MohAmed SELLAM

Interactions between shock waves and boundary layers (SWBLI) are encountered in many industrial applications dealing with supersonic flows (aircraft design, supersonic inlet, rocket nozzles...). If the shock is strong enough, those interactions may cause the boundary layer separation yielding dynamic loads, increased heat fluxes and pressure fluctuations. Even if the physics of SWBLI is not fully understood, it is well known that the separation zone as well as the reflected shock are subjected to a low-frequency streamwise motion that can spread over several tenth of the boundary layer thickness. The origin of this motion is, however, not completely elucidated. Several studies, both numerical and experimental, have linked it to the vortex shedding in the mixing layer downstream of the separation and a simple model that explains the low-frequency unsteadiness has even been developed. On the other hand, a correlation between the vortical structures in the incoming boundary layer and t...

Abstract Experiments on an axisymmetric dual-bell nozzle were performed at EDITH nozzle test facility of CNRS in Orleans, France. The main purpose of the study was to explore the possibility of controlling the flow regime transition by a secondary fluidic injection in the dual bell nozzle. The main focus of the present paper is to investigate the impact of the secondary injection parameters on the flow regimes transition in such nozzles. Secondary injection has been found to effectively control the flow regime transition and consequently to increase the propulsive performance of the device. It has also been pointed out that even a very low injected secondary mass flow rate leads to the control of the transition and contributes to reducing the lateral loads which can exist, moreover, when transitions are operated without injection.

ABSTRACT Transverse secondary gas injection into the supersonic flow of an axisymmetric convergent–divergent nozzle is investigated to describe the effects of the fluidic thrust vectoring within the framework of a small satellite launcher. Cold-flow dry-air experiments are performed in a supersonic wind tunnel using two identical supersonic conical nozzles with the different transverse injection port positions. The complex three-dimensional flow field generated by the supersonic cross-flows in these test nozzles was examined. Valuable experimental data were confronted and compared with the results obtained from the numerical simulations. Different nozzle models are numerically simulated under experimental conditions and then further investigated to determine which parameters significantly affect thrust vectoring. Effects which characterize the nozzle and thrust vectoring performances are established. The results indicate that with moderate secondary to primary mass flow rate ratios, ranging around 5 %, it is possible to achieve pertinent vector side forces. It is also revealed that injector positioning and geometry have a strong effect on the shock vector control system and nozzle performances.

A numerical investigation of low-Reynolds number flows with thermal effect around the MAV airfoils using various turbulence models, including an algebraic Baldwin-Lomax model, Spalart-Allmaras one equation, and two equation (kx and SST-kx) turbulence models, is ...

ABSTRACT The use of plasma actuators for flow control has received considerable attention in recent years. This kind of device seems to be an appropriate means of raising abilities in flow control thanks to total electric control, no moving parts and a fast response time. The experimental work presented here shows, firstly, the non-intrusive character of the visualization of the density field of an airflow around a cylinder obtained using a plasma luminescence technique. Experiments are made in a continuous supersonic wind tunnel. The static pressure in the flow is 8 Pa, the mean free path is about 0.3 mm and the airflow velocity is 510 m s−1. Pressure measurements obtained by means of glass Pitot tube without the visualization discharge are proposed. Measured and simulated pressure profiles are in good agreement in the region near the cylinder. There is good correlation between numerical simulations of the supersonic flow field, analytical model predictions and experimental flow visualizations obtained by a plasma luminescence technique. Consequently, we show that the plasma luminescence technique is non-intrusive. Secondly, the effect of a dc discharge on a supersonic rarefied air flow around a cylinder is studied. An electrode is flush mounted on the cylinder. Stagnation pressure profiles are examined for different electrode positions on the cylinder. A shock wave modification depending on the electrode location is observed. The discharge placed at the upstream stagnation point induces an upstream shift of the bow shock, whereas a modification of the shock wave shape is observed when it is placed at 45° or 90°.

