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Copyright of Machine Tool & Hydraulics is the property of Guangzhou Mechanical Engineering Research Institute (GMERI) and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

This study presents a bionic structure-based liquid cooling plate designed to address the heat generation characteristics of prismatic lithium-ion batteries. The size of the lithium-ion battery is 148 mm × 26 mm × 97 mm, the positive pole size is 20 mm × 20 mm × 3 mm, and the negative pole size is 22 mm × 20 mm × 3 mm. Experimental testing of the Li-ion battery's heat generation model parameters, in conjunction with bionic structure and micro-channel features, has led to the development of this innovative cooling system. The traditional bionic liquid cooling plate's structure is often singular; however, the flow path of the liquid cooling plate designed in this paper is based on the combination of the distribution of human blood vessel branches and the structure of insect wing veins. The external dimension of the liquid cooling plate is 152 mm × 100 mm × 6 mm (length × width × height). Utilizing numerical simulation and thermodynamic principles, we analyzed the heat transfer efficacy of the bionic liquid cooling module for power batteries. Specifically, we investigated the impact of varying coolant flow rates and the contact radius between flow channels on the thermal performance of the bionic battery modules. Our findings indicate that a liquid flow rate of 0.6 m/s achieves a stable maximum surface temperature and temperature differential across the bionic battery liquid cooling module, with a relatively low overall system power consumption, suggesting room for further enhancement of heat transfer performance. By augmenting the contact radius between flow channels, we observed an initial increase in the maximum surface temperature, temperature differential, and inlet–outlet pressure differential at a flow rate of 0.2 m/s. However, at flow rates equal to or exceeding 0.4 m/s, these parameters stabilized across different design Scenarios. Notably, the pump power consumption remained consistent across various scenarios and flow rates. This study's outcomes offer valuable insights for the development of liquid-cooled battery thermal management systems that are energy-efficient and offer superior heat transfer capabilities. [ABSTRACT FROM AUTHOR]

Purpose: With the advancements in electric vertical take-off and landing (eVTOL) aircraft technology such as batteries, mechanisms, motors, configurations and so on, designers and engineers are encouraged to create unique and unconventional configurations of eVTOL aircraft to provide better capabilities and higher efficiencies to compete in the market. The box fan-in-split-wing tiltrotor eVTOL aircraft is an innovative design that aims to address the aerodynamic inefficiencies such as propeller effects in cruise and engine mounts drag that existed in traditional eVTOL aircraft designs such as vectored thrust, rotorcraft, lift + cruise and multi-copter configurations. This paper aims to propose a multi-disciplinary design process to conceptually design the box fan-in-split-wing Tiltrotor eVTOL aircraft. Design/methodology/approach: An unconventional methodology was used to design the UAM aircraft, and the following parameters are considered: capable of vertical take-off and landing, highly aerodynamic with a high lift-to-drag ratio, low Cd0 modern and appealing, rechargeable or battery swappable and feature to minimise or negate propeller drag. A heavy emphasis on improving performance and weight based on aerodynamics was enforced during the conceptual design phase. MAPLA and XFOIL were used to identify the aerodynamic properties of the aircraft. Findings: Upon determining the key parameters and the mission requirements and objectives, a list of possible VTOL configurations was derived from theoretical and existing designs. The fan in the wing/split wing was selected, as it could stow the propellers. A tiltrotor configuration was selected because of its ability to reduce the total number of lift props/motors, reducing powerplant weight and improving aerodynamic efficiency. For the propulsion configuration, a battery–motor configuration with a hexa-rotor layout was chosen because of its ability to complement the planform of the aircraft, providing redundant motors in case of failure and because of its reliability, efficiency and lack of emissions. Coupled with the fan-in-wing / split wing concept, the box wing seamlessly combines all chosen configurations. Originality/value: The box fan-in-split-wing Tiltrotor eVTOL aircraft aims to address the aerodynamic inefficiencies of earlier designs such as propeller effects in cruise and engine mounts drag. The potential benefits of this aircraft, such as increased range, endurance and payload capacity, make it an exciting prospect in the field of Urban Air Mobility. [ABSTRACT FROM AUTHOR]

This paper investigates the effects of the chordwise fence on the spanwise change in aerodynamic characteristics of an aircraft wing with a different taper ratio for varying angles of attack. The investigation was carried out for the tapered wing with different taper ratios of 0.41, 0.6, and 0.75. The wing is tested in a subsonic, low turbulence wind tunnel at the free stream Reynolds number of 2.3×105 for various angles of attack ranging from α = 0° to 45°. The baseline wing model is attached to a fence of different diameters of 1.5 and 2.5 mm at a plane equal to the root and tip chord. There are pressure ports spread across the span of the wing, and the corresponding surface pressure is measured using the MPS 4264 miniature pressure scanner. The surface pressure measured is analyzed further for the variation of the aerodynamic characteristics of the wing. The presence of a fence on the tapered wing forms an efficient flow control device that delays the flow separation, thereby delaying stall angles and preventing the steep transition of the favorable pressure gradient to the adverse pressure gradient at the stall. The presence of a fence on the wing surface has considerably increased the lift coefficient, and the stall is significantly delayed for a least taper ratio wing. The fence has suppressed the interaction of the leading-edge vortices with the tip vortices; thereby, the spanwise flow from the root chord to the tip chord is controlled. [ABSTRACT FROM AUTHOR]

The occurrence of probabilistic common cause errors (PCCEs) in man‐machine systems can lead to multiple human errors being affected by common causes and will contribute greatly to the reliability of the system. In this paper, A reliability modeling method based on event tree (ET)‐fault tree (FT) model for man‐machine systems subjected to PCCEs is proposed. Under the influence of PCCEs, risk factors in the system are not independent, and human errors affected by the same common cause will occur with different probabilities. To describe the dependencies among risk factors, a PCCE gate is proposed to extend FTs in the ET‐FT model. To analyze the impact of common causes on the human error probabilities and quantify the extended FT, an explicit method and an implicit method based on the Cognitive Reliability and Error Analysis Method are proposed. The proposed quantification methods consider the different relationships among multiple common causes in a single model. Finally, the proposed methods are validated in the reliability assessment of the emergency response system under malfunction of the solar wing mechanism. [ABSTRACT FROM AUTHOR]

