Afterbody Flow Research Articles

Experimental study on the synergy of sweeping jets on the afterbody flows of a slanted-base cylinder

The synergistic effects of sweeping jets (SWJs) hold significant potential in controlling afterbody vortex flows of a cargo fuselage. A combination of SWJ pairs, located at various streamwise positions, are employed to investigate the efficiency of their synergy in influencing the afterbody flows of a slanted-base cylinder with a 30° angle at a Reynolds number of 200,000. The total mass flow rate is kept similar resulting in a variation of the momentum coefficient of SWJs from Cμ=4.3×10−3 to 1.0×10−1. Energy efficiency and mass flow interactions were examined through direct force measurement, stereo Particle Image Velocimetry (PIV), and surface pressure measurement. The results show that the jet impingement on the surface is stronger when there is synergy among SWJs positioned more upstream (e.g. the combination of Top and Mid pairs), particularly at low momentum coefficients (Cμ < 0.02); conversely, when considering the synergy of SWJs located more downstream (e.g. the combination of Mid and Bot pairs), a greater reduction in surface pressure is observed at higher momentum coefficients (Cμ > 0.05); a potential net energy saving is achieved in the case of the combination of Top, Mid and Bot pairs (T+M+B). The controlled vorticity distribution observed in the collaborative configurations was mostly sandwiched by curves of individual SWJ cases, which is due to the fact that downstream SWJs (located at Mid and Bot positions in case T+M+B for instance) manage to weaken the enhanced vortex induced by the SWJ at upstream (located at Top position), highlighting a compromise in the synergy effectiveness of the jets with reverse roles at different positions. However, by taking the similar advantages of SWJ at certain positions on the vortex control, at Cμ=1.7×10−2, the vorticity reduced by the case combining Mid and Bot pairs is even 7 % and 13.5 % greater than cases Mid and Bot, respectively, which could serve as a hint on the potential benefit of SWJs’ synergy in the future.

Time-resolved particle image velocimetry and pressure sensitive paint measurements of afterbody flow dynamics

A pair of counter-rotating vortices is produced in the wake of typical cargo aircraft fuselages due to their ramp shape. With a fast polymer-ceramic pressure sensitive paint (PC-PSP), we measure the global pressure fluctuations at very low velocities (15 to 75 m/s) in a surrogate cylinder with a slanted edge model. Through spectral proper orthogonal decomposition, we observe a family of convective pressure waves with peak energy at frequencies matching the vortex wandering frequencies in this flow.

Numerical workflow for scale-resolving computations of space launcher afterbody flows with and without jets

Scale resolving numerical methods are necessary to accurately predict the afterbody flows of space launchers. Numerical workflows built for such computations have to be designed in order to obtain a minimal numerical dissipation to resolve fine turbulent structures, a sufficient robustness to capture eventual shock waves and both a reasonable computational cost and an acceptable user workload enabling comparative design studies. This article presents the development of a hybrid numerical framework based on Ducros’s sensor and designed to switch from a low-dissipation formulation in presence of vortical structures to a robust formulation around high gradients. This hybrid workflow is used with ZDES, including its latest automatic mode (ZDES mode 2 (2020)), to simulate a transonic space launcher afterbody experiment with and without a cold air propulsive jet [1,2]. The salient physical properties of the base flow are investigated and the evaluation of computed results follows the extended nomenclature for validation of simulation techniques [3] from level 0 to 5. The framework is thus validated for instantaneous and mean flow visualizations, base and extension mean pressure coefficient distribution, pressure fluctuation levels, one-point and two-point spectral analyses. Such an efficient automatic RANS/LES strategy could be well suited to study realistic launcher afterbody geometries with propulsive jet(s).

