Single-stage experimental evaluation of boundary layer bleed techniques for high lift stator blades. Compressors design

Design and evaluation of stator blades for use in boundary layer bleed techniques in axial flow compresso

Single-stage experimental evaluation of boundary layer blowing techniques for high lift stator blades. 1 - Compressor design

Boundary layer blowing techniques for high lift stator blades in axial flow compressor

Body Force Modeling of Axial Turbomachines Without Calibration

This dissertation proposes a new method of modelling turbomachinery blade boundary layer and shock losses using the body force method. Body force methods are used to model fan/compressor performance at a lower computational cost than unsteady Reynolds-Averaged Navier-Stokes (URANS) computations in non-uniform inflows. Most loss modelling approaches in the literature require calibration. Some recent work has shown the use of non-calibrated methods for entropy generation calculations. However, recent non-calibrated methods cannot estimate flow losses with boundary layer separation. In this dissertation, an artificial neural network has been developed and trained to analytically relate the blade geometry and flow regime to the boundary layer momentum thickness at the trailing edge. The trailing edge momentum thickness is used in a body force loss model that accounts for the relative total pressure drop. This model is capable of predicting the loss at off-design conditions. The accuracy of the model is over 90$\%$ in 2D cascades. The model is then applied to the NASA rotor 67 compressor blade row. The model captures the high entropy generation near the tip region for uniform and non-uniform inflows. For non-uniform inflow, it predicts the isentropic efficiency to within 1$\%$ compared to a URANS computation

Unsteady aerodynamics of blade rows

The requirements placed on an unsteady aerodynamic theory intended for turbomachinery aeroelastic or aeroacoustic applications are discussed along with a brief description of the various theoretical models that are available to address these requirements. The major emphasis is placed on the description of a linearized inviscid theory which fully accounts for the affects of a nonuniform mean or steady flow on unsteady aerodynamic response. Although this linearization was developed primarily for blade flutter prediction, more general equations are presented which account for unsteady excitations due to incident external aerodynamic disturbances as well as those due to prescribed blade motions. The motivation for this linearized unsteady aerodynamic theory is focused on, its physical and mathematical formulation is outlined and examples are presented to illustrate the status of numerical solution procedures and several effects of mean flow nonuniformity on unsteady aerodynamic response

The effect of reaction on compressor performance

Reaction is the fundamental parameter by which the asymmetry of the velocity triangle of a stage is set. Little is understood about the effect that reaction has on either the efficiency or the operating range of a compressor. A particular difficulty in understanding the effect of reaction is that the rotor and stator have a natural asymmetry caused by the centrifugal effects in the rotor boundary layer, being much larger than those in the stator boundary layer. In the thesis a novel approach has been taken: McKenzie’s ‘linear repeating stage’ concept is used to remove the centrifugal force effects. The centrifugal effects are then reintroduced as a body force. This allows the velocity triangle effect and centrifugal effect to be decoupled. The ability to accurately decouple these two asymmetries has led to a number of major findings. The thesis shows the surprising result that, depending on how the solidity of the stage is set, 50% reaction can either result in the maximum, or the minimum, profile loss. When the solidity is set by the shape factor of the suction-surface boundary layer at the blade trailing-edge, and conventional levels of design work coefficient (Ψ=0.44) and flow coefficient (Φ=0.60) are set, the profile loss becomes independent of reaction. When the centrifugal effects are removed, 50% reaction is shown to minimise endwall loss, maximise stage efficiency and maximise operating range. When the centrifugal effects are reintroduced, the compressor with the maximum design efficiency is found to rise in reaction by 5% (from 50% reaction to 55% reaction) and the compressor with the maximum operating range is found to rise in reaction by 15% (from 50% reaction to 65% reaction). In a real multistage compressor there is often a requirement for axial flow at the inlet and exit the compressor. This naturally results in high reaction. In the central stages of the compressor, it is possible to maximise the stage efficiency by reducing the reaction to 55%. This is done by raising the interstage swirl through the first stage and dropping it through the last stage. It is shown that if a 10 stage compressor, which originally had a constant stage reaction of 75%, was rebladed so that the central 8 stages had 55% reaction, then the overall design efficiency would rise by 0.58%

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 $EP/L015943/1$

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

Blade row interaction in radial turbomachines

A computational study has been performed to investigate the effects of blade row interaction on the performance of radial turbomachines, which was motivated by the need to improve our understanding of the blade row interaction phenomena for further improvement in the performance. High-speed centrifugal compressor stages with three settings of radial gap are configured and simulated using a three-dimensional Navier-Stokes flow method in order to investigate the impact of blade row interaction on stage efficiency. The performance predictions show that the efficiency deteriorates if the gap between blade rows is reduced to intensify blade row interaction, which is in contradiction to the general trend for stage axial compressors, hi the compressors tested, the wake chopping by diffuser vanes, which usually benefits efficiency in axial compressor stages, causes unfavourable wake compression through the diffuser passages to deteriorate the efficiency. Similarly, hydraulic turbine stages with three settings of radial gap are simulated numerically. A new three-dimensional Navier-Stokes flow method based upon the dual-time stepping technique combined with the pseudo-compressibility method has been developed for hydraulic flow simulations. This method is validated extensively with several test cases where analytical and experimental data are available, including a centrifugal pump case with blade row interaction. Some numerical tests are conducted to examine the dependency of the flow solutions on several numerical parameters, which serve to justify the sensitivity of the solutions. Then, the method is applied to performance predictions of the hydraulic turbine stages. The numerical performance predictions for the turbines show that, by reducing the radial gap, the loss generation in the nozzle increases, which has a decisive influence on stage efficiency. The blade surface boundary layer loss and wake flow mixing loss, enhanced with a higher level of flow velocity around blading and the potential flow disturbances, are responsible for the observed trend