The object of this study is the technological preparation of the production of a light aircraft wing using reverse engineering technology. The subject of research is a quality indicator–the geometric accuracy of manufacturing the convex-concave parts of aerospace technology. Calculations of geometric accuracy were performed for the program-instrumental method of co-ordination. As the experimental part it was taken the worn out wing tip of a light aircraft. The following results were obtained. An approach for specifying the aerodynamic airfoil and cross sections of the wing tip when constructing its digital model has been proposed. A 3D scanning of the wing tip with the formation of a digital portrait in STL format, as well as its refinement into a STEP format, using organic and mechanical methods, was accomplished. A digital mock-up of the wing tip was built taking into account the geometry of the aerodynamic airfoil in cross sections as well as a digital mock-up of the form (mould) for its manufacture according to the polygonal model, which was created by the organic method due to it had the highest dimensional accuracy. It was determined that the maximum deviation of the actual wing contour from the theoretical one was as follows: the upper deviation was 0.84 mm, the lower deviation was –0.65 mm. The maximum deviation of the actual wing contour from the theoretical one was ±0.3 mm. The expected (calculated) errors did not exceed the specified value of the tolerance on the wing outer contour that equal to ±1.0 mm, thus, the adopted method of assembling the wing under the conditions of co-ordination by the program-instrumental method ensured the specified geometric accuracy. The results of experimental studies confirmed the adequacy of the proposed approach for determining the aerodynamic airfoil of the cross-sections of the digital mock-up of convex-concave parts for aerospace technology during their technological preparation for production with the use of reverse engineering. [ABSTRACT FROM AUTHOR]

Copyright of Ingenius, Revista Ciencia y Tecnología is the property of Universidad Politecnica Salesiana and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Variable camber wing technology stands out as the most promising morphing technology currently available in green aviation. Despite the ongoing advancements in smart materials and compliant structures, they still fall short in terms of driving force, power, and speed, rendering mechanical structures based on kinematics the preferred choice for large long-range civilian aircraft. In line with this principle, this paper introduces a linkage-based variable camber trailing edge design approach. Covering coordinated design, internal skeleton design, flexible skin design, and drive structure design, the method leverages a two-dimensional supercritical airfoil to craft a seamless, continuous two-dimensional wing full-size variable camber trailing edge structure, boasting a 2.7 m span and 4.3 m chord. Given the significant changes in aerodynamic load direction, ground tests under cruise load utilize a tracking-loading system based on tape and lever. Results indicate that the designed single-degree-of-freedom Watt I mechanism and Stephenson III drive mechanism adeptly accommodate the slender trailing edge of the supercritical airfoil. Under a maximum cruise vertical aerodynamic load of 17,072 N, the structure meets strength requirements when deflected to 5°. The research in this paper can provide some insights into the engineering design of variable camber wings. [ABSTRACT FROM AUTHOR]

Purpose: Endurance time is an important factor limiting the progress of flapping-wing aircraft. In this study, this paper developed a prototype of a double-wing flapping-wing micro air vehicle (FMAV) that mimics insect-scale flapping wing for flight. Besides, novel methods for optimal selection of motor, wing length and battery to achieve prolonged endurance are proposed. The purpose of this study is increasing the flight time of double-wing FMAV by optimizing the flapping mechanism, wings, power sources, and energy sources. Design/methodology/approach: The 20.4 g FMAV prototype with wingspan of 21.5 cm used an incomplete gear flapping wing mechanism. The motor parameters related to the lift-to-power ratio of the prototype were first identified and analyzed, then theoretical analysis was conducted to analyze the impact of wing length and flapping frequency on the lift-to-power ratio, followed by practical testing to validate the theoretical findings. After that, analysis and testing examined the impact of battery energy density and efficiency on endurance. Finally, the prototype's endurance duration was calculated and tested. Findings: The incomplete gear facilitated 180° symmetric flapping. The motor torque constant showed a positive correlation with the prototype's lift-to-power ratio. It was also found that the prototype achieved the best lift-to-power ratio when using 100 mm wings. Originality/value: A gear-driven flapping mechanism was designed, capable of smoothly achieving 180° symmetric flapping. Besides, factors affecting long-duration flight – motor, wings and battery – were identified and a theoretical flight duration analysis method was developed. The experimental result proves that the FMAV could achieve the longest hovering time of 705 s, outperforming other existing research on double-wing FMAV for improving endurance. [ABSTRACT FROM AUTHOR]

As the population of the world's major cities increases, urban ground transportation capability reaches its limits. Consequently, electric vertical takeoff and landing aircraft have been developed to address this issue. Modern vertical takeoff and landing designs vary in configurations, each possessing distinct aerodynamic properties. In this work, we focus on the interaction between rotors and fixed wings in a parallel layout, using a prototype with an unconventional configuration as an example. The interaction between the slipstream of rotors and the wake of fixed wings is analyzed using the vortex particle method. Actuator surface model is used to deal with the boundary problem between the wing and the rotors. The applicability of the simulation method adopted in this paper is verified through experimental tests on the thrust and torque of the rotors. The results indicate that under the condition of forward flight at zero angle of attack, the existence of rotor slipstream induces a significant increase by three to four times the local lift on the wing. Correspondingly, under vertical takeoff and hovering conditions, the existence of the rotor slipstream causes the local effective velocity and circulation of the wing to increase. On the other hand, at zero angle of attack and a forward flight, the lift coefficient of the rear rotors increases slightly due to the presence of the wing wake. The lift distribution of the rear rotors along the radial direction also changes greatly. [ABSTRACT FROM AUTHOR]

This paper introduces three new gust load alleviation control strategies, which are based on variable camber technology, to achieve the design requirements for the safety, economy, and comfort of modern civil aircraft. These strategies involve adjusting the inner and outer flap differential of wings, spoilers, and ailerons to alter the aerodynamic shape of the aircraft structure, resulting in drag reduction and gust load alleviation during stepped cruise flights. Furthermore, a logic design is provided for both the cruise drag reduction function and the gust load alleviation function. To validate the effectiveness of the proposed gust load alleviation control strategies, simulation results are presented, utilizing ride quality and wing root bending moment changes as the evaluation criteria. The presented approaches ultimately lead to the development of an optimal scheme that meets the desired design requirements. [ABSTRACT FROM AUTHOR]

