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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.

Aircraft preliminary design requires a lot of complex evaluations and assumptions related to design variables that are not completely known at a very initial stage. Didactical activity becomes unclear since students ask for precise values in the starting point. A tentative in providing a simple tool for wing weight estimation is presented devoted to overcome these common difficulties and clarifies the following points: a) the intrinsic iterative nature of the preliminary design stage, b) provide useful and realistic calculation for the wing weight with very simple assumption not covered by cumbersome calculations and formulas. The procedure is applied to the calculation of wing weight for a typical general aviation aircraft in the preliminary design stage. The effect of the main variables on the wing weight variation is also presented confirming well-known results from literature and design manuals.

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.

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 increasing utilization of green resources is driving research toward the design and development of new and more efficient energy harvesting systems. As far as the wind energy sector is concerned, aerodynamic design and the definition of new airfoils, specifically devoted to the application of wind rotors, has given a huge impulse in this direction. However, the rigid blade concept is still an important constraint in energy extraction during off-design conditions or start-up phases. In order to cope with this problem, a new system has been developed and manufactured within the specifically funded VENTURAS® (VENTo: Una Risorsa Altamente Sfruttabile or ‘wind: a highly exploitable resource’) Project. The new system applies morphing to modify blade twisting along the blade span (modifying the pitch angle of some sections) while maintaining the same chord distribution as the design configuration. This innovation offers the possibility of partially adapting the blade configuration to the operating po...

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.

Purpose – The purpose of this paper is to investigate the effect of fatigue life reduction of 2024 Al alloy for aerospace components due to the corrosive (exfoliation) environment. Both standard fatigue tests on prior corroded samples and fatigue tests conducted with the samples in corrosive solution are developed to define some guidelines for the inclusion of such effect in design and to improve aircraft life management. Design/methodology/approach – The effect of corrosion is taken into consideration, introducing specific concentration factors into the life estimation relationship. Differences between fatigue in corroded specimens and fatigue in presence of corrosive environment are emphasized. No crack propagation is considered. Two alternative procedures are considered in the analysis: “a-procedure” based on maximum stress calculated on un-corroded sample section; “b-procedure” based on stress calculated on final residual section, including corrosion. Findings – Related concentration factors are deriv...

Purpose This paper aims to present a systematic performance-oriented procedure to predict structural responses of composite layered structures. The procedure has a direct application in the preliminary design of aerospace composite structures evaluating the right and most effective material. Design/methodology/approach The aforementioned procedure is based upon the definition of stiffness invariants. In the paper, the authors briefly recall the definition and the physical explanation of the invariants, i.e. the trace; then they present the scaling procedure for the selection of the best material for a fixed geometrical shape. Findings The authors report the basic principles of the scaling procedure and several examples pertaining typical responses sought in the preliminary design of aeronautic structures Research limitations/implications Typically, during early stages, engineers had to perform the daunting task of balancing among functional requirements and constraints and give the optimum solution in terms of structural concept and material selection. Moreover, preliminary design activities require evaluating different responses as a function of as less as possible parameters, ensuring medium to high fidelity. The importance of incorporating as much physics and understanding of the problem as early as possible in the preliminary design stages is therefore fundamental. A robust and systematic procedure is necessary. Practical implications The time/effort reduction in the preliminary design of composite structures can increase the overall quality of the configuration chosen. Social implications Reduction in design costs and time. Originality/value In spite of the well-known invariant properties of composites, the application and extension to the preliminary design of composite structures by means of a scaling rule is new and original.

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.

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

In the study of specific aerospace systems, such as Solar Powered HALE-UAV, the required large structural dimensions force designers to consider modular constructions with removable connections. A specific advanced pin-hole connection has been designed and its structural characteristics and performances have been discussed, with respect to static loading and fatigue conditions. The presented solution has been introduced into the main structure connections of a solar platform prototype which was subjected to a standard static shear-bending test. The test revealed very good behaviour, as expected. This indirect validation makes the authors feel confident that this advanced solution could be applied extensively.

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.

A preliminary dynamic and aeroelastic investigation is performed in a simplified equivalent configuration derived from the HELIPLAT HAVE-UAV (high-altitude very-long endurance unmanned air vehicle) wing structure. The presence of slender wing structures introduces in the structural design a certain degree of non-linearity, forcing the designer to deal with specific phenomena, one of which is described. A simple dynamic system is used to point out the important effect due to the inclusion of the deformed equilibrium configuration into the dynamic behaviour of the system. A modified wing structure configuration is used to point out this effect in the flutter behaviour of slender structures. Some interesting design indications are presented.

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.

