CFD Laboratory

United States

CFD Laboratory

United States
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Carrion M.,University of Liverpool | Carrion M.,CFD Laboratory | Woodgate M.,University of Liverpool | Woodgate M.,CFD Laboratory | And 6 more authors.
AIAA Journal | Year: 2015

This work explores the breakdown of the wake downstream of the Model Experiments in Controlled Conditions Project (known as the MEXICO project) wind-turbine rotor and assesses the capability of computational fluid dynamics in predicting its correct physical mechanism. The wake is resolved on a fine mesh able to capture the vortices up to eight rotor radii downstream of the blades. At a wind speed of 15 m/s, the main frequency present in the computational fluid dynamics signals for up to four radii was the blade-passing frequency (21.4 Hz), where the vortex cores fall on a perfect spiral. Between four and five radii downstream, higher-frequency content was present, whichindicated the onset of instabilities and results in vortex pairing. The effect of modeling a 120 deg azimuthally periodic domain and a 360 deg three-bladed rotor domain was studied, showing similar predictions for the location of the onset of instabilities. An increased frequency content was captured in the latter case. Empirical and wake models were also explored, they were compared with computational fluid dynamics, and a combination of kinematic and field models was proposed. The obtained results are encouraging and suggest that the wake instability of wind turbines can be predicted with computational fluid dynamics methods, provided adequate mesh resolution is used. Copyright © 2014 by the authors.

Dehaeze F.,University of Liverpool | Dehaeze F.,CFD Laboratory | Barakos G.N.,University of Liverpool | Barakos G.N.,CFD Laboratory | Barakos G.N.,Kazan State Technical University
Journal of Aircraft | Year: 2012

Helicopter blades experience large in-flight deformations that affect the aerodynamics of the rotor. Consequently, computational fluid dynamics (CFD) methods applied to helicopter flows must have appropriate algorithms to account for the blade deformation without deterioration in the CFD mesh quality. In this work, a hybrid mesh deformation method, suitable for use with helicopter blades, is proposed. The method begins by accounting for the blade deformation using a modal structural model. The interpolation between the finite element and CFD meshes for the blade surface is based on the constant-volume tetrahedron method and is combined with transfinite interpolation as well as the spring analogy method. The final algorithm is efficient and resulted in deformed meshes with good qualities. A range of rotor cases were considered, and the changes in volume of the CFD cells were less than 30% of their original values. The cell skewness was also kept at acceptable levels. The mesh deformation method was coupled with the helicopter multiblock CFD solver, and computations were undertaken for rigid and deformed blades. It was found that the structural deformation affected the blade loads even for hovering rotor cases, although it had a more pronounced effect in forward flight. The mesh method was efficient and accounted for less than1%of the total central processing unit time. Copyright © 2011 by Luis Delgado.

Reid T.,Newmerical Technologies International | Baruzzi G.S.,Newmerical Technologies International | Habashi W.G.,McGill University | Habashi W.G.,CFD Laboratory
Journal of Aircraft | Year: 2012

This paper presents a truly unsteady approach for the numerical simulation of in-flight electrothermal anti-icing or de-icing, using a conjugate heat transfer technique. This numerical approach has been implemented in FENSAPICE to compute the complex heat transfer phenomena occurring during in-flight de-icing with multiple heating elements following independent cycling. At each time step, the energy fluxes through the aircraft's solid skin, the melting ice layer, the liquid water film, and the external fluid are computed. The ice shape is then modified by taking into account the opposing mass balance effects of ice accreting due to the impact of supercooled droplets and/or water runback, and the partial or total melting of the existing ice layer due to heating. The results of the verification of this phase-change conduction code are presented, followed by a study of intercycle de-icing on a wing, showing intercycle ice growth. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Veillard X.,McGill University | Veillard X.,Imperial College London | Habashi W.G.,McGill University | Habashi W.G.,CFD Laboratory | Baruzzi G.S.,Newmerical Technologies International
Journal of Propulsion and Power | Year: 2011

The present paper develops the particular methodology required to simulate icing not only on the front of a jet engine, but inside multistage ones, to respond to recent safety and performance concerns. When flying in certain weather conditions, engines have been found to ingest a mix of iced and liquid particles that can result in a dangerous buildup onthe internal components of the compressor. The ice can then shed and may cause mechanical damage and performance losses to downstream components. To cost-effectively replicate such an environment, a threedimensional quasi-steady numerical approach is developed to model both rotating and static components and their interaction. An intercomponent mixing-plane approach, along with stagnation and radial equilibrium boundary conditions, has been implemented, allowing the treatment of multistage unequal-pitch blade rows via afinite element interpolation method and proper circumferential averaging. The approach is first validated for the well-documented Aachen turbine, and then used on a NASA compressor stage to highlight impingement locations of supercooled droplets and the corresponding ice shapes. Copyright © 2011 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Aliaga C.N.,Newmerical Technologies International | Aliaga C.N.,CFD Laboratory | Aube M.S.,Newmerical Technologies International | Aube M.S.,CFD Laboratory | And 4 more authors.
Journal of Aircraft | Year: 2011