Possibilities of the fluidic thrust vector control of a supersonic rocket nozzle have been investigated in the framework of the CNES program Perseus for small satellite launcher. Cold flow experiments are performed in the hypersonic wind tunnel EDITH on several nozzle types and configurations. The complex three dimensional flow field generated by the supersonic cross-flows in these test nozzles is examined. Valuable experimental data is confronted and compared with the results of numerical simulations performed using the FANS solver of the code CPS C. Test case nozzle models are numerically simulated for experimental conditions and then further investigated for additional thrust vectoring parameters. Effects which characterize the nozzle and thrust vectoring performances are depicted. The study indicates that with moderate secondary to primary mass-flow-rate ratios, ranging around 5%, it is possible to achieve pertinent vector side force. It also reveals effects of injector geometry...

ABSTRACT A renewal of interest to study the possibilities of using the gas detonation as a thrust - producing mechanism has been observed during the last ten years. This interest can be motivated by the wide simplicity of the design, the low cost of exploitation and the acceptable performance, so that the PDE can give to this propulsion concept a place as a serious competitor device. The aim of this paper consists on an analysis of the performance of a Hydrogen-Air mixture PDE, computed by the PLEXUS code, using an accurate CFD numerical scheme, for many geometric and operating configurations. The main characteristic of this approach is the possibility of using a geometry that could be structured or unstructured with the associated properties. In particular, the modeling of the detonation propagation can make into account the hydrodynamic effects and combustion one. The interaction of the gas jet, stemming from the combustion chamber, with the surrounded air is also examined and discussed. This may lead to study the PDE in flight configuration in particular with atmosphere configuration.

ABSTRACT A renewal of interest to study the possibilities of using the gas detonation as a thrust - producing mechanism has been observed during the last ten years. This interest can be motivated by the wide simplicity of the design, the low cost of exploitation and the acceptable performance, so that the PDE can give to this propulsion concept a place as a serious competitor device. The aim of this paper consists on an analysis of the performance of a Hydrogen-Air mixture PDE, computed by the PLEXUS code, using an accurate CFD numerical scheme, for many geometric and operating configurations. The main characteristic of this approach is the possibility of using a geometry that could be structured or unstructured with the associated properties. In particular, the modeling of the detonation propagation can make into account the hydrodynamic effects and combustion one. The interaction of the gas jet, stemming from the combustion chamber, with the surrounded air is also examined and discussed. This may lead to study the PDE in flight configuration in particular with atmosphere configuration.

ABSTRACT An engineering-type analysis is proposed and used to investigate the performances of thrust vectoring by fluidic injection in the divergent of a supersonic axisymmetrical convergent-divergent nozzle. This method includes several approaches which consist mainly of a fluidic obstacle height evaluation and the prediction of the separation line that results upstream of the fluidic obstacle. The construction of the separation line is also based on some separation correlations proposed in the literature. The nozzle thrust deviation is then calculated by taking into account the injectant fluid momentum rate contribution and the integration of the pressure acting on the nozzle inner wall. The sensitivity of the model versus some separation criteria is discussed. The results of the analytical model are compared with the experiments conducted recently by the authors. The comparison shows a very good agreement for some of the separation criteria over the whole range of injected to main flow-rate ratios.

The aim of this paper is to discuss the development of new contours of axisymmetric supersonic nozzles giving a uniform and parallel flow at the exit section, to improve the aerodynamic performances compared to the minimum length nozzle, by increasing the exit Mach number and the thrust coefficient, and by reduction of the nozzle's mass, while holding the same throat section between the two nozzles. The new nozzle is named the best performance nozzle. Its form contains a cylindrical central body and an external wall for the flow redress. The study is done at high temperature, lower than the dissociation threshold of the molecules. The variation of the specific heats with the temperature is considered. The design is made by the method of characteristics. The predictor-corrector algorithm is used to make the numerical resolution of the obtained nonlinear algebraic equations. The validation of results is made by the convergence of the numerical critical sections ratio with that giv...