Investigation of the base flow of a generic space launcher with dual-bell nozzle

The typical afterbody flow of a space launcher is characterized by a strong interaction of the engine’s exhaust jet and the separated shear layer emerging from the main body. This interaction is further complicated by strong changes in the spatial and temporal behavior of the afterbody flow during the atmospheric ascent of a launcher. Theoretically, a dual-bell nozzle not only allows for a gain in payload compared to standard single-bell nozzles, but also it alters the wake flow topology due to the two nozzle modes. To predict the benefits as well as the additional risks, the afterbody flow of a generic space launcher model equipped with a cold-flow dual-bell nozzle is investigated in detail. The flow was analyzed for sub-, trans- and supersonic Mach numbers ranging from 0.3 to 2.9 for a variety of nozzle pressure ratios. Particle image velocimetry measurements and schlieren measurements with high repetition rate were performed to determine the dynamics of the separated shear layer, the nozzle jet and their interaction. It is shown that the reattachment length of the base flow decreases with increasing nozzle pressure ratio. Furthermore, the nozzle pressure ratio at which the dual-bell nozzle switches from sea-level mode to altitude mode is reduced by 15% with high subsonic outer flow and by as much as 65% for an outer flow at a Mach number of 1.6. Even for a constant nozzle pressure ratio, the nozzle flow topology depends on the Mach number of the outer flow.

Hysteresis in slanted-base-cylinder afterbody flows

Afterbody flows such as those behind cargo aircraft type configurations display rich fluid dynamic behavior, including flow separation at the base of body, complex three-dimensional vortical structures and wake turbulence. A representative configuration for study is a cylinder with axis parallel to the freestream and an upswept base. At low upsweep angles, measured relative to the axis, the flow is characterized by a pair of streamwise vortices. As the angle is increased, the flow transitions to a wake closure. Experimental evidence shows that near the critical angle, (ϕ≃45°), either of the two patterns may be obtained, depending on the history of the flow i.e., the flow exhibits hysteresis. The features associated with this phenomenon are examined numerically using a validated Large-Eddy Simulation approach. A systematic campaign varying both Reynolds number and upsweep angle confirms the presence of hysteresis at ReD=25,000 and ϕ=45°. In the vortex regime, the three-dimensional separated flow from the sides rolls up into a horseshoe-type vortex with legs representing a streamwise oriented pair displaying coherent but counter-rotating vorticity field, while a diffused wake-like regime is obtained in the other. Sectional projected streamlines highlight the key kinematic differences between the two regimes. The change in base pressure due to very different separation patterns is analyzed to quantitatively examine the large drag disparity observed in experiments. Finally, instantaneous flowfields are used to obtain the shedding characteristics as well as the nature of wake turbulence in both regimes. The present study has relevance to flows in aerodynamic vehicles as well as in automobiles, and may aid in possible control strategies for drag reduction.

Flow-radiation coupling inCO2hypersonic wakes using reduced-order non-Boltzmann models

This work presents a unified physics-based reduced-order simulation framework for describing non-Boltzmann thermochemistry and radiative heating for hypersonic flows. The new approach is used to analyze afterbody flow dynamics and concomitant carbon dioxide IR radiation for the Mars 2020 mission.

Experimental study of recombining nitrogen plasmas: I. Vibronic population distributions and nonequilibrium molecular radiation

We present optical emission spectroscopy measurements in recombining nitrogen plasma flows at atmospheric pressure. An inductively coupled plasma torch is used to create an equilibrium plasma, which is then forced to recombine by flowing through a water-cooled tube. For certain conditions, the plasma is forced out of chemical equilibrium. The emission of and is studied to measure the nonequilibrium vibrational density distributions within these electronic states. These densities are found to be highly overpopulated in comparison with their corresponding equilibrium values, which is consistent with previous results in recombining flows. The measured densities are also compared with the predictions of a two-temperature model. This two-temperature model underpredicts the measured densities but is closer to the measured distributions than the equilibrium densities. The total measured radiation is approximately 100 000 times stronger than the corresponding equilibrium radiation. The 2-T model estimate of this radiation is much closer yet still underestimates by a factor of 10 the measured radiation. These data are intended as a new dataset to test the recombining plasma models used to simulate afterbody flows during atmospheric reentry.