In-flight ice accretion, even though driven by a steady flow airstream, is an inherently unsteady phenomenon. It is, however, completely ignored in icing simulation codes (one-shot) or, at best approximated via quasi-steady modeling (multishot). The final ice shapes thus depend on the length of the total accretion time (one-shot), or of the arbitrarily prescribed time intervals (multishot), during which the impact of ice growth on both airflow and water impingement is ignored. Such a longstanding heuristic approximation is removed in this paper by coupling in time the dilute two-phase flow (air and water droplets flow) with ice accretion, and is implemented in a new code, FENSAP-ICE-Unsteady. The two-phase flow is solved using the coupled Navier-Stokes and water concentration equations, and the water film characteristics and ice shapes are obtained from laws of conservation of mass and energy within the thin film layer. To continually update the geometry of the iced surface in time, arbitrary Lagrangian-Eulerian terms are added to all governing equations to account for mesh movement in the case of stationary components. In the case of rotating/stationary interacting components, a dynamically stitched grid is used. The numerical results clearly show that unsteady modeling improves the accuracy of both rime and glaze ice shape prediction, compared with the traditional quasi-steady approach with frozen solutions. The unsteady model is shown to open the door for a unified approach to icing on fixed wings, on helicopters with blades/rotors/fuselage systems. Problems of current concern in the icing community such as the ingestion of ice crystals at high altitude become tractable with the new formulation. Copyright © 2010 by W.G. Habashi.

Bilodeau D.R.,McGill University | Bilodeau D.R.,CFD Laboratory | Habashi W.G.,McGill University | Habashi W.G.,CFD Laboratory | And 3 more authors.
Journal of Aircraft | Year: 2015

A conservative Eulerian numerical approach for modeling postimpact Supercooled Large Droplets undergoing splashing and bouncing on aircraft surfaces is presented. The approach introduces the effect of the postimpact droplets by successive solutions of the conservation equations. Two models have been selected to identify the droplet splashing and bouncing conditions, and to provide initial conditions for the reinjected water. The method has been applied to droplet impingement in Supercooled Large Droplet conditions on clean and iced NACA 23012 geometries, as well as the MS(1)-0317 airfoil, and the results have been compared to experimental data. Good agreement is observed for both impingement limits and collection efficiency. Additionally, the method has been applied to a threeelement high-lift configuration operating in one of the proposed Appendix O Supercooled Large Droplet environments to demonstrate the danger posed by the re-impingement of splashing and bouncing droplets on complex interacting aerodynamic components. © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Pourbagian M.,McGill University | Pourbagian M.,CFD Laboratory | Habashi W.G.,McGill University | Habashi W.G.,CFD Laboratory | Habashi W.G.,Bell Helicopter Textron Inc.
Journal of Aircraft | Year: 2013

The optimization of electrothermal in-flight anti-icing systems is presented by introducing a general methodology. The optimization goal was to achieve an ice-free area over the protected zone by using the lowest energy possible. The power and/or lengthof the electric pads are considered as design variables. The optimization procedure is performed via a derivative-free method that typically needs many objective-function evaluations. This would be impractical as aero-icing flow simulation remains computationally intensive when coupled with conjugate heat-transfer calculations, as in the case of ice-protection systems. The cost is even more prohibitive for an optimization process, as a large number of simulations are needed. To make it practical, this work presents a surrogate-based optimization approach using proper orthogonal decomposition, in conjunction with kriging. The results obtained show that the methodologyis efficient and reliable in optimizing electrothermal ice-protection systemsinparticular, and athermalbased one in general. © 2013 by Wagdi G. Habashi. Published by the American Institute of Aeronautics and Astronautics, Inc.,.

