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A novel approach to reduce the peak lift and pitching moment on a plunging airfoil is investigated through force, moment, and velocity measurements. This approach, unlike previous investigations of delayed flow separation and leading-edge vortex suppression, uses forced separation through deployment of a minitab near the leading edge. The device can be activated for short time intervals during a gust encounter or unsteady maneuver at the expense of short-duration drag increase. Depending on the frequency and the amplitude of the wing motion and the mean angle of attack, roll-up of vorticity and the formation of a vortex can be delayed or even prevented. This change in the vortex dynamics provides effective lift and moment alleviation for post-stall angles of attack and for low reduced frequencies. In contrast, at low angles of attack, the separated shear layer may roll up for the manipulated flow, resulting in vortex shedding, and lift and nosedown pitching moment increase. These two distinct flow regimes cause decreased or increased lift force, with the most effective frequencies scaling with the reduced frequency. In contrast, the borderline between the two regions scales with the Strouhal number based on amplitude and, in particular, with the minimum effective angle of attack during the cycle. The transient response was studied by investigating impulsively started plunging oscillations. During the first cycle, lift reduction is achieved for all frequencies within the range tested.

To investigate the effect of sweep on the leading-edge vortex of high aspect ratio plunging wings, this article presents force, bending moment, and volumetric velocity measurements for high aspect ratio AR = 10 wings with sweep angles of 0° and 40°. The effect of the sweep angle on the bending moment is the largest at the minimum effective angle of attack. This is because as the leading-edge vortex sheds it moves inboard on the unswept wing while moving outboard on the swept wing. Where the leg of the leading-edge vortex connects with the wing there is significant three-dimensional flow. The axial velocity along the vortex filament, which may be towards to the wing tip or the wing root (reversed flow), exhibits increasing magnitude as the effective angle of attack decreases and the vortex filament deforms. Reversed axial flow along the vortex filament has the largest magnitudes for the unswept wing. In the vortex core, jetlike, wakelike, and uniform axial velocity profiles were observed. Unlike the classical vortex breakdown, the transition from the jetlike to the wakelike axial flow does not appear to be abrupt. The measurements also revealed evidence of spanwise instabilities in the leading-edge vortex filament.

The peak loads experienced by aircraft of all scales will typically be during gusts, turbulence or extreme manoeuvres. Understanding the aerodynamic response to these transient disturbances is therefore crucial, particularly when Leading-Edge Vortices (LEVs) occur. This fundamental study investigates the aerodynamic response to a wide range of transient plunging motions. The peak loads exhibited a strong dependence to motion amplitude yet remained relatively insensitive to motion duration. Within the parameter range tested (motion duration of T i 20 τ , or equivalent reduced frequency k i 1, and plunge amplitude of α p l , p e a k ≤ 30°), the peak lift did not exceed that of the quasi-static thin airfoil theory prediction, permitting its use as a safe limit for structural design. The normalized peak lift change displayed weak collapse with the timescale of the motion and instead showed better correlation with the non-dimensional plunge rate. The peak pitching moment scales well with plunge rate according to the theoretical prediction due to the added-mass component for plunge-up motions, but quickly diverges for plunge-down motions. At post-stall angles of attack, large-scale vortex shedding was observed and caused decaying oscillations in the loads long after the transient motion ends. For both a NACA 0012 and flat plate airfoil, the first vortex shedding cycle after the transient motion occurs around the subharmonic of the static shedding frequency. Subsequent shedding cycles then increase in frequency and asymptotically approach the static shedding frequency in around 15 to 20 convective times. This is the first study to experimentally quantify this behavior and is an aspect currently missing in existing reduced-order models, which could be significant for the prediction of successive transient disturbances. Finally, Reynolds number insensitivity was demonstrated for transient disturbances between 20,000 and 150,000, even for post-stall angles of attack where large-scale vortex shedding can occur.

In this study lift, pitching moment and particle image velocimetry (PIV) measurements are presented for a NACA 0012 airfoil with a Reynolds number of 20,000 at a geometric angle of attack in the range 0° to 20° undergoing a transient plunging motion. The transient motion is defined by a Strouhal number based on the chord length up to 0.60 and effective plunge angle amplitude from -30° to 30°. The largest peak in both lift and pitching moment was observed during the motion for all cases. This was attributed primarily to circulatory effects for the lift and added-mass effects for pitching moment. PIV showed the presence of a coherent leading-edge vortex (LEV) at positive plunge angle amplitudes which had the biggest effect on the pitching moment during motion. Substantial peaks in lift and pitching moment were observed for post-stall geometric angles of attack after the end of motion. This is postulated to be due to strong interactions between the leading and trailing-edge vortices. An increase in plunge angle amplitude caused an increase in the magnitudes of lift/moment change. This shedding phenomena was found for both positive and negative plunge angles and showed remarkably similar peak relative timings, highlighting the insensitivity of the shedding process to the initial transient forcing.

For many embedded and partially-embedded engine systems, the complexity of the flow field associated with convoluted intakes presents an area of notable research challenges. The convolution of the intake duct introduces additional flow distortion and unsteadiness which must be understood and quantified when designing the turbo machinery components. The aim of the current work is to investigate the capabilities of modern computational methods for these types of complex flows, to study the unsteady characteristics of the flow field and to explore the use of proper orthogonal decomposition methods to understand the nature of the unsteady flow distortion. The unsteady flow field for a range of S-duct configurations has been simulated and assessed using a delayed detached eddy simulation method. The configurations encompass the effects of Mach number, Reynolds number and S-duct centre line offset distance. Analysis of the conventional distortion criteria highlights the main sensitivities to the S-duct configuration and quantifies the unsteady range of these parameters. These results illustrate the strongly dynamic nature of the flow field for both total pressure as well as swirl based distortion. Analysis of the unsteady flow field shows signature regions of unsteadiness which are postulated to be related to the classical secondary flows as well as to the stream wise flow separation. The more aggressive duct, with a larger centre line offset, shows some similar characteristics, but the unsteadiness is more broadband and the distinction between these two mechanisms is less clear. A proper orthogonal decomposition of the total pressure field at the duct exit identifies the underpinning flow modes which are associated with the overall total pressure unsteadiness distributions. For the more aggressive duct, the flow modes are notably different and highlight the reduced demarcation between the unsteady flow field mechanisms