Cables are critical in engineering structures for load‐bearing, electronic connection, and mechanical transmission. Various cable‐crawling robots (CCRs) have been developed to perform scheduled inspection or convey supplies. Most existing CCRs are often actuated by motors and used in large‐scaled engineering structures. The heavy bodies of these CCRs can cause damage or even casualties once slippage or drop occurs. A small and lightweight CCR that can crawl on thin cables is highly demanded for safety inspection in narrow and confined inner spaces of engineering structures. Herein, a soft CCR (weight, 2.1 g; length, 43 mm) is developed by utilizing multilayered dielectric elastomer actuators. Compared with existing solutions, this CCR achieves crawling on thin cables (diameter: <1 mm) while crawling fastest (horizontal: 0.72 body length per second). The CCR is also capable of transporting objects (horizontal: 3.69 times its own weight; vertical: 0.76 times its own weight), climbing upward on a vertical cable, and locomoting across the water–air interface. The CCR is also demonstrated to crawl on a slack cable and circular/spiral cables. Finally, the soft robot, equipped with an endoscope, demonstrates inspections on a tensegrity structure as well as in an airplane wing model with a preplaced cable. [ABSTRACT FROM AUTHOR]

The 4R-UAV project aims to develop aerodynamically efficient and environmentally friendly UAVs based on the 4R Circular Economy principle. In this study, the feasibility of the application of PFCDs (Passive Flow Control Devices) was investigated for the small-scale low-speed 4R-UAV wing. The application of PFCDs for small-scale UAV wings is relatively unexplored. Two PFCD types, i.e., MVGs (Micro Vortex Generators) and winglets, were considered for the investigations. In the stepwise investigations, the aerodynamic performance of the MVGs and the winglets was analyzed for the short-span 4R-UAV wing, which was developed from the aerodynamically optimized airfoil SG6043mod. MVGs enhanced the wings near stall properties (especially maximum lift coefficient) and contributed to the delayed wing stall of up to 2°. MVGs manifested the main aerodynamic advantage, which was achieved at the higher angles of attack. Winglets, on the other hand, were found to be extremely effective in cruise conditions with improved pre-stall characteristics. Extensive investigations on winglets were carried out by designing six winglet configurations for the 4R-UAV wing. Blended-type winglets performed well and enhanced pre-stall properties by decreasing the drag (up to 10%) of the wing. The main performance improvement was found in the early angles of attack. In the final step, the combined effect of the integrated PFCDs was analyzed. The final wing (integrated MVGs and winglets) also exhibited enhanced performance with nearly 6% increased lift-to-drag ratio in cruise conditions. The limited aerodynamic advantage achieved from the PFCDs in small-scale UAV applications can be useful in specific (civil or military) missions. Further verifications are planned in the future by means of experimental and flight testing. [ABSTRACT FROM AUTHOR]

The oscillatory pitch motion at the leading edge of a millimeter‐scale flexible sheet on the water surface can generate undulatory locomotion for swimming, similar to a honeybee vibrating its wings for propulsion. The influence of various parameters on such swimming strategy remains unexplored. This study uses magnetic milliswimmers to probe the propulsion mechanics and impact of different parameters. It is found that this undulatory propulsion is driven by capillary forces and added mass effects related to undulatory waves of the milliswimmers, along with radiation stress stemming from capillary waves at the interface. Modifying the parameters such as actuation frequency, pitch amplitude, bending stiffness, and hydrofoil length alters the body waveform, thus, affecting the propulsion speed and energy efficiency. Although undulatory motion is not a prerequisite for water surface propulsion, optimizing body stiffness to achieve a proper undulatory waveform is crucial for efficient swimming, balancing energy consumption, and speed. The study also reveals that the induced water flow is confined near the water surface, and the flow structures evolve with varying factors. These discoveries advance the understanding of undulatory water surface propulsion and have implications for the optimal design of small‐scale swimming soft robots in the future. [ABSTRACT FROM AUTHOR]

The extensive use of lightweight materials in aerial vehicle wings involves structural flexibility phenomena that generate non-negligible deformation effects. This influence is not restricted to big aircraft but also plays a role in smaller aeroplanes and Unmanned Aerial Vehicles (UAVs). Here, we conduct wind tunnel experiments to analyze the effect of passive deformation on the wing model lift slopes. To isolate the deformation effect, we compare rigid wings with a NACA0012 airfoil imposing a prescribed spanwise deformation. We study three levels of deformation: non-deformed, around 2% and 4.5% of tip deflection. Also, we consider the effect of the wing length by using three different semi-aspect ratios (1, 2, and 4), so a total of nine rigid wing models have been analyzed for a range of Reynolds number from 80 × 1 0 3 to 160 × 1 0 3 . Deformed wing models show an increase in lift coefficient compared to non-deformed wing cases. Both deformation levels exhibit a qualitatively similar lift increment. A correlation to predict lift coefficient slope in a flat plate is adapted for a NACA0012 airfoil and validated using our experimental results and literature data. The adjusted correlation can quantify the deformation effect on the lift slope, which is comparable to using a slightly longer wing model. [Display omitted] • The deformed NACA0012 wing models show a lift increase over the non-deformed cases. • The lift slope of NACA0012 wing models are predicted using a flat plate correlation. • The fitted correlation quantifies the effect of deformation on the lift slope. [ABSTRACT FROM AUTHOR]

The digital assembly system is widely used in the automatic positioning of large aircraft components. Accurate posture measurement of the components remains challenging for high-precision assembly of the components with various engineering constraints. Existing methods often require repeated measurements of key features and are susceptible to the pronounced coupled effect, where various constraints mutually influence the solution of pose parameters. To enhance the efficiency and precision of posture measurement, this study proposes an on-line measurement method for the Foldable-wing Digital Assembly System. The key features and several reference points are measured once by the laser tracker. According to the assembly task sequence, a novel posture evaluation model with priority-based multi-constraints is established to calibrate these reference points which will be utilized as the new datum of posture measurement instead of the traditional key features. This model separately satisfies the various engineering constraints, such as coaxiality, gap, and basic positioning accuracy, in order of priority to avoid the coupling effect between different types of constraints. Consequently, features with higher priority can achieve higher accuracy without weight setting. Next, the reference points are automatically measured to calculate the posture parameters. Considering the uncertainty characteristics of the measuring instrument, the registration algorithm based on weighted total least squares adjustment is applied to reduce measurement uncertainty. Monte Carlo simulation results demonstrate the proposed method's ability to enhance positioning accuracy and reduce posture uncertainty, outperforming conventional methods. Finally, the practical application was successfully implemented, and all key features remained within tolerance. • Establish a novel posture evaluation model with priority-based multi-constraints to avoid the constraint coupling effect. • Measurement reference points are utilized as the new datum of posture measurement instead of the traditional key features. • The weighted total least squares adjustment is utilized to reduce the posture uncertainty caused by the laser tracker. [ABSTRACT FROM AUTHOR]