Johnson C.S.,University of Liverpool | Johnson C.S.,CFD Laboratory | Barakos G.N.,University of Liverpool | Barakos G.N.,CFD Laboratory
Journal of Aircraft | Year: 2014

This work presents a framework for the optimization of certain aspects of a British Experimental Rotor Programme-like rotor blade in hover and forward flight so that maximum performance can be obtained from the blade. The proposed method employs a high-fidelity, efficient computational fluid dynamics technique that uses the harmonic balance method in conjunction with artificial neural networks as metamodels, and genetic algorithms for optimization. The approach has been previously demonstrated for the optimization of blade twist in hover and the optimization of rotor sections in forward flight, transonic aerofoils design, wing and rotor tip planforms. In this paper, a parameterization technique was devised for the British Experimental Rotor Programme-like rotor tip and its parameters were optimized for a forward flight case. A specific objective function was created using the initial computational fluid dynamics data and the metamodel was used for evaluating the objective function during the optimization using the genetic algorithms. The objective function was adapted to improve forward flight performance in terms of pitching moment and torque. The obtained results suggest optima in agreement with engineering intuition but provide precise information about the shape of the final lifting surface and its performance. The main computational cost was associated with the population of the genetic algorithms database necessary for the metamodel, especially because a full factorial method was used. The computational time of the optimization process itself, after the database has been obtained, is relatively insignificant. Therefore, the computational time was reduced with the use of the harmonic balance method as opposed to the time marching method. The novelty in this paper is twofold. Optimization methods so far have used simple aerodynamic models employing direct "calls" to the aerodynamic models within the optimization loops. Here, the optimization has been decoupled from the computational fluid dynamics data allowing the use of higher-fidelity computational fluid dynamics methods based on Navier-Stokes computational fluid dynamics. This allows a more realistic approach for more complex geometries such as the British Experimental Rotor Programme tip. In addition, the harmonic balance method has been used in the optimization process. Copyright © 2013 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Lawson S.J.,University of Liverpool | Lawson S.J.,CFD Laboratory | Barakos G.N.,University of Liverpool | Barakos G.N.,CFD Laboratory
Aerospace Science and Technology | Year: 2010

This paper demonstrates the Detached Eddy Simulation (DES) approach for the computation of flows around uninhabited combat air vehicles. One of the key features of this new family of aircraft is that weapon bays are used to enhance stealth characteristics and improve aerodynamic performance. The highly energetic flow-field within the weapon bays can change dramatically the aerodynamics of these aircraft and for this reason detailed CFD analyses are needed to provide insight in the change of loads encountered when weapon bays are exposed. In contrast to previous studies where idealised, isolated cavities are used as model problems, a realistic aircraft geometry is used in this work. Computations using DES are presented for the clean aircraft, the idealised cavity and the complete configuration. Advanced multi-block topologies are used which allow for most of the geometric details of the aircraft to be preserved and resolved by the employed CFD solver. For all cases where weapon bays are present, DES is used for the simulation while statistical turbulence models prove to be adequate for the clean aircraft cases. Comparisons against experimental data demonstrate the accuracy of the employed methods and strengthen confidence in the employed DES models. The overall loading of the aircraft is well-predicted even at high angles of attack and DES appears to offer encouraging results at low and high Mach number cases. Experiments conducted by QinetiQ and DSTL are used for validation of the DES method and results indicate that less than 3dB discrepancies in the overall sound pressure level can be obtained in comparison to the tunnel data. © 2010 Elsevier Masson SAS. All rights reserved.

Woodgate M.A.,University of Liverpool | Woodgate M.A.,CFD Laboratory | Barakos G.N.,University of Liverpool | Barakos G.N.,CFD Laboratory
AIAA Journal | Year: 2012

Computational fluid dynamics (CFD) based on the Navier-Stokes equations is by far the most useful predictive method available today for helicopter analysis and design. The main drawback of CFD, and perhaps the reason for its slow acceptance bydesign offices of helicopter manufacturers, is apparently due to the substantial requirements of CPU time and the relatively slow turnaround times in comparison to lower-order methods. However, progress with CFD algorithms and parallel computing has allowed CFD analyses to be used more routinely. Typical applications include computations of a erofoil data that feed rotor performance codes and analyses of rotors in hovering flight. The computation of unsteady flow cases is, however, still challenging. This paper presents alternative ways of tackling unsteady flow problems pertinent to rotorcraft using methods that aim to reduce the time-marching unsteady computations to more manageable steady-state solutions. The techniques investigated so far by the CFD laboratory of Liverpool include time-linearized and harmonic-balance methods. The details of the methods are presented along with their implementation in the framework of the helicopter multiblock CFD solver. Results were obtained for several flow cases, ranging from pitching/translating aerofoils to complete rotors. The results highlight some of the limitations of the time-linearized method and the potential of the harmonic-balance technique. It was found that the time-linearized method can provide adequate results for cases where the unsteady flow is a rather small perturbation of a known mean. The harmonic-balance method proved to have a larger range of applicability and provided adequate results for the analysis of unsteady flows. The required CPU time was reduced, and the required core computer memory was increased. Overall, the harmonic-balance method appears to be a possible alternative to timemarching CFD for a wide range of problems. Copyright © 2011 by M. A. Woodgate and G. N. Barakos.

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