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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]

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]

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]

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: Determining the variation law of the oxygen concentration in the ullage space of the fuel tank is the key to the design of the inert system. Among various factors affecting the oxygen concentration in the ullage space of the fuel tank, the temperature difference between day and night shows particular importance while relevant analysis and calculation are scarce. Design/methodology/approach: This study establishes a theoretical simulation model of the central wing fuel tank of an aircraft according to the relevant provisions of day-night temperature variation in FAR25 airworthiness regulations, verifies the model with the existing experimental data and discusses the corresponding relationship between the oxygen concentration in the ullage space of the fuel tank and the day-night temperature difference. The influence of day and night temperature difference, fuel type, fuel load rate, initial oxygen concentration, dissolved oxygen evolution and other factors on the oxygen concentration in the ullage space of the fuel tank were analyzed, and the limit of initial oxygen concentration of the fuel tank before the shutdown at night meeting the requirements of the airworthiness provisions was proposed. Findings: The results show that the temperature difference between day and night, fuel load rate, initial oxygen concentration and other factors have different effects on the oxygen concentration in the ullage space of fuel tank. The initial oxygen concentration limit before shutdown shall be 2% below the 12% oxygen concentration stipulated by FAA. Research limitations/implications: The research results in this paper will be of good reference value to the design of the inert system and the calculation of the flammability exposure evaluation time. This paper aims to be good reference of the design of the inert system and the calculation of the flammability exposure evaluation time. Originality/value: The research results of this paper can provide practical guidance for the current civil airworthiness certification work. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is to find the droplets impact on the airplane wing structure. Two kinds of characteristics of the droplet at different velocity and viscosity are assumed. The droplet is assumed to be spherical cubic form and it is injected from the convergent divergent nozzle with a passive control. Design/methodology/approach: This paper presents the results of a numerical simulation of droplet impact on the horizontal surface. The effects of impact parameters are studied. The splash effect of the droplet also visualized. The results are presented in form of stress, strain, displacement magnitude of the droplet. Findings: Crosswire is used as passive control. The behavior of the droplet impact is observed based on the kinetic energy and the gravitational forces. Originality/value: The results predict that smooth particle hydrodynamic designed droplet not only depend on the equation of state of the droplet but also the injection velocity from the nozzle. It also determined that droplet velocity is depending on the viscosity of the fluid. [ABSTRACT FROM AUTHOR]

Purpose: This paper aims to investigate the possibility of lowering the time taken during the aircraft design for unmanned aerial vehicles by using machine learning (ML) for the configuration selection phase. In this work, a database of unmanned aircraft is compiled and is proposed that decision tree classifiers (DTC) can understand the relations between mission and operational requirements and the resulting aircraft configuration. Design/methodology/approach: This paper presents a ML-based approach to configuration selection of unmanned aircraft. Multiple DTC are built to predict the overall configuration. The classifiers are trained with a database of 118 unmanned aircraft with 57 characteristics, 47 of which are inputs for the classification problem, and 10 are the desired outputs, such as wing configuration or engine type. Findings: This paper shows that DTC can be used for the configuration selection of unmanned aircraft with reasonable accuracy, understanding the connections between the different mission requirements and the culminating configuration. The framework is also capable of dealing with incomplete databases, maximizing the available knowledge. Originality/value: This paper increases the computational usage for the aircraft design while retaining requirements' traceability and increasing decision awareness. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is to use numerical methods early in the airframe design process and access the structural performance of wing leading edge devices made of different materials and design details, under bird strike events. Design/methodology/approach: Explicit finite element analysis was used to numerically model bird strike events. Findings: Structural performance charts related to materials and general design details were drawn to explore the design space dictated by the current applicable airworthiness requirements. Practical implications: This paper makes use of the current capability in the numerical tools available for structural simulations and exposes the existing limitations in the terms of material modelling, material properties and fracture simulation using continuum damage mechanics. Such results will always be in the need of fine-tuning with experimental testing, yet the tools can shed some light very early in the design process in a relative inexpensive manner, especially for design details down selection like materials to use, structural thicknesses and even design arrangements. Originality/value: Bird strike simulations have been successfully used on aircraft design, mainly at the manufactured articles design validation, testing and certification. This paper presents a hypothetical early design case study of leading edge devices for appropriate material and skin thickness down selection. [ABSTRACT FROM AUTHOR]

