Guillaume Grossir

Experimental Characterization of Hypersonic Nozzle Boundary Layers and Free-Stream Noise Levels

Pitot pressure surveys of the Longshot hypersonic contoured nozzle in the von Karman Institute are presented. Measurements use a rake of pitot probes at four locations along the nozzle axis from half the nozzle length down to the exit. Laminar and turbulent numerical simulations are used for comparison. A good agreement is obtained with the turbulent computations conrming the existence of turbulent boundary layers along the nozzle walls. The uniformity of the free-stream parallel ow is characterized by a standard deviation of 3.5%. The free-stream noise levels determined from pitot pressure uctuations are about 7.5% and slightly lower inside the nozzle. These are comparable with the levels recorded in other similar hypersonic wind tunnels. Some ow divergences are shown to be possibly correlated to a discontinuity in the contour of the nozzle. The test time available at various locations is also determined and is shown to depend upon the location considered along and across the nozzle.

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

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 /m ≥ Reunit,∞ ≥ 3.3× 10 /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 ReRN,∞ = 123, 000. Increasing freestream unit Reynolds number results in larger RexB,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.

Design of static pressure probes for improved free-stream characterization in hypersonic wind tunnels

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.

Nosetip bluntness effects on transition at hypersonic speeds: experimental and numerical analysis under NATO STO AVT-240

The existing database of transition measurements in hypersonic ground facilities has established that the onset of boundary layer transition over a circular cone at zero angle of attack shifts downstream as the nosetip bluntness is increased with respect to a sharp cone. However, this trend is reversed at sufficiently large values of the nosetip Reynolds number, so that the transition onset location eventually moves upstream with a further increase in nosetip bluntness. This transition reversal phenomenon, which cannot be explained on the basis of linear stability theory, was the focus of a collaborative investigation under the NATO STO group AVT-240 on Hypersonic Boundary-Layer Transition Prediction. The current paper provides an overview of that effort, which included wind tunnel measurements in three different facilities and theoretical analysis related to modal and nonmodal amplification of boundary layer disturbances. Because neither first and secondmode waves nor entropy-layer instabilities are found to be substantially amplified to initiate transition at large bluntness values, transient (i.e., nonmodal) disturbance growth has been investigated as the potential basis for a physics-based model for the transition reversal phenomenon. Results of the transient growth analysis indicate that disturbances that are initiated within the nosetip or in the vicinity of the juncture between the nosetip and the frustum can undergo relatively significant nonmodal amplification and that the maximum energy gain increases nonlinearly with the nose radius of the cone. This finding does not provide a definitive link between transient growth and the onset of transition, but it is qualitatively consistent with the experimental observations that frustum transition during the reversal regime was highly sensitive to wall roughness, and furthermore, was dominated by disturbances that originated near the nosetip.

Flow characterization and boundary layer transition studies in {VKI} hypersonic facilities

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.

Detection of nitrogen flow condensation onset in a hypersonic wind tunnel using a static pressure probe

Measurements with a slender static-pressure probe in the free-stream of the Longshot hypersonic wind-tunnel have recently been performed. They have revealed pressures larger than the theoretical values obtained with the assumption of an isentropic nozzle flow. The presence of flow condensation during the nozzle expansion is now investigated as a possible source of non-isentropicity to explain the free-stream static pressure mismatch. Different stagnation temperatures are investigated which either delay or promote flow nucleation. Standard operating conditions of the Longshot wind-tunnel are demonstrated to be free of condensation. Experiments performed with lower stagnation temperatures have successfully promoted the condensation of nitrogen which could be detected by the static pressure probe. A weak amount of flow supersaturation has been achieved in agreement with heterogeneous nucleation theory. The accurate performances of static pressure probes and their usefulness for the characterization of hypersonic flows are demonstrated.

Compilation and Analysis of Second Mode Amplitudes on Sharp Cones in Hypersonic Wind Tunnels

This research effort coordinated by the NATO AVT-240 specialists’ group compiles and analyzes second-mode amplitudes on sharp slender cones at 0 degrees angle of attack. The analysis focuses on pressure fluctuations measured with piezoelectric sensors in 11 hypersonic wind tunnels operated by 9 organizations located in 3 NATO countries (Belgium, Germany, and USA) and Japan. The measurements are at freestream Mach numbers between 5 and 14, unit Reynolds numbers Re/m between 1.5 and 16 million per meter, and wall-to total temperature ratios between 0.1 and 0.8. The study shows that second-mode growth rates can be predicted with Parabolized Stability Equations (PSE) over the wide range of conditions. The maximum second-mode amplitudes vary weakly at edge Mach number Me greater than ~5.8, but significantly decrease at lower Me. The maximum N factor envelope from PSE and the measured amplitudes are used to estimate the initial amplitudes A0. At each Mach number, A0 varies approximately as Re/m^(-1). This leads to transition N factors that increase with Re/m. This behavior is consistent with the results from Marineau (AIAA Journal, Vol. 55, No. 2, 2017).

Axisymmetric Simulation of Turbulent Hypersonic flows in contoured nozzle

Reynolds-averaged Navier-Stokes (RANS) turbulence models have been widely used for simulating hypersonic turbulent boundary layers, without any specific considerations regarding the thermochemical nonequilibrium effects. In most hypersonic solvers the turbulent models are either solved uncoupled with the governing equations of chemical reacting flows or simply used in hypersonic non-reacting flows. Development of RANS models for thermochemical nonequilibrium flows remains one of the most challenging task, because of the complex phenomena governing the coupling between turbulence and chemistry. This process is not well understood and is still under investigations. Moreover, the available RANS turbulence models were developed for incompressible flows and the modeling of compressible turbulent flows is still an active area of research. Despite all of these considerations, many efforts have been made in order to assess these models for hypersonic applications and to develop a methodology for validating these models with hypersonic ground testing experimental data. An effort has also been made, in order to provide a central and a common data base for different version of RANS turbulence models. A Turbulence Model Benchmarking Working Group (TMBWG) was created. The group provide through this data base accurate and up-to-date information on widely-used RANS turbulence models. Moreover a naming convention were established in order to help avoid confusion between models when comparing results with other codes. Testcases are also provided for code to code verification. In the current work we focus on two-equations turbulence model of Menter, well known as Menter k−ω SST model, this model is used to simulate hypersonic turbulent boundary layer in a contoured nozzle of the VKI-Lomgshot facility. The reference for the standard implementation of this model is referred by TMBWG as SST. The axisymmetric formulation of the model was derived and implemented in an unstructured parallel finite volume solver COOLFLuiD which is a VKI in-house solver. Pitot pressure measurement were performed in the VKI-contoured nozzle at four locations along the symmetry axis and also along the radial direction from the symmetry axis to the supersonic part of the boundary layer. These experimental data are used in the current work as a preliminary validation of our implementation. The effect of the turbulent boundary layer in the free stream conditions i.e the nozzle exit conditions is also investigated.