Two major types of stresses affect the crack initiation in the trailing-edge (TE) adhesive joint of a rotor blade: thermal residual stresses that develop during manufacture, and mechanical stresses due to operating loads. Although giving consideration solely to the longitudinal stress component, which contributes mainly to the fatigue, is state-of-the-art, the current design guideline encourages designers to also take account of other stress components, i.e., peel and shear stresses. Hence, this research investigates the impact of the multi-axial stress state due to cyclic loading on the tunneling crack initiation in the TE adhesive joint in order to develop an engineering approach for predicting the number of load cycles toward crack initiation. To this end, a two-dimensional (2D) finite element (FE) model of the TE adhesive joint to approximate the asymptotic stress field at the bi-material corner is introduced. The model takes account of the adhesive layer’s free-edge geometry, as well as the boundary conditions, and the multi-axial thermal and mechanical internal loads in a blade. The validity of the approach was proved in a cyclic full-scale test through the good agreement between the prediction of the approach parametrized with simulated test loads and observations on crack initiation made during the cyclic test. The thermal residual stress at the inner adhesive edge dominated the crack initiation. • Presentation of an engineering approach which combines the approximation of the asymptotic stress field by means of a 2D FE model with a probabilistic stress-life (S-N) model by means of a critical distance approach. • Validation of the approach developed to predicting crack initiation on the full-scale blade level. • Identification of fatigue-dominating load components on crack initiation. • Determination of probabilistic S-N from static and cyclic experiments of neat adhesive material.

Numerical Analysis of a Double-Lap Adhesive Joint for Strengthening Wind Turbine Blade Trailing Edges

To alleviate the adverse effects of the flow-field structure caused by interstage sealing structures on the aerodynamic characteristics of compressor cascades, a blade-root through-slot structure was designed in this study. The structure links the pressure surface to the suction surface of the blade. Numerical simulation techniques were utilized to investigate the process. In this process, the through-slot structure enhances corner separation across varying jet positions, jet heights, and jet widths. The results indicate that the high-speed fluid ejected by the through-slot configuration can suppress the accumulation of low-energy fluid at the suction root. It can also alleviate blockages in the cascade passage and reduce the range of separation vortices and recirculation zones on the suction side. Consequently, the flow loss due to separation is reduced. As the through-slot jet progresses from the blade leading edge to the trailing edge, its restraining impact on the low-energy fluid cluster gradually diminishes. This leads to a corresponding reduction in its effect on the total pressure loss. With an increase in the slot height, the restraining impact on corner separation and total pressure loss first rises and then falls. As the through-slot height increases, the suppressive effect on corner separation and loss initially intensifies and then weakens. As the through-slot width increases, the suppressive effect on corner separation and total pressure loss increases steadily. Compared to the original compressor cascade, the through-slot configuration attains peak performance at 25% chord length, with a height of 6% height and a width of 10 mm, reducing the total pressure loss coefficient by 19.22%. Furthermore, as the incoming flow incidence angle enlarges, the enhancement impact of the through-slot configuration on cascade performance initially intensifies and then diminishes. The peak enhancement impact occurs at a 0° incidence angle. At this angle, the configuration can reduce flow loss by 16.72% compared to the original, significantly improving the aerodynamic performance of the high-load compressor cascade.

Abstract Hot Flow Anomalies (HFAs) are transient foreshock structures generated by solar wind discontinuity‐bow shock interactions, with their evolution only confirmed near Earth. Using MAVEN and Tianwen‐1 joint observations, we study the evolution of a Martian HFA event detected on 17 September 2022 (dayside by MAVEN, nightside by Tianwen‐1). The HFA retains a young‐type feature with negligible core magnetic fluctuations and little thickness change, but its trailing edge peak magnetic field drops by 46% (6.86–3.67 nT). We attribute the lack of young‐to‐mature evolution to Martian small bow shock, enabling HFA propagation upstream of the quasi‐perpendicular shock, precluding full development as Earth's HFAs.

Innovative 3D printed soft actuators for morphing wing trailing edges: Design, deformation, and aerodynamic evaluation

Shark skin grooves, known to reduce hydrodynamic drag, have inspired riblet structures for flow control. This study investigates their application to airfoils, where flow separation at high angles of attack (AOA) compromises aerodynamic stability and wind turbine performance. Numerical simulations were conducted using the SST k–ω model in ANSYS Fluent to analyze riblets placed on the suction surface (SS) of an airfoil. The riblets—oriented perpendicular to the flow—have a fixed height and width of 1 mm, with total lengths varying from 0.1, 0.2, 0.5, and 0.7 of the chord length. The influence of riblet geometry on trailing-edge (TE) vortex shedding and drag reduction under stall conditions is examined in detail. The results indicate that appropriately sized riblets suppress secondary vortex formation and extend the 2S vortex-shedding regime. Conversely, poorly dimensioned riblets can advance Hopf bifurcation in the wake. Analysis of the transient boundary layer structure reveals that the suppression of vortex shedding is primarily due to riblets attenuating fluid pulsation and Reynolds stresses caused by turbulent bursts.

