Large eddy simulation of high reynolds number jets with microjet injection

Large eddy simulations of two isothermal Mach 0.75 jets have been performed, one of a clean jet and one of the same jet fitted with eight equally spaced microjets. The microjets have a pressure ratio of 2.38, with a fully expanded Mach number of 1.19. The Reynolds number of the main jet in both simulations, based on the jet core velocity and diameter, is 1.3 million. The simulations were performed on a cylindrical, structured, multiblock mesh created for the clean round jet. The microjets are introduced as pressure inlet areas within the computational domain, so avoiding the complication of modelling the microjet feed pipes. Results of the clean jet simulation agree well with experimental data. The simulation shows the microjets penetrating into the jet core and disrupting the otherwise circular nature of the shear layer in the early flow development regions, though no change in mean flow variables is noticed by the end of the potential core. Two-point two-time correlation are performed on both cases and compared. The results show the microjets reduce the second and fourth order correlation amplitudes and turbulent lengthscales even at large axial locations downstream of the nozzle exit, where the effect of the microjets on the mean flow field is not present. This gives evidence as to how the microjets are able to reduce jet noise levels

LES of high speed jet flow from convergent-divergent rectangular S-bend ducts using synthetic inlet conditions

The effect of upstream duct curvature on the exhaust plume of a jet engine is further studied. Using synthetically created turbulence, improvements are made to the flow through out the S-bend validation case previously studied. The effect of a contracting 70° S-bend duct on the over-expanded exhaust plume emanating from a rectangular nozzle of aspect ratio 5.8:1 at a nozzle pressure ratio of 2.5 and Reynolds number of 7.61×105 is then studied. A modified version of the synthetic eddy method for creating artificial turbulence is initially validated. The validation of the Hydra CFD code is then expanded upon for an S-bend duct including both RANS and LES methodologies. For the combined S-bend and nozzle cases the total pressure gradients that were previously observed at the nozzle exit plane for k-ε RANS are also similarly observed using LES with synthetically created in flow turbulence thus confirming the existence of such features. The calculations were carried out using an unstructured, median-dual CFD solver with predominantly hexahedral elements containing approximately 175 million nodes

Large eddy simulation of a complete harrier aircraft in ground effect

This paper aims to demonstrate the viability of using the large eddy simulation (LES) CFD methodology to model a representative, complete STOVL aircraft geometry at touch down. The flowfield beneath such a jet-borne vertical landing aircraft is inherently unsteady. Hence, it is argued in the present work that the LES technique is the most suitable tool to predict both the mean flow and unsteady fluctuations. and, with further development and validation testing, this approach could be a replacement, and certainly a complementary aid, to expensive rig programmes. The numerical method uses a compressible solver on a mixed element unstructured mesh. Examination of instantaneous flowfield predictions from these LES calculations indicate close similarity with many flow features identified from ground effect flow visualisations, which are well known to be difficult to model using RANS-based CFD. Whilst significant further work needs to be carried out, these calculations show that LES could be a practical tool to model, for example, Hot Gas Ingestion for the Joint Strike Fighter aircraft

Computational study of jetlet structures in perforate silencers

Reynolds Averaged Navier Stokes allow for the steady state solutions while Large eddy simulation is a useful tool for predicting the spatial and temporal behavior of flow structures generated by complex geometries. However, many predictions suffer from poor grid resolution and initial conditions resulting in poor development of the initial jet shear layer and consequent incorrect prediction of critical flow behaviour. In this work, two 1mm thick perforate plate geometries of 23% and 40% porosity with 2mm diameter holes at an overall pressure ratio of 1.45 have been investigated. Results presented in this paper show the initial jetlet and fully merged jet flow-field to be sensitive to the porosity and the presence of partial holes around the circumference of the plate. The increase in porosity reduces the available entrainment flow, and increases the local jetlet interaction and resultant turbulence levels. This interaction fundamentally changes the flow structure from coherent vortex rings (found at low porosity) to a helical structure. The 2nd and 4th order spatiotemporal correlation Rij and Rij,kl are presented as evidence of the associated impact on acoustic source modeling

