Nathalie Toso-Pentecote

Determination of delamination damage in composites under impact loads

This contribution describes recent progress on materials modelling and numerical simulation of fibre reinforced composite shell structures subjected to high velocity impact loads. A Continuum Damage Mechanics (CDM) model for fibre reinforced composites is applied to model both in-ply damage and delamination failure during impact loading. The CDM model has been implemented in a commercial explicit finite element (FE) code in which a laminate is modelled by stacked shell elements with cohesive interfaces which fracture when a delamination failure energy criterion is reached. A comparison of structural response and failure modes from numerical simulations is then conducted. In particular, emphasis is put on the influence of delamination damage on the impact behaviour of composite plates and stiffened panels. Delamination damage is measured in low velocity drop tower impact tests and also in high velocity gas gun tests on composite plates with steel projectiles carried out at the DLR. In both types of test it is seen that at lower impact energies delamination damage was often the main damage mechanism, whereas at higher impact energies there was fibre fracture and plate penetration by the hard projectiles. In this case, the presence of some delamination provides an additional energy absorption mechanism in the composite plate, which can reduce complete penetration. It follows that modelling and analysis of impact damage require composites failure models which include both in-plane ply damage and delamination damage.

Simulation of High Velocity Impact on Composite Structures - Model Implementation and Validation

High velocity impact on composite aircraft structures leads to the formation of flexural waves that can cause severe damage to the structure. Damage and failure can occur within the plies and/or in the resin rich interface layers between adjacent plies. In the present paper a modelling methodology is documented that captures intra- and inter-laminar damage and their interrelations by use of shell element layers representing sub-laminates that are connected with cohesive interface layers to simulate delamination. This approach allows the simulation of large structures while still capturing the governing damage mechanisms and their interactions. The paper describes numerical algorithms for the implementation of a Ladevèze continuum damage model for the ply and methods to derive input parameters for the cohesive zone model. By comparison with experimental results from gas gun impact tests the potential and limitations of the modelling approach are discussed.

Impact Damage in Sandwich Composite Structures From Gas Gun Tests

The paper studies the High Velocity Impact (HVI) response of aircraft structures by means of gas gun impact tests and post-test NDE evaluation. The scope of the activity comprises structural components such as stringer stiffened composite panels and a range of composite sandwich structures, with projectiles such as ice, synthetic birds, runway debris and tyre/rim debris. The tests and simulations are used to support concept design, and certification phases of new aircraft structures based on carbon fibre composites. To perform these tests DLR has installed and commissioned a new gas gun test facility comprising three different barrels having calibres of 200mm, 60mm, and 32/25mm. The guns fire into a steel target chamber where the test structures are mounted. Impact damage to structures is characterised by a number of NDE techniques, including ultrasonic C-scan, lock-in thermography, X-ray and computer tomography (CT).

Numerical modelling of impact and damage tolerance in aerospace composite structures

This chapter discusses numerical methods that are used for predicting impact damage in advanced composite structures. To support the acceptance by industry for design and certification of composite aircraft structures, composites damage models are required for implementation in commercial finite element (FE) codes and validation at specimen and substructure level. DLR experience on modelling damage progression and failure in composite structures under impact is reviewed based on meso-scale composites ply damage models and an energy-based delamination failure criteria in explicit FE codes. The numerical methods are applied to predict impact damage and damage tolerance in stiffened and unstiffened aircraft panels from hard and soft body impacts, including pre-stressed panels subjected to impact loads. Issues concerned with validation of computational methods for certification by analysis are discussed.

Design and performance of novel aircraft structures with folded composite cores

This chapter focuses on novel sandwich structures with open cellular composite cores manufactured by folding thin sheet base materials into a three-dimensional structure. Folded composite cores or ‘foldcore’ manufactured from resin impregnated aramid paper have similar densities and mechanical properties to Nomex honeycomb. Furthermore, they allow ventilation to prevent moisture build-up due to their open cell design and can be manufactured cost-efficiently in a continuous process. In aircraft fuselages, a sandwich design concept could yield significant weight savings compared to an aluminium reference fuselage, through increasing frame spacing and elimination of stringers. The foldcore concept is described with typical base sheet materials, core geometries, and fabrication technology suitable for continuous manufacturing processes. Folded core properties and design are discussed with test methods used for measuring thin base sheet properties and determining core through-thickness compression and shear failure modes and strength properties. Core design is based on micromechanics cell models used in FE methods for simulating progressive damage and collapse mechanisms to provide core properties for use in sandwich structural analysis and design of sandwich structures. Finally, the damage tolerance of sandwich panels with carbon fibre/ epoxy skins and aramid fibre/phenolic folded core is discussed, based on experimental investigations by drop tower and gas gun impact tests. Comparison between observed and computed failure behaviour for a range of impact load cases shows good agreement, indicating that the FE methods could provide the basis for design and certification of these advanced aircraft sandwich structures.