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The purpose of this paper is to analyze the aerodynamic improvements obtained in a wing section with a NACA 0018 airfoil manufactured using the fused deposition modeling (FDM) technique with regard to a smooth surface made by milling. The creation of micro-riblets on the surface of the airfoil, due to the deposition of the material layer by layer, improves the general aerodynamic performance of the parts, provided that the riblets are parallel to the flow line. The incidence of the thickness of the thread deposited in each layer – to be the variable on which the geometry of the riblets is based – was studied.

The wing section was designed using 3D software. Three different models were designed by rapid prototyping, using additive and subtractive manufacturing. Two of the profiles were manufactured using FDM varying the thickness of the layer to be able to compare the aerodynamic improvements. The third model was manufactured using a subtractive rapid prototyping machine generating a smooth surface profile. These three models were tested inside the wind tunnel to be able to quantify the aerodynamic efficiency according to the geometry and the riblets size.

The manufacture of an aerodynamic profile using FDM provides, in addition to the lightness and the ability to design parts with complex geometries, an improvement in the aerodynamic efficiency of 10 per cent compared with profiles with a smooth surface.

With the aerodynamic advantage gained through the use of FDM positions, the additive manufacturing serves as an excellent alternative for the manufacture of lightweight aerodynamic parts, with low structural loading and with low Reynolds number (∼5·105). This technological advantage would be applied to the UAV (unmanned aerial vehicle) industry.

The study carried out in this article demonstrates that the use of FDM as a manufacture process of end-used parts that are subject to movement generates an additional advantage that had not been considered. The additive manufacturing allows us to directly manufacture riblets by creating the necessary surface so as to reduce the aerodynamic drag.

This study aims to use computational fluid dynamics (CFD) to understand and quantify the overall blockage within a transonic axial flow compressor (AFC), and to develop an efficient collaborative design optimization method for compressor aerodynamic performance and stability in conjunction with a surrogate-assisted optimization technique.

A quantification method for the overall blockage is developed to integrate the effect of regional blockages on compressor aerodynamic stability and performance. A well-defined overall blockage factor combined with efficiency drives the optimizer to seek the optimum blade designs with both high efficiency and wide-range stability. An adaptive Kriging-based optimization technique is adopted to efficiently search for Pareto front solutions. Steady and unsteady numerical simulations are used for the performance and flow field analysis of the datum and optimum designs.

The proposed method not only remarkably improves the compressor efficiency but also significantly enhances the compressor operating stability with fewer CFD calls. These achievements are mainly attributed to the improvement of specific flow behaviors oriented by the objectives, including the attenuation of the shock and weakening of the tip leakage flow/shock interaction intensity.

CFD-based design optimization of AFC is inherently time-consuming, which becomes even trickier when optimizing aerodynamic stability since the stall margin relies on a complete simulation of the performance curve. The proposed method could be a good solution to the collaborative design optimization of aerodynamic performance and stability for transonic AFC.

For an axial-flow compressor rotor, the upstream inflow conditions will vary as the aircraft faces harsh flight conditions (such as taking off, landing or maneuvering) or the whole compressor operates at off-design conditions. With the increase of upstream boundary layer thickness, the rotor blade tip will be loaded and the increased blade load will deteriorate the shock/boundary layer interaction and tip leakage flows, resulting in high aerodynamic losses in the tip region. The purpose of this paper is to achieve a better flow control for tip secondary flows and provide a probable design strategy for high-load compressors to tolerate complex upstream inflow conditions.

This paper presents an analysis and application of shroud wall optimization to a typical transonic axial-flow compressor rotor by considering the inlet boundary layer (IBL). The design variables are selected to shape the shroud wall profile at the tip region with the purpose of controlling the tip leakage loss and the shock/boundary layer interaction loss. The objectives are to improve the compressor efficiency at the inlet-boundary-layer condition while keeping its aerodynamic performance at the uniform condition.

After the optimization of shroud wall contour, aerodynamic benefits are achieved mainly on two aspects. On the one hand, the shroud wall optimization has reduced the intensity of the tip leakage flow and the interaction between the leakage and main flows, thereby decreasing the leakage loss. On the other hand, the optimized shroud design changes the shock structure and redistributes the shock intensity in the spanwise direction, especially weakening the shock near the tip. In this situation, the shock/boundary layer interaction and the associated flow separations and wakes are also eliminated. On the whole, at the inlet-boundary-layer condition, the compressor with optimized shroud design has achieved a 0.8 per cent improvement of peak efficiency over that with baseline shroud design without sacrificing the total pressure ratio. Moreover, the re-designed compressor also maintains the aerodynamic performance at the uniform condition. The results indicate that the shroud wall profile has significant influences on the rotor tip losses and could be properly designed to enhance the compressor aerodynamic performance against the negative impacts of the IBL.