A numerical investigation is conducted to study an H2 - O2 rocket nozzle flow in chemical and vibrational nonequilibrium. Therefore, a 9 reactions kinetic model was implemented in our two-temperature house code. The vibrational relaxation times taken from Skreb-kov's theoretical model and utilized in our simulations are found to be better suited for the H2 - O2 mixture than those evaluated using Millikan & White semi-empiric formula. The utilized 9 reactions kinetic model demonstrates a good modeling of the chemical nonequilibrium. The results show the presence of three regions in propulsive nozzles: an equilibrium region, a nonequilibrium region followed by a return to equilibrium region. Vibrational nonequilibrium effects on flowfield parameters and nozzle performances in fuel-rich flows of H2 - O2 rocket are investigated, by comparing this baseline simulation to vibrational equilibrium simulation. Vibrational nonequilibrium effects on flowfield parameters and on nozzle performances are computed and shown to be minor. A reduction of 90% of computation time is observed when using the vibrational equilibrium configuration instead of the vibrational nonequilibrium configuration.

ABSTRACT The cross injection in a supersonic flow is an issue encountered in several aerodynamic applications such as fuel injection in scramjet combustor, missile control, drag reduction and thrust vector control. In a recent work, an analytical model has been presented to calculate the fluidic thrust vectoring performance for a supersonic axisymmetric nozzle. The model is able to take into account both the injected gas thermodynamic properties and the geometrical nozzle characteristics. The analytical model has been successfully validated following the cold air flow experimental analysis, in the case of fluidic thrust vectoring applied to conical nozzle. The aim of this work is to show how far the injected gas thermodynamic properties, different from that of the nozzle main flow, could influence the fluidic thrust vectorization parameters. In this work, the experimental performance of the fluidic thrust vectoring concept, using numbers of gases as injectant, has been qualitatively and quantitatively analyzed. Schlieren visualization, force balance and wall pressure measurements were used in the case of a truncated ideal contour nozzle. The experimental results are compared to the numerical and analytical findings. Performance analysis are conducted and basic conclusions are drawn in terms of thermodynamic gas properties effect on the fluidic thrust vector system. The primary effect was related to the gas molecular weight and its specific heat ratio product. It is observed that for fixed injection conditions, the vectoring angle is higher when the injected gas molecular weight and specific heat ratio product is less than that of the primary gas. For a given mission of the launcher, it can be concluded that the mass of the embedded gas, used for the fluidic vectorization system, can be significantly reduced, depending on its molecular weight and specific heat ratio.

ABSTRACT The aim of this paper is to present and compare two different approaches for aeroelastic stability analysis of a flexible over-expanded rocket nozzle. The first approach is based on the aeroelastic stability models developed in a previous work, while the second uses the numerical fluid–structure coupling via the transpiration method. The aeroelastic frequencies of the nozzle obtained by various stability models are compared with those extracted from the numerical coupling by the method of transpiration. Both set of results show an overall good agreement.

The aim of this study is to develop a new approach for the aeroelastic stability analysis of a flexible over-expanded rocket nozzle using numerical fluid-structure coupling via the transpiration method. The method consists of a weak coupling between two numerical codes; one is dedicated to the Computational Structural Dynamics (CSD) and the other to the Computational Fluid Dynamics (CFD). Indeed, given the context of the study, which is limited to the analysis of linear stability relative to the small displacements, here, the ALE (Arbitrary Lagrangian-Eulerian) formalism, based on the mesh dynamic aspect, widely discussed in many previous studies loses its attractiveness in favour of the transpiration method. In fact, the choice of the transpiration technique to consider the coupling at the fluid-structure interface is justified by its simple implementation and its low CPU time computation compared to those of the ALE method. It only requires a simple reformulation of the boundary c...

Abstract - In this paper, validation and assessment of various turbulence models are performed in transonic flows, including an algebraic Baldwin Lomax model, Spalart Allmaras one equation, and two equation (k-ε, k-ω and SST k-ω) turbulence models. The NACA 0012 airfoil has been chosen for the turbulence model validation studies. The three test cases selected here included both attached and separated transonic flows. This study shows that the five turbulence models provide satisfactory results for transonic attached flows. However, all models, fail to predict the location of the shock correctly when a strong shock and a shock induced flows separation are present. Compared to other models, it has been shown that the SST k-ω model is the most robust model in the prediction of lift coefficient for all cases. Computed results are performed with the CFD-FASTRAN code by using the fully implicit scheme for time integration, and the upwind Roe flux difference splitting scheme for space disc...

