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Dimensional accuracy of selective laser melting (SLM) parts is one of manufacturers’ major concerns. The additive manufacturing (AM) process is characterized by high-temperature gradients, consolidation, and thermal expansion, which induce residual stress on the part. These stresses are released by separating the part from the baseplate, leading to plastic deformation. Thermo-mechanical finite elements (FE) simulation can be adopted to determine the effect of process parameters on final geometrical accuracy and minimize non-compliant parts. In this research, a geometry for process parameter calibration is presented. The part has been manufactured and then analyzed with industrial computed tomography (iCT). An FE process simulation has been performed considering material removal during base plate separation, and the computed distortions have been compared with the results of the iCT, revealing good accordance between the final product and its digital twin.

This research presents the application of a beam finite element, specifically derived for simulating bending–torsion coupling in equivalent box-beam structures with curvilinear stiffeners. The stiffener path was simulated and optimized to obtain an expected coupling effect with respect to four typical static load cases, including geometric constraints related to the additive manufacturing production method. The selected load condition was applied to the centroid of the beam section, and the structure performance was consequently determined. A variation in load position up to one-fourth of the beam width was considered for investigating the stiffener path variation corresponding to a minimum bending–torsion coupling effect. The results demonstrated the capability of such a beam finite element to correctly represent the static behavior of beam structures with curvilinear stiffeners and show the possibility to uncouple its bending–torsion behavior using a specific stiffener orientation. The simulation of a laser powder bed fusion process showed new opportunities for the application of this technology to stiffened panel manufacturing.

Purpose Structural configurations related to new green aircraft design require high efficiency and low weight. As a consequence, moderate-to-large deformation under operating loads arise and aeroelastic instabilities different with respect to rigid counterpart are possible. Coupled structural configurations can provide the right mean to overcome such a critical situations selecting the right coupling parameters and structural performance. In this work, the dynamic behaviour of stiffened box-beam architecture with selected optimal stiffener orientation to emphasize the bending-torsion coupling characteristics has been investigated. Methods An extensive experimental activity has been performed for a validation and confirmation of the numerical results. Two cantilever beams produced with different technologies and materials have been tested. Modal performance has been determined by means of a laser Doppler vibrometer (LDV), while Finite-Element Method (FEM) numerical simulation based on solid elements and equivalent single layer approach have been applied and compared. Experimental/numerical comparison have been presented pointing out the specific coupling performance of this architecture with respect to natural frequencies and modal shapes. Results The activity demonstrates a good correlation in natural frequencies that remains mostly under 4$$\%$$ % . Modal assurance criterion (MAC) has been considered in comparing experimental and numerical modal shapes. Conclusion The proposed innovative configuration demonstrates its capability to be used in aeroelastic critical problem as a mean to reduce their influence in aircraft design. The numerical procedure used for equivalencing the stiffened parts of the box-beam has also been validated in dynamical response confirming the possibility to be used in design phase.

Experimental and numerical investigations into the linear and nonlinear aeroelastic behaviour of very flexible High Altitude Long Endurance (HALE) wings are conducted to assess the effect of geometrical nonlinearities on wings displaying moderate-to-large displacement. The study shows that the dynamic behaviour of wings under large deflection, and specifically the edgewise and torsion natural frequencies and modal characteristics, are largely affected by the presence of geometrical nonlinearities. A modular wing structure has been manufactured by rapid prototyping and it has been tested to characterise its dynamic and aeroelastic behaviour. At first, several simple isotropic cantilever beams with selected crosssections are numerically investigated to extract their modal characteristics. Experiments are subsequently conducted to validate the geometrically nonlinear dynamics behaviour due to high tip displacement and to understand the influence of the beam cross-section geometry. The structural dynamics and aeroelastic analysis of a very flexible modular selected wing is then investigated. Clean-wing wind-tunnel tests are carried out to assess flutter and dynamic response. The wind-tunnel model display interesting aeroelastic features including the substantial influence of the wing large deformation on its natural frequencies and modal characteristics.