In the ionization injection regime of laser wakefield acceleration (LWFA), the control and optimization of electron bunches is a major concern. We address this by demonstrating the generation of tunable electron bunches in LWFA using temporally asymmetric laser pulses in the ionization injection regime. The introduction of pulse asymmetry modifies the wakefield structure, which affects the accelerating field and electron trapping in the ionization injection regime. Our 2D PIC simulations show that fast trailing edge asymmetry of the laser pulse produces narrow electron bunches with improved energies, whereas slow trailing edge asymmetry favors prolonged injection and yields broader bunches with higher charge. These results reveal that the temporal shape of the driving laser pulse serves as a control parameter for tailoring the charge, duration, and quality of the injected electron beam. This approach offers a versatile pathway for tunable electron bunch generation in LWFA without requiring changes in plasma density or gas composition. This paves the way for the development of compact, high-brightness electron sources for advanced accelerator and radiation applications.

Since the early 1990s, numerous theoretical methods have been proposed to predict Mach stem height in steady supersonic shock reflections by assembling sub-models for local flow structures, including incident/reflected shocks, the triple point, the slipline, and Mach stem curvature. We constructed an updated model and employed it as a benchmark to evaluate the performance of various sub-models corresponding to typical flow regions. The results show that the curved assumption for the free part of the slipline outperforms the straight-line approximation, considering the differences in regions after the reflected shock can improve the predictive accuracy, while using compatibility relations in the interactive part of the slipline is superior to the wave reflection model and better captures the linear slope of Mach stem height with wedge trailing edge height. Nevertheless, prediction errors in the slope and systematic biases in the overall Mach stem height prediction persist. To address these shortcomings, we developed a calibrated scaling law for the coefficient of a linear Mach stem model. Grounded in asymptotic reasoning and high-fidelity numerical simulations, this law yields a compact, easy-to-implement expression that achieves substantially higher accuracy than existing analytical composite models across the full parameter space. It retains well-established limiting cases, clarifies how inadequate sub-modelling degrades prediction accuracy, and provides uncertainty estimates for practical engineering applications.

Understanding the vortex interactions and wake transitions for flapping flexible foils is important because of their increased usage in bioinspired aquatic and aerial robotic propulsors. Although wake transitions have been studied for rigid foils, we experimentally investigate how flexibility alters the transitions and vortex interactions for flexible foils, which are closer to the natural flapping foils in fish, birds and insects. We conduct the experiments in a flowing soap film on a pitching airfoil with a flexible filament at its trailing edge (TE). We find that, apart from the Strouhal number ( ${\textit{St}}$ ), flexural rigidity ( ${\textit{EI}}$ ) is important to determine the transitions. We vary ${\textit{EI}}$ of the flexible filament by three orders of magnitude and also investigate an extreme case of ${\textit{EI}} \rightarrow \infty$ . Flexibility triggers the shedding of multiple small ‘secondary vortices’ (SVs) along with big ‘primary vortices’ (PVs), unlike only PVs for the rigid foil. Continuous deformations of the flexible filament play crucial roles in determining the interaction of boundary layer vortices and trailing edge vortices and, ultimately, the generation and evolution of PVs and SVs. We identify five vortex interaction mechanisms (VIMs). Depending on how SVs interact with PVs, the wake assumes different patterns. We construct the ${\textit{St}}$ – ${\textit{EI}}$ phase maps for wake transitions and newly identified VIMs. We devise a non-dimensional parameter $\varUpsilon$ , referred to as ‘Yashavant number’. One order increase in $\varUpsilon$ reduces the number of VIMs by one. Instead of following the usual transition route, the flexible foil reveals counterintuitive transition trends that strongly depend on the filament ${\textit{EI}}$ .

We used scat analysis to study the foraging habits of kenkéknem (Ursus americanus, American Black Bear) at Sun Peaks Ski Resort in Skwelkwék’welt, south-central Secwepemcúl’ecw, from May to late August 2023. The 20 scats (three in May, 10 in June, four in July, and three in August) showed the bears consumed largely green vegetation in the spring (1 May–21 June), scwicwéye (ants, primarily wood-nesting species) in early summer (22 June–31 July), and máts̓pe7 (wasps) and berries in late summer (1–30 August). Some vertebrate predation on voles (Microtus spp.) and other rodents was found. The most common vegetation in scats in spring was xwixwyúy̓sten (Equisetum spp., horsetail), which grows well in wet disturbed environments, such as the edge of ski trails. Wood-nesting ant species provided an important food source for bears in early and late summer. Given the importance of ants to the summer diet of bears, we recommend forest management in Skwelkwék’welt consider the importance of woody debris in providing suitable ant habitat.

