Feng Wang

Lip Stall Suppression in Powered Intakes

This work describes a computational study into lip stall in subsonic civil aircraft intakes and its alleviation by action of the fan. Beyond a certain flow incidence, the lower lip of a civil aircraft intakes stalls. This phenomenon causes entropy and vortical distortions at the fan face. Consequently, it has detrimental effects on vibration levels and performance of the low-pressure compressor system. The most important parameters influencing lip separation are the flight Mach number, the Reynolds number, and altitude. Fully three-dimensional simulations have been performed on a flight intake in current service for which the experimental data are available. Steady and time-resolved simulations have been performed. Distortion coefficients have been evaluated as functions of incidence and have been compared with experimental results. A comparison between an isolated intake and a powered intake shows that the fan stage has the beneficial effect of increasing tolerance to flow incidence and decreasing distor...

Virtual Gas Turbines: Geometry and Conceptual Description

This paper documents the construction of a Virtual Engine, with particular reference to its geometry and conceptual description.. The phrase Virtual Engine denotes a system which allows simulations of whole gas-turbine engines to be undertaken at any desired level of fidelity or physical modeling. In order to be of any practical use, the system must allow the computations to be setup as automatically as possible and needs to contain provisions for the exchange of boundary data between adjacent computational domains — e.g. solid-gas interfaces. The paper illustrates the application of the system to the representation and analysis of a modern commercial turbofan engine.Copyright

Mesh Generation for Turbomachinery Blade Passages with Three-Dimensional Endwall Features

Turbomachinery blade passages are conventionally meshed by sweeping a mesh from a reference stream surface to other stream surfaces along the blade span. This approach is widely used in the gas-tur...

Simulation of Multi-Stage Compressor at Off-Design Conditions

Computational Fluid Dynamics (CFD) has been widely used for compressor design, yet the prediction of performance and stage matching for multi-stage, high-speed machines remain challenging. This paper presents the authors’ effort to improve the reliability of CFD in multistage compressor simulations. The endwall features (e.g. blade fillet and shape of the platform edge) are meshed with minimal approximations. Turbulence models with linear and non-linear eddy viscosity models are assessed. The non-linear eddy viscosity model predicts a higher production of turbulent kinetic energy in the passages, especially close to the endwall region. This results in a more accurate prediction of the choked mass flow and the shape of total pressure profiles close to the hub. The non-linear viscosity model generally shows an improvement on its linear counterparts based on the comparisons with the rig ∗Corresponding author (feng.wang@eng.ox.ac.uk) TURBO-17-1165 Wang 1 data. For geometrical details, truncated fillet leads to thicker boundary layer on the fillet and reduced mass flow and efficiency. Shroud cavities are found to be essential to predict the right blockage and the flow details close to the hub. At the part speed the computations without the shroud cavities fail to predict the major flow features in the passage and this leads to inaccurate predictions of massflow and shapes of the compressor characteristic. The paper demonstrates that an accurate representation of the endwall geometry and an effective turbulence model, together with a good quality and sufficiently refined grid result in a credible prediction of compressor matching and performance with steady state mixing planes. Introduction Three-dimensional CFD simulations of multiple blade rows were developed by Adamczyk [1] using the average passage approach. Each blade row was modelled seperately and only one passage of each blade row was required in the simulation. Blade row interactions were evaluated by including additional force terms in the equations. This approach was further simplified by Denton and the resulting approach was termed as the mixing plane approach [2]. The mixing plane assumes a mixed-out state at the bladerow interface by conserving mass, momentum and energy fluxes. The drawback of this approach is that the unsteady blade interactions are removed. Despite its simplicity, the performance of turbomachines is predicted reasonably well using mixing planes. Consequently, steady state RANS simulations have been the work-horse of turbomachinery design. The capability of RANS simulations in assisting turbomachinery design was already demonstrated by early researchers [3, 4], even if limitations in computer power, meshing techniques and physical modelling restricted the spatial resolution as well as the amount of geometric detail present in the simulations. Typical examples of features omitted from the computational models in the early days were blade fillets, stator shrouds, Variable Stator Vane (VSV) penny gaps etc. Typical grids used were H-grids with pinched tip gaps. Fully meshed tip gaps were introduced and comparisons with pinched tip gaps were summarized by Denton [5]. As the meshing techniques evolved, multiblock structured meshes were introduced into the turbomachinery design and allowed the generation of optimal grids on the blade-to-blade section [6,7]. However, geometrical approximations (e.g. truncated fillets) are still commonplace because of the limitations in the way the geometries of the endwalls and of the blades are represented in the mesh technique. The effect of endwall geometric features is important in predicting the endwall flows and contributes to setting the flow capacity. One of the most important endwall features for modern machines is the shroud cavities. A typical configuration is shown in Fig. 1. Shabbir et al [8] studied the effect of the hub leakage flow on high speed compressor rotors and the pressure deficit close to the hub was well captured by including the hub leakage flow. Wellborn and Okashi [9] reported the effect of shroud cavities on the performance of multi-stage compressor experimentally. Their data showed that the shroud cavities have considerable negative impact on the compressor performance. In their numerical study, the shroud cavities were not meshed as CFD domains, instead they were introduced through a Knife-to-Knife (K2K) model [10]. Another important element in the construction of steady models for turbomachinery is the turbulence model. The flows in multistage compressor passages are highly viscous and three-dimensional around endwall regions. Furthermore the flow is prone to separation due to strong adverse pressure gradient even at design conditions. Predicting the complex flow features TURBO-17-1165 Wang 2

