure gradient boundary layer with Mach 2.5 and Reynolds number based on momentum thickness of approximately 2835. The single and multi-point statistics reported include the wall pressure fluctuation intensities, frequency spectra, space-time correlations, and convection velocities. Single and multi-point statistics of surface pressure fluctuations show good agreement with measured data and previously published simulations of turbulent boundary layers under similar flow conditions. Spectral analysis shows that the acoustic fluctuations outside the boundary layer region have much lower energy content within the high-frequency region. The space-time correlations reflect the convective nature of the pressure field both at the wall and in the freestream, which is characterized by the downstream propagation of pressure-carrying eddies. Relative to those at the wall, the pressure-carrying eddies associated with the freestream signal are larger and convect at a significantly lower speed. The preliminary DNS results of a Mach 6 boundary layer show that the pressure rms in the freestream region is significantly higher than that of the lower Mach number case.

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“…These sensors typically have flat frequency response up to about 30 to 40% of their roughly 270 to 285 kHz resonant frequency. 87 For the current experiments, five instrumentation holes were available in the model. Table 4 shows the locations of the instrumentation holes.…”

Section: Figure 79 Hifire-5 Wind Tunnel Modelmentioning

confidence: 99%

HIFiRE-1 and HIFiRE-5 Test Results

Kimmel¹,

Adamczak²,

Borg³

et al. 2014

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The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Department of Defense, Washington Headquarters Services, Directorate for Information AFRL/RQHF SPONSORING/MONITORING AGENCY REPORT NUMBER(S) AFRL-RQ-WP-TR-2014-0038 DISTRIBUTION/AVAILABILITY STATEMENTApproved for public release; distribution unlimited SUPPLEMENTARY NOTESPA Case Number: 88ABW-2014-3421; Clearance Date: 17 Jul 2014. ABSTRACTThe primary experiment for HIFiRE-1 was to measure boundary-layer transition in hypersonic flight on a 7 degree halfangle, axisymmetric cone with a 2.5 mm radius bluntness. A secondary experiment measured mean and fluctuating pressure, and heat transfer on a cylinder-33-degreee-flare geometry. The ascent provided smooth-body, low angle-ofattack, boundary-layer transition at Mach numbers greater than 5. The end of turbulent-to-laminar transition occurred at Reynolds numbers between 10.3 to 12.2 million. Second-mode N-factors of approximately 14 correlated transition. A diamond-shaped trip retained turbulent flow until reaching a roughness Reynolds number Rekk = 2200. During re-entry, transition was measured around the circumference of the cone at Reynolds numbers ranging from 3 million on the leeside to 5 million on the windward side. In the shock-boundary layer interaction, maximum sound pressure levels of 173 dB were recorded upstream of the flare, and 185 dB on the flare itself. The principal goal of HIFiRE-5 was to measure hypersonic boundary layer transition on an elliptic cone. The second stage booster failed to ignite, so the experiment reached a maximum Mach number of only 3. Nevertheless, supersonic pressure and temperature data were obtained under laminar and turbulent flow, and flight systems were validated. Traveling and stationary crossflow instabilities were clearly measured for HIFiRE-5 in ground test in a quiet flow tunnel. The frequency, phase speed, and wave angle were all in good agreement with computations.

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Fluctuating wall pressures measured beneath a supersonic turbulent boundary layer

Cited by 97 publications

References 36 publications

Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

Numerical Study of Pressure Fluctuations due to a Mach 6 Turbulent Boundary Layer

Numerical Study of Pressure Fluctuations Due to High-Speed Turbulent Boundary Layers

HIFiRE-1 and HIFiRE-5 Test Results

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