In this study, a control scheme is proposed based on Chaotic Particle Swarm Optimization (CPSO) to enhance the Linear Auto-Disturbance Rejection Controller (LADRC). The focus is on addressing the challenge of high-precision variations in angle-of-attack through dual-motor cooperative control within the lifting wing of a high-speed train. The scheme initiates with the design of a dual-loop structure for LADRC, integrating position and current control. The position loop is further refined. Subsequently, the CPSO algorithm is employed to optimize the parameters of the LADRC controller. Ultimately, the loop is closed by feeding back the position error in the cross-coupled structure to the current loop, thereby achieving high-precision control. The performance of the proposed structure is validated through both Matlab/Simulink simulations and an experimental platform. The experimental results demonstrate that CPSO-LADRC, in comparison to traditional LADRC and Proportion-Integration-Differentiation (PID) control, exhibits an increase in the maximum response time by 3.76 s and 3.3 s, respectively, a reduction in overshoot by 1.12% and 0.8%, as well as a decrease in the maximum synchronization error by 0.45 cm and 1 cm, respectively. These findings validate the effectiveness of the proposed synchronous loop controller method in simplifying computational complexity, enhancing system responsiveness, robustness, and synchronization performance. Additionally, our approach facilitates precise angle-of-attack transformation for the lifting wings of high-speed trains effectively. [ABSTRACT FROM AUTHOR]

Lightweight structures with a high stiffness-to-weight ratio always play a significant role in weight reduction in the aerospace sector. The exploration of non-conventional structures for aerospace applications has been a point of interest over the past few decades. The adaptation of lattice structure and additive manufacturing in the design can lead to improvement in mechanical properties and significant weight reduction. The practicality of the non-conventional wing structure with lattices infilled as a replacement for the conventional spar–ribs wing is determined through finite element analysis. The optimal lattice-infilled wing structures are obtained via an automated iterative method using the commercial implicit modeling tool nTop and an ANSYS workbench. Among five different types of optimized lattice-infilled structures, the Kelvin lattice structure is considered the best choice for current applications, with comparatively minimal wing-tip deflection, weight, and stress. Furthermore, the stress distribution dependency on the lattice-unit cell type and arrangement is also established. Conclusively, the lattice-infilled structures have shown an alternative innovative design approach for lightweight wing structures. [ABSTRACT FROM AUTHOR]

This paper studies static test methods and test loads of Carbon Fiber Reinforced Polymer (CFRP) wing-box, analyzes the test results, and compares the test results of T800 CFRP wing-box from HENGSHEN Co., LTD and TORAY. It comes out that the stiffness of the T800 CFRP wing-box from HENGSHEN is equivalent with that of TORAY, while the strength meets the structural design requirements, which can be applied on the scheduled aircraft structure for further verification. [ABSTRACT FROM AUTHOR]

Purpose: This paper aims to propose a structural sizing approach of an unmanned aerial vehicle (UAV) wing that takes into account the aeroelasticity effects through a fluid–structure interaction analysis. Design/methodology/approach: The sizing approach proposed in this study is an iterative process, each iteration of which consists of two sub-loops, a multidisciplinary analysis (MDA) loop followed by a structural optimization loop. The MDA loop seeks the aeroelastic equilibrium between aerodynamic forces and structural displacements using a fixed-point iteration scheme. Once the equilibrium is reached, the converged pressure loads are used for the structural optimization, which aims to find the structural thicknesses that minimize the wing weight under failure criteria. The two sub-loops are run sequentially in an iterative process until the mass is converged. The analysis models are implemented in open-source software, namely, PANUKL for aerodynamics and MYSTRAN for structures, while the whole process is automated with Python and integrated in the open-source optimization framework OpenMDAO. Findings: The approach was applied to the design of the Predator MQ-1 wing. The results of the MDAs show the convergence of the wing deformations to the flight shape after few iterations. At the end of the aeroelastic sizing loop, the result is a structurally sized wing with minimal weight considering the aeroelasticity effects. Originality/value: The approach proposed takes into account the wing aero-structural coupling effects while sizing its structure instead of a fixed load distribution. In addition, the approach is fully based on open-source codes, which are freely available for public use and can be fully reproducible. [ABSTRACT FROM AUTHOR]

Purpose: This study aims to discuss the mathematical modelling of a compliance-assisted flapping mechanism and morphable structures for an UAV. Design/methodology/approach: A compliance-assisted flapping wing was designed and modelled mathematically, and signals for the corresponding curves were calculated. The actual wing tip trace of a hummingbird was taken, and variables a, b, h and k were calculated from the image. This data was given to the mathematical model for plotting the graph, and the curve was compared with the input curve. The wing frame and mechanism for control surfaces using morphing is modelled along with single pivoted spine for centre of gravity augmentation and flight orientation control. Findings: The model efficiently approximates the 2D path of the wing using line segments using the muscle and compliance mechanism. Practical implications: Using a compliance-assisted flapping mechanism offers practical advantages. It allows us to synchronize the flapping frequency with the input signal frequency, ensuring efficient operation. Additionally, the authors can enhance the torque output by using multiple muscle strands, resulting in a substantial increase in the system's torque-to-weight ratio. This approach proves to be more favourable when compared to conventional methods involving motors or servos, ultimately offering a more efficient and robust solution for practical application. Social implications: This model focuses on creating a flexible and tunable mechanism that can at least trace four types of wing traces from the same design, for shifting from one mode of flight to another. Originality/value: Conventional ornithopter flapping mechanisms are gear or servo driven and cannot trace a wing tip, but some can trace complicated curves, but only one at a time. This model can trace multiple curves using the same hardware, allowing the user to program the curve based on their needs or bird. The authors may vary the shape of the wing tip trace to switch between forward flight, hovering, backward flying, etc., which is not conceivable with any traditional flapping mechanism. [ABSTRACT FROM AUTHOR]