Electric vertical takeoff and landing (eVTOL) aircraft of the lift+cruise configuration are significantly impacted by the drag of the lift propellers even though they are only briefly used for takeoff and landing. This work aims to quantify the drag penalty of these high lift propellers, its impact on the performance of a lift+cruise eVTOL aircraft, and potential benefits from retracting the propellers during cruise. An eVTOL model is tested in the wind tunnel with propellers extended and retracted. Results show that retracting the propellers reduces parasitic drag by 38%. This measured drag reduction is used to estimate the performance of a passenger eVTOL and a surveillance drone employing a propeller retraction system. A passenger carrying eVTOL lift+cruise aircraft with a propeller retraction system could cruise 21% faster, at the same mission range, or increase range by 13% at the same flight speed. When endurance is the main mission objective, the benefits are smaller, in the order of 7%, and the potential gains are more sensitive to the weight of the retraction system.

This note proposes a novel element removal method for ground structured based topology optimization that utilizes a double-filtering scheme based on a graph representation of the topology. Two types of systems: planar resonators and compliant mechanisms, are used to illustrate the method.

A procedure for the characterization of low velocity impact damage and its effect on residual tensile and buckling behaviour of composite structures is presented. A simplified analytical method is used to enhance the damage simulation and a degradation model is introduced in the subsequent finite element analysis. Different degradation coefficients are defined for three impact energy stages. Two different types of analysis are performed using MD.Nastran software: the non-linear progressive failure analysis to estimate the residual tensile strength and the linear analysis to predict the critical uniaxial and shear buckling loads after impact. Tensile and buckling experiments are used to validate the present methodology. A good correlation is obtained for all the cases under investigation.

Purpose The purpose of the research activity is to identify the best configuration of piezoelectric (PZT) elements for a typical condition of wing aeroelastic instability. The attention is mainly focused on the flutter behavior of the structure. However, the model can be extended with low-impact adjustments to other loading conditions. Design/methodology/approach The dynamic system consists of a thin-walled beam, whose longitudinal faces are perfectly bonded by two PZT layers and it is excited by the aerodynamic forces to assume a simple harmonic oscillation motion. The equations of motion are obtained using an energy approach by applying the extended Hamilton principle in conjunction with the Ritz method for modal approximation. The external forces acting on the system are modeled according to the Theodorsen derivation. Findings The flutter speed and the power generated from flutter oscillations can be increased by acting on the length of the PZT elements. The results show that the model with the beam substrate totally covered by the PZT in its longitudinal direction is more effective for low electrical resistance, whereas for high resistance values, the beam substrate that is partially covered provides the best results. Furthermore, both flutter postponement and energy harvesting functions can be maximized by properly choosing the beam stiffness ratio. Practical Implications Depending on the parameter we want to maximize, that is, the flutter speed or the energy harvested, it is possible to identify the best system configuration from the analysis presented in this paper. Originality/value The originality of the work appears in the sensitivity study performed on a three-dimensional piezo-aeroelastic fluttering wing, whose optimal behavior in terms of flutter postponement and power generation is analyzed using two distinct parameters, the beam stiffness ratio and the PZT length.

The recent advances in battery energy density and electric propulsion systems for automotive applications are enabling the development of the electric vertical take-off and landing (VTOL) aircraft. The electric VTOL is a new means of transport that can fly like an aircraft and take off and land vertically like a helicopter, sometimes called personal aerial vehicle. This paper compares it to the existing vehicles that may compete with it and addresses the estimation of its performances in hover, cruise flight, and the transition phase. The main parameters affecting performances are then discussed. Considerable space is dedicated to the battery mass to total mass ratio.

Curvilinear stiffened panels are being developed for aerospace structures and other applications. This paper presents an experimental and numerical study into the vibration response of three curvilinear stiffened square plates clamped at the edges. The experimental modal analysis was performed using laser Doppler vibrometry (LDV) and an impact hammer test. Two of the three plates was constructed with curvilinear stiffener geometry, while the third was a straight stiffened plate with the stiffeners oriented at an angle to the edge of the plate. Even though the natural frequencies of the plates are similar, the different stiffening patterns was sufficient in providing unique directional properties and therefore distinct mode shapes. The numerical modal analysis of the plates was performed using finite element analysis (FEA). A comparison of the experimental and numerical results was carried out in terms of natural frequencies and mode shapes, using relative differences and modal assurance criterion (MAC), respectively. The experimental and numerical results were in good agreement for all the three plates. The difference between experimental and numerical natural frequencies was typically less than 5% and the diagonal MAC values typically ranged from 0.8 to 1. This is in line with previously published results in the literature. The reasons for differences between the experimental and numerical results, and the practical significance of the findings, are also discussed.

