Purpose -This study aims to develop an innovative actuator for improving the performance of future aircraft, by adapting the airfoil shape according to the flight conditions. The flap's camber of a civil regional transportation aircraft's trailing edge actuated and morphed with the use of shape memory alloys (SMA) actuator technology, instead of the conventional split flap mechanism is studied. Design/methodology/approach -For the flap's members sizing an efficient methodology is utilised based on finite element (FE) stress analysis combined to analytically formulated design criteria. A mechanical simulation within an FE approach simulated the performance of the moving rib, integrating both aerodynamic loads and SMA phenomenology, implementing Lagouda's constitutive model. Aim of this numerical simulation is to provide guidelines for further development of the flap. A three-dimensional assembly of the flap is constructed to produce manufacturing drawing and to ensure that during its morphing no interference between the members occurrs. Eventually, the manufactured flap is integrated on a test rig and the experimental characterisations under no and static loads, and dynamic excitation are performed. Findings -Experimental results showed that the rib's SMA mechanism can adequate function under load providing satisfactory morphing capabilities. Originality/value -The investigated approach is an internal into the flap mechanism based on the shape memory effect of thin wires. In the developed mechanism, SMA wires are attached to the wing structure, where they function as actuating elements.

When catastrophic failure phenomena in aircraft structures, such as debonding, are numerically analyzed during their design process in the frame of “Damage Tolerance” philosophy, extreme requirements in terms of time and computational resources arise. Here, a decrease in these requirements is achieved by developing a numerical model that efficiently treats the debonding phenomena that occur due to the buckling behavior of composite stiffened panels under compressive loads. The Finite Element (FE) models developed in the ANSYS© software (Canonsburg, PA, USA) are calibrated and validated by using published experimental and numerical results of single-stringer compression specimens (SSCS). Different model features, such as the type of the element used (solid and solid shell) and Cohesive Zone Modeling (CZM) parameters are examined for their impact on the efficiency of the model regarding the accuracy versus computational cost. It is proved that a significant reduction in computational time is achieved, and the accuracy is not compromised when the proposed FE model is adopted. The outcome of the present work leads to guidelines for the development of FE models of stiffened panels, accurately predicting the buckling and post-buckling behavior leading to debonding phenomena, with minimized computational and time cost. The methodology is proved to be a tool for the generation of a universal parametric numerical model for the analysis of debonding phenomena of any stiffened panel configuration by modifying the corresponding geometric, material and damage properties.

Bolted joints are widely used in composite aircraft structures, for their assembly. The appropriate bolted joint configuration (hole/bolt diameter, pitch, etc.) is carefully selected during the detail design phase, where high fidelity numerical models are required with substantial computational cost and time. This work presents a design criterion, which allows the selection of the bolted joint configuration during the preliminary design phase with less computational time. The developed design criterion is based on a fully parametric finite element (FE) model, built in ANSYS V19 (Canonsburg, PA, USA), of a bolted joint with progressive damage modelling (PDM) capabilities, so that the failure of the joint can be predicted. From the numerical analyses, the bearing load and the load that bypasses the hole are calculated, up to failure, for a variety of joint configurations and loading conditions. The results of each analysis are used for plotting the failure envelope for the investigated bolted-joint configuration. Consequently, a design criterion is generated for the bolted joint. The availability of these failure envelopes, as design criterion, permit the appropriate selection of the bolted-joint configuration in an earlier design phase saving valuable time and computational cost.

The focus of this paper is the development of a digital escape room for pilots’ training in flight safety procedures. To this end, we will discuss the methodology we used in order to make sure that each stage of the evolution of our digital platform is safe and suitable for educational use. Therefore, we will analyse the first stages towards the construction and evaluation of the scenarios incorporated in our entirely digital escape room that is intended for the T-6A Texan II pilots’ education. Because of the educational character of the digital escape room, the theoretical background of our research is extremely important, since it provides the escape room with the educational aspect. As a result, for the narratives of our simulations and the development of our emergency cases, we used as our baseline the flight manual, the boldface procedures and the operating limitations of T-6A aircraft and we selected the categories of incidents/emergencies to be used as part of our virtual escape room. We, then, constructed the trial-run scenarios that we asked pilots to solve. Finally, we conducted extensive field research with the development and use of appropriate questionnaires and interviews to evaluate the realism of our scenarios. Based on the feedback we received from the field research, we will conclude this paper by discussing the limitations of this study, but also our ideas for future improvements and a timeline on our research progress, as well as on further developing our virtual escape room.

An energy-based solution for calculating the buckling loads of partially anisotropic stiffened plates is presented, such as antisymmetric cross-ply and angle-ply laminations. A discrete approach, for the mathematical modelling and formulations of the stiffened plates, is followed. The developed formulations extend the Rayleigh–Ritz method and explore the available anisotropic unstiffened plate buckling solutions to the interesting cases of stiffened plates with some degree of material anisotropy. The examined cases consider simply supported unstiffened and stiffened plates under uniform and linearly varying compressive loading. Additionally, a reference finite element (FE) model is developed to compare the calculated buckling loads and validate the modelling approach for its accuracy. The results of the developed method are also compared with the respective experimental results for the cases where they were available in the literature. Finally, an extended discussion regarding the assumptions and restrictions of the applied Rayleigh–Ritz method is made, so that the limitations of the developed method are identified and documented.