The main objective of this paper is to describe a methodology to be applied in the preliminary design of a tiltrotor wing based on previously developed conceptual design methods. The reference vehicle is the Next-Generation Civil Tiltrotor Technology Demonstrator (NGCTR-TD) developed by Leonardo Helicopters within the Clean Sky research program framework. In a previous work by the authors, based on the specific requirements (i.e., dynamics, strength, buckling, functional), the first iteration of design was aimed at finding a wing structure with a minimized structural weight but at the same time strong and stiff enough to comply with sizing loads and aeroelastic stability in the flight envelope. Now, the outcome from the first design loop is used to build a global Finite Element Model (FEM), to be used for a multi-objective optimization performed by using a commercial software environment. In other words, the design strategy, aimed at finding a first optimal solution in terms of the thickness of composite components, is based on a two-level optimization. The first-level optimization is performed with engineering models (non-FEA-based), and the second-level optimization, discussed in this paper, within an FEA environment. The latter is shown to provide satisfactory results in terms of overall wing weight, and a zonal optimization of the composite parts, which is the starting point of an engineered model and a detailed FEM (beyond the scope of the present work), which will also take into account manufacturing, assembly, installation, accessibility and maintenance constraints.

The tiltrotor wing structure is one of the most critical and heavily investigated structures in design due to the fundamental need to consider the interactions between the wing, pylon, and rotor systems to achieve aircraft aeroelastic stability. Indeed, in high-speed forward flight, wing flexural and torsional stiffness have fundamental roles in pitch-whirl stability. Another specific concern of tiltrotors is dynamic mode placement; it is necessary to properly place wing bending modes away from prop-rotor forcing frequencies. The main aeroelastic stability and dynamics requirements and the wing design process flow for the next generation civil tiltrotor are presented in this work. In this context, the use of composite materials plays a fundamental role in the attempt to satisfy the requirements, with the perpetual aim of minimizing the structural weight. An overview of the idealized and adopted models for strength, aeroelasticity, and whirl flutter analysis is provided. The primary focus was on the aeroelastic tailoring process. To satisfy, at the same time, all of the structural dynamic and aeroelastic stability requirements, the best compromise, with an acceptable weight penalty, was the mixture of two methodological solutions: adding unidirectional tape in the zones of the upper and lower skins for flexural out-of-plane frequency and adding a proper number of ±45° fabric layers at the locations of the skin with the highest value of strain energy for in-plane torsional modes. The results show that the proposed method based on modal strain energy analysis enables a tiltrotor aeroelastic tailored wing design. It can be easily employed in similar applications (e.g., vehicle scale-up/down) with the advantage of using the stiffness requirements derived directly from the aeroelastic ones (i.e., structural frequencies). The specific wing achieved aeroelastic clearance by adding only 2.7% of extra mass.

The T-WING project, a CS2-CPW (Clean Sky 2 call for core partner waves) research initiative within FRC IADP (Fast Rotor-Craft Innovative Aircraft Demonstrator Platform), focuses on developing, qualifying and testing the new wing of the Next-Generation Civil Tilt-Rotor (NGCTR). This paper introduces a case study about a methodology for refining the stick-beam model for the NGCTR wing, aligning it with the GFEM (Global Finite Element Model) wing’s dynamic characteristics in terms of modal frequencies and mode shapes. The initial stick-beam model was generated through the static condensation of the GFEM wing. The tuning process was formulated as an optimization problem, adjusting beam properties to minimize the sum of weighted quadratic errors in modal frequencies and Modal Assurance Criterion (MAC) values. Throughout the optimization, the MAC analysis ensured that the target modes were tracked, and, at each iteration, a new set of variable estimates were determined based on the gradient vector and Hessian matrix of the objective function. This methodology effectively fine-tunes the stick-beam model for various mass cases, such as maximum take-off weight (MTOW) and maximum zero-fuel weight (MZFW).

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