For meeting the requirements of tactical missiles seeking miniaturized launch devices for storage, transportation, and launch, a tube-launched missile wing is adopted, which folds before launch and quickly unfolds after launch. As a structure installed on the missile body to generate the required aerodynamic force for manipulating the missile, the tube-launched missile wing can effectively stabilize the missile's flight attitude. At present, most research on the unfolding mechanism of missile folding wings is focused on one-time folding. When the wingspan is large, multiple folding is required to meet the launch requirements of modern tube-launched missiles. Therefore, this article designs a dual-joint folding wing deployment mechanism and studies the rigid–flexible coupling dynamic modeling and related technologies of folding wings based on this structure. Based on the inertial coupling between large-scale rigid body motion and structural flexible deformation, the folding wing breaks through the element convergence of the model and achieves the applicability of the structural model through zero-order approximation model analysis and other technologies. Simulation results show that the hybrid coordinate method can fully and accurately display the vibration information of flexible folding wings. At different speeds, the first-order coupling model is more advanced than the zero-order coupling model. In addition, increasing rotational speed, increasing wing thickness, and reducing wing span length can effectively increase the fundamental frequency of wing flutter. The structural design of folding wings has shown important reference significance. [ABSTRACT FROM AUTHOR]

This study investigates the aerodynamic behavior of a high-performance vehicle and the interaction with its rear wing in straight-line and steady-state cornering conditions. Analyses are performed with Reynolds-averaged Navier–Stokes based computational fluid dynamics simulations using a moving reference frame and overset mesh technique, validated against moving ground wind tunnel experiments. The results indicate a significant 20% decrease in downforce and 35% increase in drag compared to straight-line conditions at the smallest considered corner radius of 2.9 car-lengths. Downforce losses primarily stem from performance deficits on the underbody and rear wing, alongside elevated upper body lift. Drag penalties mainly result from additional pressure drag induced by a recirculation wake vortex generated behind the vehicle's inboard side. The vehicle's lateral pressure distribution is also affected, introducing a centripetal force that increases with smaller corner radii. Additionally, analyses of the rear wing reveal alternations of its aerodynamic characteristics in cornering, particularly impacting vortical flow and suction on the lower surface. Throughout the operating conditions, the rear wing's individual downforce contribution falls off beyond its stall angle. At higher angles of attack, the rear wing primarily generates downforce by pressurizing the vehicle's upper surfaces, but its interaction with the near-wake leads to a substantially increased pressure drag. Overall, these findings provide crucial insights into the intricate aerodynamic interactions of high-performance vehicles in diverse operating conditions as well as form an essential foundation for future research on static and active aerodynamic designs in the pursuit to optimize vehicle performance in dynamic driving conditions. [ABSTRACT FROM AUTHOR]

The evolution of aircraft wing development has seen significant progress since the early days of aviation, with static testing emerging as a crucial aspect for ensuring safety and reliability. This study focused specifically on the engineering phase of static testing for the Clean Sky 2 T-WING project, which is dedicated to testing the innovative composite wing of the Next-Generation Civil Tiltrotor Technology Demonstrator. During the design phase, critical load cases were identified through shear force/bending moment (SFBM) and failure mode analyses. To qualify the wing, an engineering team designed a dedicated test rig equipped with hydraulic jacks to mirror the SFBM diagrams. Adhering to specifications and geometric constraints due to several factors, the jacks aimed to minimize the errors (within 5%) in replicating the diagrams. An effective algorithm, spanning five phases, was employed to pinpoint the optimal configuration. This involved analyzing significant components, conducting least square linear optimizations, selecting solutions that met the directional constraints, analyzing the Pareto front solutions, and evaluating the external jack forces. The outcome was a test rig setup with a viable set of hydraulic jack forces, achieving precise SFBM replication on the wing with minimal jacks and overall applied forces. [ABSTRACT FROM AUTHOR]

This study addresses the crucial role of post-buckling behavior analysis in the structural design of composite aeronautical structures. Traditional engineering practices tend to result in oversized composite components, increasing structural weight. EASA AMC 20-29's Building Block Approach suggests phased testing, but its time and cost challenges necessitate a shift to high-fidelity post-buckling analyses, exemplified by MSC NASTRAN SOL 400. This approach, showcased in the analysis of the Next Generation Civil Tilt Rotor Technology Demonstrator's wing (NGTCTR-TD), effectively de-risks static tests, contributing to a more efficient certification process. The study demonstrates how advanced simulations provide detailed insights into local buckling phenomena, allowing precise stress distribution analysis. These analyses eliminate the risk of structural failure, paving the way for safer, more efficient, and cost-effective airframe structures. Future developments aim to validate numerical analyses with experimental data, further emphasizing the reliability and benefits of high-fidelity simulations. [ABSTRACT FROM AUTHOR]

The parametric design method is widely utilized in the preliminary design stage for hypersonic vehicles; it ensures the fast iteration of configuration, generation, and optimization. This study proposes a novel parametric method for a wide-range, wing-morphing glide vehicle. The whole configuration, including a waverider fuselage, a rotating wing, a blunt leading edge, rudders, etc., can be easily described using 27 key parameters. In contrast to the typical parametric method, the new method takes internal payloads into consideration during the shape optimization process. That is, the vehicle configuration can be flexibly adjusted depending on the internal payloads; these payloads may be of random amounts and have different shapes. The code for the new parametric design method is developed using the secondary development tools of UG (UG 10.0) commercial software. The lift and drag characteristics over a wide operational range (H = 6–25 km, M = 2.5–8.5, AOA = 0–10°) were numerically investigated, as was the influence of the retracting angle of the morphing wings. It was found that, for the mode of the fully deployed wings, the lift-to-drag ratio (L/D) remained at a high level (≥4.7) over a Mach range of 4.0–8.5 and an AOA range of 4–7°. For the mode of the fully retracted wings, the drag coefficient remained smaller than 0.02 over a Mach range of 4.0–8.5 and an AOA range of 0–5°. A wide L/D of 0.3–4.7 could be achieved by controlling the retracting angle of the wings, thus demonstrating a good potential for flight maneuverability. The flexible change in L/D proved to be a combined result of varying pressure distribution and edge-flow spillage. This will aid in the further optimization of lift/drag characteristics. [ABSTRACT FROM AUTHOR]

For a flexible wing system, two different exosystems are used to generate unknown references and output disturbances, respectively. The wing system is further perturbed by unknown distributed and boundary disturbances that have the same frequencies to the references and output disturbances. This article proposes two robust control methodologies to address the robust output regulation of the wing. At first, the observer based output feedback controls are designed based on state feedback controls and the following observers: (1) two exosystem observers based on tracking errors; (2) two exosystem observers addressing the uncertainties of output disturbances; (3) the PDE observer estimating the wing states. For the closed‐loop system, the tracking errors are proved to be exponentially convergent toward zero. Let the output disturbances be addressed through two adaption laws, not the observers in the above item (2), which together with other two exosystem observers and the PDE observer, gives observer based adaptive controls. From both cases, we obtain that the exosystem observers based on the tracking errors are necessary for the robust output regulation. By comparison, the former robust controls have a lower dimension and concise structure. Simulation examples have been provided to show the effectiveness of the robust controls. [ABSTRACT FROM AUTHOR]