Purpose: Wingtip loss is an existing type of transport aircraft hazard which is a real threat to flight safety caused by a missile strike, underwing engine explosion or impact with obstructions when performing near-ground operations. The primary effect of the wingtip loss is an asymmetric rolling moment, which may result in the fatal loss of control for the aircraft. This study aims to assess whether aerodynamic degradation will cause a wing-damaged transport aircraft to lose its balance under a certain level of wing damage and if a pilot can compensate for the loss of aerodynamic force and regain the balance of the aircraft. Design/methodology/approach: In this paper, experimental and numerical studies were conducted to investigate the aerodynamic characteristics of a wingtip-lost transport aircraft in landing configuration. Various levels of wing damages including wingtip, slat and flap loss were considered. The numerical simulations were performed with ANSYS Fluent. The computational fluid dynamics calculation was validated by wind tunnel tests. Findings: The aerodynamic performance of the aircraft with wing-damaged condition was presented. It was revealed that the wingtip loss leads to an asymmetric rolling moment and a reduction of the lift force, which affects the balance of the transport aircraft. The methods to compensate for the lift force and the asymmetric rolling moment were investigated for a safe landing. The lateral balance cannot be maintained in cases with serious damage on the wing (larger than 53% of the semi-span) or moderate damage on the wing with loss of slats and flaps. Originality/value: The nonlinear results indicate the importance of aerodynamic assessment for the sake of training pilots to properly handle the hazard situation and explore the critical facts leading to the air crash. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is to investigate the equivalent control authority of the conventional and circulation control (CC) wing of the aircraft and assess the energy expenditure and aerodynamic efficiency of the CC wing. Design/methodology/approach: Four target cases with different flap deflection angles θ are set in advance for the conventional wing, and then a series of cases with different jet momentum coefficients Cμ are set for the CC wing. The lift, drag and momentum coefficient curves of the CC wing are compared to those of the four conventional wing cases. The curves with the best agreement are selected to establish the corresponding relation between θ and Cμ. The energy expenditure of the CC system is analyzed. The concept of equivalent lift-to-drag ratio for the CC wing is introduced to compare the aerodynamic efficiency with the conventional wing Findings: The control authority of the conventional wing at θ = 0º, 10º, 20º, 30º are equivalent to the control authority of the CC wing with Cµ = 0.0, 0.005, 0.009 and 0.012. The CC system is more efficient at small Cµ than large Cµ. Practical implications: This study could contribute to the application of the CC system on flapless aircrafts. Originality/value: The corresponding relation between θ and Cµ is established by matching the equivalent control authority between the conventional wing and CC wing. [ABSTRACT FROM AUTHOR]

Purpose: This paper aims to present a preliminary study on a disruptive vertical take-off and landing (VTOL) configuration based on the best wing system concept by L. Prandtl. Design/methodology/approach: A preliminary design has been addressed from several points of views: a conceptual design has been carried out thanks to in-house optimization tool; aerodynamic performances, propulsion design and mechanical design have been addressed to make the first prototype for preliminary vertical flight tests. Findings: The study shows the feasibility of box-wing configuration for VTOL aircraft. Practical implications: The work shows a general design procedure for box-wing unmanned air vehicle (UAV) configuration. The study of this configuration can be easily adopted in wider range, from UAV to general aviation. In the last category, it can be a promising configuration for the future of urban air mobility. Originality/value: This work lays the foundation for studying and testing box-wing configuration for unmanned VTOL aircraft. The design procedure can be scaled to manned aircraft belonging to general aviation aircraft. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is to present a novel approach based on the artificial bee colony (ABC) algorithm aiming to achieve maximum acceleration and maximum endurance for morphing unmanned aerial vehicle (UAV) design. Design/methodology/approach: Some of the most important issues in the design of UAV are the design of thrust system and determination of the endurance of the UAV. Although propeller selection is very important for the thrust system design, battery selection has the utmost importance for the determination of UAV endurance. In this study, the calculations of maximum acceleration and endurance required by ZANKA-II during the flight are considered simultaneously. For this purpose, a model based on the ABC algorithm is proposed for the morphing UAV design, aiming to achieve the maximum acceleration and endurance. In the proposed model, the propeller diameter, propeller pitch and battery values used in morphing UAV's power system design are selected as the input parameters; maximum acceleration and endurance are selected as the output parameters. To obtain the maximum acceleration and endurance, the optimum input parameters are determined through the ABC algorithm-based model. Findings: Considerable improvements on maximum acceleration and endurance of morphing UAV with ABC algorithm-based model are obtained. Research limitations/implications: The endurance and acceleration due to the thrust are two separate parameters that are not normally proportional to each other. In this study, optimization of UAV's endurance and acceleration is considered with equal importance. Practical implications: Using artificial intelligence techniques causes fast and simple optimization for determination of UAV's endurance and acceleration with equal importance. In the simulation studies with ABC algorithm, satisfactory results are obtained. Social implications: The results of the study have showed that the proposed approach could be an alternative method for UAV designers. Originality/value: Providing a new and efficient method saves time and reduces cost in calculations of maximum acceleration and endurance of the UAV. [ABSTRACT FROM AUTHOR]