The flutter instability of high aspect ratio wings is numerically investigated by means of finite element aeroelastic analysis. The effect of static equilibrium deflection as well as the additional differential stiffness was introduced in the calculation by a new procedure based on MSC Nastran™ commercial code. The difference between the classic linear flutter calculation and the new procedure presented has been demonstrated numerically and also experimentally. The main parameters that influence the aeroelastic phenomenon were also highlighted and used to define the aeroelastic similarity model related to a typical aircraft of the high altitude long endurance aircraft category.

Despite the ubiquity of both linear and nonlinear multimember resonators in MEMS and kinetic energy harvesting devices very few research efforts examine the orientation of members in the resonator on its dynamic behavior. Previous efforts to design this type of resonator constrains the members to have relative orientations that are 0 ○ or 90○ to each other, i.e., the elements are connected inline with adjoining members or are perpendicular to adjoining members. The work expands upon the existing body of research by considering the effect of the relative orientation between members on the dynamic behavior of the system. In this manuscript, we derive a generalized reduced-order model for the design of a multi-member planar resonator that has integer multiple modal frequencies. The model is based on a Rayleigh Ritz approximation where the number of degrees of freedom equals the number of structural members in the resonator. The analysis allows the generation of design curves, representing all the possible solutions for modal frequencies that are commensurate. The generalized model, valid for an N-DOF structure, is then restricted for a 2- and 3-DOF system/member resonator, where the linear dynamic behavior of the resonator is investigated in depth. Furthermore, this analysis demonstrates a rule of thumb; relaxing restrictions on the relative orientation of members in a planar structure, allows the structure to exhibit exactly N commensurable frequencies if it contains N members.

The aim of this work is to illustrate a new wing morphing concept and to compare its performance with the conventional wing-flap configuration. While there is no formal definition for the word "morphing", it is usually considered to mean large shape change or transfiguration. In the field of aeronautics, "shape morphing" has been used to identify those aircraft that undergo significant geometrical changes to enhance or adapt to their mission profiles. This investigation aims to find out more about non conventional solutions over the classical wing structural design. The main goal in this research is to replace the ordinary wing structure adopted on a model aircraft with a new morphing wing. In this way, aerodynamic performances during different flight conditions may be improved thanks to the prevention of typical aerodynamic losses caused by geometrical discontinuities in conventional design.

In this manuscript, we investigate the use topology optimization to design planar resonators with modal fre- quencies that occur at 1 : n ratios for kinetic energy scavenging of ambient vibrations that exhibit at least two frequency components. Furthermore, we are interested in excitations with a fundamental component containing large amounts of energy and secondary component with smaller energy content. This phenomenon is often seen in rotary machines; their frequency spectrum exhibits peaks on multiple harmonics, where the energy is primarily contained in the rotation frequency of the device. Several theoretical resonators are known to exhibit modal frequencies that at integer multiples 1:2 or 1:3. However, designing manufacturable resonators for other geometries is still a daunting task. With this goal in mind, we utilize topology optimization to determine the layout of the resonator. We formulate the problem in its non-dimensional form, eliminating the constraint on the allowable frequency. The frequency can be obtained a posteriori by means of linear scaling. Conversely, to previous research, which use the clamped beam as initial guess, we synthesize the final shape starting from a ground structure (or structural universe) and remove of the unnecessary beams from the initial guess by means of a graph-based filtering scheme. The algorithm determines the simplest structure that gives the desired frequency’s ratio. Within the optimization, the structural design is accomplished by a linear FE analysis. The optimization reveals several trends, the most notable being that having members connected orthogonally as in the L-shaped resonator is not the preferred topology of this devices. In order to fully explore the angle of orientation of connected members on the modal characteristics of the device; we derive a reduced-order model that allows a bifurcation analysis on the effect of member orientation on modal frequency. Furthermore, the reduced order approximation is used solve the coupled electro-mechanical equation of a vibration based energy harvester (VEH). Finally, we present the performance of the VEH under various base excitations. These results show an infinite number of topologies that can have integer ratio modal frequencies, and in some cases harvest more power than a nominal L shaped harvester, operating in the linear regime.

An accurate aeroelastic assessment of powered HALE aircraft is of paramount importance considering that their behaviour contrasts the one of conventional aircraft mainly due to the use of high aspect-ratio wings with distributed propulsion systems. This particular configuration shows strong dependency of the wing natural frequencies to the propulsion distribution and operating conditions. Numerical and experimental investigations are carried out to better understand the behaviour of flexible wings, focusing on the effect of distributed electric propulsion systems. Several configurations are investigated, including a single propulsion system using an engine pod (a weight with embedded electric motor, a propeller, and the wing-attached structure) installed at selected spanwise positions, and configurations with two and three propellers.