On the coupling of a zonal body-fitted/immersed boundary method with ZDES: Application to the interactions on a realistic space launcher afterbody flow

One of the next foreseen challenges in CFD consists in the capability to simulate quantitatively the spectral content of the turbulent flow around realistic geometries. In this context, the present work focuses on a new methodology named ZIBC standing for Zonal Immersed Boundary Conditions enabling to account for complex configurations at high Reynolds numbers. The numerical strategy allowing the coupling between a turbulence modeling method (e.g. RANS, URANS, ZDES, LES or DNS) and IBC (Immersed Boundary Conditions) is detailed. In this paper, the modeling method retained is the Zonal Detached Eddy Simulation (ZDES) which has reached a high level of maturity on turbulent separated flow simulations. This methodology is applied to a full space launcher configuration to assess its capability to return the interactions between the technological details, modeled with IBC, and the simplified afterbody, modeled with a body-fitted (BF) approach consisting in classical no-slip boundary conditions, in the turbulent flow field surrounding the main stage of the space launcher afterbody. The proposed method is thoroughly assessed on a realistic geometry of the European Ariane 5 launcher and the ZIBC simulation is successfully compared with the available experiments.

Compressible Turbulence Model Consistency for Separated High-Speed Flow Regimes

Fluid compressibility can remarkably modify the behavior of turbulent flow with respect to the incompressible case. The presence of strong pressure-induced density changes adds to the turbulence phenomenology new specific processes and mechanisms, such as new pathways for energy exchanges and a strong coupling between momentum and energy exchanges, significantly altering the flow behavior. Many research groups are working on methods to include these compressibility effects in conventional turbulence models to predict the flowfield for compressible conditions more realistically. This work aims at the assessment and validation of several compressible corrections to an in-house hypersonic multigrid reactive Navier–Stokes code within the “baseline” Spalart–Allmaras turbulence model. The base pressure enhancement for supersonic afterbody flow computations is first addressed. Then, the shock/turbulent boundary-layer interaction at the junction of a cone/flare configuration is dealt with. The analysis of the numerical results and their comparison with the experimental data show that, in most cases, the proposed corrections produce a significant simulation improvement. Finally, an in-house hypersonic wind-tunnel test campaign is performed on generic blunt-cone/flare configurations. The pressure distributions along the model as well as the forces and moments are used to calibrate the numerical calculations in the range of Mach numbers of 4.8 to 6.2.

Propulsive jet influence on generic launcher base flow

Afterbody flow phenomena represent a significant source of uncertainties in the design of a launcher. Therefore, there is a demand for measuring such flows in wind tunnels. For propulsive jet simulation a new jet facility was integrated into a hypersonic/supersonic wind tunnel. The jet simulation resembles the generic model of a staged rocket launcher. The design and the qualification of the facility are reported. This includes measurements of pressure, temperature and Mach number distribution. Pressure and Schlieren measurements are conducted in the wake of the generic launcher. The unsteady pressure characteristics at the generic rocket base and fairing are analyzed for supersonic and hypersonic freestream. The influence of the under-expanded jet is reported and the jet temperatures are varied. On the base fluctuations at a Strouhal number around 0.25 dominates supersonic freestream flows. Additionally, a fluctuation level increase on the base is observed for Strouhal numbers above 0.75 in hypersonic flow regime, which is attributed to the interactions of wake flow and jet.

Zonal Immersed Boundary Conditions: Application to a High-Reynolds-Number Afterbody Flow

The present paper focuses on a new numerical strategy consisting of the zonal use of the immersed boundary conditions combined with the ability of zonal detached-eddy simulation to simulate high-Reynolds-number separated flows. The validation test case is a simplified space launcher afterbody with a control device in the form of a short cylindrical serrated skirt. “Zonal immersed boundary conditions” are used to handle the introduction of the skirt into the preexisting structured curvilinear grid that describes the noncontrolled configuration. The governing equations are solved using a standard body-fitted finite volume technique over the whole grid. A direct forcing source term is added when cells are internal to the skirt (that is, solid) to drive the velocity and the turbulence variables to the chosen values. Numerical simulations are performed at a Reynolds number of and a freestream Mach number of 0.702. The numerical results demonstrate the ability of the zonal immersed boundary conditions to successfully impose the desired values at solid nodes. The first- and second-order moments, as well as the fluctuating pressure field, illustrate a fairly good agreement between the experiment and the numerical simulation. It is shown that the serrated skirt acts as a smooth device, which does not affect the global dynamics of the flow and only delays the location of the separation. Finally, the zonal immersed boundary conditions appear to successfully reproduce the effect of the skirt while drastically reducing the time of the grid generation.