Purpose: The finite element model developed for a new-designed aircraft was used to solve some problems of structural dynamics. The key purpose of the task was to estimate the critical flutter velocities of the light airplane by performing numerical analysis with application of MSC Software. Design/methodology/approach: Flutter analyses processed by Nastran require application of some complex aeroelastic model integrating two separate components – structural model and aerodynamic model. These sub-models are necessary for determining stiffness, mass and aerodynamic matrices, which are involved in the flutter equation. The aircraft structural model with its non-structural masses was developed in Patran. To determine the aerodynamic coefficient matrix, some simplified aerodynamic body-panel geometries were developed. The flutter equation was solved with the PK method. Findings: The verified aircraft model was used to determine its normal modes in the range of 0-30 Hz. Then, some critical velocities of flutter were calculated within the range of operational velocities. As there is no certainty that the computed modes are in accordance with the natural ones, some parametric calculations are recommended. Modal frequencies depend on structural parameters that are quite difficult to identify. Adopting their values from the reasonable range, it is possible to assign the range of possible frequencies. The frequencies of rudder or elevator modes are dependent on their mass moments of inertia and rigidity of controls. The critical speeds of tail flutter were calculated for various combinations of stiffness or mass values. Practical implications: The task described here is a preliminary calculational study of normal modes and flutter vibrations. It is necessary to prove the new airplane is free from flutter to fulfil the requirement considered in the type certification process. Originality/value: The described approach takes into account the uncertainty of results caused by the indeterminacy of selected constructional parameters. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper presents the effects of actively morphing root chord and taper on the energy of the flight control system (i.e. FCS). Design/methodology/approach: Via regarding previously mentioned purposes, sophisticated and realistic helicopter models are benefitted to examine the energy of the FCS. Findings: Helicopters having actively morphing blade root chord length and blade taper consume less control energy than the ones having one of or any of passively morphing blade root chord length and blade taper. Practical implications: Actively morphing blade root chord length and blade taper can be used for cheaper helicopter operations. Originality/value: The main originality of this paper is applying active morphing strategy on helicopter blade root chord and blade taper. In this paper, it is also found that using active morphing strategy on helicopter blade root chord and blade taper reasons less energy consumption than using either passively morphing blade root chord length plus blade taper or not any. This causes also less fuel consumption and green environment. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this study is to examine the effect of passive and active morphing of blade root chord length and blade taper on the control effort of the flight control system (FCS) of a helicopter. Design/methodology/approach: Physics-based helicopter models, which are functions of passive and active morphing, are created and applied in helicopter FCS design to determine the control effort. Findings: Helicopters, having both passively and actively morphing blade root chord length and blade taper, experience less control effort than the ones having either only passively morphing blade root chord length or only blade taper or only actively morphing blade root chord length and blade taper. Practical implications: Both passively and actively morphing blade root chord length and blade taper can be implemented for more economical autonomous helicopter flights. Originality/value: Main novelty of our article is simultaneous application of passive and active morphing ideas on helicopter root chord length and blade taper. It is also proved in this study that using both passive and active morphing ideas on helicopter blade root chord and blade taper causes much less energy consumption than using either only passive morphing idea on helicopter blade root chord and blade taper or only active morphing idea on helicopter blade root chord and blade taper. This also reduces fuel consumption and also makes environment cleaner. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is to present a novel approach based on the differential search (DS) algorithm integrated with the adaptive network-based fuzzy inference system (ANFIS) for unmanned aerial vehicle (UAV) winglet design. Design/methodology/approach: The winglet design of UAV, which was produced at Faculty of Aeronautics and Astronautics in Erciyes University, was redesigned using artificial intelligence techniques. This approach proposed for winglet redesign is based on the integration of ANFIS into the DS algorithm. For this purpose, the cant angle (c), the twist angle (t) and taper ratio (λ) of winglet are selected as input parameters; the maximum value of lift/drag ratio (Emax) is selected as the output parameter for ANFIS. For the selected input and output parameters, the optimum ANFIS parameters are determined by the DS algorithm. Then the objective function based on optimum ANFIS structure is integrated with the DS algorithm. With this integration, the input parameters for the Emax value are obtained by the DS algorithm. That is, the DS algorithm is used to obtain both the optimization of the ANFIS structure and the necessary parameters for the winglet design. Thus, the UAV was reshaped and the maximum value of lift/drag ratio was calculated based on new design. Findings: Considerable improvements on the max E are obtained through winglet redesign on morphing UAVs with artificial intelligence techniques. Research limitations/implications: It takes a long time to obtain the optimum Emax value by the computational fluid dynamics method. Practical implications: Using artificial intelligence techniques saves time and reduces cost in maximizing Emax value. The simulation results showed that satisfactory Emax values were obtained, and an optimum winglet design was achieved. Thus, the presented method based on ANFIS and DS algorithm is faster and simpler. Social implications: The application of artificial intelligence methods could be used in designing more efficient aircrafts. Originality/value: The study provides a new and efficient method that saves time and reduces cost in redesigning UAV winglets. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is finding the optimal geometric parameters and developing of a method for optimizing a light unmanned aerial vehicle (UAV) wing, maximizing, at the same time, its endurance with the assumed parameters of aircraft mission. Design/methodology/approach: The research is based on the experience gained by the author's contribution to the project of building medium-altitude, long-endurance class, light UAV called "Samonit". The author was responsible for the structure design, wind tunnel tests and flight tests of the "Samonit" aircraft. Based on the experience, the author was able to develop an optimization process considering various disciplines involved in the whole aircraft design topics such as aerodynamics, flight mechanics, structural stiffness and weight, aircraft stability and maneuverability. The presented methodology has a multidisciplinary nature, as in the process of optimization both aerodynamic aspects and the influence of wing geometric parameters on the wing structure and weight and the aircraft payload were taken into account. The optimal wing configuration was obtained using the genetic algorithms. Findings: As a result, a set of wing geometrical parameters has been obtained that allowed for achieving twice as long endurance as compared with the initial one. Practical implications: Using the methodology presented in the paper, an aircraft designer can easily find the optimum wing configuration of a designed aircraft, satisfying the mission requirements in a best way. Originality/value: An original procedure has been developed, based on the actual design, wind tunnel tests and numerical calculations of "Samonit" aircraft, enabling the determination of optimum wing configuration for a small unmanned aircraft. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is to investigate the possibility of manufacturing fused deposition modelling (FDM) 3D printed structures such as wings or fuselages for small remote control (RC) air craft and mini unmaned aerial vehicles (UAVs). Design/methodology/approach: Material tests, design assumptions and calculations were verified by designing and manufacturing a small radio-controlled motor-glider using as many printed parts as possible and performing test flights. Findings: It is possible to create an aircraft with good flight characteristics using FDM 3D printed parts. Current level of technology allows for reasonably fast manufacturing of 3D printed aircraft with good reliability and high success ratio of prints; however, only some of the materials are suitable for printing thin wall structures such as wings. Practical implications: The paper proves that apart from currently popular small RC aircraft structural materials such as composites, wood and foam, there is also printed plastic. Moreover, 3D printing is highly competitive in some aspects such as first unit production time or production cost. Originality/value: The presented manufacturing technique can be useful for quick and cost-effective creating scale prototypes of the aircraft for performing test flights. [ABSTRACT FROM AUTHOR]