In this paper a revised algorithm, derived by the Stud GA, to optimize the stacking sequence of layered composite structures is presented. The constrained combinatorial problem is formulated and hence solved by means of the algorithm presented herein. Unlike the Stud GA, the proposed algorithm uses a fitness-proportionate selection to find potential mates to the Stud and a self-adaptive extinction to avoid premature convergence. Elitism is also allowed. Standard benchmark problems have been used to fully assess the validity of the proposed algorithm. Results in terms of reliability and efficiency are compared against three GAs and a Permutation Search (PS) algorithm. It has been demonstrated, by means of a parametric study, that a low probability of extinction is sufficient to ensure fast convergence as well as reliable results. Finally the optimization of a more complex aeronautic structure is addressed aiming to prove the validity of the proposed algorithm in a two level optimization strategy. Results in terms of efficiency are compared against those of an in-house standard GA.

This article evaluates the amount of energy that can be extracted from a gust using an aeroelastic energy harvester composed of a flexible wing with attached piezoelectric elements. The harvester operates in a subcritical flow region. It is modeled as a linear Euler–Bernoulli beam sandwiched between two piezoceramics. The extended Hamilton’s principle is used to derive the harvester’s equations of motion and an eigenfunction expansion is used to form a three-degree-of-freedom reduced-order model. The degrees of freedom retained in the model are two flexural degrees for the in-plane and out-of-plane displacements, and a torsional degree for the rotational displacement. Wagner and Küssner functions are used to represent the unsteady aerodynamic and gust loading, respectively. The amount of energy extracted from the system is then compared for two different deterministic gust profiles, 1-COSINE and two sharp-edged gusts forming a square gust, for various magnitudes and durations. The results show that the harvester is able to extract more energy from the square gust profile, although for both profiles the harvester extracts more power after the gust has subsided.

The application of fuel cell (FC) technology to aircraft propulsion and/or energy supply is becoming of great interest for undoubted advantages in terms of pollution emissions and noise reduction. A better understanding of problems related to fuel cells applied to aeronautics is sought by the European Commission (EC) funded project Environmentally Friendly Inter-City Aircraft Powered by Fuel Cells (ENFICA-FC). The main objective of the ENFICA-FC project was to develop and validate the use of a fuel cell-based power system for the propulsion of more-electric/all-electric aircraft. The fuel cell system was installed in the light sport aircraft Rapid 200, which was flight and performance tested. One of the key items under investigation is the simulation of the cooling system and the evaluation of fuel cell temperature. The polymer electrolyte membrane fuel cell (PEMFC) is considered to be the best candidate for the fuel cell vehicle because it has high power density, solid membrane electrolyte, and as it operates at low temperatures, it has a fast start-up. However, to generate a reliable and efficient power response and to prevent membrane degradation or damage with hydrogen and oxygen depletion, a sophisticated control technique becomes vitally important. In particular, as the ionic conduction of the polymeric membrane is a function of its degree of humidification, the stack temperature has to be carefully controlled to avoid phenomena of water evaporation, causing an increase of ohmic drop and a decrease of stack performances. The output voltage and hence the power of the fuel cell system is affected considerably by the change of the stack temperature. A simplified fluid-dynamic model has been developed and validated by computational fluid dynamics (CFD) analysis and it is used to compute the air flow to the fuel cell heat-exchanger inlet. Propeller effects are included referring to an optimal propeller specifically designed for the ENFICA-FC project. A mathematical model of the fuel cell system dynamics coupled with the fluid-dynamic model was studied in detail and experimentally validated during two flight tests of the Rapid 200-FC.

CubeSats primary structures are usually made with aluminium alloys, with few examples of CFRP primary structures under study. Power system battery arrays usually occupy spacecraft internal volume and mass that should be available to the payload. A CFRP structural/battery array configuration has been designed, allowing to integrate the electrical power system in the bus primary structure. Its configuration has been developed with the modular design philosophy of the AraMiS CubeSat. It is sized as a tile, mounted on an external face of the 1U CubeSat. It accommodates two solar cells, while the opposite face accommodates power system circuitry. Following a cellular structure concept, a set of commercial LiPo batteries has been placed between two CFRP panels and spaced out with CFRP ribs. Compliance with launch mechanical loads has been evaluated with a finite element analysis. A preliminary thermal analysis has been performed to simulate a LEO orbit environment. The results indicate that even with a low degree of structural integration, more volume and mass can be allocated to the payload, with respect to traditional, functionally separated designs in aluminium alloy. The low degree of integration allows to employ relatively cheap, commercial off-the-shelf components.