Experimental analysis of the pressure–velocity correlations of external unsteady flow over rocket launchers

Based on simultaneous wall pressure and velocity measurements, the aerodynamic load of a launcher body model is investigated. Two different configurations are considered in order to study the influence of geometrical artifacts on the after-body flow and consequently on the aerodynamic load. After a brief presentation of the experimental setup used to get the unsteady wall pressure and external velocity field, a global analysis of the integrated pressure along the nozzle is presented for both configurations. It is shown that the unsteady load induced on a configuration with attachment device involves characteristic frequencies which fits the mechanical response of the structure whereas no particular behavior is observed on the configuration without attachment device. Then, a Proper Orthogonal Decomposition (POD) is successively performed from the wall pressure field and from the external velocity field highlighting the relationship between the most energetic structures of the flow and the involved phenomenon. Finally, a pressure–velocity correlation of the POD modes is presented and the most energetic structures of the velocity field are linked to the unsteady load observed on the nozzle. It is then confirmed the structural influence of the attachment device and its contribution to the unsteady loads acting on the nozzle.

Effect of compressibility on the global stability of axisymmetric wake flows

We study the linear dynamics of global eigenmodes in compressible axisymmetric wake flows, up to the high subsonic regime. We consider both an afterbody flow at zero angle of attack and a sphere, and find that the sequence of bifurcations destabilizing the axisymmetric steady flow is independent of the Mach number and reminiscent of that documented in the incompressible wake past a sphere and a disk (Natarajan &amp; Acrivos, J. Fluid Mech., vol. 254, 1993, p. 323), hence suggesting that the onset of unsteadiness in this class of flows results from a global instability. We determine the boundary separating the stable and unstable domains in the (M, Re) plane, and show that an increase in the Mach number yields a stabilization of the afterbody flow, but a destabilization of the sphere flow. These compressible effects are further investigated by means of adjoint-based sensitivity analyses relying on the computation of gradients or sensitivity functions. Using this theoretical formalism, we show that they do not act through specific compressibility effects at the disturbance level but mainly through implicit base flow modifications, an effect that had not been taken into consideration by previous studies based on prescribed parallel base flow profiles. We propose a physical interpretation for the observed compressible effects, based on the competition between advection and production of disturbances, and provide evidence linking the stabilizing/destabilizing effect observed when varying the Mach number to a strengthening/weakening of the disturbance advection mechanism. We show, in particular, that the destabilizing effect of compressibility observed in the case of the sphere results from a significant increase of the backflow velocity in the whole recirculating bubble, which opposes the downstream advection of disturbances.

Elephant modes and low frequency unsteadiness in a high Reynolds number, transonic afterbody wake

Experiments and large eddy numerical simulation of a fully turbulent afterbody flow in the high subsonic regime, typical of that developing in the wake of a space launcher, exhibit a large-scale low frequency oscillation of the wake. In the present paper, we investigate to what extent the existence of the synchronized oscillations can be interpreted, at the high Reynolds numbers prevailing in this class of flows, by a local stability analysis of the mean flow, as measured in experiments or computed in numerical simulations. This analysis shows the presence of a pocket of absolute instability in the near wake, slightly detached from the body. The global frequency is strikingly well predicted by the absolute frequency at the upstream station of marginal absolute instability, this frequency selection being in agreement with the theory of nonlinear global modes. This result strongly suggests that a so-called elephant mode is responsible for the intense oscillations observed in the lee of space launcher configurations.

Numerical simulation of axisymmetric afterbody flows with jet exhaust

A fifth-order accurate compact difference scheme was used to compute the flow over an axisymmetric afterbody with jet exhaust. The solution was based on the mass-averaged Navier-Stokes equations combined with a two-parameter differential model of turbulence. The computations were performed on a specially generated mesh such that the flow in the exterior and interior of the nozzle could be described simultaneously. Numerical results are presented for various external flow conditions and various chamber pressures.

