Xuejun Fan

Large-Eddy Simulation of Time Evolution and Instability of Highly Underexpanded Sonic Jets

High-pressure jet injection into quiescent air is a challenging fluid dynamics problem in the field of aerospace engineering. Although plenty of experimental, theoretical, and numerical studies have been conducted to explore this flow, there is a dearth of literature detailing the flow evolution and instability characteristics, which is vital to the mixing enhancement design and jet noise reduction. In this paper, a density-based solver for compressible supersonic flow, astroFoam, is developed based on the OpenFOAM library. Large-eddy simulations of highly underexpanded jets with nozzle pressure ratios from 5.60 to 11.21 at a Reynolds number around 105 are carried out with a high-resolution grid. A grid-convergence study has been conducted to confirm the fidelity of the large-eddy simulation results. The large-eddy simulation results have also been validated against available literature data in terms of the time-averaged near-field properties of underexpanded jets. The turbulent transition processes are r...

Full-scale Detached Eddy Simulation of kerosene fueled scramjet combustor based on skeletal mechanism

Large Eddy Simulation (LES) of kerosene fueled scramjet combustor is generally scare in the literature, due mainly to the formidable computational cost arisen by complex kerosene mechanism. In this study, the skeletal reduction of a detailed reaction mechanism (2185 species/8217 steps) of aviation kerosene is conducted using directed relation graph with error propagation and sensitivity analysis (DRGEPSA) method, resulting a skeletal mechanism consisting of 39 species/153 elemental reactions for China Daqing RP-3 aviation kerosene. The comparisons of adiabatic flame temperature, total heat release, ignition delay and laminar flame speed predicted by the skeletal mechanism show an overall good accordance with the original detailed reaction mechanism. Then a three-dimensional Detached Eddy Simulation (DES) modeling based on the skeletal kerosene mechanism is employed for the numerical analysis of a full-scale scramjet combustor, which has been experimentally tested in a long-time direct connect supersonic combustion test platform (abbreviated as DTZ) assembled in Chinese Academy of Sciences (CAS). Pressure and heat flux measurement systems are attached to the combustor assembly to monitor the real-time combustion performance and provide validation data for the numerical modeling. Three cases with fuel equivalence ratios from 0.8, 1.0 to 1.2 and the same crossflow conditions at Mach 2.0 are modeled. The time-averaged static pressure and heat flux are in generally good agreement with the experiment with the peak heat flux slightly underpredicted. The instantaneous and/or time-averaged pressure, momentum, temperature and turbulence fields, which are difficult to be measured, are analyzed to reveal the main flow and combustion physics, especially those related to the flame distribution and holding. The combustion is identified as in ramjet mode for the investigated cases. With the increasing of fuel equivalence ratio, the shock train propagates upstream in the isolator and the interaction between the upstream and downstream combustion assumes different patterns. c 2015, AIAA American Institute of Aeronautics and Astronautics. All rights reserved.(132 refs)

Numerical Investigation on Flame Stabilization in DLR Hydrogen Supersonic Combustor with Strut Injection

ABSTRACT Flame stabilization in the DLR hydrogen supersonic combustor with strut injection was numerically investigated by using an in-house large eddy simulation code developed on the OpenFoam platform. To facilitate the comparison and analysis of various hydrogen oxidation mechanisms with different levels of mechanism reduction, the proposed 2D calculation model was validated against both the 3D simulation and the experimental data. The results show that the 2D model can capture the DLR flow and combustion characteristics with satisfactorily quantitative accuracy and significantly less computational load. By virtue of the flow visualization and the analyses of species evolution and heat release, the supersonic combustion in the DLR combustor can be divided into three stages along the streamwise direction: the induction stage where ignition occurs and active radicals are produced, the transition stage through which radicals are advected to the downstream, and the intense combustion stage where most heat release occurs. Furthermore, the sensitivity analysis of key reaction steps identifies the important role of chain carrying and heat release reactions in numerically reproducing the three-stage combustion stabilization mode in the DLR combustor.

Comparative Study of Elliptic and Round Scramjet Combustors Fueled by RP-3

To explore the combustion performance of nonrectangular supersonic combustors, the flow and combustion characteristics in the round and round-to-elliptic shape-transition scramjet combustors were c...

A low-dissipation scheme based on OpenFoam designed for large eddy simulation in compressible flow

To reduce the numerical dissipation in compressible flow modeling, a low-dissipation compressible solver is developed for large eddy simulation based on the original compressible solver rhoCentralFoam within the framework of an open source computational fluid dynamics package OpenFOAM. In rhoCentralFoam, the central-upwind scheme of Kurganov and Tadmor is applied to capture flow discontinuities, but its dissipation is too strong to resolve fine turbulence structures under finite mesh resolutions. The current lowdissipation solver adopts a new hybrid scheme, which combines the dissipative KurganovTadmor scheme with the nondissipative central scheme. By aid of a shock sensor, the dissipative scheme is used to capture the flow discontinuities near shock waves and the central scheme is used to resolve the turbulence structures in the smooth flow area. In the framework of unstructured mesh, the central scheme is extended from second order to forth order, which greatly reduces the dispersion error and weakens the oscillations near flow discontinuities. To improve the numerical stability of the central scheme, the skewsymmetric form of the convective term is adopted to preserve the local kinetic energy and maintain the self-stability of central scheme without adding an explicit dissipative term. In addition, a low-storage third-order TVD Runge-Kutta method for temporal discretization is newly implemented in the low-dissipation solver to further reduce the numerical dissipation. A series of benchmark flow problems, such as Sod shock tube test, Shu-Osher problem, Green-Taylor vortex evolution, and wall-bounded turbulence generation based on synthetic eddy method, are computed and compared to examine the low-dissipation solver’s capability in capturing flow discontinuities as well as resolving fine turbulence structures. The accuracy and stability of the low-dissipation solver are further validated against experimental data for a scramjet model with a supersonic airstream passing over the flame holder structure.