The application of fuel-cell technology to aircraft propulsion and/or energy supply is becoming of great interest for undoubted advantages in terms of pollution emissions and noise, features particularly important for commuter airplanes that usually take off and land from and in urban areas. The conversion of conventional aircraft into more/all-electric one tends to be based on the installation of such innovative systems. A better understanding of problems related to fuel cells applied to aeronautics is sought by the European Commission funded project environmentally friendly intercity aircraft powered by fuel cells ENFICA-FC. The main objective of the ENFICA-FC project is to develop and to validate the use of a fuel-cell based power system for propulsion of an all-electric aircraft. The fuel-cell system will be installed in a light sport aircraft Rapid 200 that will be flown and performance tested as a proof of functionality and future applicability. Specific aspects have to be investigated in the onboard installation of the innovative system and new design indications have to be pointed out in order to fulfill the conversion. One of the key items under investigation is the simulation of existing cooling system and the evaluation of motor and fuel-cell temperature; the temperature has to be maintained within the limits established by manufacturers of critical systems during all likely operating conditions as indicated by aeronautic regulations for general aviation. The computational problem addressed in this paper is the numerical computational fluid dynamics CFD simulation of the existing air-cooling system that satisfies air request for cooling and venting. An engineering model has been developed and it is used to support air inlet and outlet design. The flow is reasonably approximated by a potential flow plus boundary layer; hence, total upstream pressure losses are neglected except for those within the thin boundary layer. Pressure recovery of incoming cooling air and pressure coefficient distribution have been studied by using VSAERO panel code, and the optimal position of cooling and venting intakes is defined according to aerodynamic results. Propeller effects are included by referring to an optimal propeller specifically designed for the ENFICA-FC project. Propeller slip-stream wake is modeled by an actuator disk plus a swirl membrane and used to study the problem of motor, electronic, and fuel cells' cooling during takeoff.

Flutter-critical velocity is usually estimated from deterministic analyses by assuming that physical and geometrical parameters are perfectly known. However, aleatory and epistemic uncertainties, especially those associated with the definition of material properties, are intrinsically present partly because of the random nature of every physical system and partly because those properties are evaluated from a finite number of observations. The paper focuses on the combination of structural reliability analysis with aeroelastic simulation to give a correct flutter speed evaluation for design purposes. Two typical aeronautical test cases have been considered: (1) an isotropic material structure and (2) a composite material structure. Different computational methodologies (classical and developed by authors) have been coupled with a simple two-dimensional aeroelastic model to study the qualitative consequences of uncertainty in determination of critical flutter speed and to provide a comparison between differ...

The next generation of slender, flexible aircraft wings requires extremely lightweight structures capable of carrying a considerableamountof non-structuralweight.With the increased slenderness and flexibility, possible with the advent of advanced composites, these wings can exhibit aeroelastic instabilities quite different from their rigid counterparts. The design of highly flexible aircraft, such as high-altitude long endurance (HALE) configurations, must include phenomena that are not usually considered in traditional aircraft design, such that an alternative design philosophy has been proposed for this class of vehicles. The discussion in this article is restricted, among the various aeroelastic phenomena, to the flutter condition. Classical procedures usually refer to aero-structural systems where the undeformedstate is taken as the reference point. This is not the case with slender wing configuration where, due to the high structural flexibility, a proper beam model, capable of describing the structural flight deflections, should be adopted. Consequently, the flutter analysis has to be performed considering the deflected state as a reference point. Herein, an approximate procedure is proposed and flutter is evaluated by means of Galerkin's approach applied to the perturbed small motions of the aero-structural system. The effect of typical parameters, including stiffness ratio, mass eccentricity, store pod, deflection amplitude, as well as the wing aspect ratio, are considered. For a simplified wing configuration, comparisons between analytical and experimental findings are presented along with discussions and suggestions for new design criteria. Keywords: advanced aircraft design, linear/non-linear aeroelasticity, flutter, experimental aeroelasticity

