Fabio Pinna

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.

Transition prediction for oblique breakdown in supersonic boundary layers with uncertain disturbance spectrum

Laminar to turbulent transition in supersonic boundary layer is numerically investigated by combining linear stability theory and Uncertainty Quantication. Linear stability theory is used to determine the N factor for the e N transition prediction method for a Mach 6 at plate test case. Transition onset location is determined by using the N factor experimentally obtained in the facility where the test was carried out. Uncertainty quantication is used to compute the probability of transition within the intermittency region downstream of the transition onset. The stochastic approach allows to model the transition region as in the experimental cases since a gradual passage from the laminar to the turbulent ow is obtained. The probability of transition resembles the shape of the skin friction or the heat ux distribution generally observed in the experiments within the transition region. Here we focus on the transition zone rather than on the onset location.

Numerical computations of hypersonic boundary layer roughness induced transition on a at plate

The work is focused on numerical simulations of roughness induced transition for hypersonic flow on a flat plate wall mounted roughness element. Numerical simulations are compared to experimental results in order to resemble the physics highlighted in the tests. In particular, stress has been placed on the detection of the vortices in the wake behind the roughness element and on the onset of transition.

Effect of uneven wall blowing on hypersonic boundary-layer stability and transition

The understanding of aerothermodynamics in complex multiphysics environments such as atmospheric entry is of vital importance for the design of optimized vehicles. A better insight into the interaction between the thermal protection system ablation and the flow’s stability and transition would significantly increase the available payload and reduce the overall mission costs. In order to improve the modeling of the effect of ablation-induced outgassing, a continuously blowing boundary condition for linear stability theory is developed and tested against existing homogeneous and porous conditions in cold wind-tunnel calorically perfect gas conditions. The new model predicts a substantially higher destabilization of the boundary layer than the classic one. An exhaustive parametric study addresses the effect of the most important parameters influencing the boundary condition characteristics. Porous layer stabilization/destabilization maps for different porous layers and flow conditions are built, serving as a starting point for a stabilizing-porous-coating design methodology. Previous controversial literature findings are clarified in the light of the current work, showing a consistent way of addressing similar future tests. The semiempirical eN method is used to predict the location of transition onset in noisy wind-tunnel conditions, obtaining substantially different predictions with different wall models.The understanding of aerothermodynamics in complex multiphysics environments such as atmospheric entry is of vital importance for the design of optimized vehicles. A better insight into the interaction between the thermal protection system ablation and the flow’s stability and transition would significantly increase the available payload and reduce the overall mission costs. In order to improve the modeling of the effect of ablation-induced outgassing, a continuously blowing boundary condition for linear stability theory is developed and tested against existing homogeneous and porous conditions in cold wind-tunnel calorically perfect gas conditions. The new model predicts a substantially higher destabilization of the boundary layer than the classic one. An exhaustive parametric study addresses the effect of the most important parameters influencing the boundary condition characteristics. Porous layer stabilization/destabilization maps for different porous layers and flow conditions are built, serving as a...

DEKAF: spectral multi-regime basic-state solver for boundary layer stability

As the community investigates more complex flows with stronger streamwise variations and uses more physically inclusive stability techniques, such as BiGlobal theory, there is a perceived need for more accuracy in the base flows. To this end, the implication is that using these more advanced techniques, we are now including previously neglected terms of O(Re2). Two corresponding questions follow: (1) how much accuracy can one reasonably achieve from a given set of basic-state equations and (2) how much accuracy does one need to converge more advanced stability techniques? The purpose of this paper is to generate base flow solutions to successively higher levels of accuracy and assess how inaccuracies ultimately affect the stability results. Basic states are obtained from solving the self-similar boundary-layer equations, and stability analyses with LST, which both share O(1/Re) accuracy. This is the first step toward tackling the same problem for more complex basic states and more advanced stability theories. Detailed convergence analyses are performed, allowing to conclude on how numerical inaccuracies from the basic state ultimately propagate into the stability results for different numerical schemes and instability mechanisms at different Mach numbers.

