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This paper aims to achieve an optimum flow separation control over the airfoil using a passive flow control method by introducing a bio-inspired nose near the leading edge of the National Advisory Committee for Aeronautics (NACA) 4 and 6 series airfoil. In addition, to find the optimised leading edge nose design for NACA 4 and 6 series airfoils for flow separation control.

Different bio-inspired noses that are inspired by the cetacean species have been analysed for different NACA 4 and 6 series airfoils. Bio-inspired nose with different nose length, nose depth and nose circle diameter have been analysed on airfoils with different thicknesses, camber and camber locations to understand the aerodynamic flow properties such as vortex formation, flow separation, aerodynamic efficiency and moment.

The porpoise nose design that has a leading edge with depth = 2.25% of chord, length = 0.75% of chord and nose diameter = 2% of chord, delays the flow separation and improves the aerodynamic efficiency. Average increments of 5.5% to 6° in the lift values and decrements in parasitic drag (without affecting the pitching moment) for all the NACA 4 and 6 series airfoils were observed irrespective of airfoil geometry such as different thicknesses, camber and camber location.

The two-dimensional computational analysis is done for different NACA 4 and 6 series airfoils at low subsonic speed.

This design improves aerodynamic performance and increases the structural strength of the aircraft wing compared to other conventional high lift devices and flow control devices. This universal leading edge flow control device can be adapted to aircraft wings incorporated with any NACA 4 and 6 series airfoil.

The results would be of significant interest in the fields of aircraft design and wind turbine design, lowering the cost of energy and air travel for social benefits.

Different bio-inspired nose designs that are inspired by the cetacean species have been analysed for NACA 4 and 6 series airfoils and universal optimum nose design (porpoise airfoil) is found for NACA 4 and 6 series airfoils.

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.

The purpose of this paper is to conduct a comparative analysis between a straight blade (SB) and a curved caudal-fin tidal turbine blade (CB) and to examine the aspects relating to geometry, turbulence modelling, non-dimensional forces lift and power coefficients.

The comparison utilises results obtained from a default horizontal axis tidal turbine with turbine models available from the literature. A computational design method was then developed and implemented for “horizontal axis tidal turbine blade”. Computational fluid dynamics (CFD) results for the blade design are presented in terms of lift coefficient distribution at mid-height blades, power coefficients and blade surface pressure distributions. Moving the CB back towards the SB ensures that the total blade height stays constant for all geometries. A 3D mesh independency study of a “straight blade horizontal axis tidal turbine blade” modelled using CFD was carried out. The grid convergence study was produced by employing two turbulence models, the standard k-ε model and shear stress transport (SST) in ANSYS CFX. Three parameters were investigated: mesh resolution, turbulence model, and power coefficient in the initial CFD, analysis.

It was found that the mesh resolution and the turbulence model affect the power coefficient results. The power coefficients obtained from the standard k-ε model are 15 to 20 per cent lower than the accuracy of the SST model. Further analysis was performed on both the designed blades using ANSYS CFX and SST turbulence model. The variation in pressure distributions yields to the varying lift coefficient distribution across blade spans. The lift coefficient reached its peak between 0.75 and 0.8 of the blade span where the total lift accelerates with increasing pressure before drastically dropping down at 0.9 onwards due to the escalating rotational velocity of the blades.

The work presents a computational design methodological approach that is entirely original. While this numerical method has proven to be accurate and robust for many traditional tidal turbines, it has now been verified further for CB tidal turbines.

This paper aims to design an optimum vertical axis wind turbine (VAWT) and assess its techno-economic performance for wind energy harvesting at high-speed railway in Malaysia.

This project adopted AutoCAD and ANSYS modeling tools to design and optimize the blade of the turbine. The site selected has a railway of 30 km with six stops. The vertical turbines are placed 1 m apart from each other considering the optimum tip speed ratio. The power produced and net present value had been analyzed to evaluate its techno-economic viability.

Computational fluid dynamics (CFD) analysis of National Advisory Committee for Aeronautics (NACA) 0020 blade has been carried out. For a turbine with wind speed of 50 m/s and swept area of 8 m², the power generated is 245 kW. For eight trains that operate for 19 h/day with an interval of 30 min in nonpeak hours and 15 min in peak hours, total energy generated is 66 MWh/day. The average cost saved by the train stations is RM 16.7 mil/year with battery charging capacity of 12 h/day.