This paper addresses stability analysis, control design, and simulation of high-altitude, long-endurance unmanned aerial vehicles. These aircraft are highly flexible and have very low structural frequencies. Hence, the separation between the elastic and rigid body motions no longer exists, and an approach that unifies these two motions must be used. The available data of the aircraft considered include the geometry of the aircraft, aerodynamic data, and mass, flexural rigidity, and torsional rigidity distributions. The wing of the aircraft is modeled as a flexible beam undergoing bending and torsion, whereas the remaining members are assumed to be rigid. The equations of motion are obtained by means of the Lagrangian equations in quasi coordinates. A perturbation approach separates the problem into nominal dynamics and perturbation dynamics. The equations for nominal dynamics are used to design desired maneuvers and to determine the corresponding structural deformations. The equations for perturbation dynamics are used to address stability of the aircraft on the desired flight paths, to design feedback controls to maintain stable flights, and to simulate the motion of the aircraft.

The application of fuel cell technology to aircraft propulsion and/or energy supply is attracting a great deal of interest because of the undoubted advantages that can be attained in terms of pollution emissions and noise reduction. The goal of the ENFICA-FC project (Environmentally Friendly Inter City Aircraft powered by Fuel Cells) has been to develop and validate new fuel cell-based power system concepts for more/all electric aircrafts pertaining to the "inter-city" segment of the market. As a result of the project, a technology demonstrator has been designed, manufactured, and tested: a proton exchange membrane (PEM) fuel cell-based power system was installed in the Skyleader RAPID200 aircraft which underwent its demonstrative flights in 2010. On the basis of the flight experiences gained during the ENFICA-FC program, the main topic of this paper is the description and validation of the multi-disciplinary dynamic model of the PEMFC-based powertrain of the ultra-light aircraft, fed with a pure hydrogen flow. Within this paper, emphasis is given to the model validation and dynamic system behavior. In order to validate the predicted results, data from real fuel cell flight (two 40-minute missions performed during the ENFICA-FC test campaign) have been used and are presented in this paper. Fil: Correa Perelmuter, Gabriel. Consejo Nacional de Investigaciones Científicas y Técnicas; Argentina. Universidad Nacional de Catamarca. Facultad de Ciencias Exactas y Naturales; Argentina Fil: Santarelli, Massimo. Politecnico di Torino; Italia Fil: Borello, F.. Politecnico di Torino; Italia Fil: Cestino, E.. Politecnico di Torino; Italia Fil: Romeo, G.. Politecnico di Torino; Italia

Research is being carried out at the Turin Polytechnic University with the aim of designing a HAVE/UAV (High Altitude Very-long Endurance/Unmanned Air Vehicle). The vehicle should be able to climb to an altitude of 17–20 km by taking advantage of direct sun radiation and maintaining a level flight; during the night, a fuel cells energy storage system would be used. A computer program has been developed to carry out a parametric study for the platform design. The solar radiation change over one year, altitude, masses and efficiencies of the solar and fuel cells, as well as the aerodynamic performances have all been taken into account. The parametric studies have shown how the efficiency of the fuel and solar cells and mass have the most influence on the platform dimensions. High modulus CFRP has been used in designing the structure in order to minimize the airframe weight. A Blended Wing Body (BWB) configuration of Solar HALE Aircraft Multi Payload & Operation (SHAMPO) with 8 brushless electric motors has been developed, as a result of the parametric study. The BWB solution, compared with conventional designs, seems to provide the best compromise between performance, availability of surfaces for solar-cells, and volume for multi-payload purposes. Several profiles and wing plans have been analyzed using the CFD software Xfoil and Vsaero. The airfoil coordinates at the root and along the wing span as well as the wing planform were optimised to achieve the best efficiency. A FEM analysis was carried out using the Msc/Patran/Nastran code to predict the static and dynamic behaviour of the UAV structure.

The model of a geometrically nonlinear wing hosting piezoelectric patches with the dual purpose of suppressing aeroelastic vibration and harvesting vibrational energy is presented in this paper. The nonlinearities are introduced in order to consistently reproduce the behavior of the flexible structure, since moderate to large displacements can occur in response of external loading conditions. A nonlinear shear underfomable 3-D Euler-Bernoulli beam theory is used to model the displacements field and structural nonlinearities up to the third order are retained in the model of a straight untapered composite wing. A linear indicial functional representation of the unsteady aerodynamic loads in an incompressible flowfield is adopted. The extended Hamilton principle is used to derive the aeroelastic equations of motion. The composite cantilever wing includes two piezoelectric elements, perfectly bonded on its lower and upper longitudinal surfaces in the proximity of the wing root, and electrically connected by a resistive load, functioning as energy harvesting devices. During the state of deformation the piezoelectric components induce electric charges to be stored for future use as a supplementary power source. The piezoelectric layers also function as damping elements with desired load alleviation properties. The effectiveness of such a solution, both in terms of the amount of energy harvested and load alleviation characteristics, for a well defined wing configuration have been evaluated in this paper. Numerical results and discussions are followed by pertinent conclusions and directions for future work.

