Temporally and spatially resolved flow in a two-stage axial compressor. Part 2: Computational assessment

Fluid dynamics of turbomachines are complicated due to aerodynamic interactions between rotors and stators. It is necessary to understand the aerodynamics associated with these interactions in order to design turbomachines that are both light and compact as well as reliable and efficient. The current study uses an unsteady, thin-layer Navier-Stokes zonal approach to investigate the unsteady aerodynamics of a multi-stage compressor. Relative motion between rotors and stators is made possible by use of systems of patched and overlaid grids. Results have been computed for a 2 1/2-stage compressor configuration. The numerical data compares well with experimental data for surface pressures and wake data. In addition, the effect of grid refinement on the solution is studied

High Fidelity Simulation of Loss Mechanisms in Compressors

Further improvements in aero-engine efficiencies require better understanding of loss mechanisms. The rise of high performance computing is unlocking the potential of scale-resolving simulations for industrially relevant cases thus allowing new levels of simulation fidelity. As a result, previously unexplored effects of unsteadiness can be simulated and their impact on loss understood. The work undertaken in this thesis aims to establish a framework for accurate loss predictions using scale-resolving simulations and inform the field with regards to the effects of unsteadiness on loss mechanisms within the multi-stage compressors. The lack of computational requirements for accurate industrial simulations lead to inconsistent loss predictions even for scale-resolving simulations depending on the chosen convergence criteria. This work studies aspects of loss generation by employing two test-cases: Taylor-Green vortex and compressor cascade subjected to freestream turbulence. The results show that both resolving local entropy generation rate and capturing the inception and growth of instabilities are critical to accuracy of loss prediction. In particular, the interaction of free-stream turbulence at the leading-edge and development of instabilities in the laminar region of the boundary layer are found to be important. These two outcomes allow for a formulation of resolution criteria that ensure accurate loss predictions for compressor flows. One of the major sources of uncertainty in the current simulation methods for compressor flows is the level of unsteadiness and its impact on loss This work makes a series of steps towards understanding the nature and the origin of unsteadiness within multi-stage machines and investiages the impact of gapping on mid-span compressor loss. It is found that freestream turbulence levels rise significantly as the size of the rotor-stator axial gap is reduced. This is because of the effect of axial gap on turbulence production mechanisms, which amplify at smaller axial gaps and drive increases in dissipation and loss. This effect is found to raise loss by between 5.5 - 9.5\% over the range of conditions tested here. This effect was found to significantly outweigh the beneficial effects of wake recovery on loss.Financial support for this work was provided by the Whittle Laboratory and the University of Cambridge through the Denton Scholarship fund and the CDT in Gas Turbine Aerodynamics, funded by the EPSRC $EP/L015943/1$

Prediction of the unsteady turbulent flow in an axial compressor stage. Part 2: Analysis of unsteady RANS and LES data

This paper presents the analysis of URANS and LES database in a stage of an axial subsonic compressor. Details about numerical methods and comparison with experiments can be found in a companion paper. The analysis here focuses on the transition processes that take place in the rotor and stator rows. In the rotor, LES and URANS show that transition develops at mid-chord and is induced by the adverse pressure gradient. In the stator, the flow behavior is more complex since the transition is influenced by the rotor passing wakes, a laminar separation bubble on the suction side and the accumulation of rotor wakes on the pressure side. The analysis also investigates the unsteady flow patterns at the rotor/stator interface, from mid-span to the casing. In the tip region, LES shows the development of frequencies that are not correlated to the blade passing frequency, while URANS only predicts multiples of the blade passing frequency

Evaluation de la méthode SAS sur un rotor de compresseur haute-pression

International audienceThe main design objectives of a high pressure compressor are the aerodynamic efficiency and the operating range (e.g. the surge margin). Those quantities are impacted by secondary and leakage flows occurring in the blade passage such as corner separation or stall and tip leakage flows. The turbulence modeling influences strongly the prediction of the overall performances. The aims of the present study were (i) the validation of the combination of the SAS approach with the DRSM turbulence model by comparison to experimental data, especially to laser measurements in the tip of a rotor of a high pressure compressor and (ii) the discussion of the flow prediction improvements with respect to turbulence approaches classically used in CFD and industry: URANS simulations and standard SAS simulation i.e. combined with SST turbulence model. The SAS results are compared to experimental data and to URANS results (SST and DRSM). Only the simulations with IGV wakes predict the velocity fluctuations near tip gap, from the leading edge. Concerning the time-averaged performances, the stagnation pressure losses are slightly overestimated by SAS, especially with DRSM model. This is due to an amplification of the hub corner separation. Moreover, the isentropic efficiency is very sensitive to the SAS approach and to the turbulence model. The spectral analysis shows that the prediction of the amplitude and frequencies of the power spectral density of static pressure is improved using the SAS approach instead of URANS one. The SAS approach leads to PSD similar to ZDES, especially with the DRSM model. Thus, the SAS-DRSM is able to well predict the tip leakage flow with the fine mesh. Nevertheless, this approach amplifies the hub corner separation leading to a strong underestimation of overall performances

Performance Analysis of an Annular Diffuser Under the Influence of a Gas Turbine Stage Exit Flow