Bird wings vary in size and morphology in terms of both size and number of feathers and the underlying skeletal anatomy. The number of primary remiges does not seem to vary much between bird species but, by contrast, the number of secondary remiges is reported to range between 6 and 40 depending on bird size. Given that the primaries are attached to the manus, and the secondaries are attached to the ulna, it was predicted that as bone lengths increased with increasing size of the bird, then feather count would increase. Data were collected for 268 species from 25 different orders, and phylogenetically controlled analysis explored the allometry between feather count and bone size. The number of primaries was typically 10 or 11 and did not vary with manus size. By contrast, the number of secondaries increased with ulna length, but only in some orders. For example, in Gruiformes, the number of secondary feathers increased concomitantly with ulna length but despite a two orders of magnitude range in body mass, almost all species in the Passeriformes had nine secondary remiges. It is unclear why, for instance, species with an ulna length of 70 mm can have between 9 and 24 secondaries depending on their order. This variation in secondary feather number can be added to variation in relative wing bone lengths, flight feather lengths, flight feather mechanical properties, and flight feather vane densities as another potential mechanism of adaptation to flight requirements. The apparent constraint of wingspan is scaling as approximately body mass1/3. Further research is needed to explore whether changes in secondary feather number relative to ulna length are accompanied by changes in feather vane width or the overlap of adjacent feathers and how this relates to wing aerodynamics. [ABSTRACT FROM AUTHOR]

The foundation of an offshore wind turbine (OWT) accounts for up to 20% of the total cost; therefore, investigations into reliable and efficient foundations are important for the offshore wind energy industry. This article describes an innovative winged caisson foundation for OWTs, which adapts the conventional caisson foundation to produce higher overturning capacities. The behaviour of this proposed novel foundation in the sand was investigated through experimental and numerical modelling using a series of 1 g tests (1/70th scale), centrifuge model tests (70 g ), and finite element (FE) simulations. Models with various wing sizes installed in different sand densities were tested in the laboratory by applying an overturning load. In this article, the moment–rotation performance of the foundation under monotonic loading is compared against that of conventional caissons to assess the potential benefits of adding wings to caisson foundations. Results show that there is a significant increase in overturning capacity provided by the novel foundation. FE models were first validated against experimental results, then, used to conduct a parametric study to better understand the behaviour of the winged caissons beyond the scaled model tests. In addition, the FE models were used to evaluate the drained response of the proposed foundation under combined loading. [ABSTRACT FROM AUTHOR]

Within the double-sided countersunk riveting process of aircraft wings with a composite wedge structure, riveting consistency is poor, and composite damage is severe, which seriously affects the performance and reliability of the aircraft structure. This paper used the principal stress method to establish a stress model of countersunk riveting, and, based on the analysis of the stress on the structure during the pressure-riveting process, a composite structure rivet was designed. A finite element simulation model of the double-sided countersunk riveting of composite wedge structures' composite rivets was established. The influences of the structure and the matching parameters of composite rivets on both the plastic flow of pressure riveting and the compressive stress of the structure during the pressure-riveting process were analyzed. The structural parameters and riveting process of composite rivets were optimized. The results show that the composite rivet structure could significantly reduce the contact-compressive stress at the riveting joint by more than 20%, thereby reducing the damage caused by the riveting to the composite material. For 4 mm rivets, an aperture of 4.04~4.06 mm can achieve precise relative interference riveting at 0.6% to 1.0%. Employing a 2.6 mm rivet elongation can exactly fill the countersunk hole of the wedge. [ABSTRACT FROM AUTHOR]

The aeroelastic response of lightweight low-speed aircrafts with slender wings under extreme flow turbulence intensity is not well understood. Experiments on a commercial unmanned aerial vehicle (UAV) with a 3 m wingspan and aspect ratio of 13.6 were performed in a large open-return wind tunnel with extreme flow turbulence intensity of ≈10%. The wing bending displacement and the flow beneath the wing were measured by using laser-displacement sensors and tomographic particle image velocimetry (PIV) techniques, respectively. The unsteady lift produced by the wing was also measured by using a high-capacity load cell at an angle of attack of two degrees for three freestream velocities of 13.4 m/s, 17.9 m/s, and 26.8 m/s, representing the UAV's stall speed, a speed approximately equal to the cruise speed, and a speed considerably higher than the cruise speed, respectively. It was found that a high flow turbulence intensity with large integral length scales relative to the wing chord plays a dominant role in the large unsteady lift and wing displacements measured. The power spectral density (PSD) of the wing structural vibration shows that flow shedding from the wing and the integral length scales have a significant impact on the overall power inherent in the bending vibration of the wing. Computations of the vorticity isosurfaces in the flow measurement volume surrounding the aileron reveal a striking observation: an aileron deflection of 10° becomes less effective in producing additional spanwise vorticity, which is proportional to circulation and lift at 26.8 m/s since the freestream already has elevated levels of vorticity. A paradigm shift in design is suggested for light aircraft structures with slender wings operating in highly turbulent flow, which is to employ multiple control surfaces in order to respond to this flow and mitigate large bending or torsion displacements and the probability of structural failure. [ABSTRACT FROM AUTHOR]

