Browse Topic: Thermoplastics
With performance advances proposed for the Future Vertical Lift suite of aircraft and advancements in the electronic battlefield, it is imperative that advanced materials and concepts be included in the vehicle designs to meet the aggressive weight reduction objectives, structural requirements, and operational environment capabilities. Integrating electromagnetic (EM) shielding during the design process offers an opportunity to make progress towards the performance goals. To this end, efforts must be made to minimize the impact of this shielding to platform weight and structural performance. This article presents work to develop a hybrid multifunctional composite material technology that incorporates copper mesh into a carbon fiber and thermoplastic matrix structural composite material to achieve required levels of EM shielding and high levels of structural efficiency while reducing the overall weight of the system. This article focuses on the design of a representative helicopter tailcone as means of illustrating expected improvements in weight and manufacturing cost. Baselines that include structure fabricated from aluminum and thermoset composite are established and predicted improvements quantified.
Carbon/epoxy stiffened panels are being increasingly used in transport rotorcraft. The reduced mass density and high stiffness of carbon/epoxy composites can lead to higher levels of vibration relative to comparable metallic structures, which themselves can have vibrations and interior noise high enough to damage the hearing of crew and passengers. The current investigation explores a method to reduce the vibration of carbon/epoxy stiffened panels by introducing thickness tapers known as acoustic black holes (ABHs). The ABH feature is integrated into either the stiffeners or plate of a representative stiffened panel configuration. A finite element (FE) parametric study was used to guide designs that reduce the vibration of the panel without compromising the compressive buckling capability or mass of the panel. FE studies showed that a 30 ply to 12 ply thickness taper longitudinally oriented in the blade stiffener can reduce vibrations and increase compressive buckling capability. Carbon/epoxy panels were manufactured using a low-cost out-of-autoclave material with simple molding. Experimental testing concluded that integrating the ABH into the stiffeners longitudinally helped to reduce the broadband vibration by 5 dB and increase the buckling load (+4.3%) and collapse load (+16.5%) without increasing the mass greatly compared to a traditional baseline design.
This work proposes an experimental and numerical activity aimed at developing methods to evaluate the strength and toughness of Kevlar/Epoxy composite fastened joints used in aeronautical structures and exposed to high energy impacts. Experiments were conducted using an Arcan rig that allowed applying various loading conditions, ranging from pull-through to bearing. A non-linear model of the material based on a bi-phasic decomposition and hybrid meshing technique was built and calibrated. The material model was used to develop a high-fidelity model of the junction to simulate the pull-through test with the Abaqus/Explicit finite element solver. The results of the analysis point out that the implemented progressive damage laws are capable of achieving an appreciable experimental-numerical correlation, both from the qualitative and the quantitative standpoint. Therefore, the combined experimental-numerical approach is promising for developing a validated numerical tool capable of predicting the overall response of different junctions with minimal experimental effort, so to provide data for simulating impact scenarios at the rotorcraft structural scale and designing critical joints.
Carbon fiber reinforced epoxy composite stiffened panels are increasingly being used for structural components in large transport rotorcraft. However, problems are arising with high levels of vibration and interior noise due to the increased stiffness-to-density ratio of composites. The current investigation explores the potential of reducing vibrations in carbon/epoxy stiffened panels with the integration of acoustic black holes (ABH), namely features that incorporate a power law thickness taper. The proposed approach involves designing a taper into the thickness of the blade stiffeners as well as the thin plate. Integration of ABHs into the fuselage structure has the potential to reduce broadband vibrations. Multiple parametric studies with either an ABH integrated into the blade stiffener or a grid of ABHs integrated into the plate were conducted, and the tradeoffs between vibration amplitudes, panel mass, and compressive buckling load were examined. Carbon/epoxy panels were fabricated using vacuum-bag-oven processing with out-of-autoclave prepreg and verified to be of good quality. The integrated velocity response, a proxy for the radiated noise from a panel, and compressive buckling were simulated using finite elements. Comparisons were made to experimentally measured data from modal testing and compression buckling testing. Experimental results indicated that when an ABH is integrated into the blade stiffener and 15 ABHs are integrated into the plate in a grid configuration, the panel mass was unchanged, the integrated velocity response decreased by 2.82 dB, and the buckling load increased by 2.9% compared to a baseline non-tapered design.
Thermoplastic composites are serious competitor for classic epoxy composites. They have comparable properties to epoxy composites, but characterize much lower processing costs. There are several methods of manufacturing the components from thermoplastic composites. One of the most interesting method in terms of efficiency is thermoforming on a press. This technology allows to product of the aircraft parts such as: ribs, brackets, covers, stiffeners. Thermoplastic composites are resistant to most solvents such as grease, oil and aviation fuel. They are also non-flammable and heat-resistant. This all makes them suitable for use in aircraft as upholstery, casing or elements around the tank. PZL Mielec has been developing press thermoforming technologies since 2016 and is the owner of the several patents in this area.
Rotorcraft components, which are often made with reinforced fiber composites, are subjected to severe fatigue loadings due to increased performance demands. Therefore, considerable research interest exists in improving fatigue life of conventional fiber reinforced composites. Nanocomposites are a new class of materials which seek to improve mechanical performance of materials by creating nanoscale crack-nanofiller interactions. In this study we demonstrate the fatigue life improvement of conventional composites by addition of SiO2 nanofillers. The epoxy resin was initially modified with nanofillers to test the static fracture toughness. Once the improvement in static facture toughness was confirmed, three phase modified fiber reinforced composites were made using the modified resin. Cyclic tests were performed at various stress level which demonstrate that three phase nanocomposites perform better than conventional fiber reinforced composites. Fractographic analysis suggests that nanofiller de-bonding from the matrix as well as crack deflection around nanofiller clusters contributes to the improved fracture toughness and fatigue life.