Virtual Gas Turbine: Pre-Processing and Numerical Simulations

The design process of a gas turbine engine involves interrelated multi-disciplinary and multi-fidelity designs of engine components. Traditional component-based design process is not always able to capture the complicated physical phenomenon caused by component interactions. It is likely that such interactions are not resolved until hardware is built and tests are conducted. Component interactions can be captured by assembling all these components into one computational model. Nowadays, numerical solvers are fairly easy to use and the most time-consuming (in terms of man-hours) step for large scale gas turbine simulations is the preprocessing process. In this paper, a method is proposed to reduce its time-cost and make large scale gas turbine numerical simulations affordable in the design process. The method is based on a novel featured-based in-house geometry database. It allows the meshing modules to not only extract geometrical shapes of a computational model and additional attributes attached to the geometrical shapes as well, such as rotational frames, boundary types, materials, etc. This will considerably reduce the time-cost in setting up the boundary conditions for the models in a correct and consistent manner. Furthermore, since all the geometrical modules access to the same geometrical database, geometrical consistency is satisfied implicitly. This will remove the time-consuming process of checking possible mismatching in geometrical models when many components are present. The capability of the proposed method is demonstrated by meshing the whole gas path of a modern three-shaft engine and the Reynold’s Averaged Navier-Stokes (RANS) simulation of the whole gas path.© 2016 ASME

Automated Hex Meshing for Turbomachinery Secondary Air System

In this paper, we present a novel process of creating hexahedral meshes for the turbomachinery secondary air system. The meshing process is automated from the geometry import to the mesh setup and requires minimum human interventions. The core of the process is a hexahedral meshing algorithm with a boundary layer mesh automatically created. The hex meshing algorithm combines the pave-sweep, general sheet-insertion and a novel technique which creates the boundary layer mesh by carefully placing, maintaining and dicing a buffer layer around a geometry. After the mesh is created, relevant boundary conditions for the mesh are also assigned automatically. The whole meshing process is systematically automated and has the potential to considerably reduce the time cost in meshing the turbomachinery secondary air system.

Fan Similarity Model for the Fan-Intake Interaction Problem

Very high bypass ratio turbofans with large fan tip diameter are an effective way of improving the propulsive efficiency of civil aero-engines. Such engines, however, require larger and heavier nacelles, which partially offset any gains in specific fuel consumptions. This drawback can be mitigated by adopting thinner walls for the nacelle and by shortening the intake section. This binds the success of very high bypass ratio technologies to the problem of designing an intake with thin lips and short diffuser section, which is well matched to a low speed fan. Consequently, the prediction of the mutual influence between the fan and the intake flow represents a crucial step in the design process. Considerable effort has been devoted in recent years to the study of models for the effects of the fan on the lip stall characteristics and the operability of the whole installation. The study of such models is motivated by the wish to avoid the costs incurred by full, three-dimensional (3D) computational fluid dynamics (CFD) computations. The present contribution documents a fan model for fan–intake computations based on the solution of the double linearization problem for unsteady, transonic flow past a cascade of aerofoils with finite mean load. The computation of the flow in the intake is reduced to a steady problem, whereas the computation of the flow in the fan is reduced to one steady problem and a set of solutions of the linearized model in the frequency domain. The nature of the approximations introduced in the fan representation is such that numerical solutions can be computed inexpensively, while the main feature of the flow in the fan passage, namely the shock system and an approximation of the unsteady flow encountered by the fan are retained. The model is applied to a well-documented test case and compares favorably with much more expensive 3D, time-domain computations.