Wind energy harvesting is not commonly used in Malaysia due to its low wind speed ranging from 1.5 to 4.5 m/s. Conventional wind turbine requires a minimum cut-in wind speed of 11 m/s to overcome the inertia and starts generating power. Hence, this paper proposes an optimum design of VAWT to harvest an unconventional untapped wind sources from railway. The research finding complements the alternate energy harvesting technologies which can serve as reference for countries which experienced similar geographic constraints.

This paper aims to deal with the numerical investigation of laminar separation bubble (LSB) characteristics (length and height of the bubble) of SS007 airfoil at the chord Reynolds number of Re_c = 0.68 × 10⁵ to 10.28 × 10⁵.

The numerical simulations of the flow around SS007 airfoil were carried out by using the commercial fluid dynamics (CFD) software, ANalysis system (ANSYS) 15. To solve the governing equations of the flow, a cell-centred control volume space discretisation approach is used. Wind tunnel experiments were conducted at the chord-based Reynolds number of Re_c = 1.6 × 10⁵ to validate the aerodynamic characteristics over SS007 airfoil.

The numerical results revealed that the LSB characteristics of a SS007 airfoil, and the aerodynamic performances are validated with experimental results. The lift and drag coefficients for both numerical and experimental results show very good correlation at Reynolds number 1.6 × 10⁵. The lift coefficient linearly increases with the increasing angle of attack (AOA) is relatively small. The corresponding drag coefficient was found to be very small. After the formation of LSB which leads to burst to cause airfoil stall, the lift coefficient decreases and increases the drag coefficient.

Low Reynolds number and LSB characteristics concept in aerodynamics is predominant for both civilian and military applications. These include high altitude devices, wind turbines, human powered vehicles, remotely piloted vehicles, sailplanes, unmanned aerial vehicle and micro aerial vehicle. In this paper, the micro aerial vehicle flight conditions considered and investigated the LSB characteristics for different Reynolds number. To have better aerodynamic performances, it is strongly recommended to micro aerial vehicle (MAV) design engineers that the MAV is to fly at 12 m/s (cruise speed).

MAVs and unmanned aerial vehicles seem to give some of the technical challenges of nature conservation monitoring and law enforcement a versatile, reliable and inexpensive solution.

The SS007 airfoil delays the flow separation and improves the aerodynamic efficiency by increasing the lift and decreasing the drag. The maximum increase in aerodynamic efficiency is 12.5% at stall angle of attack compared to the reference airfoil at Re = 2 × 10⁵. The results are encouraging and this airfoil could have better aerodynamic performance for the development of MAV.

Active flow control on the NACA 0024 airfoil defined as suction-injection jet at the chord-based Reynolds number of 1.5 × 1e + 5 is studied.

The three-dimensional incompressible unsteady Reynolds-averaged Navier–Stokes equations with the SST k-ω turbulence model are used to study the effects of coflow-jet (CFJ) on the dynamic and static stall phenomena. CFJ implementation is conducted with several momentum coefficients to investigate their turnover. Furthermore, the current work intends to analyze the CFJ performance by varying the Reynolds number and jet momentum coefficient and comparing all states to the baseline airfoil, which has not been studied in prior research investigations.

It is observed that at the momentum coefficient (Cµ) of 0.06, the lift coefficients at low attack angles (up to a = 15) dramatically increase. Furthermore, the dynamic stall at the given Reynolds number and with the lowered frequency of 0.15 is explored. In the instance of Cµ = 0.07, the lift coefficient curve does not show a noticeable stall feature compared to Cµ = 0.05, suggesting that a more powerful stronger jet can entirely control the dynamic stall.

Furthermore, the current work intends to analyze the CFJ performance by varying the jet momentum coefficient and comparing all states to the baseline airfoil, which has not been studied in prior research investigations.

Aircraft pitch control is fundamental for the performance of micro aerial vehicles (MAVs). The purpose of this paper is to establish a simple experimental procedure to calibrate pitch instrumentation and classical control algorithms. This includes developing an efficient pitch angle observer with optimal estimation and evaluating controllers under uncertainty and external disturbances.