Wing design in the last decades directed efforts and studies to improve the aircraft performance allowing fuel reduction, which implies more ecological and sustainable solutions and cost reduction for air transportation. Efficient design with a morph wing can fulfill such objectives. Different concepts of morph wing have been numerically investigated in this study proposing different wing configurations with the aim to improve aircraft performance in particular for the rolling maneuver. The assessment is performed by using an aerodynamic analysis based on a low fidelity 2D method such as the panel method, in combination with a 3D analysis using vortex lattice method, without considering elastic structural effects. The main morph wing concept is based on the constant morph of a part of the wing tip which has the aim of acting as a control surface hence as a substitution of the aileron for the rolling maneuver. The twist will then be applied linearly decreasing to the center of the wing. Within this concept three main configurations were evaluated: (a) the twist of the fixed airfoil section (Morph type M1 – rib twist), (b) the twist of the aileron which profile is changing according to a curvilinear law based on two morph angles (Morph type M2) and (c) both trailing edge and leading edge are morphed (Morph type M3). A comparison with a correspondent conventional aileron configuration at the same rolling moment coefficient showed an advantage of the morph wing in terms of drag coefficient with reductions that go up to 30%. It was also observed that in some cases such as M1 and M2C a morph deflection lower than 10° produce the same rolling moment coefficient of a typical small aircraft aileron deflection of 25°. Moreover, a parametric evaluation showed an optimum of the span wise parameter y rib to be 40% of the semi span b . Furthermore, a smoother distribution of the lift along the span wise direction will be determined in comparison with a conventional aileron. This also implies a smoother approach to stall conditions that can be beneficial for the pilot.

Airframe noise generated at wing trailing edges and high-lift devices, such as flaps, remains a major challenge during landing, with significant contributions in the low-frequency band of 500-1500 Hz. While solid surfaces reflect this acoustic energy, metallic porous materials can effectively absorb it through viscous and thermal dissipation within their internal pore structure. To address this, the present study examines the acoustic absorption characteristics of open-cell AlSi porous cylinders featuring controlled pore diameters between 0.3 mm and 2.25 mm. Measurements were conducted in an acoustic impedance tube according to the ISO 10534-2:2023 standard, using six cylindrical samples (28 mm diameter, 70 mm length). Two sets of measurements were performed for each sample (front and rear faces), and the average values were used. The findings indicate that the normal-incidence sound absorption coefficient α rises as pore size increases, reaching 0.93-0.97 at low frequencies of 500-700 Hz for the samples with the largest pores (1.8-2.25 mm). These results indicate that open-cell AlSi alloys offer strong low-frequencies sound absorption, positioning them as promising options for aeroacoustic noise mitigation, including applications such as porous trailing edge and hybrid flap designs.

The interaction of wings with vertical gusts is still not well understood, especially at the lower Reynolds numbers relevant to small Uncrewed Aerial Systems. Studies at higher Reynolds numbers with relatively slowly developing gusts have shown that morphing wing camber is effective at mitigating gusts, while at low Reynolds number, gusts are understood to be a leading edge phenomena relatively unaffected by the trailing edge. This study evaluates one-way trailing edge morphing and oscillating trailing edge motions with respect to the ability to mitigate vertical gust impacts. The results show that, for the gust condition tested, one-way trailing edge morphing at low Reynolds number acts similarly to higher Reynolds number studies. However, the long duration of the gust tested showed that one-way trailing edge morphing can only mitigate gusts for a certain length of time, after which stall occurs. Oscillating the trailing edge is shown to mitigate gusts in a similar fashion to prior studies on oscillating wings, where predictable oscillatory lift is possible. However, oscillating the trailing edge is only able to generate a predictable lift behavior for a certain length of time, after which the leading edge effects of the gust dominate over trailing edge motion. Higher oscillation frequency and Strouhal number are shown to have a detrimental effect on the predictability of lift during gust interactions, but have no effect altering the time at which highly unsteady stall dynamics occur.

Serration manufacturing effects on propeller trailing edge noise mechanisms

Particle image velocity of the gust-induced flowfield and lift for the elastic wing are measured simultaneously in the wind tunnel test. The experimental results indicate that the wing’s pronounced elastic motion alters the flow acceleration performance of the gust, which brings a challenge to control at varying gust frequencies. To address this, a deep reinforcement learning (DRL) control is presented for the first time in a gust alleviation wind tunnel test, which is not only trained but also tested using only experimental data. A GLA wind tunnel test is conducted with a seamless morphing trailing edge (TE). From the perspective of the phase offset between the gust and the morphing TE, the controller can adjust the TE morphing to its optimal phase across different gust frequencies, varying inflow velocities, and gust amplitudes. Hence, the aerodynamics induced by the morphing TE can cancel out the gust-induced lift, which can be suppressed significantly. Specifically, at U∞=10 m/s, GR=0.08, and fg=3.0 Hz, the gust-induced lift coefficient is alleviated by 76.0%. Furthermore, the presented DRL controller can also take effect under superimposed multifrequency sinusoidal gusts.