The innovative highly flexible wings made of extremely light structures, yet still capable of carrying a considerable amount of non- structural weights, requires significant effort in structural simulations. The complexity involved in such design demands for simplified mathematical tools based on appropriate nonlinear structural schemes combined with reduced order models capable of predicting accurately their aero-structural behaviour. The model presented in this paper is based on a consistent nonlinear beam-wise scheme, capable of simulating the unconventional aeroelastic behaviour of flexible composite wings. The partial differential equations describing the wing dynamics are expanded up to the third order and can be used to explore the effect of static deflection imposed by external trim, the effect of gust loads and the one of nonlinear aerodynamic stall. As to provide a rationale evaluation of the important nonlinear contributions in aeroelastic wing simulations, the aeroelastic governing equations in linear, quadratic, and cubic forms, are considered and compared. Critical conditions will be investigated. The governing equations will be solved based on a discretization technique, along with Galerkin's method starting with a reduced order procedure and passing through a multi-modal approach. Three degrees of freedom in edgewise, flapwise, and torsion, are needed to describe efficiently the dynamics of the wing. Interesting design indicators will be highlighted.

Purpose – This study aims to analyse slender thin-walled anisotropic box-beams. Fiber-reinforced laminated composites could play an important role in the design of current and future generations of innovative civil aircrafts and unconventional unmanned configurations. The tailoring characteristics of these composites not only improve the structural performance, and thus reduce the structural weight, but also allow possible material couplings to be made. Static and dynamic aeroelastic stability can be altered by these couplings. It is, therefore, necessary to use an accurate and computationally efficient beam model during the preliminary design phase. Design/methodology/approach – A proper structural beam scheme, which is a modification of a previous first-level approximation scheme, has been adopted. The effect of local laminate stiffness has been investigated to check the possibility of extending the analytical approximation to different structural configurations. The equivalent stiffness has been evalua...

The lifting surfaces of next generation of flying vehicle exhibits enhanced flexibility, particularly for high aspect-ratio solutions needed for high-altitude long endurance aircrafts and for 24/7 operations. Often the wing of these vehicles is designed as slender body with an aeroelastic behavior distinctive of cantilever beam. Based on this typical assumption the governing equations of a thin-walled beam modified to account for surface mount piezoelectric elements and subjected to deterministic external gust loads have been derived and its dynamic behavior examined. This paper assesses the effectiveness of piezoelectric elements to carry out load alleviation function over the slender structure invested by atmospheric disturbances along with the evaluation of the amount of the power density harvested via a suitable electric circuit connected to the piezoelectric elements.Copyright © 2014 by ASME

Long-endurance, unmanned air vehicles (UAVs) usually employ high-aspect-ratio, flexible wings to reduce weight and power required to achieve extended flight operation. There are currently several advanced concept designs and prototypes that will sustain autonomous operations for months or even years. The increased wing flexibility can lead to large structural deflections for trimmed flight states. These large elastic deformations within the static equilibrium state of the vehicle can induce aeroelastic instabilities quite different from their rigid counterparts. Therefore, it is imperative to perform flutter analysis, which considers the trimmed, large amplitude deformed state as a reference condition. Considering the complexity of the system under investigation, the aeroelastic flutter analysis process usually includes an experimental test phase for verification purposes. In this respect, wind tunnel test models provide the opportunity to modify and calibrate theoretical models, through a process of model updating. This paper proposes the development of an experimental, aeroelastic, slender-wing model that can be designed using a scaling procedure. The model is based on a consistent, aeroelastic, beam-wise scheme, capable of simulating the unconventional aeroelastic behavior of flexible wings.