A wind tunnel test bench is designed to simulate fixed-wing aircraft dynamics. Key elements of the instrumentation commonly found in MAVs are characterized in a gyroscopic test bench. A data fusion algorithm is calibrated to match the gyroscopic test bench measurements and is then integrated into the autopilot platform. The elevator-angle to pitch-angle dynamic model is obtained experimentally. Two different control algorithms, based on model-free and model-based approaches, are designed. These controllers are analyzed in terms of parametric uncertainties due to wind speed variations and external perturbation because of sudden weight distribution changes. A series of experimental tests is performed in wind-tunnel facilities to highlight the main features of each control approach.

With regard to the instrumentation algorithms, a simple experimental methodology for the design of optimal pitch angle observer is presented and validated experimentally. In the context of the platform design and identification, the similitude among the theoretical and experimental responses shows that the platform is suitable for typical pitch control assessment. The wind tunnel experiments show that a fixed linear controller, designed using classical frequency domain concepts, is able to provide adequate responses in scenarios that approximate the operation of MAVs.

The aircraft orientation observer can be used for both pitch and roll angles. However, for simultaneousyaw angle estimation the proposed design method requires further research. The model analysis considers a wind speed range of 6-18 m/s, with a nominal operation of 12 m/s. The maximum experimentally tested reference for the pitch angle controller was 20°. Further operating conditions may require more complex control approaches (e.g. scheduling, non-linear, etc.). However, this operating range is enough for typical MAV missions.

The study shows the design of an effective pitch angle observer, based on a simple experimental approach, which achieved locally optimum estimates at the test conditions. Additionally, the instrumentation and design of a test bench for typical pitch control assessment in wind tunnel facilities is presented. Finally, the study presents the development of a simple controller that provides adequate responses in scenarios that approximate the operation of MAVs, including perturbations that resemble package delivery and parametric uncertainty due to wind speed variations.

The purpose of this paper is to get the topology shape and material distribution of composite rotor beam under the requirement of cross-sectional characteristics.

A new multi-material topology optimization method is given. Designated shear center (SC) position and stiffness terms are combined as the objective function. Multi-material model including isotropic and anisotropic materials are employed. Sensitivity analysis is given based on gradient-based algorithm, and density filtering scheme is adopted to avoid checkerboard problem.

The topology design method of composite rotor beam provides innovative cross-sectional shape and material distribution method. The final topology shape like “ > ” is given for different material types and cross-sectional shape under SC position requirement. The coefficient of stiffness components has great influence on the cross-sectional final topology shape.

The proposed method is just to give cross-sectional topology shape. To obtain final actual composite rotor beam structure, shape and size optimization should be conducted if the topology shape is given.

This method is suitable for the preliminary design of helicopter rotor beam to get designated SC position and stiffness terms.

The proposed method provides a new gradient-based algorithm for multi-material topology optimization design of composite rotor beam.

The purpose of this paper is to study the structure and dynamic development of a pair of co-rotating trailing vortices, during their formation, interaction and merging, using detailed experimental measurements of the velocity and vorticity fields.

The vortices were generated using two half wings (NACA0030) positioned at equal and opposite angles of attack at the entrance of the test section of an open-circuit, subsonic, wind tunnel. Velocity vector measurements were obtained at Re_c = 133,000, on cross-plane grids at several distances from the trailing edges of the wings, using an in-house developed four-sensor hot wire anemometer probe.

The results include cross-plane contour plots of the mean and fluctuating velocity as well as mean vorticity fields. Each of these variables is affected in a different way, providing complementary information on the development of the flow field. After shedding, the two vortices are swept along the stream-wise direction and spiral around each other, thereby developing a braid of two vortices, which then deforms the external flow field. Gradually, the interaction with the external flow field links both vortices together until the final merging and the formation of a new stable linear vortex emerges.

Trailing vortices have been rendered particularly important during the past decades, because of increasing traffic density of very heavy aircrafts and several plane “incidents”, which were attributed to the action of the vortex wake.

The presented results provide information on the evolution and merging of a pair of vortices formed by a closely spaced differential wing configuration. The vortices interact almost immediately after shedding as expected in flap–flap or flap–wing vortices interaction.

Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical Research Council, Reports and Technical Notes of the United States National Advisory Committee for Aeronautics and publications of other similar Research Bodies as issued.