Purpose: This paper aims to present a novel structural reliability analysis scheme with considering the structural strength degradation for the wing spar of a generic hypersonic aircraft to guarantee flight safety and structural reliability. Design/methodology/approach: A logarithmic model with strength degradation for the wing spar is constructed, and a reliability model of the wing spar is established based on stress-strength interference theory and total probability theorem. Findings: It is demonstrated that the proposed reliability analysis scheme can obtain more accurate structural reliability and failure results for the wing spar, and the strength degradation cannot be neglected. Furthermore, the obtained results will provide an important reference for the structural safety of hypersonic aircraft. Research limitations/implications: The proposed reliability analysis scheme has not implemented in actual flight, as all the simulations are conducted according to the actual experiment data. Practical implications: The proposed reliability analysis scheme can solve the structural life problem of the wing spar for hypersonic aircraft and meet engineering practice requirements, and it also provides an important reference to guarantee the flight safety and structural reliability for hypersonic aircraft. Originality/value: To describe the damage evolution more accurately, with consideration of strength degradation, flight dynamics and material characteristics of the hypersonic aircraft, the stress-strength interference method is first applied to analyze the structural reliability of the wing spar for the hypersonic aircraft. The proposed analysis scheme is implemented on the dynamic model of the hypersonic aircraft, and the simulation demonstrates that a more reasonable reliability result can be achieved. [ABSTRACT FROM AUTHOR]

Purpose: The purpose of this paper is to present a method for analysis and optimization of morphing wing. Moreover, a numerical advantage of morphing airfoil wing, typically assessed in simplified two-dimensional analysis is found using higher fidelity methods. Design/methodology/approach: Because of multi-point nature of morphing wing optimization, an approach for optimization by analysis is presented. Starting from naïve parametrization, multi-fidelity aerodynamic data are used to construct response surface model. From the model, many significant information are extracted related to parameters effect on objective; hence, design sensitivity and, ultimately, optimal solution can be found. Findings: The method was tested on benchmark problem, with some easy-to-predict results. All of them were confirmed, along with additional information on morphing trailing edge wings. It was found that wing with morphing trailing edge has around 10 per cent lower drag for the same lift requirement when compared to conventional design. Practical implications: It is demonstrated that providing a smooth surface on wing gives substantial improvement in multi-purpose aircrafts. Details on how this is achieved are described. The metodology and results presented in current paper can be used in further development of morphing wing. Originality/value: Most of literature describing morphing airfoil design, optimization or calculations, performs only 2D analysis. Furthermore, the comparison is often based on low-fidelity aerodynamic models. This paper uses 3D, multi-fidelity aerodynamic models. The results confirm that this approach reveals information unavailable with simplified models. [ABSTRACT FROM AUTHOR]

Purpose: The present paper aims to assess the reliability and the limitations of analysing flight stability of a box-wing aircraft configuration known as PrandtlPlane by means of methods conceived for conventional aircraft and well known in the literature. Design/methodology/approach: Results obtained by applying vortex lattice methods to PrandtlPlane configuration, validated previously with wind tunnel tests, are compared to the output of a "Roskam-like" method, here defined to model the PrandtlPlane features. Findings: The comparisons have shown that the "Roskam-like" model gives accurate predictions for both the longitudinal stability margin and dihedral effect, whereas the directional stability is always overestimated. Research limitations/implications: The method here proposed and related achievements are valid only for subsonic conditions. The poor reliability related to lateral-directional derivatives estimations may be improved implementing different models known from the literature. Practical implications: The possibility of applying a faster method as the "Roskam-like" one here presented has two main implications: it allows to implement faster analyses in the conceptual and preliminary design of PrandtlPlane, providing also a tool for the definition of the design space in case of optimization approaches and it allows to implement a scaling procedure, to study families of PrandtlPlanes or different aircraft categories. Social implications: This paper is part of the activities carried out during the PARSIFAL project, which aims to demonstrate that the introduction of PrandtlPlane as air transport mean can fuel consumption and noise impact, providing a sustainable answer to the growing air passenger demand envisaged for the next decades. Originality/value: The originality of this paper lies in the attempt of adopting analysis method conceived for conventional airplanes for the analysis of a novel configuration. The value of the work is represented by the knowledge concerning experimental results and design methods on the PrandtlPlane configuration, here made available to define a new analysis tool. [ABSTRACT FROM AUTHOR]