The aim of the present study is to analyze the aeroelastic stability of a supersonic nozzle in over-expanded conditions, by using an aeroelastic stability model. To reach this objective, a research software written in Fortran, has been developed for 2D and 3D nozzle configurations. The obtained results are compared and validated for the 2D and 3D cases with those of previously studies.

Reignition phenomena occurring in a supersonic nozzle flow may present a crucial safety issue for rocket propulsion systems. These phenomena concern mainly rocket engines which use H2gas (GH2) in the film cooling device, particularly when the nozzle operates under over expanded flow conditions at sea level or at low altitudes. Consequently, the induced wall thermal loads can lead to the nozzle geometry alteration, which in turn, leads to the appearance of strong side loads that may be detrimental to the rocket engine structural integrity. It is therefore necessary to understand both aerodynamic and chemical mechanisms that are at the origin of these processes. This paper is a numerical contribution which reports results from CFD analysis carried out for supersonic reactive flows in a planar nozzle cooled with GH2film. Like the experimental observations, CFD simulations showed their ability to highlight these phenomena for the same nozzle flow conditions. Induced thermal load are als...

ABSTRACT The aim of this paper is to present a newaeroelastic stability model taking into account the viscous effects for a supersonic nozzle flow in overexpanded regimes. This model is inspired by the Pekkari model which was developed initially for perfect fluid flow. The new model called the “Modified Pekkari Model” (MPM) considers a more realistic wall pressure profile for the case of a free shock separation inside the supersonic nozzle using the free interaction theory of Chapman. To reach this objective, a code for structure computation coupled with aerodynamic excitation effects is developed that allows the analysis of aeroelastic stability for the overexpanded nozzles. The main results are presented in a comparative manner using existing models (Pekkari model and its extended version) and the modified Pekkari model developed in this work.

ABSTRACT The flowfield resulting from the transverse gas injection into the supersonic cross-flow is the problematic of many aerospace applications ranging from the scram-jet fuel injection to the reaction jets and fluidic thrust vectoring(FTV). The prominent case of FTV by the use of the secondary injection represents promisingly attractive and effective way of control for small aerospace vehicles. Main advantages of the FTV are light-weightiness, simplicity and potential efficiency of such system comparing to the conventional mechanical TVC and mechanical deflectors. [5] Elimination of heavy and robust actuators and their replacement with only the fast-opening valves leads to the very significant reduction in mass. Fast dynamic response (~500Hz) to the conventional (~30Hz) [8] , very small losses in specific impulse and thus thrust are promising efficiency benefits. The CNES ”Perseus” project which this study is part of and design concept have aim of incorporating fluidic TVC on the future ”micro” launcher. Some of the results from the ongoing investigation are presented in this article.

ABSTRACT A Numerical and modeling study has been accomplished to investigate the effects of a secondary injection in an axisymmetrical,convergent-divergent nozzle for fluidic thrust vectoring purpose. The annular secondary gas injection in the axisymmetrical nozzle causes complex effects (boundary layer separation, shock wave interaction,…). The present paper focuses on the results of some computational and modeling investigations, where the influence of some parameters (pressure ratios, injection slot size and location, injected mass flow rates…) are studied. To perform this work, a 3D Navier-Stokes calculations, with several turbulence models were used. Previously, a theoretical model of a secondary injection in a primary jet had been constructed. To characterize the separation zone caused by the injection, different correlations have been tested. The results indicate that fluidic annular injection in an axisymmetrical nozzle can produce significant thrust-vector angles up to 16 °. The nozzle pressure ratio and the mass flow rate ratio were in the range of 2 to 10 and 2 to 7% respectively. Some results were validated on NASA experiments in both 2D and axisymmetric tests.