ABSTRACT The ability to construct a composite, semimonocoque, damage-resistant, cargo floor for a rotary wing application using an IM7 graphite/polyetheretherketone (PEEK) composite with in-situ tape-placement fabrication technology has been demonstrated. Through an evolutionary process, a damage-tolerant thermoplastic composite cargo floor was designed according to realistic requirements, and subelement representative structures were developed to verify the design viability and approach. The fabricated and tested structural composite floor subelements demonstrated the feasibility of the technology, illustrated the ability to customize the design to meet unique cargo floor properties (e.g., cargo-loading features), and validated the maturity of the approach and fabrication technology for rotary-wing applications.
ABSTRACT A model defined at the ply scale to predict the failure of laminated composites for static or fatigue loading is proposed. The model describes the loss of strength in the fiber direction for a significant level of transverse damage. This meso-scale model has been characterized on woven ply laminates used for rotorcraft dynamic components, such as glass/epoxy of Starflex®, carbon fiber/epoxy, and carbon fiber/PEEK of H160 main rotor hub. Failure behavior prediction at coupon level has been validated regarding static and fatigue failure mode in tension for epoxy resin woven ply laminates. Characterizations have been also provided for PEEK resin in balanced woven ply laminate, regarding static or fatigue failure mode. Those activities are crucial to increase the level of confidence in failure model, to rely on virtual testing at coupons level, and to better predict damage and failure at component level. This work intends to support the building block approach during development and certification of such critical applications for rotorcraft.
ABSTRACT Fiber reinforced polymer composites can save weight in rotorcraft structures, but have not been widely used in driveshafts due in part to their limited impact tolerance. The objective of the current investigation is to evaluate the effects of incorporating variable amounts of nanosilica (NS) particles in the matrix on the ballistic impact tolerance of carbon/epoxy tubes loaded in torsion. Tubes manufactured with these matrix materials were ballistically impacted using a round steel projectile aimed at normal incidence across the major diameter. After impact, the tubes were nondestructively inspected and subjected to mechanical tests to determine the axial and shear stiffness and the residual shear strength in torsion. In the best material formulations, which were 15 and 25 weight percent NS in epoxy, the use of NS decreased the impact damage area by 50%, increased the residual shear strength by 38%, and increased the energy absorbed per unit damage area by 120% versus the control material with no NS. Overall, the addition of NS significantly improved the impact tolerance of carbon/epoxy tubes loaded in torsion, with little change in the mass density, glass transition temperature, and elastic modulus.
ABSTRACT This paper shows the history and background of the motivation for the recently modified rules concerning the evaluation of fatigue and damage tolerance of metallic and composite structures. The most significant modifications which were made compared to the previous ones are discussed, specifically the split into two paragraphs for metallic and composite structure and the change towards performance based regulations. Examples to demonstrate the application on Airbus Helicopters products are the H145 Fenestron™ hub and blade and the H160 main rotor hub plate. For the metallic parts of the Fenestron™ hub a substantiation according to the flaw tolerant fatigue evaluation requirements of the FAR 29.571 was chosen. The thermoset carbon composite blade of the H145 Fenestron™ as the second example illustrates the application of the newly introduced composite requirements. Finally the thermoplastic carbon composite main rotor hub plate for the H160 is presented as a new technology in the first case of full applicability of the new rules to complete new helicopter model.
ABSTRACT It is a great challenge to perform an accurate and efficient fatigue life prediction of a bonded composite structure with the presence of geometry and material heterogeneity induced stress concentration. The present fatigue damage characterization of composite structures is still dominated by the use of a phenomenological stress-life (S-N) approach due to the availability of extensive S-N data and lower cost in generation of S-N data from fatigue tests at different applied stress ratios. Because of the inaccurate life prediction using the S-N approach for the structure with stress concentrators, a more rational fracture mechanics approach based on a Paris type crack growth law can be applied to compute the crack growth driving force provided that an initial flaw has to be introduced. In order to simulate both the crack initiation and propagation, a dual spring model is implemented at each nodal point where the static failure is simulated using springs of a cohesive type material model while fatigue crack propagation is calculated using springs of an elastic penalty stiffness coupled with a virtual crack closure technique (VCCT). In order to validate the dual spring model for the fatigue damage prediction, two types of Tee-joints are fabricated and tested by the National Institute for Aviation Research (NIAR) with and without a Teflon insertion. A calibration analysis is performed to determine the fatigue crack growth parameters using Tee-joints with a Teflon insert followed by the blind fatigue prediction of the specimens without a Teflon insert.
ABSTRACT This paper presents the methodology and results for ballistic impact testing of thermoplastic composite materials. Ten different materials are investigated. The impact behavior of Aluminum 6082 is used as a reference to compare the results. The impact tests are performed with a gas cannon. Force - time, displacement - time as well as velocity data are recorded. Analytic suggestions for the calculation of the penetration speed of the materials are compared with measured results. It can be seen that it is possible to calculate the penetration speed within a certain percentage of the measured value. Also, the absorbed energies during the penetration process are compared. The results show that glass fiber composites have a better impact material behavior than carbon fiber-reinforced composites (CFRP). Thermoplastic matrix systems are a cheap option for composites but they do not have a significant better high speed impact and damage behavior than duroplastic resins.
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