Purpose The complex flow behavior over an oscillating aerodynamic body, e.g. a helicopter rotor blade, a rotating wind turbine blade or the wing of a maneuvering airplane involves combinations of pitching and plunging motions. As the parameters of the problem (Re, St and phase difference between these two motions) vary, a quasi-steady analysis fails to provide realistic results for the aerodynamic response of the moving body, whereas this study aims to provide reliable experimental data.Design/methodology/approach In the present study, a pitching and plunging mechanism was designed and built in a subsonic closed-circuit wind tunnel as well as a rectangular aluminum wing of a 2:1 aspect-ratio with a NACA64-418 airfoil, used in wind turbine blades. To measure the pressure distribution along the wing chord, a number of fast responding transducers were embedded into the mid span wing surface. Simultaneous pressure measurements were conducted along the wing chord for the Reynolds number of 0.85 × 106 for both steady and unsteady cases (pitching and plunging). A flow visualization technique was used to detect the flow separation line under steady conditions.Findings Elevated pressure fluctuations coincide with the flow separation line having been detected through surface flow visualization and flattened pressure distributions appear downstream of the flow separation line. Closed hysteresis loops of the lift coefficient versus angle of attack were measured for combined pitching and plunging motions.Practical implications The experimental data can be used for improvement of unsteady fluid mechanics problem solvers.Originality/value In the present study, a new installation was built allowing the aerodynamic study of oscillating wings performing pitching and plunging motions with prescribed frequencies and phase lags between the two motions. The experimental data can be used for improvement of computational fluid dynamics codes in case that the examined aerodynamic body is oscillating. [ABSTRACT FROM AUTHOR]

Purpose The purpose of this paper is to present the methodology and approach adapted to conduct a wind tunnel experiment on the inverted joined-wing airplane flying model together with the results obtained.Design/methodology/approach General assumptions underlying the dual-use model design are presented in this paper. The model was supposed to be used for both wind tunnel tests and flight tests that significantly drive its size and internal structure. Wind tunnel tests results compared with the outcome of computational fluid dynamics (CFD) were used to assess airplane flying qualities before the maiden flight was performed.Findings Extensive data about the aerodynamic characteristics of the airplane were collected. Clean configurations in symmetric and asymmetric cases and also configurations with various control surface deflections were tested.Practical implications The data obtained experimentally made it possible to predict the performance and stability properties of the unconventional airplane and to draw conclusions on improvements in further designs of this configuration.Originality/value The airplane described in this paper differs from frequently analyzed joined-wing configurations, as it boasts a front lifting surface attached at the top of the fuselage, whereas the aft one is attached at the bottom. The testing technique involving the application of a dual-use model is also innovative. [ABSTRACT FROM AUTHOR]

Purpose The traditional numerical methods to predict the interaction between the wing and propeller are too complex and time-consuming for computation to a certain extent. Therefore, they are not applicable for a real-time integrated turboprop aircraft model. This paper aims to present a simplified model capable of high-precision and real-time computing.Design/methodology/approach A wing model based on the lifting line theory coupled with a propeller model based on the strip theory is used to predict the propeller-wing interaction. To meet the requirement of real-time computing, a novel decoupling parameter is presented to replace lifting line model (LLM) applied for wings with a simplified fitting model (FM).Findings The comparison between the LLM and the simplified FM demonstrates that the results of the FM have a good agreement with the results of the LLM, which means that the simplified FM has the advantages of both high-accuracy and real-time computation.Practical implications After simplification, the propeller-wing interaction model is suitable for a real-time integrated turboprop aircraft model.Originality/value A novel decoupling parameter is presented to replace LLM applied for wings with a simplified FM, which has the advantages of both high-accuracy and real-time computation. [ABSTRACT FROM AUTHOR]

Purpose The area behind the engine/wing junction of conventional civil aircraft configurations with underwing-mounted turbofans is susceptible to local flow separation at high angles of attack, which potentially impacts maximum lift performance of the aircraft. This paper aims to present the design, testing and optimization of two distinct systems of fluidic actuation dedicated to reduce separation at the engine/wing junction.Design/methodology/approach Active flow control applied at the unprotected leading edge inboard of the engine pylon has shown considerable potential to alleviate or even eliminate local flow separation, and consequently regain maximum lift performance. Two actuator systems, pulsed jet actuators with and without net mass flux, are tested and optimized with respect to an upcoming large-scale wind tunnel test to assess the effect of active flow control on the flow behavior. The requirements and parameters of the flow control hardware are set by numerical simulations of project partners.Findings The results of ground test show that full modulation of the jets of the non-zero mass flux actuator is achieved. In addition, it could be shown that the required parameters can be satisfied at design mass flow, and that pressure levels are within bounds. Furthermore, a new generation of zero-net mass flux actuators with improved performance is presented and described. This flow control system includes the actuator devices, their integration, as well as the drive and control electronics system that is used to drive groups of actuators.Originality/value The originality is given by the application of the two flow control systems in a scheduled large-scale wind tunnel test. [ABSTRACT FROM AUTHOR]

Purpose The aim of this study is to assess and describe possible consequences of a bird strike on a Polish-designed PZL-106 Kruk agricultural aircraft. Due to its susceptibility to such events, a wing slat has been chosen for analysis.Design/methodology/approach Smooth particle hydrodynamics (SPH) formulation has been used for generation of the bird finite element model. The simulations were performed by the LS-Dyna explicit finite element analysis software. Several test cases have been analysed with differing parameters such as impact velocity, initial velocity vector direction, place of impact and bird mass.Findings Results of this study reveal that the structure remains safe after an impact at the velocity of 25 m/s. The influence of bird mass on slat damage is clearly observable when the impact velocity rises to 60 m/s. Another important finding was that in each case where the part did not withstand the applied load, it was the lug where first failure occurred. Some of the analysed cases indicated the possibility a consequent wing box damage.Practical implications This finding provides the manufacturer an important insight into the behaviour of the slat and suggests that more detailed analysis of the current lug design might improve the safety of the structure.Originality/value Even though similar analyses have been performed, they tended to focus on large transport aircraft components. This investigation will enhance our understanding of structural response of small, low-speed aircraft to a bird impact, which is a realistic scenario for the chosen case of an agricultural plane. [ABSTRACT FROM AUTHOR]

