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Many Small Unmanned Aerial Vehicles (SUAV) are driven by small scale, fixed blade propellers. Flow produced by the propeller can have a significant impact on the aerodynamics of a SUAV. Therefore, in Computational Fluid Dynamic (CFD) simulations, it is often necessary to simulate the SUAV and propeller coupled together. For computational efficiency, the propeller can be modeled in a steady-state view by using momentum source terms to add the thrust and swirl produced by the propeller to the flow field. Many momentum source term models are based on blade element theory. Blade element theory divides the blade into element sections in the spanwise direction and assumes each element to operate independently as a two-dimensional (2D) airfoil.

Accurate turbulence modeling remains a critical problem in the prediction capability of computational fluid dynamics. One particular area lacking accurate simulations is separated turbulent flows. In this paper, the recently developed one - equation Rahman-Agarwal-Siikonen (RAS) turbulence model is used to simulate the flow of several canonical turbulent flow cases. The commercially available software ANSYS Fluent and the open source software OpenFOAM are used for the flow field calculations. It is shown that the RAS model improves the accuracy of flow simulations compared to the commonly used one - equation Spalart-Allmaras (SA) and two-equation SST k-ω turbulence models.

The purpose of this study is to investigate the effects of the size and location of wind fences on the rotor performance to avoid as well as quantify any adverse effects of outside wind on the measurement accuracy. To this end, firstly, a novel actuator disk method is developed which couples open source CFD code named OpenFOAM with blade element method and rotor flapping equations. Secondly, the parametric studies are conducted with respect to wind fence configurations which contain fence radius, inlet/outlet duct size and location etc. For quantitative evaluation of rotor performance variation according to inlet/outlet duct size, the mass flow and momentum rate on the ducts are. The rotor performance variation depending on the wind fence configuration is examined and consequently parametric study cases are classified according as calculated rotor thrust coefficient. Moreover, it is explained the difference of flow field in wind fence by demonstrating the pressure coefficient.

Liquid ring pumps are used in aircraft fuel systems in conjunction with main impeller pumps. These pumps are used for priming the pump system as well as to remove fuel vapor and air from the fuel. Prediction of cavitation in liquid ring pumps is important as cavitation degrades the performance of these pumps and leads to their failure. As test based assessment of cavitation risk in liquid ring pump is expensive and time consuming, recent approaches have been to assess and predict the risk of cavitation using Computational Fluid Dynamics (CFD) methods with the goal to quicken the design process and optimize the performance of these pumps. The present study deals with the development and assessment of a CFD methodology to simulate cavitation for a liquid fuel pump used in aircraft fuel systems. The study simulates the cavitation phenomena using a multi-phase flow model consisting of fuel vapor, air, and liquid fuel phases.

This paper looks into with the aerodynamic properties and stability of the feeder airship in the framework of MAAT project. FP7 MAAT project is based on the concept of two different types of airships (the cruiser and the feeder) working together as a transportation system. The feeder considered in this paper is a rigid airship with an unconventional envelope shape. Aerodynamic forces and moments acting on the airship during the horizontal and vertical flight modes are of special interest in this study, because the aerodynamic performance of the aircraft directly influences its general dynamic behavior and, thus, its in-flight stability. A set of CFD simulations was conducted for vertical and horizontal flights of the feeder airship. Drag and lift forces and pitching moment together with their coefficients, were obtained for different altitudes and velocities from the proposed operational ranges of the airship.

Ice accretion is a phenomenon in which supercooled water droplets impinge and accrete on a body. In the present study, we focus on a jet engine because it is well known that ice accretion on blades and vanes leads to performance degradation and has caused severe accidents. Although various anti-icing and deicing systems have been developed, such accidents still occur. Therefore, it is important to clarify the phenomenon of ice accretion in a jet engine. However, flight tests for ice accretion are very expensive, and in the wind tunnel it is difficult to reproduce every climate condition where ice accretion occurs. Therefore, it is expected that computational fluid dynamics (CFD), which can estimate ice accretion in various climate conditions, will be a useful way to predict the ice accretion phenomenon. The characteristic phenomena of supercooled large droplets (SLD) are splash and bounce.