Tidal intrusion fronts are surface convergences that occur at constrictions in estuaries during the flood tide, separating incoming higher‐salinity water from lower‐salinity, stratified estuarine water. Previous observations of tidal intrusion fronts describe a V‐shaped planform, with the apex of the V pointing into the estuary, however the significance of this structure has not been previously explained. Observations near the mouth of the James River estuary during the flood tide reveal the development of a quasi‐steady, V‐shaped front. Considering a reference frame oriented normal to the front, the velocity and density structure are consistent with gravity‐current dynamics, but the oblique orientation of the front relative to the impinging flow indicates strong, along‐front shear, which results from vorticity produced by flow separation at the lateral boundaries as well as topographic torque from upstream. The combination of convergence and along‐front shear leads to enhanced mixing, as revealed by acoustic backscatter images of shear instability and persistent subcritical gradient Richardson number in the frontal zone. Oblique fronts such as this tidal intrusion front are common features of estuaries, and they play an important role in vertical exchange due to subduction and mixing of surface water. Plain Language Summary: Surface fronts are among the most notable features of estuarine and coastal environments. They have important consequences for the physics as well as the biology, due to the strong gradients, aggregation of particles and organisms, and mixing that occur at fronts. One of the most conspicuous types of estuarine fronts is the tidal intrusion front, which forms near the mouth during the incoming tide and usually exhibits a V shape pointing into the estuary. This study explains the processes responsible for the V‐shaped configuration of the front and their implications. The density contrast across the front is one important ingredient of the dynamics, resulting in a strong convergence of surface flow. The other key ingredient is the shear across the front, with faster inflow on the inside of the V and weaker flow on the outside. This shear is due in large part to "flow separation" at the sides of the estuary—the same process that causes stall of airplane wings or the formation of eddies behind rocks at the edges of a stream. The paper shows that this combination of processes leads to strong mixing in the frontal zone—much stronger than would occur in the absence of the cross‐front shear. Key Points: The "V"‐shaped tidal intrusion front exhibits a combination of density‐driven, convergent flow and along‐front shear resulting from lateral flow separationThe along‐estuary flow at the front is supercritical (with respect to gravity current propagation), but the velocity in the front‐normal direction is criticalThe along‐front shear significantly augments vertical mixing within the frontal zone relative to the mixing in normal gravity currents [ABSTRACT FROM AUTHOR]

Lightning is one of the natural hazards that any aircraft may encounter while navigating. If adequate precautions are not taken against lightning, structural damage, operational disruptions, and loss of life and property can occur. Thus, studying the mechanism of damage caused by lightning strikes in an aircraft's structural material is necessary to optimize the structure, minimize the damage, and reduce the cost caused by lightning. In the present article, the lightning-induced damage behavior of an aircraft structural material was investigated from an analytical perspective. For this purpose, two analytical-based models were developed: an improved electromagnetic pressure impact model (IEPIM) and the damage model in an aircraft wing. For the IEPIM, the findings of the article showed that the proposed pressure model is in good agreement with the experimental studies, borrowed from the open literature, for 100 and 200 kA lightning current. For the damage model, the findings of the article indicated that (i) even though lightning strikes to the regions with the same characteristics on an aircraft wing in terms of the lightning strike zone, the amount of deflection in the wing increases as the impact point approaches the wing tip and decreases as it approaches the wing root, (ii) without changing the lightning strike point (x 0), when the damping coefficient (ξ) is increased in the range of 0 , 2 ξ , the amount of deflection decreases as the amount of damping coefficient increases, and (iii) when lightning with a current of 100 kA hits to the wing root of an aircraft, the pressure impact of the lightning causes more torsion deflection than bending deflection at the wing root; however, when it hits to the mid-wing or wing tip of an aircraft, the pressure impact of the lightning causes more bending deflection than torsion deflection at the mid-wing or wing tip. [ABSTRACT FROM AUTHOR]

There are complex coupling relationships between flight motion and energy management for hybrid electric fixed wing UAVs with photovoltaic/fuel cell/battery. Aiming at realizing high-energy-efficiency long endurance autonomous flight, this paper proposes a coordinated optimization method of trajectory optimization and energy management for hybrid electric UAVs. Numerical simulation and experiment are carried out. Separated by surplus demand energy except that produced by PV, this method includes an optimal energy flight trajectory optimization layer and an energy management layer. For the former, fuzzy neural network sequential convex optimization (FNNSCP) is proposed to handle the flight trajectory optimization problem for maximizing utilization solar energy. For the latter, an online energy management strategy based on convex quadratic programming MPC (CQPMPC) is proposed to realize the power allocation. Numerical simulation results show that FNNSCP reduced 2% energy requirement compared with Radau pseudo-spectral method (RPM) from flight trajectory optimization level. From energy management level, CQPMPC saves hydrogen consumption by 3.1% and 16.3% compared with nonlinear model predictive control (NMPC) and fuzzy logic state machine (FLSM), respectively. In the experiment, less hydrogen fuel is consumed by proposed method by 5.2% and 15.6% compared with NMPC and FLSM, respectively. Experimental results show that the proposed method has better energy-saving performance and excellent real time performance. It indicates that optimizing flight trajectory and improving energy management efficiency can realize the high-energy-efficiency flight of hybrid electric UAVs. This method is conducive to promoting the application of hydrogen energy in the field of UAVs. [Display omitted] • A coordinated optimization method is proposed for the high-energy-efficiency flight. • Fuzzy neural network adaptive sequential convex optimization method is proposed. • Online energy management strategy with convex quadratic programming MPC is proposed. • The proposed method is verified by simulation and energy management experiment. [ABSTRACT FROM AUTHOR]

This study focused on designing a single-degree-of-freedom (1-DoF) mechanism emulating the wings of rock pigeons. Three wing models were created: one with REAL feathers from a pigeon, and the other two models with 3D-printed artificial remiges made using different strengths of material, PLA and PETG. Aerodynamic performance was assessed in a wind tunnel under both stationary (0 m/s) and cruising speed (16 m/s) with flapping frequencies from 3.0 to 6.0 Hz. The stiffness of remiges was examined through three-point bending tests. The artificial feathers made of PLA have greater rigidity than REAL feathers, while PETG, on the other hand, exhibits the weakest strength. At cruising speed, although the artificial feathers exhibit more noticeable feather splitting and more pronounced fluctuations in lift during the flapping process compared to REAL feathers due to the differences in weight and stiffness distribution, the PETG feathered wing showed the highest lift enhancement (28% of pigeon body weight), while the PLA feathered wing had high thrust but doubled drag, making them inefficient in cruising. The PETG feathered wing provided better propulsion efficiency than the REAL feathered wing. Despite their weight, artificial feathered wings outperformed REAL feathers in 1-DoF flapping motion. This study shows the potential for artificial feathers in improving the flight performance of Flapping Wing Micro Air Vehicles (FWMAVs). [ABSTRACT FROM AUTHOR]