Purpose This paper aims to compare three closed non-planar wing configurations with a reference conventional wing-plus-horizontal tail aircraft, considering structural aspects, weights and aerodynamic characteristics, as well as operational issues, such as cruise performance.Design/methodology/approach A vortex lattice code is used and coupled with an in-house code for structural beam calculation subroutine to evaluate the configurations as a function of the four main parameters identified in the study.Findings The study concludes that the non-planar wing configurations have better performances than a conventional aircraft. Moreover, the joined-wing configuration seems to be better than the others, including the box-wing configuration, achieving an increase of 17 per cent in the range for maximum payload compared to the reference aircraft and a 3 per cent reduction of maximum take-off weight.Research limitations/implications In the study, characteristic tools for a conceptual design are used, and, thus, absolute results should be considered with caution. Nonetheless, as all the cases are studied in the same way, there is a good precision in comparative or relative results.Practical implications The work shows that the non-planar wing configurations can be used as an alternative to the conventional aircraft to meet the objectives for the future of the aviation industry.Social implications Non-planar wing configurations are able to reduce fuel consumption. Their use could lead to reductions in pollutant emissions and the impact on the environment of commercial aviation.Originality/value This study considers aerodynamic and structural aspects at the same time, as well as several non-planar wing configurations, making possible to obtain a more realistic comparison between them. [ABSTRACT FROM AUTHOR]

Purpose – The purpose of this paper is to present a structural design and optimization module for aircraft structures that can be used stand-alone or in a high-fidelity multidisciplinary design optimization (MDO) process. The module is capable of dealing with different design concepts and novel materials properly. The functionality of the module is also demonstrated. Design/methodology/approach – For fast sizing and optimization, linear static finite element (FE) models are used to obtain inner loads of the structural components. The inner loads and the geometry are passed to a software, where a comprehensive set of analytical failure criteria is applied for the design of the structure. In addition to conventional design processes, the objects of stiffened panels like skin and stringer are not optimized separately and discrete layups can be considered for composites. The module is connected to a design environment, where an automated steering of the overall process and the generation of the FE models is implemented. Findings – The exemplary application on a transport aircraft wing shows the functionality of the developed module. Originality/value – The weight benefit of not optimizing skin and stringer separately was shown. Furthermore, with the applied approach, a fast investigation of different aircraft configurations is possible without constraining too many design variables as it often occurs in other optimization processes. The flexibility of the module allows numerous investigations on influence of design concepts and failure criteria on the mass and layout of aircraft wings. [ABSTRACT FROM AUTHOR]

Purpose – The purpose of this paper is to identifying ways to reduce the effects of wing-vortex interaction by applying surface porosity on selected areas of the exposed surface. A number of papers recently have investigated the aerodynamic implication of free-stream vortices impinging upon airfoils. Design/methodology/approach – The free-stream disturbance in these studies were represented by planting a vortex ahead of the wing or using some other disturbance invoking mechanism like von-Karman vortices in the wake of a cylinder or using a flipping plate to invoke a discrete vortex. In the present work, a well-defined method was used to germinate a system of controlled vortices of known strength, size and frequency ahead of the wing, and the impact of the subsequent interaction was studied with and without the presence of the surface porosity. The simulations tackled a number of cases when porosities of up to 20 and 22 per cent were applied to selected regions near the leading edge, with vortices of controlled strengths directed at the wing surface. Findings – The results showed that the effects of large vortices spanning the entire lengths of the wing can indeed be damped when porosity is selectively applied at strategic regions. Practical implications – Surface porosity application at strategic regions of a wing may dampen the effects of the unsteadiness of the incoming flow. This has profound implications on flight safety and structural damage prevention. Further implications could possibly be extended to UAV and wind turbines that operate at heavy gusting environment. Originality/value – Implementation of this particular method resolves some of the issues arisen when an airplane encounters atmospheric turbulence. [ABSTRACT FROM AUTHOR]

Purpose - The purpose of this paper is to describe the research performed on flexible-wing micro air vehicle (MAV). Typical attributes associated with the aerodynamics of MAVs are low Reynolds number, low altitude flying environments and low aspect ratio platforms. These attributes give birth to several challenges such as poor aerodynamic performance, nonlinear lift patterns and reduced gust tolerance. Flexible-wing MAV is renowned for improved aerodynamic characteristics such as smooth flight in gusty conditions than its rigid-wing counterpart. Design/methodology/approach -- The wind-tunnel experiments are carried out for various configurations to determine the ways of further enhancing lift. The baseline geometric description for all MAVs includes 15-cm box dimension and an aspect ratio of 1. The experimental results of the baseline configuration are compared with other experimental results available in literature. After due validation, the effects of following parameters are quantized and compared with the rigid-wing counterpart: underlying skeleton; wing membrane extension; wing membrane relaxation; and wing membrane material (latex, silk, poly-vinyl chloride plastic sheet and nylon). Findings -- It is found that the skeleton layout significantly governs the lift characteristics. The effect of membrane extension and relaxation proved to be of little advantage. Latex sheets are found to be the best choice for membrane material. The aerodynamic assessment at low Reynolds number has demonstrated significant improvement of lift characteristics for flexible wings over rigid-wing counterparts. Research limitations/implications -- The results presented in this paper are based on wind-tunnel experimentation. Further experimentation through flight test may be needed to reveal the true aerodynamic performance under unsteady maneuvers. Practical implications -- The material properties vary significantly during fabrication. A technique to standardize the properties of flexible membranes is a missing link in literature and warrants further investigation. Originality/value - This concept of flexible wing has shown high potential. The primary objective of this paper is to experimentally investigate ways of further enhancing the lift of flexible-wing MAVs by controlling flexibility passively. While various researchers have spent many years on developing the optimum wing frame for the flexible wing, research on different wing materials has been limited. This is the first paper of its kind covering all aspects of wing-frame design, material, effects of extension and relaxation on wing membrane. [ABSTRACT FROM AUTHOR]