Code to code comparison on hypersonic high enthalpy transitional boundary layers

In the present study three boundary layer stability codes are compared based on hypersonic high enthalpy boundary layer flows around a 7° blunted cone. The code to code comparison is conducted between the following codes: the NOnLocal Transition analysis code (NOLOT) of the German Aerospace Center (DLR), the Stability and Transition Analysis for hypersonic Boundary Layers code (STABL) of University of Minnesota and the VKI Extensible Stability and Transition Analysis code (VESTA) of the von Karman Institute. The comparison focuses on the role of real gas effects on the second mode instability, in particular the disturbance frequency. The experimental test cases for the code to code comparison are provided by the DLR High Enthalpy Shock Tunnel Gottingen (HEG) and the JAXA High Enthalpy Shock tunnel (HIEST).

Statistical inverse analysis and stochastic modeling of transition

A computational method is introduced to infer statistical information on boundarylayer perturbations upstream of laminar-turbulent transition in a supersonic boundary layer. The method uses the intermittency function as a basis, which specifies the amount of time a flow is turbulent at a given streamwise location. The methods yields a joint probability density function for amplitude and frequency of boundary-layer perturbations upstream of transition. It relies on linear stability theory to link the probability density function with the intermittency. In order to infer parameters governing the function, we perform a statistical inverse analysis using the Markov chain Monte Carlo method. The approach is applied to a synthetic test case and to experimental data. The knowledge of the location of laminar to turbulent transition is essential for numerous engineering applications. For a hypersonic vehicle, an accurate prediction of the transition location may allow to precisely define the dimensions of the vehicle’s thermal protection system. Transition does not only affect the performance of an atmospheric re-entry vehicle, it also plays a key role in the safety of the vehicle and its payload. The development of credible engineering models for the prediction of laminar-turbulent transition is therefore an important task. The transition location depends on a large number of parameters and is, consequently, difficult to predict. In particular, perturbations in the boundary layer upstream of transition are known to affect the transition process. Except in studies performed within a carefully controlled laboratory environment, transition is a highly random process due to its strong sensitivity to these perturbations. Instead of trying to determine fixed values of the most important governing parameters, in particular those relating to boundary-layer perturbations, we should characterize them in a statistical way in order to account for the stochastic nature of transition. The laminar-turbulent transition process may be divided into three stages: receptivity, linear disturbance evolution and nonlinear breakdown to turbulence. Most transition modeling approaches are deterministic and rely on empirical data (Langtry & Menter 2009), but they do not capture the stochastic nature of the physical processes active during these stages. Only the nonlinear breakdown stage, including the formation of turbulent spots, has been modeled in a probabilistic way (Vinod & Govindarajan 2004; Pyecnik et al. 2011).

Simulation of induced transition in hypersonic regime: Validation of foot print of the vortical structures

The main goal of this research is to propose a numerical methodology able to predict the flow before the onset of transition occurring in a hypersonic flow. Presently CFD computations are carried out and compared with a huge database of wall experimental results obtained at VKI [4]. The numerical results are expected to provide a 3D definition of the structures letting their footprint at the wall and to support understanding of future hypersonic experiments at VKI. The study focuses on an isolated roughness immersed in the wall boundary layer with a flow at Mach 6 and a Reynolds number based on the ramp height of 20806. The used solver is a VKI in-house code named COOLFluiD [2]. This object oriented framework for fluid dynamic is suited for multiphysics computation and allows the use of different numerical techniques. Presently, the use of perfect gas thermodynamics is adopted. Two different numerical techniques both in a steady and laminar configuration will be compared (Finite Volume Method versus Residual Distribution Method). A critical comparison/validation of the two type of calculations is done in front of the qualitative wall measurements (time averaged Heat flux and oil visualizations) . The tridimensional topology of the flow will be then revealed using vortical surfaces extracted thanks to the Q-criterion but also wall streamlines and streamlines extracted from the symmetry plane. The computations carried on so far with a 2nd order FVM does not allow to match completely the flow features, even if behind the ramp the Q criterion shows the appearance of a couple of streamwise vortices. The solution with other numerical techniques and possibly with a finer mesh will bring more insight in the flow physics. A further step will be achieved by using transition modeling as PSE [1] and bi-global stability analysis [3]. The steady laminar flow, obtained by CFD computations, will be used as mean flow to compute the flow stability downstream of the roughness.