Formulas for the thermodynamic properties of dense nitrogen.

53rd AIAA Aerospace Sciences Meeting

Masutti

Chazot

2015

A summary of the activities performed over the last years at the von Karman Institute for Fluid Dynamics in the frame of hypersonic boundary layer transition studies is presented. Free-stream noise levels have been determined in the H3 Mach 6 conventional wind tunnel using double hot-wires and modal analysis. In the Longshot wind tunnel at Mach 10, an improved free-stream characterization method, based on the use of free-stream static pressure probes, has been applied, alleviating the needs for the limiting adiabatic and isentropic nozzle flow assumptions. Based on these improved flow characterization, natural transition experiments have been performed in both wind tunnels on 7 • half-angle conical geometries at 0 • angle of attack and with different nosetip radii. Measurements techniques include either infrared thermography or flush-mounted fast response thermocouples in order to determine the transition onset location. Boundary layer instabilities are visualized using a LIF-based Schlieren technique at Mach 10, revealing rope-shape structures typical of the second mode disturbances. Wall measurements using fast-response pressure sensors complete the investigations. Dominant boundary layer disturbances at various locations along the cone are determined and compared with theoretical predictions. The corresponding N-factor is inferred for each wind tunnel. A comparison of the different measurement techniques is finally reported. Nomenclature Symbols c Specific heat, J/(kg.K)

Section: Ivb Flow Characterization In the Longshot Wind Tunnelmentioning

confidence: 99%

Flow characterization and boundary layer transition studies in VKI hypersonic facilities (Invited)

53rd AIAA Aerospace Sciences Meeting

Masutti

Chazot

2015

“…The nozzle was designed by the method of characteristics 18 assuming an equilibrium flow and taking into account the dense gas phenomena present in the reservoir using an equation of state described in Ref. 21,22. The inviscid contour obtained was then corrected for the boundary layer displacement thickness based upon the correlation of Edenfield.…”

Section: Iia Wind-tunnelmentioning

confidence: 99%

Design of Static Pressure Probes for Improved Free-Stream Characterization in Hypersonic Wind-Tunnels

52nd Aerospace Sciences Meeting

Paris

Rambaud

et al. 2014

Slender fast-response static pressure probes are tested in the Longshot hypersonic windtunnel with free-stream Mach number larger than 10 and Reynolds number 30000 ReD 60000 based on the probe diameter. They aim at improving the characterization of the freestream. Numerical simulations predict the wall pressure measured to be within 5% of the free-stream static pressure. The viscous effects along the probes are found experimentally to be limited for L /D > 16.5 in agreement with numerical results. Limited influence of the angle of attack of the probes is found for α < 2 • thanks to the probing geometry used. The free-stream quantities at the nozzle exit (Mach, Reynolds...) are derived from the static pressure measurements and compared to those estimated assuming an isentropic nozzle flow expansion. The usefulness of static pressure probes is demonstrated through their ability to detect possible non-isentropic nozzle flow expansions which do not influence traditional measurements such as the test section stagnation pressure. This direct access to free-stream quantities allows to significantly improve the accuracy on the free-stream quantities and benefits to the characterization of hypersonic wind-tunnel flowfields.

“…The nozzle was designed by the method of characteristics 28 assuming an equilibrium flow and taking into account the dense gas phenomena present in the reservoir using an equation of state described in Refs. 29,30. The inviscid contour obtained was then corrected for the boundary layer displacement thickness based upon the correlation of Edenfield.…”

Section: Iia Wind Tunnelmentioning

confidence: 99%

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

7th AIAA Theoretical Fluid Mechanics Conference

Regert

Pinna

et al. 2014

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.