Bruno Chanetz

Experimental and Numerical Study of the Laminar Separation in Hypersonic Flow

Abstract This article is devoted to an experimental and numerical study of shock wave/boundary layer interaction in hypersonic laminar flow (M = 10). The experimental was performed in the ONERA R5Ch wind tunnel on a hollow cylinder flare. The flow stream delivered by the R5Ch wind tunnel produces physical conditions which justify both the theoretical approach using classical Navier-Stokes equations and the approach by Direct Simulation Monte-Carlo. So the aim of this study is to improve the capacity of Navier-Stokes and DSMC codes to predict high Mach number interactions. Pressure and heat flux have been measured at the wall and compared with the results obtained with Navier-Stokes and DSMC codes. Two different meshes have been considered with the three Navier-Stokes codes used. The conclusion is that the codes can give a good evaluation of the physical quantities at the wall.

Shock Wave/Transitional Boundary-Layer Interactions in Hypersonic Flow

The experimental and numerical transitional interactions in hypersonic flow are studied. The experiments were performed on a hollow cylinder-flare model in the ONERA R2Ch wind tunnel at a Mach number of 5 and for varying stagnation pressure. Wall pressure and heat-flux measurements, laser Doppler velocimetry, pitot boundary-layer surveys, surface flow visualizations, and schlieren photographs provide a precise and complete description of the flowfield. In all of the cases examined here, grid-converged axisymmetric mathematical solutions of the problem were obtained by use of the two-dimensional numerical simulation, but it was found that these solutions do not fit experiments when the Reynolds number is increased. A purely three-dimensional organization of the flow then appears, characterized by the Gortler vortices. Two families of solutions were thus evidenced, and the precise calculation of the physical one remains a numerical challenge. The prediction of transition by use of stability calculations is only partly possible because the waves used do not have a sufficiently general form to model such a complex physical problem. New information on the true nature of what is commonly called a transition mechanism in this kind of flow is deduced from these results.

Interference Between a Cylindrical Bow Shock and a Plane Oblique Shock

Hypersonic flows are prone to intense shock waves whose intersection, or interference, gives rise to a system of waves and slip surfaces that can have a large influence on the aerodynamics of a vehicle. It is convenient to distinguish six types of interference associated with very distinct flow structures, depending on the intensity and relative direction of the intersecting shock waves. Among these classes of interference, those of types III and IV are the origin of shear layers or supersonic jets whose impact on a nearby surface creates potentially destructive peaks of pressure and heat flux locally. Type III and IV interference and the corresponding heat transfer distributions were investigated experimentally. The model consisted of a cylindrically blunted plate and a wedge serving as an oblique shock generator. These experiments were carried out in a short-duration wind tunnel at Mach numbers 6 and 16 in air and at Mach number 6.6 in carbon dioxide. The Reynolds number based on the plate bluntness diameter was varied in the range from 2.2 x 10 4 to 1.6 x 10 6 . The influence of the impinging shock location on the interference heat transfer was carefully investigated

Optimization of Viscous Waveriders Derived from Axisymmetric Power-Law Blunt Body Flows

A method based on a Euler code is used to study waveriders generated from the flowfield around cones or axisymmetric power-law blunt bodies. In the cone case, the results are compared with those given by the Taylor-Maccoll system and inviscid hypersonic small-disturbance theory. This last theory shows its limits at a Mach number of 5, for cone angles providing the best lift-to-drag ratios. In the axisymmetric power-law blunt body flows case, the optimization is led using a nonlinear simplex method to get the upper surface that enables the waverider to have the best lift-to-drag ratio, the thickness-to-length ratio being fixed. Compared with the cone-derived waverider, the proposed blunt-body-derived model allows a 20% gain in volume for near equal optimized lift-to-drag ratios.

Shock-Wave/Turbulent Boundary-Layer Interactions in Nonequilibrium Flows

Shock-wave/turbulent boundary-layer interactions in the presence of thermal and chemical nonequilibrium phenomena are analyzed. The approach relies on a linear eddy viscosity two-equation turbulence modeling that accounts for the coupling of turbulence with chemistry and vibration, and it employs a total variation diminishing e nite volume numerical methodology. The capability of the model has e rst been assessed for a cylinder e are cone guration, and results have been compared with experiments. The model has then been applied to assess the aerodynamic performance of control surfaces of a reusable launch vehicle. In particular, the effects of wall temperature and e ap dee ection on the separation, aerothermal loads, and e ap efe ciency have been studied. The analysisshowsthatturbulencebecomesimportantfore apdee ectionanglesgreaterthanacriticalvalue[ O (15deg)], thus avoiding thecrisis of thee ap efe ciency that is observed under laminarconditions and extending the operating capabilitiesofthecontrolsurface.Thestudy also showsthatthewalltemperatureaffectssignie cantly theefe ciency and the operating envelope of the e ap primarily under turbulent conditions.

