Hypersonic Wind-Tunnel Measurements of Boundary-Layer Pressure Fluctuations

Casper, Katya M.

doi:10.2514/6.2009-4054

Cited by 107 publications

(66 citation statements)

References 36 publications

(46 reference statements)

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“…Pressure fluctuations measured beneath turbulent boundary layers on cones range from 1 to 3 % in conventional wind tunnels. 39 Similar ground measurements under quiet conditions are not available. The rise in pressure fluctuations near 20 seconds is typical of measurements of transitional boundary layers made in wind tunnels.…”

Section: Insight From High-bandwidth Instrumentationmentioning

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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Section: Insight From High-bandwidth Instrumentationmentioning

confidence: 99%

HIFiRE-1 and HIFiRE-5 Test Results

Kimmel¹,

Adamczak²,

Borg³

et al. 2014

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“…The two axial locations annotated in the figure at x = 208 and 490 mm, correspond to those where measured wall-pressure fluctuation data are available for comparison. 16 At these locations the maximum N-factors are 3.5 and 7.0 respectively, and the corresponding frequencies are 350 and 220 kHz.…”

Section: Linear Stabilitymentioning

confidence: 99%

“…This model geometry was tested in the Boeing/AFOSR Mach 6 Quiet Tunnel at the same freestream conditions. [15][16] The second model is a circular flared cone with a spherical blunt nose region (Fig. 3).…”

Section: Models and Flow Conditionsmentioning

confidence: 99%

“…Simulations were performed on three axisymmetric geometries for which experimentally measured instability data are available for comparison: 1) a straight cone, 2) a circular flared cone with the flare starting from the nose, and 3) a circular flared cone with an initial straight nose region tangent to a flare downstream. The first two geometries were tested in the Boeing/AFOSR Mach 6 Quiet Tunnel at Purdue University [15][16] and the third geometry was tested in the NASA Langley Mach 6 Quiet Tunnel. [17][18] In each case, we first solved for the mean flow field and then performed linear local and non-local stability analyses for the computed boundary-layer profiles.…”

Section: Introductionmentioning

confidence: 99%

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Receptivity of Hypersonic Boundary Layers over Straight and Flared Cones

Balakumar

Kegerise

2015

AIAA Journal

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The effects of adverse pressure gradients on the receptivity and stability of hypersonic boundary layers were numerically investigated. Simulations were performed for boundary layer flows over a straight cone and two flared cones. The steady and the unsteady flow fields were obtained by solving the twodimensional Navier-Stokes equations in axi-symmetric coordinates using the 5 th -order accurate weighted essentially non-oscillatory (WENO) scheme for space discretization and using third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. The mean boundary layer profiles were analyzed using local stability and non-local parabolized stability equations (PSE) methods. After the most amplified disturbances were identified, two-dimensional plane acoustic waves were introduced at the outer boundary of the computational domain and time accurate simulations were performed. The adverse pressure gradient was found to affect the boundary layer stability in two important ways. Firstly, the frequency of the most amplified second-mode disturbance was increased relative to the zero pressure gradient case. Secondly, the amplification of first-and second-mode disturbances was increased. Although an adverse pressure gradient enhances instability wave growth rates, small nose-tip bluntness was found to delay transition due to the low receptivity coefficient and the resulting weak initial amplitude of the instability waves. The computed and measured amplitudefrequency spectrums in all three cases agree very well in terms of frequency and the shape except for the amplitude.

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“…This has been achieved thanks to the improved bandwidth of the measurements techniques (several hundreds of kilohertz are typically required), using hot-wires, [9][10][11] pressure sensors, [12][13][14][15][16] heat flux sensors 17,18 or optical diagnostics. 19,20 These investigations at hypersonic Mach numbers between 6 and 8, have shown the dominance of the so-called second mode of instabilities on conical geometries.…”

Section: Nomenclaturementioning

confidence: 99%

Hypersonic Boundary Layer Transition on a 7 Degree Half-Angle Cone at Mach 10

Grossir

Regert

Pinna

et al. 2014

7th AIAA Theoretical Fluid Mechanics Conference

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Hypersonic boundary layer transition experiments are performed in the low-enthalpy Longshot wind tunnel with a free-stream Mach number ranging between 12 ≥ M∞ ≥ 9.5 and Reynolds number between 12 × 10 6 /m ≥ Reunit,∞ ≥ 3.3 × 10 6 /m. The model is an 800 mm long 7• half-angle cone with nosetip radii between 0.2 and 10 mm. Instrumentation includes flushmounted fast-response thermocouples and pressure sensors. Boundary layer transition onset location is determined from the wall heat flux distribution. Nose bluntness has a strong stabilizing effect. No transition reversal could be observed at RB = 10RN for a Reynolds number based on the nosetip radius of ReR N ,∞ = 123, 000. Increasing freestream unit Reynolds number results in larger Rex B ,e. Wavelet analysis of the boundary layer fluctuations shows that numerous wave packets are present during the transition process. Comparison with Linear Stability Theory results for second mode waves shows an excellent agreement for the most amplified frequencies. The N -factor of the wind-tunnel is 5 based on these computations and on the transition location measured experimentally. The convection velocity of the disturbances is closely approximated by the local boundary layer edge velocity for all conditions investigated. Schlieren flow visualization of the instabilities exhibits the typical rope shape of second mode disturbances for the sharpest nosetips. For nose bluntness larger than 4.75 mm, disturbances are mainly present at the edge of the boundary layer and within the inviscid shock layer. Their shape no longer presents the second mode typical structure although a frequency analysis of the disturbances is still compatible with second mode instabilities. Present results confirm the dominance of second mode waves in the transition process along a conical geometry for Mach numbers larger than 10.

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Hypersonic Wind-Tunnel Measurements of Boundary-Layer Pressure Fluctuations

Cited by 107 publications

References 36 publications

HIFiRE-1 and HIFiRE-5 Test Results

HIFiRE-1 and HIFiRE-5 Test Results

Receptivity of Hypersonic Boundary Layers over Straight and Flared Cones

Hypersonic Boundary Layer Transition on a 7 Degree Half-Angle Cone at Mach 10

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