A framework for the bi-level optimization of a generic transport aircraft fuselage using aeroelastic loads

The aeroelastic loads and design processes at the German Aerospace Center, Institute of Aeroelasticity in the framework of multi-disciplinary optimization are constantly evolving. New developments have been made in the in-house model generation tool ModGen, which allow us to create detailed fuselage models for preliminary design. As a part of the subsequent developments to integrate the fuselage structure in our aeroelastic design process, a new framework for optimizing the fuselage structure has been developed. The process is based on a bi-level optimization approach which follows a global-local optimization methodology to simplify a large optimization problem. A sub-structuring procedure is used to define stiffened panels as independent structures for local optimization. The panels are sized with stress and buckling constraints with consideration of several aeroelastic load cases. Furthermore, in this paper, we present a physical sub-structure grouping process which enables reduced number of panel optimizations and saves considerable computational effort with little compromise in the solution accuracy

Aeroelastic Design of a Highly-Flexible Wing Using a Simplified Composite Optimization Approach within cpacs-MONA

Since composite structures have shown a growing utilization in the aircraft structural design, as a first step, a simplified implementation for composite optimization has been implemented into cpacs-MONA. Thereby the material properties of a composite lay-up are converted to a linear anisotropic material definition to be used for two-dimensional elements (MSC Nastran PSHELL/MAT2-combination). This calculation is done by using the generically known ABD-matrix formulation. For this publication cpacs-MONA sets up an aircraft configuration with three different material specifications, resulting in different flexibilities of the load-carrying wing structures: - Model A: Aluminum material definition for the structural wing components (ribs, spars and skin) - Model B: Composite material definitions for the wing components (reference model) - Model C: Model B with increased strain allowables (highly-flexible wing) Since cpacs-MONA comprises of an iterative loads analysis and structural optimization loop, the question on how the different wing flexibilities effect the loads, the structural mass, the dynamic behavior and the resulting aeroelastic stability of the complete aircraft is addressed within this paper

A Validation Study for the Computation of Nonlinear Modal Frequency using a Hamiltonian Reduced Order Model

A common structural design verification is to conduct modal analysis and to ensure that the vibration modes of the structure do not get harmonically excited during operation. Since the modal frequencies are dependent on the mass and stiffness properties of the structure, they are an influencing factor in the design optimization. Modal frequencies are obtainable using eigenvalue analysis after linearization approximations in the restoring force terms of the governing equation of motion. While this approximation is valid for lower loads, it is not acceptable for strong nonlinear vibrations. A measurable shift in the modal frequency is observable when the amplitude of vibrations is in the order of the thickness of the vibrating structure. This phenomenon is more pronounced in thin walled and flexible structures where the vibration amplitudes can exceed the linear threshold relatively more easily. A credible approach in computation of nonlinear modes is the utilization of parametric continuation schemes for generating nonlinear frequency-amplitude response. However, utilization of this scheme with a large degree of freedom model is a computationally intensive approach which necessitate the development of reduced order models. A model reduction method in the finite element framework, termed as Hamiltonian Reduced Order Model (ROM), has been recently developed at Delft University of Technology. The present study is conducted for the experimental validation of the ROM. Governing equations of motion of a stiffened plate have been derived using the Hamiltonian ROM and the nonlinear frequency response has been generated using a continuation scheme. For validation, experiments have been conducted using the Laser Doppler Vibrometer to obtain similar frequency response curves. A comparison of the numerical and the experimental results shows excellent agreement. Furthermore, accurate numerical response is obtainable using only a single degree of freedom in the reduced order model which proves the effectiveness of the Hamiltonian ROM. The results also demonstrate the necessity of a nonlinear damping model for obtaining comparable results.Aerospace Engineerin

A Coupled Global-Local Optimization Process for Fuselage Structures

The aeroelastic loads and design processes at DLR-AE in the framework of multi-disciplinary optimization are constantly evolving. New developments have been made in the in-house model generation tool ModGen which allow us to create fuselage models for preliminary design. As a part of the subsequent developments to integrate the fuselage structure in our aeroelastic design process, a new framework for optimizing the fuselage structure has been developed. The process is based on a multi-level optimization approach which follows a coupled global-local optimization methodology to simplify a very large optimization problem. In this paper, we present a physical sub-structure grouping process which enables reduced number of panel optimizations and saves considerable computational effort with little compromise in the solution accuracy

Model validation of an aeroelastically-tailored forward swept wing using fibre-optical strain measurements

The Institute of Aeroelasticity of the German Aerospace Center (DLR) in Göttingen has analysed means to passively minimise gust load encounters within the DLR project ALLEGRA. One goal was to develop and test a highly flexible forward-swept wind tunnel model which is aeroelastically optimised for maximum tip deflection while constraining the tip twist. Stiffness optimisation tools offer the possibility to design wings with very specific aeroelastic constraints. From the optimised stiffness distribution a feasible stacking sequence needs to be derived to reach the full potential of the stiffness optimisation process. After manufacturing, the wind tunnel model needs to be validated thoroughly. Here, a two-step validation process will be shown. First, static tests obtaining deflection and strain measurements are utilised to adjust the finite element model by a computational model updating (CMU) approach [1-3]. Second, experimental eigenfrequencies and strain shapes are used to check the validity of the updated model

Koiter-Newton Based Model Reduction for Large Deflection Analysis of Wing Structures

Wing structures subjected to large deflections are prone to nonlinear load-deflection behavior. Geometric nonlinearities can arise due to the accompanying large rotations and in-plane deflections that manifest in the form of stiffening effects in the nonlinear static response. To account for these nonlinearities, reduced-order modeling techniques in combination with nonlinear finite element formulations have been previously proposed. However, these methods often have a limited range of validity due to linear eigenmode-based formulations with assumptions of small rotations. In this paper, a large deflection analysis framework based on the Koiter-Newton model reduction technique is presented. It is demonstrated that the reduced model in its basic form is ineffective for large deflection analysis. To resolve this, an incremental updating procedure is used for the reduced-order model that incorporates the necessary nonlinear effects. The model updating enables the computation of nonlinear response for a large range of deflection

A momentum subspace method for the model-order reduction in nonlinear structural dynamics: Theory and experiments

The article proposes a method developed for model order reduction in a Finite Element (FE) framework that is capable of computing higher order stiffness tensors. In the method, a truncated third order asymptotic expansion is used for transformation of an FE model to a reduced system. The basis matrix in the formulation of the reduced-order model (ROM) is derived from linear mode shapes of the structure. The governing equations are derived using Hamilton's principle and the method is applied to geometrically nonlinear vibration problems to test its effectiveness. An initial validation of the numerical formulation is obtained by comparison of results from time domain nonlinear vibration analyses of a rectangular plate using Abaqus. Subsequently, a stiffened plate is modeled to test a more complex structure and a continuation algorithm is used in combination with the ROM to compute nonlinear frequency response curves. The validation of the stiffened plate has been performed through comparisons with the results of nonlinear vibration experiments. The experiments are conducted with Polytec Laser Doppler Vibrometer and PAK MK-II measurement systems for large amplitude vibrations to validate the nonlinear vibration analyses.Aerospace Structures & Computational MechanicsDynamics of Micro and Nano System