Large eddy simulation of a compressor cascade and the influence of spanwise domain

A controlled diffusion compressor cascade is studied using large eddy simulation (LES). The aim of this study is to assess the capability of LES to be used in an industrial context. The Reynolds number is approximately 700 000 based on chord length and inlet velocity. A 'thin-slice' representation of the cascade is used as the reference grid, and the influence of a narrow span is studied by comparison simulations with a domain that has a span five times larger than the thin-slice grid. While the instantaneous flow-fields of the thin-slice and wide-domain simulations are qualitatively similar, the thin-slice simulations suffer from flow confinement problems caused by the imposition of the narrow span. The non-unity axial velocity density ratio of the flow enforces the use of inviscid wall spanwise boundaries, which have a parasitic influence on the development of the flow in the thin-slice simulations. The resultant data obtained from the thin-slice simulations are therefore compromised and the computed loss estimation is considered unreliable. However, when comparing mean quantities such as surface pressure and boundary layer growth the narrow does give reasonable predictions. While the inviscid spanwise walls also affect the flow near the boundaries in the wide domain simulations, there is sufficient region of span from which reliable flow data and loss estimations can be obtained. For blade flows at off-design conditions, a span of 20 per cent of the blade chord is sufficient to give good agreement with experimental data. This incurs a computational cost that may be too high to incorporate parametric LES studies into the design cycle of turbomachinery components with current computers

Large eddy simulation of a controlled diffusion compressor cascade

In this research a Controlled Diffusion (CD) compressor cascade stator blade is simulated at a Reynolds number of similar to 700,000, based on inflow velocity and chord length, using Large Eddy Simulation (LES). A wide range of flow inlet angles are computed, including conditions near the design angle, and at high negative and positive incidence. At all inlet angles the surface pressure distributions are well-predicted by the LES. Near the design angle the computed suction side boundary layer thickness agrees well with experimental data, whilst the pressure side boundary layer is poorly predicted due to the inability of LES to capture natural boundary layer transition on the present grid. A good estimation of the loss is computed near the design angle, whilst at both high positive and negative incidences the loss is less well predicted owing to discrepancies between the computed and experimental boundary layer thickness. At incidences above the design angle a laminar separation bubble forms near the leading edge of the suction surface, which undergoes a transition to turbulence. Similar behaviour is noted on the pressure surface at negative incidence. At high negative incidence contra-rotating vortex pairs are found to form around the leading edge in response to an unsteady stagnation line across the span of the blade. Such structures are not apparent in time-averaged statistical data due to their highly-transient nature

The sub-grid-scale approach for modelling the impingement cooling flow in the combustor pedestal tile

The widely used gas turbine combustor double-walled cooling scheme relies on very small pedestals. In a combustor it is impractical for CFD to resolve each pedestal individually as that would require a very large amount of grid points and consequent excessive computation time. These pedestals can be omitted from the mesh and their effects captured on the fluid via a pedestal sub-grid scale (SGS) model. The aim is to apply the SGS approach, which takes into account the effects on pressure, velocity, turbulence and heat transfer, in an unstructured CFD code. The flow inside a 2-D plain duct is simulated to validate the pedestal SGS model and the results for pressure, velocity and heat transfer are in good agreement with the measured data. The conjugate heat transfer inside a 3-D duct is also studied to calibrate the heat source term of the SGS model due to the pedestals. The resolved flow in the combustor pedestal tile geometry is numerically investigated using RANS and LES in order to first assess the viability of the RANS and LES to predict the impinging flow and second to provide more validation data for the development of the SGS pedestal correlations. It is found that the complexity of such a flow, with high levels of curvature, impingement and heat transfer, poses a challenge to the standard RANS models. The LES provides more details of the impinging flow features. The pedestal model is then applied to the complete tile to replace the pedestals. The results are close to both the fully resolved CFD and the measurements. To improve the flow features in the impingement zone the first two rows were resolved with the mesh and combined with the SGS modelling for the rest of the tile, this gave optimum results of pressure, velocity and turbulence kinetic energy distribution inside the pedestal cooling tile

