Passive Aeroelastic Tailoring

2015

This work explores the use of alternative internal structural designs within a full-scale wing box structure for aeroelastic tailoring, with a focus on curvilinear spars, ribs, and stringers. The baseline wing model is a fully-populated, cantilevered wing box structure of the Common Research Model (CRM). Metrics of interest include the wing weight, the onset of dynamic flutter, and the static aeroelastic stresses. Twelve parametric studies alter the number of internal structural members along with their location, orientation, and curvature. Additional evaluation metrics are considered to identify design trends that lead to lighter-weight, aeroelastically stable wing designs. The best designs of the individual studies are compared and discussed, with a focus on weight reduction and flutter resistance. The largest weight reductions were obtained by removing the inner spar, and performance was maintained by shifting stringers forward and/or using curvilinear ribs: 5.6% weight reduction, a 13.9% improvement in flutter speed, but a 3.0% increase in stress levels. Flutter resistance was also maintained using straight-rotated ribs although the design had a 4.2% lower flutter speed than the curved ribs of similar weight and stress levels were higher. For some configurations, the differences between curved and straight ribs were smaller, which provides motivation for future optimization-based studies to fully exploit the trade-offs. Nomenclature ε Distance between two structural members η Fraction that determines control line endpoint locations in a wing section CG Center of gravity CG root Center of gravity at the root region (first 1/8 of wing semi-span) CG tip Center of gravity at the tip region (last 1/8 of wing semi-span) CRM Common Research Model FGM Functionally graded materials/metals i or IBD Inboard wing section KS Kreisselmeier-Steinhauser function LE Leading edge o or OBD Outboard wing section p1 i , p1 o Number of structural members within a wing section p1* Vector defining spanwise structural members as spars or stringers (p2 i , p3 i) , (p2 o , p3 o) Control line parameters in a wing section [p4 i , p5 i , p6 i ], [p4 o , p5 o , p6 o ] Curvature definition parameters in a wing section p7 i , p7 o Rib rotation parameter TE Trailing edge x Direction parallel to the aircraft fuselage centerline y Direction parallel to ground and perpendicular to the aircraft fuselage centerline

2017

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52nd Aerospace Sciences Meeting, 2014

This work explores the use of tow steered composite laminates, functionally graded metals (FGM), thickness distributions, and curvilinear rib/spar/stringer topologies for aeroelastic tailoring. Parameterized models of the Common Research Model (CRM) wing box have been developed for passive aeroelastic tailoring trade studies. Metrics of interest include the wing weight, the onset of dynamic flutter, and the static aeroelastic stresses. Compared to a baseline structure, the lowest aggregate static wing stresses could be obtained with tow steered skins (47% improvement), and many of these designs could reduce weight as well (up to 14%). For these structures, the trade-off between flutter speed and weight is generally strong, although one case showed both a 100% flutter improvement and a 3.5% weight reduction. Material grading showed no benefit in the skins, but moderate flutter speed improvements (with no weight or stress increase) could be obtained by grading the spars (4.8%) or ribs (3.2%), where the best flutter results were obtained by grading both thickness and material. For the topology work, large weight reductions were obtained by removing an inner spar, and performance was maintained by shifting stringers forward and/or using curvilinear ribs: 5.6% weight reduction, a 13.9% improvement in flutter speed, but a 3.0% increase in stress levels. Flutter resistance was also maintained using straightrotated ribs although the design had a 4.2% lower flutter speed than the curved ribs of similar weight and stress levels were higher. These results will guide the development of a future design optimization scheme established to exploit and combine the individual attributes of these technologies. The general concept of aeroelastic tailoring is not new [1]; and a large number of papers have been published on the subject, some of which are referenced in this section. It is found, however, that many of these papers utilize a simplified wing structure: a beam or a flat plate. This is with good reason: a simplified treatment allows trends and trade-offs to be clearly seen and understood. In this work, however, a fully-populated wing box structure (within the Common Research Model (CRM) wing [3]) is used to take advantage of more complex configurations and tailoring schemes that modify the internal structural arrangement and material properties. It is also found in the literature that relatively few papers consider multiple aeroelastic metrics, and those that do, use typically conflicting metrics, so this will be emphasized here as well. The current work is focused on tailoring schemes that have been recently enabled by advances in materials or manufacturing methods. Three are considered, as shown in Figure 1: fiber tow steered composites, functionally graded materials (FGM), and curvilinear ribs/spars/stringers. The first technology is enabled by an automated fiber tow placement machine (where composite fibers are applied along curvilinear paths within the plane of a laminate), the latter two by metal additive manufacturing processes like electron beam freeform fabrication (EBF 3) [4]. None of these methods have been extensively used for aeroelastic tailoring and/or aircraft structural design (largely due to their novelty), but some existing works can be found in the literature, as will be noted below. Many papers that consider the design of tow steered aerospace structures are confined to single panels subject to buckling [5], frequency [6], or strength [7] considerations. Panels with cutouts are of particular interest [7], as tailored load paths can potentially alleviate the stress concentrations around the cutout (window, bay door, etc.). Similar cutout based considerations are applicable to wing design as well, and these strategies may also prove useful at the junctions of spars and skins (and the resulting stress concentration), for example. Expanding from single panels to entire wing structures, few papers exist in the literature. Haddadpour and Zamani [8] modeled a subsonic cantilevered wing as a thin-walled beam and altered the steering paths of the composite skins to maximize the aeroelastic flutter speed. Stodieck et al. [9] modeled the wing as a plate-like structure and demonstrated the relationship between the tow steering pattern and various aeroelastic metrics of interest: flexural axis, divergence, flutter, and gust loads.