Within the research project HELINET (network of stratospheric platforms for traffic monitoring, environmental surveillance and broadband services), a preliminary reliability design is performed in conjunction with the platform design. The aim of the discussion consists in determining the configuration of the propulsion system and in the choice of the number of propulsion units (a propulsion unit consists of one propeller and one or more motors). The reliability and the weight of different very simple systems are evaluated and compared with one another in order to indicate the best one. The analysis of the fleet reliability and a study of its standby platforms is also taken into consideration. The analysis of the reliability of the solar cell array is also an interesting issue. A parametric analysis based on well-known formulae is used. Some results are verified by means of a numerical simulation based on the Monte Carlo methodology. Some important design indications could be obtained from this effort.

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.

Among renewable energy sources, wind is one of the most exploited, due to the relative low cost and independence from the sun. However, to harvest the highest amount of energy, it is important to maximize the global efficiency of the wind generator. To do so it is necessary to understand the behavior of each mechanical and electrical component and specifically how they interact. In this work we present WindDesigner, an open tool for wind generators analysis and design. It allows to dynamically configure the composition of the generator and to estimate important parameters, like efficiency and energy production, according to the variation of wind speed. It is written in Matlab and its modular and open structure can be easily expanded and improved. The tool has been applied to the design of a specific wind generator, enabling design space exploration and selection of alternative implementation options. Thanks to WindDesigner it will be possible to design the optimum wind generator structure for every site condition and load needs.

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.

Additional loads are created on wing-box panels when bending loads are applied to the box beam. Bending curvature, associated with longitudinal load, causes a distributed load perpendicular to both upper and lower wing-box panels resulting in a lateral pressure. The nonlinear analysis of anisotropic panels with initial imperfections has been developed for symmetric panels under combined biaxial compression, shear loads, and lateral pressure in order to obtain the out-of-plane panel deflection in the pre- and postbuckling range. The nonlinear differential equations are expressed in terms of the out-of-plane displacement and the Airy function. They are solved with the Galerkin method for various boundary conditions. Only by taking into account the nonlinear effects of the lateral pressure on the skin panels is it possible to obtain a good correlation between theoretical and experimental behavior. Experimental results on graphite/epoxy wing-box beams under pure bending showed remarkable deformations of the panels (particularly the compressed panel) in a direction normal to their centroidal surface, affecting both local strain values and the failure load value of the structures. Due to the effect of the lateral pressure, the blade-stiffened panel wing box showed a remarkable deformation also in the prebuckling Held.

This article reports both the theoretical analysis and experimental results performed on advanced composite wing boxes under pure torsion. By taking into account the nonlinear effective shear modulus of skin panels operating during postbuckling, it is possible to obtain good correlation between the theoretical and the experimental behavior. An incomplete diagonal shear stress field is used to calculate the effective shear modulus that can be reduced by up to 50% in comparison to the unbuckled panels. This causes a drastic reduction in the wing box's torsional stiffness. At a shear load triple that of buckling, our results revealed angles of twist up to 50% greater than those predicted by the linear analysis.

The postbuckling behaviour of anisotropic panels with initial imperfections is investigated. Nonlinear analysis is developed for symmetric panels under combined biaxial compression and shear loads in order to obtain the out-of-plane panel deflection in the postbuckling range. The nonlinear differential equations are obtained by using the principle of the stationary value of the total potential energy and are expressed in terms of the out-of-plane displacement and the Airy function. They are solved with the Galerkin method for various boundary conditions. The theoretical results are in good agreement with the few results concerning isotropic plates and the simply supported anisotropic results found in literature. A new and original test facility was built in order to apply simultaneously both biaxial compression and shear loads. An anisotropic panel, clamped along the four edges, has been tested under different combined loads; the correlation between the experimental and analytical results has been quite good. The results demonstrate the influence of initial imperfections on panel deflection in the postbuckling range: the curves obtained experimentally in the presence of imperfections cross those obtained theoretically, where imperfections are ignored (producing lower out-of-plane deflection).

Next generation of composite civil aircrafts and unconventional configurations, such as High Altitude Long Endurance HALE-UAV, exhibit aeroelastic instabilities quite different from their rigid counterparts. Consequently, one has to deal with phenomena not usually considered in classical aircraft design. Alternative design criteria are needed in order to maintain the safety levels imposed by the regulations and required for certification. The A2-Net-Team project aims to build a multi-disciplinary network of researchers with complementary expertise to develop analytical methods used for a better understanding and assessment of the factors contributing to the occurrence of critical aeroservoelastic instabilities. Along with modeling and numerical investigations a test article will also provide the opportunity to modify and calibrate theoretical models, to highlight and explore their limits, to recommend the necessary modifications and future pertinent investigations. The objective of this paper is to provide the authors' point of view on a consistent scaling procedure applicable when investigating the flutter stability of a full-size high aspect-ratio composite wing structure by means of an economical laboratory wind tunnel test model. Sizing and technological development aspects are discussed along with a detailed design procedure to gather the necessary data for the model manufacture. To exemplify the procedure a wooden model is fabricated and used to provide insight into the scaling procedure and verify the acquired data. Finally a dynamically scaled composite model is defined for a wind tunnel aeroelastic test.