A comparative study of elliptical and round scramjet combustors by Improved Delayed Detached Eddy Simulation

To explore the combustion performance of non-rectangular type supersonic combustors, the flow and combustion characteristics in round and round-to-elliptic shape-transition (RdEST) supersonic combustors under the same configurations of flight Mach number and fuel equivalence ratio were compared based on modeling results. The fuel equivalence ratio is maintained the same as 0.8, while two inlet Mach numbers of 2.5 and 3.0 both corresponding to a real flight Mach number of 6.5 are tested. To alleviate the strict requirements on wall-normal and -parallel grid spacing, Improved Delayed Detached Eddy Simulation (IDDES) is employed in this study to enable an automatic choice of RANS or LES mode depending on the local boundary layer thickness and turbulent viscosity. To reduce the computational cost of stiff kerosene oxidation chemistry, a total of four versions of skeletal mechanisms (respectively 48s/197r, 39s/153r, 28s/92r and current 19s/54r) have been developed based on the detailed 2815s/8217r Dagaut mechanism by using a highly efficient and reliable directed relation graph with error propagation and sensitivity analysis (DRGEPSA) method together with manual path analysis. Although the mechanism size has been significantly reduced, key kinetic properties such as adiabatic flame temperature, heat release rate, ignition delay and laminar flame speed all agree well with the original detailed mechanism. The static pressure along streamwise direction is compared with the measurement to validate the modeling results. Two key aspects are well predicted, i.e. the pressure ratio and the initial pressure rise location, indicating that the flame anchoring location and the distribution of wave structures inside the combustor are close to the actual situation. Then the aerodynamic fields are analyzed for the round and elliptic combustors to compare their flow, mixing and combustion related flow structures. The three-dimensional wave structures inside the elliptic combustor are firstly shown to reveal the influence of non-axisymmetric cross-section on the shock train and Mach field. Especially the time evolution of the flame region is analyzed, and dominant flame modes are extracted by the aid of proper orthogonal decomposition (POD) method. 漏 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

Development of a radiative heating facility for studying flow and heat transfer in hydrocarbon-cooled structures

An experimental facility was designed to simulate the heat exchange between the hot gas and the fuel-cooled wall in a scramjet combustor. Thermal radiation from an electrically heated graphite plate is employed to unilaterally heat up a multi-channeled cooling plate. A maximum heat flux of over 0.8 MW/m2 was achieved for an effective heating area up to 1000 mm × 40 mm. Precise control of the back pressure of a coolant (up to 5 MPa) in a unique way was also demonstrated. With this facility, studies of flow and heat transfer in hydrocarbon-cooled structures can be performed under a well-controlled manner.

Development of zone flamelet model for scramjet combustor modeling

To alleviate the huge computational cost in supersonic combustor modeling and to improve the accuracy of traditional unsteady flamelet model, a zone flamelet is proposed. The main idea of zone flamelet is to divide the whole turbulent combustion field into a finite number of control zones and the chemical status in each zone is represented by a single flamelet. With proper zone division, the scattering of variables over the mixture fraction space is in controllable small, thus the representative flamelet approaches the real scalar distribution. The flamelets exchange information through flux-conserved convection when across the zone boundary, thus the flamelet variables can be transport from upstream to downstream in a flow manner. Although one additional mixture fraction space is resolved, great computational cost is still saved because the zone division in physical space is much coarser than the flow simulation mesh. A simple historical statistics approach is proposed to estimate the representative temperature, in order to further alleviate the computational cost in solving the flamelet temperature equation usually with numerous sub-models for non-adiabatic terms, e.g. radiation and wall heat loss. The zone flamelet model is then applied to model a scramjet combustor operated at a flight Mach number of 6.5 and a fuel equivalence ratio of 0.8. The performance of zone flamelet model in highly non-equilibrium supersonic combustion is compared with the traditional PaSR model.

Experimental Studies of Film Cooling in Supersonic Combustors

The current work is an experimental investigation of the dependence of film-cooling effectiveness on the injection angle, mass flux and injection temperature in supersonic combustors. The mainstream Mach number is 2.5, and the coolant was injected with sonic speed. The total temperature of the mainstream is 1500K, and for the injection it ranges from 300K to 1060K. Three injection angle is respectively 0, 43, 137 . The coolant mass flux ranges from 1% to 6% of the mainstream mass flux, and the mainstream mass flux is 2.5kg/s. The results show that a smaller injection angle has a better performance in film cooling effectiveness and indicate that an increase in film-cooling effectiveness with the increase of the coolant mass flux. The influence of coolant temperature is more complex. The rising of coolant temperature reduces the film-cooling effectiveness without kerosene combustion, while has not significant difference with kerosene combustion.

Development of a radiative heating system for studies of heat transfer in fuel-cooled structures

A radiative heating system was designed to simulate the heat exchange between the hot gas and the fuel-cooled structure in a scramjet combustor. A flat-plate cooling structure was heated unilaterally using thermal radiation from an electrically-heated graphite plate. The system was designed to work at a maximum heat flux of 1.5 MW/M-2 for the effective heating area up to 1000mmX40mm.