Purpose -- The purpose of this paper is to explore some of the challenges associated with the integration of an LH2-fuelled advanced hybrid-electric distributed propulsion system with the airframe. The airframe chosen as a case study is an ultra-high-capacity blended wing body configuration. It is designed to represent an A-380 class vehicle but in the 2025-2030 timeframe. The distributed propulsion system is a hybrid-electric concept that utilizes high-temperature superconducting technologies. The focus of the study is the application of LH2 as a fuel, with comment being given to kerosene and LCH4. Design/methodology/approach -- The study consists of a conceptual design developed through the preliminary design phase and part way into the detailed design phase. Findings -- The relationship between passenger capacity and fuel capacity is developed. Some remaining challenges are identified. Practical implications -- The study supports further conceptual design studies and more detailed system studies. Social implications -- The study contributes to the development of more environmentally benign aviation technologies. The study may assist the development of solutions to the peak oil challenge. Originality/value -- The study explores the integration of a number of complex systems into an advanced airframe to an unusual depth of engineering detail. [ABSTRACT FROM AUTHOR]

Purpose - The purpose of this paper is to investigate the effect of spanwise shape of the leading edge on unsteady aerodynamic characteristics of wings during forward flapping and gliding flight. Design/methodology/approach - A computational fluid dynamics approach was conducted to analyze the flow around airfoils with sinusoidal-like protuberances at a Reynolds number of 10[sup 4]. Three-dimensional time-dependent incompressible Navier-Stokes equations are numerically solved by using finite volume method. A multigrid mesh method, which was applied to the situation of fluid across the heaving models is used to simulate this type of flow. The simulations are performed for the wavelength between neighbouring peaks of 0.25c and 0.5c. For each wavelength, two heights of the tubercles which are 5 per cent and 10 per cent of the chordwise length of wing, are employed on the leading edge of wings. The aerodynamic forces and flow structure around airfoils are presented and compared in detail. Special attention is paid to investigate the effect of leading-edge shape on the fluid dynamic forces. Findings - Present results reveal that the wings with leading-edge tubercles have an aerodynamic advantage during gliding flight and also have the potential advantages during flapping forward flight. Originality/value - On the basis of computational study, an improved scenario for flapping wing microaviation vehicle has been originally proposed. [ABSTRACT FROM AUTHOR]

Purpose - This is the first of two companion papers presenting the results of research into a contra-rotating propeller designed to drive a super manoeuvrable micro air vehicle (MAV). The purpose of this first paper is to describe the design process and numerical analyses. The second paper is devoted to the experimental results verifying the computations. Design/methodology/approach - Software based on the analytical formulas derived by Theodore Theodorsen was used in the design procedure. Three-dimensional finite-volume simulation, performed with the use of commercial software verified the results. Finally, two-dimensional simulation was conducted to explore the effect of the propeller-wing interaction. The meshes applied in these analyses are described. Findings - Propeller geometry received as a result of the design procedure is presented. The computation results for different turbulence models applied are discussed. Time dependent characteristics of contra-rotating propeller are presented as well as conclusions regarding propeller-wing interaction. Research limitations/implications - Propeller was designed for a fixed wing aeroplane, not for helicopter rotor. Therefore, conditions characteristic for fixed wing aeroplane flight are analysed only. Reynolds numbers below 50000 are considered. Practical implications - Designed contra-rotating propeller can be used in fixed wing aeroplane if torque equal to zero is required. Software based on the formulas derived by T. Theodorsen can be used to design the propellers. Originality/value - Software applied in the design procedure was originally developed by one of authors although it is based on the formulas derived by T. Theodorsen. Contra-rotating propeller simulation results for different turbulence models are discussed for the first time. Moreover, unique time dependent characteristics of contra-rotating propeller are presented. [ABSTRACT FROM AUTHOR]

Purpose – Understanding the performance and flight envelope of a damaged aircraft is a preliminary requirement to recover the aircraft after damage. This paper aims to provide a comprehensive understanding of wing damage effect on airplane performance, local stability, and flying quality of each trim state inside the achievable flight envelope. Design/methodology/approach – This paper demonstrates the use of attainable equilibrium points which are referred as trim states in order to estimate a damaged airplane manoeuvring flight envelope using a numerical computation method. Findings – Wing damaged airplane manoeuvring flight envelope is estimated for different portions of the wing tip loss. Local stability at each trim condition inside the estimated flight envelope is analysed, and also motion flight modes and flying quality sensitivity to the wing damage are explored. Originality/value – Local stability and flying quality analysis at each trim condition inside the flight envelope which demonstrate the effect of damage provides a criterion to prioritize the choice of trimmed flight condition as motion primitives for the airplane post-damage flight and safe landing. [ABSTRACT FROM AUTHOR]