Electronic systems play a key role in the high reliability and safety of modern aircrafts. Components have become smaller and faster leading to an increase in power density and making thermal management a more vital part of the overall design. This paper will demonstrate how thermal transient testing combined with computational fluid dynamics (CFD) can help balance these design constraints and ensure that critical devices will work safely within their prescribed temperature limits. The first part of the paper will focus on the thermal characterization of a component using a continuous measurement method known as the static method per JEDEC51-1 standard. This data provides an insight of the overall component thermal performance and it can also be translated to structure function which helps with detecting potential internal thermal bottlenecks, such as a die attach material.

The focus of this paper is on the numerical simulation of compressible flow in a diffusing S-duct inlet; this flow is characterized by secondary flow as well as regions of boundary layer separation. The S-Duct geometry produces streamline curvature and an adverse pressure gradient resulting in these flow characteristics. The geometry used in this investigation is based on a NASA Glenn Research Center experimental diffusing S-Duct which was studied in the early 1990's. The CFD flow solver ANSYS - Fluent is employed in the investigation of compressible flow through the S-duct. A second-order accurate, steady, density-based solver is employed in a finite-volume framework. The three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations are solved on a structured mesh with a number of turbulence models, namely the Spalart - Allmaras (SA), k-ε, k-ω SST, and Transition SST models, and the results are compared with the experimental data.

This paper is dedicated to the study and improvement of the aerodynamic properties of the feeder airship in the context of MAAT project. FP7 MAAT project is based on the concept of two different types of airships (the cruiser and the feeder) working together as a transportation system. The current feeder concept includes unconventional shape changing envelope. Two problems are considered in this paper. The first problem is to find a condition of the effective vertical ascent for the feeder (from the ground to the altitude of the cruiser). A series of CFD simulations were carried out for the top flow for a range of altitudes from 0 to 16 km and velocities between 2 and 10 m/s. The results confirm the appearance of some negative effects, including high drag during the vertical ascent, especially, at low altitudes. The second problem is to study and reduce the side wind effects on the ascending feeder airship.

Cyclogiros have the possibility to provide hovering and achieve forward speeds higher than present day helicopters. Yet, due to the inherent unsteady aerodynamics, most of the research approaches have been based on limited experimental testing of computational fluid dynamics. However, and for a parametric analysis, it is important to provide analytic tools that can help in the preliminary design stage. In this work the complexity of the forward flight modeling will be described using an innovative approach. The present proposed model will be benchmarked against experimental and CFD results, further it will be used to propose rotor geometries for 3 working conditions.

There has been an increased interest in low speed aerodynamics for Unmanned Aerial Vehicles (UAVs) and Micro Aerial Vehicles (MAVs). These vehicles which are increasingly being used for reconnaissance purposes operate in the Root Chord Reynolds number range of 104 to around 105 and thus, the flow regime encountered is in the low Reynolds number transitional flow range. Computational Fluid Dynamics (CFD) methods which employ eddy viscosity based RANS turbulence models that are formulated for high Reynolds number flow are not well suited for such low Reynolds number range as they cannot accurately model the formation of laminar separation bubble and subsequent onset of transition. In this paper, the transition k-ω SST model is assessed for aerodynamics prediction for the SD7003 airfoil for Reynolds number ranging from 104 to 9 × 104 and angle of attack ranging from 0° to 8°. The assessment is carried out against available experimental and Large Eddy Simulation results.

In this work, the Hydra Technologies Unmanned Aerial System (UAS) wing was simulated in several flight conditions. The simulations were carried out with the FLUENT Computational Fluid Dynamics code in 2D, and the grid was drawn with ICEM. Simulations were conducted using the turbulent models Spalart-Allmaras, SST k-ω and transition SST k-ω. The calculated results were compared with Xfoil. This comparative study allowed increasing the accuracy of the results as for the aerodynamic parameters of the wing CL, CD, CM and CP. The parallel objective was to determine the transition point. All the turbulent models were able to accurately simulate the correct behavior of the CP but transition SST k-ω appeared to be the only one able to find the transition point. In this study, we simulate, in 2D, a specific section of the wing located at the wing tip.