1996

Design-oriented analysis has become increasingly important as more and more problems traditionally solved in isolation are being approached from a multidisciplinary point of view. One such problem is the aeroelastic optimization of supersonic transport wings. Whereas simplified analytical techniques may not be sophisticated enough, and complex numerical models may be too cumbersome, this paper puts forward a new approach to achieving a balance between modeling fidelity and required accuracy. Higher fidelity analysis techniques, usually associated with design stages where key geometric variables have been fixed, are used to model a design space consisting of these important geometric variables. This is accomplished through the combined use of a Design of Experiment/Response Surface Method technique and parametric analysis tools (including an automated finite element grid generation procedure). The result is a prediction method for the structural weight of an aeroelastically optimized wing for use in an Integrated Product and Process Development environment, where cost, performance, and manufacturing trades can be accomplished. The technique is to be demonstrated on the aeroelastic design of a wing for a generic High Speed Civil Transport, based on a select set of planform and airfoil design variables. Finally, a framework for evaluating new technologies within the aeroelastic optimization is outlined.

29th Structures, Structural Dynamics and Materials Conference, 1988

ITM Web of Conferences, 2019

Nowadays the composite materials have become the materials of choice to be used in the new aerospace structures that need to be not only larger and larger in size but also to be better performing in terms of aeroelastic responses inherent to thin-walled, slender structures. The advantage of composite materials airframes stems from their low structural weight which determines lower fuel consumption while preserving at the same time the airworthiness of the designed aircraft. But more important than the fuel consumption, the composite materials allow for the optimal tailoring of its layers in terms of specific design objectives. The paper presents such an aeroelastically tailored load carrying wing model which can passively control specific aeroelastic effects. The article focuses on the bend-twist coupling of the structural response to aerodynamic forces and on the parameter estimation/model updating techniques used to characterize the finite element model of the composite wing. Resu...

58th AIAA/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 2017

The aeroelastic performance of a wing, including static aeroelastic shape, flutter/divergence speed and gust load response, has a significant influence on aircraft design. The tailoring of aeroelastic responses therefore offers potential weight savings. In this paper, the spars and stringers planform geometry (i.e. shape and root/tip chord wise location) on a representative wind tunnel model aircraft wing are used to modify the wing aeroelastic performance. Several optimisations are performed to illustrate the ability of the spars and stringers planform geometrical features to change the wing vibrational mode natural frequencies, deformation under a static tip load and aerodynamic load, gust response and aeroelastic instability speed. Changing the stringers planform geometry is shown to offer minor variation in the wing deformation and loads. Changing the spars planform geometry is shown to enable a reduction in root bending moment under static aerodynamic loading greater than 10%, a reduction in maximum root bending moment encounter during a worst case scenario gust event greater than 10% and a 25% increase in flutter speed. The improvements due to a change in the spar planform geometry are compared to the effect of changing the wing sweep angle. A framework to characterise Euler-Bernoulli beam properties on wings with geometric coupling is then developed and validated to relate the stiffness and bend/twist coupling parameter to the full 3D FE models.

