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For flow around elongated bluff bodies, flow separations would occur over both leading and trailing edges. Interactions between these two separations can be established through acoustic perturbation. In this paper, the flow and the acoustic fields of a D-shaped bluff body (length-to-height ratio L/H = 3.64) are investigated at height-based Reynolds number Re = 23,000 by experimental and numerical methods. The purpose of this paper is to study the acoustic feedback in the interaction of these two separated flows.

The flow field is measured by particle image velocimetry, hotwire velocimetry and surface oil flow visualization. The acoustic field is modeled in two dimensions by direct aeroacoustic simulation, which solves the compressible Navier–Stokes equations. The simulation is validated against the experimental results.

Separations occur at both the leading and the trailing edges. The leading-edge separation point and the reattaching flow oscillate in accordance with the trailing-edge vortex shedding. Significant pressure waves are generated at the trailing edge by the vortex shedding rather than the leading-edge vortices. Pressure-based cross-correlation analysis is conducted to clarify the effect of the pressure waves on the leading-edge flow structures.

The understanding of interactions of separated flows over elongated bluff bodies helps to predict aerodynamic drag, structural vibration and noise in engineering applications, such as the aerodynamics of buildings, bridges and road vehicles.

This paper clarifies the influence of acoustic perturbations in the interaction of separated flows over a D-shaped bluff body. The contribution of the leading- and the trailing-edge vortex in generating acoustic perturbations is investigated as well.

In this paper, experimental and numerical results of a lambda wing have been compared. The purpose of this paper is to study the behaviour of lambda wings using a CFD tool and to consider different numerical models to obtain the most accurate results. As far as the consideration of numerical methods is concerned, the main focus is on the evaluation of computational methods for an accurate prediction of contingent leading edge vortices’ path and the flow separation occurring because of the burst of these vortices on the wing.

Experimental tests are performed in a closed-circuit wind tunnel at the Reynolds number of 6 × 105 and angles of attack (AOA) ranging from 0 to 10 degrees. Investigated turbulence models in this study are Reynolds Averaged Navior–Stokes (RANS) models in a steady state. To compare the accuracy of the turbulence models with respect to experimental results, sensitivity study of these models has been plotted in bar charts.

The results illustrate that the leading edge vortex on this lambda wing is unstable and disappears soon. The effect of this disappearance is obvious by an increase in local drag coefficient in the junction of inner and outer wings. Streamlines on the upper surface of the wing show that at AOA higher than 8 degrees, the absence of an intense leading edge vortex leads to a local flow separation on the outer wing and a reverse in the flow.

Results obtained from the behaviour study of transition (TSS) turbulence model are more compatible with experimental findings. This model predicts the drag coefficient of the wing with the highest accuracy. Of all considered turbulence models, the Spalart model was not able to accurately predict the non-linearity of drag and pitching moment coefficients. Except for the TSS turbulence model, all other models are unable to predict the aerodynamic coefficients corresponding to AOA higher than 10 degrees.

The presented results in this paper include lift, drag and pitching moment coefficients in various AOA and also the distribution of aerodynamic coefficients along the span.

The presented results include lift, drag and pitching moment coefficients in various AOA and also aerodynamic coefficients distribution along the span.

The purpose of this exhaustive experimental study is to investigate the fluid-structure interaction in the flexible membrane wings over a range of angles of attack for various Reynolds numbers.

In this paper, an experimental study on fluid-structure interaction of flexible membrane wings was presented at Reynolds numbers of 2.5 × 10⁴, 5 × 10⁴ and 7.5 × 10⁴. In the experimental studies, flow visualization, velocity and deformation measurements for flexible membrane wings were performed by the smoke-wire technique, multichannel constant temperature anemometer and digital image correlation system, respectively. All experimental results were combined and fluid-structure interaction was discussed.