Fuel cells could become the main power source for small general aviation aircraft or could replace APU and internal sub-systems on larger aircraft, to obtain all-electric or more-electric air vehicles. There are several potential advantages of using such a power source, that range from environmental and economic issues to performance and operability aspects. A preliminary design is reported. Also, the paper contains a description of testing activities related to experimental flights of an all-electric general aviation aircraft fueled by hydrogen. Great importance has been given to the testing phase of the prototype and examples of each testing stage are shown ranging from the single components to the final test flights. During the 6 experimental flights a rotation speed of 84 km/h was obtained in 184 m of taxi at power of 35 kW. Level flight was attained at 135 km/h and endurance of 39 min by mean of only a fuel cell power setting (speed world record for the FAI Sporting Code Category C – airplane).

The aeroelastic design of highly flexible wings, made of extremely light structures yet still capable of carrying a considerable amount of non-structural weights, requires significant effort. The complexity involved in such design demands for simplified mathematical tools based on appropriate reduced order models capable of predicting the accurate aeroelastic behaviour. The model presented in this paper is based on a consistent nonlinear beam model, capable of simulating the unconventional aeroelastic behaviour of flexible composite wings. The partial differential equations describing the wing dynamics are reduced to a dimensionless form in terms of three ordinary differential equations using a discretization technique, along with Galerkin's method. Within this approach the nonlinear structural model an unsteady indicial based aerodynamic model with dynamic stall are coupled. Only three degrees of freedom in edgewise, flapwise, and torsion, are needed to describe efficiently the dynamics of the wing and to evaluate the sensitivity to system parameters, such as stiffness ratio, aspect ratio, and root angle of attack. Interesting design indicators will be highlighted. In addition to analytical results, a wind tunnel test article will be introduced to assess the validity of the proposed model.

The stability analysis of plates and shells in high speed flow deals with the determination of the flutter instability boundary. A linear analysis is made using the basic principles of the theory of aero-elasticity of isotropic bodies, the theories of flexible plates, the stability equations and associated boundary conditions obtained through a linear formulation. Herein, the nonlinear stability of flexible plate immersed in a high speed gas flow is considered. The model takes into account quadratic and cubic aerodynamic nonlinearities as well as cubic geometric nonlinearities. It is shown that the inclusion of quadratic aerodynamic nonlinear components can lead to the appearance of "amplitudefrequency" phenomena in both the pre-critical as well as in the post-critical flow speed regimes. The influence of the free stream flow speed on the "amplitude-frequency" dependence phenomena is also presented.

Many of the catastrophic errors in health care are related to inadequate procedures. Robust preventative actions are therefore required to minimize the risks inherent to the prescribing, dispensing and administration of medicines. The redesigning of subsystem processes is a goal that should be undertaken to improve the overall safety of hospitals. Many organizations are beginning to apply traditional aerospace engineering methodologies to the study of patient safety. In a hospital setting, the pharmacy department is responsible for the procurement, distribution, and control of all medicines used within the organization. Pharmacists should ensure that medicines are delivered to patient care areas in a safe and secure manner and that they are available to the administration within a time frame that meets the essential needs of patients. A Failure mode analysis applied to pharmaceutical distribution has been carried out in the S.G.Battista Hospital in Turin (Italy), from the receiving of the goods stage to the delivery to the departments. Two main high risk activities have been identified from risk matrices and risk priority number analysis. The first activity concerns the picking and packing phases and the second one is due to the low efficiency of the random controls of pharmacists. The analysis has also pointed out some medium risk activities in the refill request control, storage temperature control and consignment to delivery service departments. The present analysis has offered the opportunity of quantifying safety within a specific hospital environment and of designing feasible corrective actions. The real effectiveness of the proposed actions will be verified during a subsequent experimental phase.

The aim of this work is to apply an innovative adaptive ℒ1 techniques to control flutter phenomena affecting highly flexible wings and to evaluate the efficiency of this control algorithm and architecture by performing the following tasks: i) adaptation and analysis of an existing simplified nonlinear plunging/pitching 2D aeroelastic model accounting for structural nonlinearities and a quasi-steady aerodynamics capable of describing flutter and post-flutter limit cycle oscillations, ii) implement the ℒ1 adaptive control on the developed aeroelastic system to perform initial control testing and evaluate the sensitivity to system parameters, and iii) perform model validation and calibration by comparing the performance of the proposed control strategy with an adaptive back-stepping algorithm. The effectiveness and robustness of the ℒ1 adaptive control in flutter and post-flutter suppression is demonstrated. Results and discussion will follow with pertinent conclusions and future outlooks.