This paper develops a 2D steady flow hydrodynamic analysis model for dual foils via potential theory-coupled boundary element method. Each foil embeds a vortex sheet to achieve required circulations, with the total potential split into incoming flow potential, source-induced disturbed potential, and vortex sheet-driven disturbed potential. In conjunction with two auxiliary equations obtained from the Kutta condition at each foil's trailing edge, the boundary integral equation is incorporated into a unified numerical matrix; these equations are solved simultaneously to determine both the potential distribution on the foils and the vortex strengths. Comprehensive convergence analyses have been carried out, while the interaction effects between the two foils are discussed in detail. Additionally, the hydrodynamic performances of foils equipped with slats or flaps are also presented.

A numerical study of a generic projectile with a solid base at Mach 2 is used to highlight the impact of fins on the wake across a range of roll and pitch angles from zero to 12 deg. To accommodate this wide range of parameters, with and without fins, the Reynolds-averaged Navier–Stokes equations are solved with a realizable k-ϵ turbulence model and validated against available experimental data. The effects of fins on the wake are characterized by a detailed analysis of the emergent shock and vortical structures, as well as the consequent aerodynamic loads. At a zero angle of attack, the fin trailing edge introduces new shocks/expansions that fundamentally alter the shape and strength of the recompression shock. Pitch and roll further modify the shocks and vortices, which can now be distinguished as forebody, wingtip, and trailing-edge components. At certain pitch and roll conditions, wingtip vortices bound and split the recompression shock into multiple sections, changing their inception and subsequent trajectory. Under pitch conditions, fins effectively flip the trailing-edge vortices, which now form on the lower side of the base surface rather than the upper side, with a reversal of streamwise vorticity direction. The results identify the main features of interest for further exploration with scale-resolving simulations.

The vortex structure formed by the trailing edge cutback lip induces intense unsteady effects on the cutback surface, thereby significantly influencing the distribution of coolant films, which in turn affects cooling performance. This study experimentally investigates the unsteady film cooling behavior on a trailing edge cutback surface under blowing ratios (BR) ranging from 0.5 to 1.5. Using hot-wire anemometry (HWA) and fast-response pressure-sensitive paint (fast-PSP), the velocity fluctuation spectra and instantaneous cooling effectiveness are analyzed. Spectral analysis identifies increasing vortex shedding frequencies with the rising blowing ratio, while a transition from mainstream-dominated to coolant-dominated flow occurs above BR = 0.7. Time-averaged cooling effectiveness remains excellent (η > 0.9) upstream of X/L = 6, but declines thereafter, with higher blowing ratios improving coolant coverage. However, instantaneous film cooling effectiveness distributions reveal significant unsteadiness, characterized by fragmented film patterns. Spectral proper orthogonal decomposition (SPOD) and cross correlation analysis indicate an association between the degradation of cooling performance and the presence of Brown–Roshko (B–R) vortex structures. Furthermore, the dominant frequencies identified via SPOD agree with the HWA results. At lower blowing ratios, coherent mainstream vortices induce strong fluctuations, while higher blowing ratios produce finer coolant vortex structures that enhance film cooling performance. These findings provide essential insights not only into steady-state conditions but more importantly into the unsteady behavior of trailing-edge cutback cooling, thereby informing the optimization of blade cooling design.

Numerical study on mixing and ignition enhancement characteristics in an air-intake strut supersonic combustor

Shock wave interference is a common phenomenon in hypersonic vehicles. It modifies the load distribution across the wing surface, thereby affecting the aeroelastic characteristics. This paper investigates the influence of shock wave interference location and intensity on the flutter characteristics of a hypersonic airfoil using a method based on a reduced-order model. The flutter mechanism under shock wave interference is revealed through root locus analysis and the unsteady aerodynamic work method. When the shock wave interference moves from the leading edge to the midpoint of the airfoil, the critical flutter velocity gradually increases but remains lower than the critical flutter velocity in the absence of shock wave interference. When the shock wave moves from the midpoint to the trailing edge, the critical flutter velocity reaches its maximum value and then gradually decreases. The influence of shock wave interference intensity on the flutter characteristics exhibits monotonic behavior. Finally, this paper finds that, in order to improve the aeroelastic stability of hypersonic airfoils, the first step is to move the shock wave interference location away from the front area of the airfoils. Second, the flutter mode can be changed to increase the critical flutter velocity.