Reynolds-stress modelling of transonic afterbody flows

AbstractSeveral afterbody flows, involving shock-boundary-layer interaction, are used to evaluate recent developments in a realizable low-Reynolds-number, second-moment closure of turbulence. The model considered is a compressibility-adapted variant of the recent incompressible-flow form of Craft and Launder. This includes a tensorially cubic model for the influential pressure-strain process, ϕij, which satisfies the two-component-turbulence limit at the wall, is directly applicable to low-Reynolds-number flow regions and does not rely on or use surface-topography parameters, such as wall-normal distance or direction. Improved predictions for afterbody flows are demonstrated, relative to existing low-Reynolds-number two-equation models and the most elaborate form of Reynolds-stress closure incorporating a linear approximation for the pressure-strain process.

Modeling of Afterbody Flows with Realistic Propulsive Gases: A Prospective Study

The organization of afterbody e ows of missiles or launchers determines the base pressure, which is itself a key factor ine uencing the thrust budget. The structure of the afterbody is submitted to aerothermal loads due to the mixing of the external e ow with the high-temperature gases exiting from the nozzle. The general feature of the afterbody e ow, determining energy transport phenomena, directly ine uences the thermal e uxes to which the structure is submitted. The importance of establishing precise methods of calculation to predict transport phenomena in the mixing of realistic gaseous species appears now at the design level. To realize this program, fundamental works on afterbody e ows have been performed in the e eld of modeling, where the NASCA code was developed for these purposes, and also in the experimental e eld, where data benchmarks were constituted for cold supersonic e ows. For the case of realistic temperature and chemical compositions, benchmarks do not currently exist, and the aim is to establish a model of transport phenomena in nonreacting multicomponent gases that is consistent with the more commonly used turbulence models. After validating thecode on a referencecone guration with supersonic cold e ow and a cold propulsive jet, prospective calculations are performed with more realistic high-temperature jet conditions and chemical compositions. Some divergences between the model’ s predictions of aerothermal e elds are observed. Nomenclature Cp, ˆ Cp, Cpi = heat capacities at constant pressure Cv, Cvi = heat capacities at constant volume Di, DT, Dij = molecular, turbulent, and binary diffusion

Modeling of afterbody flows with realistic propulsive gases - A prospective study

Modeling of afterbody flows with realistic propulsive gases - A prospective study

Supersonic base-flow computation using higher-order closure turbulence models

Three higher-order turbulence models—the low-Reynolds-number form of a standard two-equation model, the two-layer algebraic stress model, and the Reynolds-stress model—are combined with Navier-Stokes equations to compute the supersonic flows behind a missile-type afterbody and two-dimensional two-stream flow behind a thick base. The governing equations are numerically solved by an implicit total variation diminishing algorithm for high accuracy. The preconditioned bi-conjugate-gradient stable scheme is also employed to accelerate and stabilize the convergence. Those results demonstrate that the Reynolds-stress model is better than the other two models in predicting the mean flow, turbulent quantities, and surface pressure distribution in the recirculating regions/Nevertheless, improvements in the Reynolds-stress model are suggested so as to more accurately predict the turbulent properties of shock-shear-layer interaction around the recompression region. All three turbulence models fail to achieve a sufficient velocity recovery rate in the redevelopment region of a supersonic base flow.

Numerical Analysis of Nozzle and Afterbody Flow of Hypersonic Transport Systems

Using an implicit Finite-Volume Navier–Stokes code, the flow field in a Single Expansion Ramp Nozzle (SERN) for a hypersonic aircraft is studied. Comparisons between experimental data and CFD calculations for certain components of the integrated exhaust system (cold two-dimensional nozzle flow, high temperature reacting three-dimensional combustion chamber flow, and two-dimensional nozzle flow with external flow) are presented. To show the sensitivity of the considered components to off-design operating conditions, comprehensive numerical studies have been carried out. For the determination of nozzle performance a detailed two-dimensional analysis from transonic to hypersonic flight Mach numbers has been performed. A direct optimization method has been used to investigate the influence of the lower nozzle flap shape on the thrust vector.