In this investigation the performance of a gas turbine exhaust diffuser subject to the outlet flow conditions of a turbine stage is evaluated. Towards that goal, a fully three-dimensional computational analysis has been performed where several turbine stage-exhaust diffuser configurations have been studied: a turbine stage with a shrouded rotor coupled to a diffuser with increasing divergence angle in the diffuser, and a turbine stage with an unshrouded rotor was also considered for the exhaust diffuser performance analysis. The large load of this investigation was evaluated using a steady state numerical analysis utilizing the "mixing plane" algorithm between the rotating rotor and stationary stator and diffuser rows. Finally, an unsteady analysis is performed on a turbine stage with an unsrhouded rotor coupled to an annular exhaust diffuser with an outer wall opening angle of 18°. It has been found that the over the tip leakage flow in the unshrouded rotor emerges as a swirling wall jet at the upper wall of the diffuser. When using the turbine with the shrouded rotor no wall jet was observed, making the flow at the entrance to the diffuser "quasi-uniform". The maximum opening angle of the diffuser upper wall achieved before the diffuser stalls was 12° with a static pressure recovery coefficient of Cp = 0.293. When the wall jet was observed, diffuser opening angles of 18° were possible with a static pressure recovery of Cp = 0.365. Consequently the wall jet energizes the diffuser upper wall boundary layer flow, allows for higher static pressure recovery levels and postpones diffuser stall. By altering the speed of the rotor the effect of the swirl in the turbine exit plane on the performance of the diffuser was explored. In the case where the wall jet was absent the diffuser recovers more pressure when the inlet is swirl-free. In this case the performance of the diffuser is independent on whether the turbine exit flow has co or counter swirl. In the presence of the wall jet, higher static pressure recovery was achieved when the wall jet was in co-swirl and the core flow at a slightly counter-swirl direction. This observation was more pronounced when larger diffuser upper wall opening angles were considered. In the unsteady analysis it was found that the wall jet axial velocity and swirl intensities pulsate with the relative position of the rotor to the stator. The wall jet is always co-swirling while the core flow is counter-swirling. Moreover, the wall jet does not penetrate the diffuser boundary layer as deeply as was observed in the steady state case and flow separation occurs at the upper endwall corner of the diffuser. Furthermore the performance of the diffuser shows a periodic variation that seems to depend on the relative position of the rotor to the stator. The averaged pressure recovery coefficient is Cp = 0.321 which is 11.0 % less than predicted in the steady state case

Near-Wall Flow in Turbomachinery Cascades—Results of a German Collaborative Project

This article provides a summarizing account of the results obtained in the current collabora-tive work of four research institutes concerning near-wall flow in turbomachinery. Specific questions regarding the influences of boundary layer development on blades and endwalls as well as loss mech-anisms due to secondary flow are investigated. These address skewness, periodical distortion, wake interaction and heat transfer, among others. Several test rigs with modifiable configurations are used for the experimental investigations including an axial low speed compressor, an axial high-speed wind tunnel, and an axial low-speed turbine. Approved stationary and time resolving measurements techniques are applied in combination with custom hot-film sensor-arrays. The experiments are complemented by URANS simulations, and one group focusses on turbulence-resolving simulations to elucidate the specific impact of rotation. Juxtaposing and interlacing their results the four groups provide a broad picture of the underlying phenomena, ranging from compressors to turbines, from isothermal to non-adiabatic, and from incompressible to compressible flows.The investigations reported in this article were conducted within the framework of the joint research project “Near-Wall Flow in Turbomachinery Cascades” which was funded and supported by the Deutsche Forschungsgemeinschaft (DFG) under grant number PAK 948. The responsibility for the contents of this publication lies entirely by the authors.Peer ReviewedPostprint (published version

Influence of cavity flow on turbine aerodynamics

In order to deal with high temperatures faced by the components downstream of the combustion chamber, some relatively cold air is bled at the compressor. This air feeds the cavities under the turbine main annulus and cool down the rotor disks ensuring a proper and safe operation of the turbine. This thesis manuscript introduces a numerical study of the effect of the cavity flow close to the turbine hub on its aerodynamic performance. The interaction phenomena between the cavity and main annulus flow are not currently fully understood. The study of these phenomena is performed based on different numerical approaches (RANS, LES and LES-LBM) applied to two configurations for which experimental results are available. A linear cascade configuration with an upstream cavity and various rim seal geometries (interface between rotor and stator platform) and cavity flow rate available. A rotating configuration that is a two stage turbine including cavities close to realistic industrial configurations. Additional losses incurred by the cavity flow are measured and studied using a method based on exergy (energy balance in the purpose to generate work)

Computational methods for internal flows with emphasis on turbomachinery

Current computational methods for analyzing flows in turbomachinery and other related internal propulsion components are presented. The methods are divided into two classes. The inviscid methods deal specifically with turbomachinery applications. Viscous methods, deal with generalized duct flows as well as flows in turbomachinery passages. Inviscid methods are categorized into the potential, stream function, and Euler aproaches. Viscous methods are treated in terms of parabolic, partially parabolic, and elliptic procedures. Various grids used in association with these procedures are also discussed