Laurent Selle

Large-Eddy Simulation of Supercritical-Pressure Round Jets

This paper presents the numerical computation of a turbulent jet of nitrogen into nitrogen under supercritical pressure. The large-eddy simulation framework for turbulence modeling is used and real-gas effects are accounted for through a cubic equation of state and appropriate viscosity and conductivity coefficients. The purpose of this paper is to evaluate how low-pressure large-eddy simulation equations coupled with real-gas thermodynamics and transport compare with experiments. Although this approach does not take into account the impact of high density gradients and nonlinear thermodynamics on turbulence modeling, the results show reasonable agreement with available experimental data and reveal the importance of numerics for such computations. The simulations indicate a limited influence of the density ratio and the thermodynamic conditions on the jets spreading rate and pseudosimilarity behavior.

Numerical Benchmark for High-Reynolds-Number Supercritical Flows with Large Density Gradients

Because of the extreme complexity of physical phenomena at high pressure, only limited data are available for solver validation at device-relevant conditions such as liquid rocket engines, gas turbines, or diesel engines. In the present study, a two-dimensional direct numerical simulation is used to establish a benchmark for supercritical flow at a high Reynolds number and high-density ratio at conditions typically encountered in liquid rocket engines. Emphasis has been placed on maintaining the flow characteristics of actual systems with simple boundary conditions, grid spacing, and geometry. Results from two different state-of-the-art codes, with markedly different numerical formalisms, are compared using this benchmark. The strong similarity between the two numerical predictions lends confidence to the physical accuracy of the results. The established database can be used for solver benchmarking and model development at conditions relevant to many propulsion and power systems.

Analysis of the impact of heat losses on an unstable model rocket-engine combustor using Large-Eddy Simulation

This paper presents the numerical simulation of a model rocket combustor with continuously variable acoustic properties thanks to a variable-length injector tube. An unstable operating point is chosen and simulated with fully compressible three-dimensional unsteady simulations. Two wall conditions are used for the simulations, one is an adiabatic condition and the second one models heat losses through the chamber’s wall. The mechanism of the instability is studied for the adiabatic case, the structure of the acoustic mode is determined, and a balance on acoustic energy are performed. Then, the damping of the instability due to heat losses is observed.

Using Large Eddy Simulations to Understand Combustion Instabilities in Gas Turbines

This paper presents a study of the stability of a swirled premixed combustion chamber both with and without reaction using Large Eddy Simulation and a numerical solver able to handle complex geometries. It is shown that cold flow instabilities, which are observed in the non-reacting case (especially precessing vortex cores), are strongly damped in the reacting case, suggesting that these natural flow instabilities are not the first source of combustion instabilities in such configurations.

Numerical Investigation of Flow and Combustion in a Single Element GCH4/GOx Rocket Combustor

The flow and combustion in a GCH4/GOX single-element rocket combustor is analysed by several groups using different numerical models and tools. The tools and simulation setups vary with respect to modeling fidelity and computational expense. A short overview of the tools and the individual simulation setups is given. The focus of the paper is the comparison of the results obtained by the different groups as well as with experimental data. This encompasses the study of specific features of the combustor flow among the different simulations, as well as the validation with typical rocket engine design and performance parameters, such as wall heat flux and combsution pressure, gained from hot firing tests.

INFLUENCE OF HEAT TRANSFER AND MATERIAL TEMPERATURE ON COMBUSTION INSTABILITIES IN A SWIRL BURNER

The current work focuses on the large eddy simulation (LES) of combustion instability in a laboratory-scale swirl burner. Air and fuel are injected at ambient conditions. Heat conduction from the combustion chamber to the plenums results in a preheating of the air and fuel flows above ambient conditions. The paper compares two computations: In the first computation, the temperature of the injected reactants is 300 K (equivalent to the experiment) and the combustor walls are treated as adiabatic. The frequency of the unstable mode ( 635 Hz) deviates significantly from the measured frequency ( 750 Hz). In the second computation, the preheating effect observed in the experiment and the heat losses at the combustion chamber walls are taken into account. The frequency ( 725 Hz) of the unstable mode agrees well with the experiment. These results illustrate the impor- tance of accounting for heat transfer/losses when applying LES for the prediction of com- bustion instabilities. Uncertainties caused by unsuitable modeling strategies when using computational fluid dynamics for the prediction of combustion instabilities can lead to an improper design of passive control methods (such as Helmholtz resonators) as these are often only effective in a limited frequency range.

Delayed-time domain impedance boundary conditions (D-TDIBC)

Defining acoustically well-posed boundary conditions is one of the major numerical and theoretical challenges in compressible Navier–Stokes simulations. We present the novel Delayed-Time Domain Impedance Boundary Condition (D-TDIBC) technique developed to impose a time delay to acoustic wave reflection. Unlike previous similar TDIBC derivations (Fung and Ju, 2001–2004 [1], [2], Scalo et al., 2015 [3] and Lin et al., 2016 [4]), D-TDIBC relies on the modeling of the reflection coefficient. An iterative fit is used to determine the model constants along with a low-pass filtering strategy to limit the model to the frequency range of interest. D-TDIBC can be used to truncate portions of the domain by introducing a time delay in the acoustic response of the boundary accounting for the travel time of inviscid planar acoustic waves in the truncated sections: it gives the opportunity to save computational resources and to study several geometries without the need to regenerate computational grids. The D-TDIBC method is applied here to time-delayed fully reflective conditions. D-TDIBC simulations of inviscid planar acoustic-wave propagating in truncated ducts demonstrate that the time delay is correctly reproduced, preserving wave amplitude and phase. A 2D thermoacoustically unstable combustion setup is used as a final test case: Direct Numerical Simulation (DNS) of an unstable laminar flame is performed using a reduced domain along with D-TDIBC to model the truncated portion. Results are in excellent agreement with the same calculation performed over the full domain. The unstable modes frequencies, amplitudes and shapes are accurately predicted. The results demonstrate that D-TDIBC offers a flexible and cost-effective approach for numerical investigations of problems in aeroacoustics and thermoacoustics.

Numerical Investigation of Flow and Combustion in a Single-Element GCH4/GOx Rocket Combustor: Chemistry Modeling and Turbulence-Combustion Interaction

The flow and combustion in a GCH4/GOX single-element rocket combustor is analysed by several groups using different numerical models and tools. The tools and simulation setups vary with respect to modeling fidelity and computational expense. A short overview of the tools and the individual simulation setups is given. The focus of the paper is the comparison of the results obtained by the different groups as well as with experimental data. This encompasses the study of specific features of the combustor flow among the different simulations, as well as the validation with typical rocket engine design and performance parameters, such as wall heat flux and combsution pressure, gained from hot firing tests.