The originality of this paper lies in developing a shroud wall contour optimization design strategy to control the tip leakage loss and the shock/boundary layer interaction loss in a transonic compressor rotor. The obtained results could be beneficial for transonic compressors to tolerate the complex upstream inflow conditions.

The purpose of this study is to increase maximum lift/drag ratio (E_max) of tactical unmanned aerial vehicles (TUAVs) via applying novel small aerodynamic modifications.

A TUAV is manufactured in Erciyes University, Faculty of Aeronautics and Astronautics, Model Aircraft Laboratory. It has both passive and active morphing capabilities. Its nosecone and tailcone shapes are redesigned to improve E_max. Moreover, active flow control is also built on its wing for improving E_max.

Using these novel small aerodynamic modifications, considerable improvement on E_max is obtained.

Permission of Directorate General of Civil Aviation in Turkey is required for testing TUAVs in real-time applications.

Small aerodynamic modifications such as nosecone-tailcone shape modifications and building active flow control on wing are very beneficial for improving E_max of TUAVs.

Small aerodynamic modifications satisfy confidence, high performance and easy utility demands of TUAV users.

The study will enable the creation of novel approaches to improve E_max value and therefore aerodynamic performance of TUAVs.

In a modern commercial airliner programme the ability to produce a technically good product at the right time is not enough. Efficient programme management, in all its facets, is absolutely essential to ensure a proper return on the big investments required and the smooth and progressive deployment of financial, plant and manpower resources through each programme to the next.

The purpose of this paper is to describe the effects of characteristic geometric parameters on parafoil aerodynamic performance by using computational fluid dynamics (CFD) technique.

The main characteristic geometric parameters cover the planform geometry, arc‐anhedral angle, basic airfoil and leading‐edge cut. By using the CFD technique, a large number of numerical parafoil models with different geometric parameters are developed to study the correlations between these parameters and parafoil aerodynamic performance.

The CFD technique is feasible and effective to study the effects of characteristic geometric parameters on parafoil aerodynamic performance in three‐dimensional (3‐D) flowfield condition. The planform geometry can affect the aerodynamic performance obviously. An increase in arc‐anhedral angle decreases the lift of a parafoil but has little effect on lift‐drag ratio. The model with smaller leading‐edge radius and thinner thickness of parafoil section achieves larger lift‐drag ratio. The leading‐edge cut has little effect on lift but increase drag dramatically; meanwhile, its effect on flowfield is confined to the nearby region of leading edge.

The presented 3‐D numerical simulation results of parafoil models are shown to have good agreement with the tunnel test data in general trend; meanwhile, considering its relatively low‐cost, the CFD method could be further used to predict coefficients in pre‐research or at non‐experimental conditions.

The paper can form the foundation of further studies on parafoil aerodynamic performance with different geometric parameters.

The present paper aims to address the description of a numerical optimization procedure, based on mesh morphing, and its application for the improvement of the aerodynamic performance of an industrial glider which suffers of a large separation occurring in the wing–fuselage junction region at high incidence angles.

Shape variations were applied to the baseline configuration through a mesh morphing technique founded on the mathematical framework of radial basis functions (RBF). The aerodynamic solutions were obtained coupling an RANS code with the mesh morphing tool RBF Morph™. Two shape modifiers were set up to generate a parametric numerical model. An optimization procedure, based on a design of experiment sampling, was set up implementing the fully automated workflow within a high performance computing (HPC) environment. The optimal candidates maximizing the aerodynamic efficiency were identified by means of a cubic RBF response surface approach.

The separation was significantly reduced, modifying the local geometry of fuselage and fairing and maintaining the wing aerofoil unchanged. A relevant aerodynamic efficiency improvement was finally gained.

The developed procedure proved to be a very powerful and efficient tool in facing aerodynamic design problems. However, it might be computationally very expensive if a large number of design variables are adopted and, in those cases, the method can be suitably used only within the HPC environment.

Such an optimization study is part of an explorative set of analyses that focused on better addressing the numerical strategies to be used in the development of the EU FP7 Project RBF4AERO.

This study aims to investigate the morphing concepts able to manipulate the dynamics of the downstream unsteadiness in the separated shear layers and, in the wake, be able to modify the upstream shock–boundary layer interaction (SBLI) around an A320 morphing prototype to control these instabilities, with emphasis to the attenuation or even suppression of the transonic buffet. The modification of the aerodynamic performances according to a large parametric study carried out at Reynolds number of 4.5 × 10⁶, Mach number of 0.78 and various angles of attack in the range of (0, 2.4)° according to two morphing concepts (travelling waves and trailing edge vibration) are discussed, and the final benefits in aerodynamic performance increase are evaluated.