In this paper, a numerical analysis of a biomimetic unmanned aerial vehicle (UAV) is presented. Its wings feature three grids at the tip similar to the primary feathers of birds in order to modify the lift distribution over the wing and help in reducing the induced drag. Numerical analysis using computational fluid dynamics (CFD) is presented to analyze the aerodynamic effects of the changes in dihedral and angle of attack (with respect of the rest of the wing) of these small grids at the tip. The aerodynamic performances (lift, drag, and efficiency) and rolling capabilities are obtained under different flight conditions. The effects of changing the dihedral are small. However, the change in the grid angle of attack increases aerodynamic efficiency by up to 2.5 times when the UAV is under cruise flight conditions. Changes to the angle of attack of the grids also provide increased capabilities for rolling. Finally, boundary values of the pressure coefficient and non-dimensional velocity contours are presented on the surfaces of the UAV, in order to relate the aerodynamic results to the aerodynamic patterns observed over the wing. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this research paper is to recover the autonomous flight performance of a mini unmanned aerial vehicle (UAV) via stochastically optimizing the wing over certain parameters (i.e. wing taper ratio and wing aspect ratio) while there are lower and upper constraints on these redesign parameters. Design/methodology/approach: A mini UAV is produced in the Iskenderun Technical University (ISTE) Unmanned Aerial Vehicle Laboratory. Its complete wing can vary passively before the flight with respect to the result of the stochastic redesign of the wing while maximizing autonomous flight performance. Flight control system (FCS) parameters (i.e. gains of longitudinal and lateral proportional-integral-derivative controllers) and wing redesign parameters mentioned before are simultaneously designed to maximize autonomous flight performance index using a certain stochastic optimization strategy named as simultaneous perturbation stochastic approximation (SPSA). Found results are used while composing UAV flight simulations. Findings: Using stochastic redesign of mini UAV and simultaneously designing mini ISTE UAV over previously mentioned wing parameters and FCS, it obtained a maximum UAV autonomous flight performance. Research limitations/implications: Permission of the directorate general of civil aviation in the Republic of Türkiye is essential for real-time UAV autonomous flights. Practical implications: Stochastic redesign of mini UAV and simultaneously designing mini ISTE UAV wing parameters and FCS approach is very useful for improving any mini UAV autonomous flight performance cost index. Social implications: Stochastic redesign of mini UAV and simultaneously designing mini ISTE UAV wing parameters and FCS approach succeeds confidence, highly improved autonomous flight performance cost index and easy service demands of mini UAV operators. Originality/value: Creating a new approach to recover autonomous flight performance cost index (e.g. satisfying less settling time and less rise time, less overshoot during flight trajectory tracking) of a mini UAV and composing a novel procedure performing simultaneous mini UAV having passively morphing wing over certain parameters while there are upper and lower constraints and FCS design idea. [ABSTRACT FROM AUTHOR]

Purpose: This paper aims to assess the feasibility of additively manufactured wind tunnel models. The additively manufactured model was used to validate a computational framework allowing the evaluation of the performance of five wing models. Design/methodology/approach: An optimized fighter wing was additively manufactured and tested in a low-speed wind tunnel to obtain the aerodynamic coefficients and deflections at different speeds and angles of attack. The flexible wing model with optimized curvilinear spars and ribs was used to validate a finite element framework that was used to study the aeroelastic performance of five wing models. As a computationally efficient optimization method, homogenization-based topology optimization was used to generate four different lattice internal structures for the wing in this study. The efficiency of the spline-based optimization used for the spar-rib model and the lattice-based optimization used for the other four wings were compared. Findings: The aerodynamic loads and displacements obtained experimentally and computationally were in good agreement, proving that additive manufacture can be used to create complex accurate models. The study also shows the efficiency of the homogenization-based topology optimization framework in generating designs with superior stiffness. Originality/value: To the best of the authors' knowledge, this is the first time a wing model with curvilinear spars and ribs was additively manufactured as a single piece and tested in a wind tunnel. This research also demonstrates the efficiency of homogenization-based topology optimization in generating enhanced models of different complexity. [ABSTRACT FROM AUTHOR]

The design and development of a wing for a completely brand-new aircraft represents, in aeronautics, one of the highest challenges from an engineering point of view. The present work describes a novel methodology devoted to execute numeric simulation in a non-linear post-buckling regime to verify the composite wing compliance under the design load conditions. The procedure was developed as part of a wing design and research activity and was motivated by the need to have more realistic results, without standard conservatisms like the no-buckling up to ultimate load, to be of use for achieving further weight savings. To carry this out, it was obviously necessary to ensure that the structural integrity was also guaranteed in the post-buckling regime, above the limit load, and therefore in a highly non-linear regime. The present work illustrates a numerical approach based on non-linear finite element analysis which uses the inertia relief option in order to have a more realistic representation of the structural response of the wing in its real context. All that represents a novelty since, at present, the commercial FE codes allow us to use the inertia relief option exclusively for linear analysis. Obviously, the approach can be applied to any other structural component with similar needs. The obtained results show that the differences between linear and non-linear regime are not negligible and, above all, that it is possible to design a wing (or other structural components) considering, at the same time, the large deformation due to the post-buckling regime, the material non-linearities due to the failures and any other non-linearities in order to achieve the challenging weight requirement of the new aircraft generation. [ABSTRACT FROM AUTHOR]

The filtered lifting line theory is an analytical approach used to solve the equations of flow subjected to body forces with a Gaussian distribution, such as used in the actuator line model. In the original formulation, the changes in chord length along the blade were assumed to be small. This assumption can lead to errors in the induced velocities predicted by the theory compared to full solutions of the equations. In this work, we revisit the original derivation and provide a more general formulation that can account for significant changes in chord along the blade. The revised formulation can be applied to wings with significant changes in chord along the span, such as wind turbine blades. [ABSTRACT FROM AUTHOR]

Distributed electric propulsion (DEP) with four propellers distributed along the rear edge of the wing (pusher DEP configuration) promote aerodynamic interactions to a higher level. To study the aerodynamic performance of DEP with the rear wing through simulations and experiments, the multi-reference frame (MRF) with sliding grid is combined with wind tunnel tests. The obtained results demonstrate that the lift and drag of DEP increase with the angle of attack (AoA) and are related to the relative position of the propellers and wing. The propeller has no significant effect on the lift of the wing, and the lift and the AoA remain linear when the AoA is less than 16°. By contrast, the lift coefficient is much higher than the baseline (isolated wing), and the lift is greatly improved with the increasing drag when the AoA is greater than 16°. This is because the flow around the wing of the pusher configuration remains attached due to the suction of the inflow of the propeller on the trailing edge vortex. In addition, the acceleration effect on the free flow improves the kinetic energy of the airflow, which effectively delays the separation of the airflow in the slipstream region. [ABSTRACT FROM AUTHOR]