51st AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference<BR> 18th AIAA/ASME/AHS Adaptive Structures Conference<BR> 12th, 2010

The study presents a multilevel optimization methodology for the preliminary structural design of transportation aircraft wings. A global level is defined by taking into account the primary wing structural components (i.e., ribs, spars and skin) which are explicitly modeled by shell layered finite elements. Wing substructures such as stringers are implicitly represented by an equivalent formulation of the structural properties. The global level is analyzed and optimized for minimum mass under element stress constraints. Selected wing skin panels are extracted from the global wing and further remodeled with detailed stringers. Boundary conditions are transferred from the finite element (FE) global level solution to the FE detailed stiffened panel models. A finite element analysis is performed on the selected local level panels, which are mass optimized under additional stability constraints, providing a new optimal skin-stringer layout. The global model is then updated with the local level optimum results, and a number of iterative global-local optimization loops are executed. In the current study the DLR in-house tools are used for the structural modeling and sizing of the wing global level, and a new stiffened panel generator is introduced for the local level modeling. A local optimization module which includes instability failure criteria is implemented to redesign the stiffened panels for minimum mass. The global and the local levels communicate through a framework developed to assist an automated and flexible multilevel optimization, and to minimize the time consuming activities required to generate detailed finite element models. The methodology is tested and demonstrated using a transportation DLR aircraft wing geometry as global level, and a variable number of upper skin blade stiffened panels remodeled in detail as local level.

The objective of this paper is to develop an accurate model for optimal design through design the structure of wing that combine the composite (Skins) and isotropic materials (all other structures) and compare this with the same wing made by changing the orientation of composite ply orientation in skin. The optimum design for each wing with different ply orientation can be obtained by comparing stress and displacement. Structural modelling is completed with the help of CATIA V5, each components moddeled separately and assembled using Assembly workbench of CATIAV5, this assembly is then converted to IGS file. Finite element modelling is completed in MSc Patran using the IGS file as geometry, the element type used for meshing was 2D shell elements with QUAD4 element topology and different parts are connected using RBE2 connection. Static analysis done using MSc Nastran. The finite element model obtained is analysed by applying an inertia force of 1g and then aerodynamic result (lift) is used to simulate the wing loading on the wings. Optimum design is found by tabulating stress and displacement for each ply combination Keywords: Composite Wing, Modelling in CATIA V5, Finite element Analysis in Nastran, Optimum ply orientation. І.INTRODUCTION The critical element of aircraft is the design of the wings. Several factors influence the selection of material of which strength allied to lightness is the most important. Composite materials are well known for their excellent combination of high structural stiffness and low weight. Because of higher stiffness-to-weight or strength-to-weight ratios compared to isotropic materials, composite laminates are becoming more popular. Composite structures typically consist of laminates stacked from layers with different fiber orientation angles. The layer thickness is normally fixed, and fiber orientation angles are often limited to a discrete set such as 0°, ±30°, ±45°, ±75°, and 90°. This leads to different combinations of ply orientation and among that one will gives the better results , that is the optimized design for composite structures. A unidirectional laminate is a laminate in which all fibers are oriented in the same direction, cross-ply laminate is a laminate in which the layers of unidirectional lamina are oriented at right angles to each other and quasi-isotropic laminate behaves similarly to an isotropic material; that is, the elastic properties are same in all direction. Unidirectional composite structures are acceptable only for carrying simple loads such as uniaxial tension or pure bending. In structures with complex requirements of loading and stiffness, composite structures including angle plies will be necessary. Since each laminate in the composite material can have distinct fibre orientations which may vary from the adjoining laminates, the optimum ply orientation is also obtained as a result of the parametric study conducted using NASTRAN finite element package by varying the orientation sequence in the composite. II. GEOMETRICAL CONFIGURATIONS The wing design is an iterative process and the selections or calculations are usually repeated several times. A variety of tools and software based on aerodynamics and numerical methods have been developed in the past decades, there by a reduction in the number of iterations is observed. Normally two spar construction is common in transport aircraft wing design. The spar near to the leading edge of the wing is called as front spar and the spar closer to the aft portion of the wing is called as rear spar of the wing. One end of the spar near the root of the wing is connected to the fuselage called root of wing, the other end towards the tip of the wing is a free end. This configuration is very similar to the cantilever beam arrangement in any engineering structure. Spars and Ribs are connected using L angle fittings. Figure 1 below shows the Location of Spar and Ribs from root of wing and Figure 2 shows the complete wing structure modelled in CATIA V5.