Research Topic: Fatigue, Damage Tolerance and Crash Worthiness

The objective of this research project was to investigate candidate damage tolerance test methodologies for sandwich composites and to propose specific methodologies and configurations for standardization. Three candidate test configurations are currently under evaluation. The first methodology utilizes an end-loaded Compression After Impact (CAI) test configuration. Initial evaluations have been performed using sandwich configurations with both carbon/epoxy and glass/epoxy facesheets and Nomex honeycomb core. Research has focused on establishing the required specimen size to prevent interactions between the damage present and the boundary conditions during loading as well as special requirements to promote proper alignment and end loading for a variety of sandwich configurations. Second, a four-point flexure test methodology has been identified for evaluating post-impact performance. The composite sandwich panel may be oriented such that the damaged facesheet is loaded in compression or tension. Similar to the CAI test configuration, research has focused on the required size of the gage section between the inner loading points as well as additional sandwich panel requirements to prevent outer-span core shear as well as loading-point compression failures. The third candidate methodolgy, the “hydromat” test configuration, utilizes a water-filled bladder to apply a distributed load to the surface of the composite sandwich panel while supporting the panel along its edges. Initial evaluation using composite sandwich panels with carbon/epoxy and glass/epoxy facesheets and Nomex honeycomb core has exposed difficulties when using this test configuration for assessing strength reductions associated with damage. All three candidate test methodologies are being examined for their limits of applicability and to establish recommended testing procedures. Expected benefits to aviation include standardized test methodologies for use in assessing the damage tolerance of sandwich composites used in aircraft structures.

The purpose of this research is to determine if Fourier transform infrared (FTIR) spectroscopy can be used to quantify incipient thermal damage (ITD) in aerospace composites. ITD is the early stages of thermal damage that cannot be detected by visual or ultrasonic methods but can lead to significant reductions in matrix-dominated properties of the composite. This project involves the detection of thermal damage with FTIR and performing a simulated repair guided by FTIR. The first task is the development of calibration curves for FTIR data on short-beam strength (SBS) samples with varying levels of ITD. This effort focuses on constructing calibration curves from SBS samples that have varying levels of ITD using an ExoScan FITR configured with diffuse reflectance. The SBS samples are 177 °C cure carbon fiber reinforced epoxy laminates that were carefully heat soaked to achieve uniform ITD in the samples. FTIR spectra and calibration curves relating the FTIR spectra to levels of ITD are reported and compared to previous results. These calibration curves will be used to guide mapping of thermal damage both on the surface and during the scarfing repair process with localized heat damage.

The utilization of active control systems for gust alleviation, load redistribution, flight control, and ride comfort improvements has matured over the last two decades to levels of reliability and safety that allow implementation and certification on both military and civil aircraft. Active Flutter Suppression (AFS) – the reliance on active control systems to stabilize flutter-unstable vehicles throughout their flight envelopes – while studied extensively and demonstrated in theoretical, numerical, wind tunnel tests, and flight tests, is not approved for wide spread implementation, mainly because of the catastrophic nature of vehicle response to the failure of such systems. This paper and talk will present a general introduction to active flutter suppression in the context of flight vehicle aeroelasticity and aeroservoelasticity. First steps taken in the effort to assess the state of the art and the implementation level of readiness of the technology will be described.

The design of aircraft structures made of composite materials is heavily influenced by damage tolerance requirements. The problem of predicting failure in notched laminates has been the subjected of numerous studies. In general, these investigations have focused on the response of laminates to in-plane tension, compression or shear. In spite of the fact that out-of-plane bending, twisting, or shear can be an important load situation, very little research has been devoted to this topic. The overall goal of this research is to develop analysis techniques that are useful for the design of composite aircraft structure subjected to general out-of-plane loading. For this project we will limit ourselves to the out-of-plane bending case and focus on some very basic experiments and modeling efforts involving simple structures (center-notched, unstiffened laminates) under pure bending. In partnership with the Boeing Commercial Airplane Company, we will determine the modes of failure of the laminates and evaluate the capability of some currently existing analysis techniques for predicting these failures. Accomplishing our objective will require both experimental and computational efforts.