Some main activities have been developed as part of the HELINET research project which was financed in January 2000, by the European Commission, (Fifth Framework Program in the IST action), to develop a European project in the field of stratospheric platforms. From the aeronautical point of view the following items were studied: (1) the design of an autonomous high altitude long-endurance unmanned air vehicle (HALE-UAV) platform capable of remaining aloft for very long periods of time (between 6 and 9 months) using a solar-powered and fuel cell system and to gain a complete understanding of the feasibility of a near-term aerodynamic HALE concept, in particular as far as stratospheric platforms are concerned (mainly dependent on high efficiency and reliability of solar cells, fuel cells and electric motors), (2) an evaluation of the production and service costs and an assessment of the safety and regulatory aspects of the platform, and (3) the manufacturing of a scale-sized technological demonstrator and the execution of static tests on it up to the ultimate load. The first HELIPLAT® (HELIos PLATform) configuration was worked out, on the basis of a preliminary parametric study. The platform was a twin-boom tail type monoplane with eight brushless motors, a long horizontal stabilizer and two rudders. A scaled-prototype was designed to demonstrate the feasibility of this configuration and to perform some structural static and dynamic tests on it. The main CFRP structures were manufactured by CASA (Spain): the principal wing and horizontal tail tubular spars, booms, vertical tail spars and some reinforced ribs. These parts were delivered to the Aerospace Engineering Dept. (DIASP) at the Politecnico di Torino (POLITO) and assembled using special joints while some other necessary parts were manufactured by POLITO-DIASP. A parallel activity was performed to define the structural test configurations and structural test frame. The manufacturing activities and the development of the structural test system is described in the first part of the paper. Static and dynamic experimental tests were performed in two phases (2003 and 2004) on the prototype and the results of the static tests are presented in this paper and compared with numerical and theoretical computations.

High Altitude Long Endurance (HALE) aircraft exhibit aeroelastic behavior that is quite different from conventional flight vehicles. For such aircraft, large static structural deformations that occur during standard flight conditions could severely reduce the flutter speed. Limit cycle oscillations may occur, reducing their operational life. In this paper, a non-linear aeroelastic model is proposed as a design/diagnostic analysis tool. A geometrically exact non-linear structural model for large deformations by Hodges and Dowell, modified according to Da Silva second order geometrical non-linear terms, has been adopted. The structural model has been coupled with an unsteady aerodynamic model for an incompressible flow field, based on the Wagner aerodynamic indicial function, in order to obtain a non-linear aeroelastic model. Within the analysis, linear and non-linear static and dynamic aeroelastic responses have been considered. The approach enables one to expedite the calculation process. In addition to analytical and computational results, a high aspectratio wing model has been selected as a candidate for a feasibility study case and to validate the theoretical model. Preliminary tests have been conducted and more experiments are underway.

Research is at present being carried out at the Turin Polytechnic University with the aim of designing an HAVE/UAV (high altitude very-long endurance/unmanned air vehicle). The vehicle should climb to 17-20km by mainly taking advantage of direct Sun radiation and thereafter maintain 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, the altitude, masses and efficiencies of the solar and fuel cells, and the aerodynamic performances have all been taken into account. The parametric studies have shown how fuel cells and solar cells efficiency and mass have the most influence on the platform dimensions. A wide use of high modulus CFRP has been made in designing the structure in order to minimise the airframe weight. A first configuration of HELIPLAT® (HELIos PLATform) was worked out, following a preliminary parametric study. The platform is a monoplane with eight brushless electric motors, a twin-boom tail type with an oversized horizontal stabiliser and two rudders. The co-ordinates at the root and along the wing span as well as the wing planform were optimised to achieve the best efficiency. Several profiles and wing plans have been analysed using the CFD software Xfoil and Vsaero. Several wind-tunnel tests were carried out to compare the analytically predicted performances. A preliminary design of a scale-sized technological demonstrator was completed with the aim of manufacturing a proof-of-concept structure. A FEM analysis was carried by using the Msc/Patran/Nastran code to predict the static and dynamic behaviour of the UAV structure.