Purpose – Computational efficiency is always the major concern in aircraft design. The purpose of this paper is to investigate an efficient aeroelasticity optimization design method. Analysis of composite wing elastic axis is presented in the current study and its application on aeroelasticity optimization design is discussed. Design/methodology/approach – Elastic axis consists of stiffness centers. The stiffness centers of eight cross sections are analyzed and the wing elastic axis is obtained through least-squares procedure. In the analysis of the cross section stiffness center, the wing model is approximated by assuming the wing cross section as a thin walled structure with a single cell closed section and assuming the composite material to be a 3D anisotropic material. In aeroelasticity optimization design, objective functions are taken to be the wing weight and elastic axis position. Design variables are the thickness and area of wing components. Findings – After aeroelasticity optimization design, the wing weight decreases while the divergent velocity increases. Meanwhile, it can achieve an expected result but costs much less computational time than the conventional method. Practical implications – The results can be used for aircraft design or as an initial value for the next detailed optimization design. Originality/value – The computational time can be dramatically reduced through the aeroelasticity optimization design based on the elastic axis. It is suitable for engineering applications. [ABSTRACT FROM AUTHOR]

The article presents a survey of the structural problems associated with variable sweep winged aircraft. Some of the advantages of wing sweep in an aircraft design are examined, including reduced fatigue damage. The two types of structural problems associated with variable geometry aircraft are discussed. Also provided are examples of different concepts for wing pivot mechanisms.

Purpose – The purpose of this paper is to investigate the aerodynamics of wing in ground effect with tiltable endplates for a new type wing-in-ground effect (WIG) craft. Design/methodology/approach – The concept of tiltable endplates was implemented into the design of a WIG craft. Numerical investigation on aerodynamics of the tiltable endplate was carried out. The endplate effect on aerodynamics was deeply investigated with a rectangular wing at given angle of attack and flight height. The size of endplate relative to whole wing was then studied based on given endplate deflection angle and flight height. Finally, aerodynamics and flow of tiltable endplate in various flight heights and endplate deflection angles were analyzed. Aerodynamics, pressure and wingtip vortex were recorded in the study. Findings – Endplate influences development of wingtip vortex and improves aerodynamics. Tiltable endplate can enable WIG craft to yield improved aerodynamic performance and worthwhile economy improvements on long-distance flights in and out of ground effect (OGE). Research limitations/implications – The results are entirely based on computational fluid dynamics (CFD). The gap between "numerical world" and "real world" depends on development and appropriate application of CFD. The current work shows further understanding of ground effect and aerodynamics of wing in ground effect. Practical implications – The aerodynamics and aerodynamic optimization of wing in ground effect are of the great importance for WIG craft. The work improves the design and research on aerodynamics of WIG craft. Originality/value – The concept of tiltable endplate for a new type wing in ground effect allows WIG craft to achieve good aerodynamic performance not only in ground effect but also in OGE. This was studied and proved in the current work. [ABSTRACT FROM AUTHOR]

Purpose - This paper seeks to develop an approach for the unsteady inverse problem of two-dimensional oscillating airfoils based on the finite difference method (FDM) solution of the transient Euler equations. Design/methodology/approach - The solution strategies are determined according to the mathematical model for the inverse-problem of oscillating airfoils. Then the unsteady nonreflecting far field boundary condition and the permeable wall boundary condition are employed to treat the boundary conditions. The applications are carried out for the modification of an oscillating airfoil according to the design targets of the unsteady pressure distribution in an oscillating period. Findings - The results show that the pressure distributions over the new airfoils coincide with the design objects indicating that the mathematical model and solution strategy developed in this paper is rational and reliable. Research limitations/implications - This method is limited to frictionless flow. Originality/value - The paper provides a new FDM solution of unsteady inverse problem for oscillating airfoils, which can be extended to treat the multipoint problem of airfoil design. [ABSTRACT FROM AUTHOR]

The methodology backed by relevant analysis results, flutter flight test (FFT) results, and their interpretation is presented for the flutter qualification of a transport aircraft with wing mounted external store suspension. Before the flight tests, extensive flutter analyses are performed by MSC/Nastran for different weight and CG configurations to reduce the number of costly FFTs based on the results of flutter analysis and decide on the aircraft test configurations. Analysis results and test results indicate that although the flutter speeds of the modified aircraft are lowered compared to the basic aircraft, the overall damping of the wing vibration modes stay above the values required by the flutter regulation MIL-A-8870C. [ABSTRACT FROM AUTHOR]

Highlights from the article: It is my pleasure to present this special issue of the I Aircraft Engineering and Aerospace Technology Journal i , dedicated to the topics of unmanned aerial vehicle (UAV) configurations, unmanned aircraft systems (UAS) architectures, performance of remotely control vehicles and systems for pilot support, discussed at the 7th European Aeronautics Science Network (EASN) International Conference on "Innovation in European Aeronautics Research"[[1]]. Christophe Hermans, President of the Council of European Aerospace Societies (CEAS) with keynote lecture "Aerospace Europe: Strengthening Collaboration and Knowledge Dissemination; Dr Fay Collier, Associate Director for Flight Strategy, Integrated Aviation Systems Program, NASA Langley Research Center with keynote lecture "Accelerating Market Introduction of Emerging Innovations through Integrated Technology Demonstrations";.