This article examines through high fidelity (Hi-Fi) numerical simulation the effects of the trailing edge (TE) actuation and of travelling waves along a specific area of the suction side starting from practically the most downstream position of the shock wave motion according to the buffet and extending up to nearly the TE. The present paper studies through spectral analysis the coherent structures development in the near wake and the comparison of the aerodynamic forces to the non-actuated case. Thus, the physical mechanisms of the morphing leading to the increase of the lift-to-drag ratio and the drag and noise sources reduction are identified.

This study investigates the influence of shear-layer and near-wake vortices on the SBLI around an A320 aerofoil and attenuation of the related instabilities thanks to novel morphing: travelling waves generated along the suction side and trailing-edge vibration. A drag reduction of 14% and a lift-to-drag increase in the order of 8% are obtained. The morphing has shown a lift increase in the range of (1.8, 2.5)% for angle of attack of 1.8° and 2.4°, where a significant lift increase of 7.7% is obtained for the angle of incidence of 0° with a drag reduction of 3.66% yielding an aerodynamic efficiency of 11.8%.

This paper presents results of morphing A320 aerofoil, with a chord of 70cm and subjected to two actuation kinds, original in the state of the art at M = 0.78 and Re = 4.5 million. These Hi-Fi simulations are rather rare; a majority of existing ones concern smaller dimensions. This study showed for the first time a modified buffet mode, displaying periodic high-lift “plateaus” interspersed by shorter lift-decrease intervals. Through trailing-edge vibration, this pattern is modified towards a sinusoidal-like buffet, with a considerable amplitude decrease. Lock-in of buffet frequency to the actuation is obtained, leading to this amplitude reduction and a drastic aerodynamic performance increase.

One of the best methods to improve wind turbine aerodynamic performance is modification of the blade’s airfoil. The purpose of this paper is to investigate the effects of gurney flap geometry and its oscillation parameters on the pitching NACA0012 airfoil.

This numerical solution has been carried out for different cases of gurney flap mounting angles, heights, reduced frequencies and oscillation amplitudes, then the results were compared to each other. The finite volume method was used for the discretization of the governing equations, and the PISO algorithm was used to solve the equations. Also, the “SST” was adopted as the turbulence model in the simulation.

In this paper, the different parameters of gurney flap were investigated. The results showed that the best range of gurney flap height are between 1 and 3.2% of chord and the best ratio of lifting to drag coefficient is achieved in gurney flap with an angle of 90° relative to the chord direction. The dynamic stall angle of the airfoil with gurney flap decreases were compared to without gurney flap. Earlier LEV formation can be one of the main reasons for decreasing the dynamic stall angle of the airfoil with gurney flap. Increasing the reduced frequency and oscillation amplitude causes rising of maximum lift coefficient and consequently lift curve slope. Moreover, gurney flap with mounting angle has a lower hinge moment than the gurney flap without mounting angle but with the same vertical axis length. So, there is more complexity in structural design concerning the gurney flap without mounting angle.

Improving aerodynamic efficiency of airfoils is vital for obtaining more output power in VAWTs. Gurney flaps are one of the best mechanisms to increase the aerodynamic performance of the airfoil and increases the efficiency of VAWTs.

Investigating the hinge moment on the connection point of the airfoil, gurney flap and try to compare the gurney flap with and without angle.

This paper aims to study the extended trailing edge airfoil for a range of angle of attack at different intensities of turbulence.

In this paper, an experimental study on NACA 0020 airfoil with thin extended trailing edge modification of amplitude of h = 0.1c, 0.2c and 0.3c at the Reynolds number of 2.14 × 10⁵ are tested. The research was carried out for an angle of attack ranging from 0° = α = 35° for the turbulence intensity of 0.3%, 3%, 5%, 7% and 12%. From the experimental readings, the surface pressures are scanned using a Scanivalve (MPS2464) pressure scanner for a sampling frequency of 700 Hz. The scanned pressures are converted to aerodynamic force coefficient and the results are combined and discussed.

The airfoil with the extended trailing edge will convert the adverse pressure gradient to a plateau pressure zone, indicating the delayed flow separation. The CL value at higher turbulence intensity (TI = 12%) for the extended trailing edge over perform the base airfoil at the post-stall region. The maintenance of flow stability is observed from the spectral graph.

A thin elongated trailing edge attached to the conventional airfoil serves as a flow control device by delaying the stall and improving the lift characteristics. Additionally, extending the airfoil's trailing edge helps to manage the performance of the airfoil even at a high level of turbulence.

Distinct from existing studies, the presented results reveals how the extended trailing edge attached to the airfoil performs in the turbulence zone ranging from 0.3% to 12% of TI. The displayed pressure distribution explains the need for increasing trailing edge amplitude (h) and its impact on flow behaviour. The observation is that extended trailing edge airfoil bears to maintain the performance even at higher turbulence region.