Computational study of aero-acoustic sources in perforate silencers

Reynolds Averaged Navier Stokes and Large Eddy Simulations of two perforate plates at a overall pressure ratio of 1.45 have been performed to allow analysis of the sensitivity of acoustic noise sources to porosity. Two geometries are presented: A 23% porosity and a 40% porosity 1mm plate with 2mm diameter holes. Results presented in this paper show the initial jetlet and fully merged jet flow-field to be sensitive to the porosity and the presence of partial holes around the circumference of the plate. The increase in porosity reduces the available entrainment flow, and increases the local jetlet interaction and resultant turbulence levels. This interaction fundamentally changes the flow structure from coherent vortex rings (found at low porosity) to a helical structure. The 2nd and 4th order spatio- temporal correlation Rij and Rij,kl are presented as suggested validation data for acoustic source modeling together with far-field noise spectra obtained via a Ffowcs-Williams & Hawkings surface integral method

Numerical predictions of turbulent underexpanded sonic jets using a pressure-based methodology

The objective of this work is to model underexpanded turbulent sonic jets. A pressure-based computational fluid dynamics methodology has been employed, incorporating extensions to handle high speed flows. A standard two-equation turbulence model is used, with an optional compressibility correction. Comparison with experimental jet centre-line Mach number showed the correct shock cell wavelength but a too rapid decay. The compressibility correction had no effect on the shock cell decay but increased the potential core length to give better agreement with experiment. Calculations for nozzle pressure ratios up to 30 showed the variation of Mach disc location in good agreement with experiment. For nozzle pressure ratios above 6, unsteady solutions were observed, emanating from the intersection of the Mach disc with the shear layer. Experimental work has identified similar large-scale instabilities; the peak mode of the prediction had a Strouhal number of 0.16, close to experimental values

High resolution simulations of high Reynolds number jets with microjet injection

Large eddy simulations have been performed of a Mach 0.9 jet at a Reynolds number of 1.3 million for both a clean and microjet injected configurations. Two numerical grids were used for the simulations differing in the number of azimuthal cells. The first with 720 cells and an azimuthal clustering near the microjet injection locations, and the other with 1440 cells and a uniform azimuthal cell spacing that matches the finest cell for the clustered case. The grids contained 100 million and 200 million cells respectively. A standard Smagorinsky sub-grid scale model was used together with a synthetic trip in the nozzle shear layer. The non-uniform grid with 720 cells azimuthally showed a variation in laminar to turbulent transition location that was a function of the clustering, with later transition in the coarser regions. However, this had little detrimental impact on mean velocity distributions further downstream. The results of the simulations were compared with PIV experimental data and good agreement of mean radial velocity and turbulent kinetic energy profiles were obtained. The microjets caused a deformation of the shear layer, reducing the radial location of peak turbulence kinetic energy in-line with the microjets. Additionally, the shear layer is translated away from the jet centreline between the microjets and becomes flat in the regions between the microjets. A Ffowcs Williams-Hawking technique was used to propagate the unsteady pressure fluctuations to the far-field. Spectral data at downstream and sideline observer locations indicated the presence of a high-frequency peak in the microjet case which is consistent with the size of the microjets. The microjets provide a blockage effect to the main jet and the peak is probably a shedding like behaviour. Overprediction of overall sound pressure levels by 6-8 dB was found when compared to the experimental data, however, the correct behaviour with observer angle was captured and more importantly the microjets showed a reduction in OASPL of around 2 dB across a range of angles, similar to the experiment results