In the flexible wings with the higher aspect ratio, higher vibration modes were noticed because the leading-edge separation was dominant at lower angles of attack. As both Reynolds number and the aspect ratio increased, the maximum membrane deformations increased and the vibrations became visible, secondary vibration modes were observed with growing the leading-edge vortices at moderate angles of attack. Moreover, in the graphs of the spectral analysis of the membrane displacement and the velocity; the dominant frequencies coincided because of the interaction of the flow over the wings and the membrane deformations.

Unlike available literature, obtained results were presented comparatively using the sketches of the smoke-wire photographs with deformation measurement or turbulence statistics from the velocity measurements. In this study, fluid-structure interaction and leading-edge vortices of membrane wings were investigated in detail with increasing both Reynolds number and the aspect ratio.

The purpose of this paper is to identifying ways to reduce the effects of wing-vortex interaction by applying surface porosity on selected areas of the exposed surface. A number of papers recently have investigated the aerodynamic implication of free-stream vortices impinging upon airfoils.

The free-stream disturbance in these studies were represented by planting a vortex ahead of the wing or using some other disturbance invoking mechanism like von-Karman vortices in the wake of a cylinder or using a flipping plate to invoke a discrete vortex. In the present work, a well-defined method was used to germinate a system of controlled vortices of known strength, size and frequency ahead of the wing, and the impact of the subsequent interaction was studied with and without the presence of the surface porosity. The simulations tackled a number of cases when porosities of up to 20 and 22 per cent were applied to selected regions near the leading edge, with vortices of controlled strengths directed at the wing surface.

The results showed that the effects of large vortices spanning the entire lengths of the wing can indeed be damped when porosity is selectively applied at strategic regions.

Surface porosity application at strategic regions of a wing may dampen the effects of the unsteadiness of the incoming flow. This has profound implications on flight safety and structural damage prevention. Further implications could possibly be extended to UAV and wind turbines that operate at heavy gusting environment.

Implementation of this particular method resolves some of the issues arisen when an airplane encounters atmospheric turbulence.

The large‐scale numerical simulation of fluid flow is described as a discipline within the field of software engineering. As an example of such work, a vortex flowfield is analysed for its essential physical flow features, an appropriate mathematical description is presented (the Euler equations with an artificial viscosity model), a numerical algorithm to solve the mathematical equations is described, and the programming methodology which allows us to attain a very high degree of vectorization on the CYBER 205 is discussed. Four simulated flowfields with vorticity shed from wing edges are computed with up to as many as one million grid points and verify the realism of the simulation model. The computed solutions show all the qualitative features that are expected in these flows. The twisted cranked‐and‐cropped delta case is one where the leading‐edge vortex is highly stretched and unstable, displaying ultimately inviscid large‐scale turbulent‐like phenomena.

The purpose of this study is to investigate the aerodynamic characteristics of a low-to-moderate-aspect-ratio, tapered, untwisted, unswept wing, equipped of sheared wing tips.

In this work, wind tunnel tests were made to study the influence in aerodynamic characteristics over a typical low-to-moderate-aspect-ratio wing of a general aviation aircraft, equipped with sheared – swept and tapered planar – wing tips. An experimental parametric study of different wing tips was tested. Variations in its leading and trailing edge sweep angle as well as variations in wing tip taper ratio were considered. Sheared wing tips modify the flow pattern in the outboard region of the wing producing a vortex flow at the wing tip leading edge, enhancing lift at high angles of attack.

The induced drag is responsible for nearly 50% of aircraft total drag and can be reduced through modifications to the wing tip. Some wing tip models present complex geometries and many of them present benefits in particular flight conditions. Results have demonstrated that sweeping the wing tip leading edge between 60 and 65 degrees offers an increment in wing aerodynamic efficiency, especially at high lift conditions. However, results have demonstrated that moderate wing tip taper ratio (0.50) has better aerodynamic benefits than highly tapered wing tips (from 0.25 to 0.15), even with little less wing tip leading edge sweep angle (from 57 to 62 degrees). The moderate wing tip taper ratio (0.50) offers more wing area and wing span than the wings with highly tapered wing tips, for the same aspect ratio wing.