Code verification/validation with respect to experimental data banks

Abstract The past 40 years have known a spectacular development of CFD capabilities. It is now possible to compute complex three-dimensional unsteady flows even at the design stage by solving the Unsteady Averaged Navier–Stokes Equations (URANS approach) and progress are made every day in still more advanced approaches such as LES and DNS. However, the confidence in CFD methods is still limited because of uncertainties in the numerical accuracy of the codes and of the inadequacy of the turbulence models they use. Thus, there is still a need for well made and well documented experiments to validate the codes and to help in their improvement. Such experiments must also fulfil quality criteria to be considered as safe enough and really useful for code validations. The article presents a discussion of the strategy to be followed to ensure the reliability and accuracy of a code by placing emphasis on the experimental aspects of code validation. The purpose is illustrated by considering recent examples of CFD validation operations based on basic – or building block – experiments. The first case considers an experiment on a purely laminar shock wave/boundary layer interaction used to assess the numerical accuracy of several codes. Other examples deal with the crucial problem of the validation of turbulence models in strongly interacting flows. The conclusion stresses the importance to constitute high quality data banks on typical flows still difficult to predict. The problem of data dissemination is also briefly addressed.

Study of the Effect of Glow Discharges Near a M = 3 Bow Shock

This paper aims at investigating the effect of discharges generated near steady bow shocks and their possible use for localized heat deposition leading to drag reduction. We report the observation of steady discharges generated in front of a blunted model in a M = 3 supersonic flow. The test model is designed as a pair of coaxial electrodes, and set up in the ONERA R1Ch M = 3 blowdown wind tunnel. A high voltage power supply is used to generate negative discharges. A corona regime, a glow regime, and a filamentary arc regime are observed. The negative corona regime consists of Trichel pulses dissipating less than 200 mW of electrical power. The glow discharge absorbs powers up to 0.5 kW. It displays a strong light emission in the vicinity of the shock front, followed by a darker region downstream of the shock. The drag coefficient is measured and shows no measurable change when a glow discharge is switched on. To explain this and further investigate the effect of localized heat deposition at the shock front, modified Rankine-Hugoniot jump relations are computed, taking into account a volumetric heat source. This allows one to compute a fair estimate of the drag coefficient and shows that drag reduction by localized heating in the shock front is possible. However, it also shows that in our experiment, the plasma thermal power is too small to appreciably reduce the drag, possibly because of the role of the electron impact excitation of N 2 vibrations, whose fairly long relaxation time could shift downstream the effective gas heating. More generally, the model shows that power-efficient plasma-induced drag reduction requires high plasma heating efficiency.

Mach 3 shock-wave unsteadiness alleviation using a negative corona discharge

This paper reports experimental data obtained in ONERA M = 3 R1Ch blow-down wind tunnel. A truncated conical body with a central spike mounted on the truncated face is tested in a M = 3 airflow. Several runs have been performed, with a freestream unit Reynolds varying in the range Reu 2 [0.9 10 7 m 1 ; 2.1 10 7 m 1 ]. Depending on the spike length LS, the flow around the body is steady or unsteady. Following Kuo, 1 the model is also a pair of electrodes and is supplied with AC or DC power. Thus on-board discharges can be generated. Although different discharge regimes have been obtained on steady or unsteady flowfield, we simply report here the effect of corona discharges on an unsteady flowfield. This unsteadiness is characterized by a shock pulsation at the reduced frequency St = fc.LS/u0 = 3.10 2 . We observe a complete alleviation of the shock unsteadiness when a negative discharge is generated at the spike tip. With positive discharges, the effect is weaker. In both cases the input power is less than 1W. Discharge imaging and current measurements suggest that the negative discharge is a Trichel pulse corona discharge, whereas the positive discharge possibly involves streamers. The possible channels of interaction between the discharge and the airflow are briefly discussed.

Experimental Analysis of Aerodynamic Interactions Occurring on Hypersonic Spacecraft

Experimentalinvestigationshavebeenachievedin theONERAresearch facilitiesof theChalais ‐Meudon Center in order to validate design tools used for the dee nition of a scramjet for a hypersonic spacecraft. Aerodynamic interactions occurring on such a spacecraft have been thoroughly studied, notably 1 ) the shock/shock interactions ahead of the air inlets, 2 ) the shock-wave/boundary-layer interaction and its control inside the air inlet, and 3)the e ow cone uence between an external e ow and an aerospike nozzle jet in an underexpansion condition. These studies contribute to the knowledge of the e ow physics of aerodynamic interactions that occur on hypersonic spacecraft, compensate for the lack of published detailed measurements of such e owe eld patterns, and provide well-documented test cases to validate computer codes.

High-enthalpy hypersonic project at ONERA

Abstract This article deals with the research performed in the framework of the internal ONERA Hyperenthalpic Hypersonic Project. This project involved fifteen researchers specialised in different domains, which are the following: fundamental studies on shock/wave and shock/boundary layers interactions; study of the laminar/turbulent transition; real gas effects and associated physical modelling; the development of non-intrusive optical diagnostic methods usable in cold and hot wind tunnels; flow rebuilding in a hot shot wind tunnel such as F4; conception, realisation and validation of computational fluid dynamics (CFD) solvers.