Although many studies have been reported on the aerodynamics of wing tips, most of them presented complex non-planar geometries and were developed for cruise flight in high subsonic regime (low lift coefficient). In this work, an exploration and parametric study through wind tunnel tests were made, to evaluate the influence in aerodynamic characteristics of a low-to-moderate-aspect-ratio, tapered, untwisted, unswept wing, equipped of sheared wing tips (wing tips highly swept and tapered).

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.

This study aims to investigate the effects of propeller locations on the aerodynamic characteristics of a generic 55° swept angle sharp-edged delta wing unmanned aerial vehicle (UAV) model.

A generic delta-winged UAV model has been designed and fabricated to investigate the aerodynamic properties of the model when the propeller is placed at three different locations. In this research, the propeller has been placed at three different positions on the wing, namely, front, middle and rear. The experiments were conducted in a closed-circuit low-speed wind tunnel at speeds of 20 and 25 m/s corresponding to 0.6 × 10⁶ and 0.8 × 10⁶ Reynolds numbers, respectively. The propeller speed was set at constant 6,000 RPM and the angles of attack were varied from 0° to 20° for all cases. During the experiment, two measurement techniques were used on the wing, which were the steady balance measurement and surface pressure measurement.

The results show that the locations of the propeller have significant influence on the lift, drag and pitching moment of the UAV. Another important observation obtained from this study is that the location of the propeller can affect the development of the vortex and vortex breakdown. The results also show that the propeller advance ratio can also influence the characteristics of the primary vortex developed on the wing. Another main observation was that the size of the primary vortex decreases if the propeller advance ratio is increased.

There are various forms of UAVs, one of them is in the delta-shaped planform. The data obtained from this experiment can be used to understand the aerodynamic properties and best propeller locations for the similar UAV aircrafts.

To the best of the author’s knowledge, the surface pressure data available for a non-slender delta-shaped UAV model is limited. The data presented in this paper would provide a better insight into the flow characteristics of generic delta winged UAV at three different propeller locations.

A numerical method that solves the Euler equations for compressible flow is used to study vortex stretching. The particular case simulated is supersonic flow M∞=1.2 α=10 degrees around the twisted and cambered cranked‐and‐cropped TKF delta wing of MBB. This geometry induces multiple leading‐edge vortices in a straining velocity field that brings about flow instabilities but the result is a state of statistical equilibrium. The discretization contains over 600,000 cells and offers sufficient degrees of freedom in the solution to resolve the small‐scale unstable modes that lead to disordered vortex flow.

The goal of this research is to demonstrate micro‐electro‐mechanical systems (MEMS)‐based transducers for aircraft maneuvering. Research in wind tunnels have shown that micro‐actuators can be used to manipulate leading edge vortices found on aerodynamic surfaces with moderate to highly swept leading edges, such as a delta wing. This has been labeled as the MEMS vortex shift control (MEMS‐VSC). The work presented in this paper seeks to detail the evolution of real‐world flight tests of this research using remotely piloted vehicles (RPVs).

Four different RPVs were constructed and used for flight tests to demonstrate the ability of using MEMS devices to provide flight control, primarily in the rolling axis.

MEMS devices for high angle‐of‐attack (AOA) turning flights have been demonstrated and the paper finds that the success of a complex project like the MEMS‐VSC requires the marriage of basic science expertise found in academia and the technical expertise found in industry.

Owing to the need to test fly the RPVs at low altitudes for video documentation while performing high AOA maneuvers, the attrition of the RPVs becomes the dominant factor to the pace of research.

MEMS sensors and actuators can be used to augment flight control at high AOA, where conventional control surfaces typically experiences reduced effectiveness. Separately, the lessons learned from the integration efforts of this research provide a potentially near parallel case study to the development of ornithopter‐based micro aerial vehicles.

This is the only research‐to‐date involving the demonstration of the MEMS‐VSC on real‐world flight vehicles.