Multi-site fatigue damage, hidden cracks in hard-to-reach locations, disbonded joints, erosion, impact, and corrosion are among the major flaws encountered in today's extensive fleet of aging aircraft and space vehicles. The use of in-situ sensors for real-time health monitoring of aircraft structures are a viable option to overcome inspection impediments stemming from accessibility limitations, complex geometries, and the location and depth of hidden damage. Reliable, structural health monitoring systems can automatically process data, assess structural condition, and signal the need for human intervention. Prevention of unexpected flaw growth and structural failure can be improved if on-board health monitoring systems are used to continuously assess structural integrity. Such systems are able to detect incipient damage before catastrophic failures occurs. Condition-based maintenance practices could be substituted for the current time-based maintenance approach. Other advantages of on-board distributed sensor systems are that they can eliminate costly, and potentially damaging, disassembly, improve sensitivity by producing optimum placement of sensors and decrease maintenance costs by eliminating more time- consuming manual inspections. This report presents a Sandia Labs-aviation industry effort to move SHM into routine use for aircraft maintenance. This program addressed formal SHM technology validation and certification issues so that the full spectrum of concerns, including design, deployment, performance and certification were appropriately considered. The Airworthiness Assurance NDI Validation Center (AANC) at Sandia Labs, in conjunction with Boeing, Delta Air Lines, Structural Monitoring Systems Ltd., Anodyne Electronics Manufacturing Corp. and the Federal Aviation Administration (FAA) carried out a certification program to formally introduce Comparative Vacuum Monitoring (CVM) as a structural health monitoring solution to a specific aircraft wing box application. Validation tasks were designed to address the SHM equipment, the health monitoring task, the resolution required, the sensor interrogation procedures, the conditions under which the monitoring will occur, the potential inspector population, adoption of CVM into an airline maintenance program and the document revisions necessary to allow for routine use of CVM as an alternate means of performing periodic structural inspects. To carry out the validation process, knowledge of aircraft maintenance practices was coupled with an unbiased, independent evaluation. Sandia Labs designed, implemented, and analyzed the results from a focused and statistically-relevant experimental effort to quantify the reliability of the CVM system applied to the Boeing 737 Wing Box fitting application. All factors that affect SHM sensitivity were included in this program: flaw size, shape, orientation and location relative to the sensors, as well as operational and environmental variables. Statistical methods were applied to performance data to derive Probability of Detection (POD) values for CVM sensors in a manner that agrees with current nondestructive inspection (NDI) validation requirements and also is acceptable to both the aviation industry and regulatory bodies. This report presents the use of several different statistical methods, some of them adapted from NDI performance assessments and some proposed to address the unique nature of damage detection via SHM systems, and discusses how they can converge to produce a confident quantification of SHM performance An important element in developing SHM validation processes is a clear understanding of the regulatory measures needed to adopt SHM solutions along with the knowledge of the structural and maintenance characteristics that may impact the operational performance of an SHM system. This report describes the major elements of an SHM validation approach and differentiates the SHM elements from those found in NDI validation. The activities conducted in this program demonstrated the feasibility of routine SHM usage in general and CVM in particular for the application selected. They also helped establish an optimum OEM-airline-regulator process and determined how to safely adopt SHM solutions. This formal SHM validation will allow aircraft manufacturers and airlines to confidently make informed decisions about the proper utilization of CVM technology. It will also streamline the regulatory actions and formal certification measures needed to assure the safe application of SHM solutions.

Reliable structural health monitoring (SHM) systems can automatically process data, assess structural condition and signal the need for human intervention. There is a significant need for formal SHM technology validation and quantitative performance assessment processes to uniformly and comprehensively support the evolution and adoption of SHM systems. In recent years, the SHM community has made significant advances in its efforts to evolve statistical methods for analyzing data from in-situ sensors. Several statistical approaches have been demonstrated using real data from multiple SHM technologies to produce Probability of Detection (POD) performance measures. Furthermore, limited comparisons of these methods - utilizing different simplification assumptions and data types - have shown them to produce similar POD values. Given these encouraging results, it is important to understand the circumstances under which the data was acquired. Thus far, the statistical analyses have assumed the viability of the data outright and focused on the performance quantification process once acceptable data has been compiled. This paper will address the array of parameters that must be considered when conducting tests to acquire representative SHM data. For some SHM applications, it may not be possible to simulate all environments in one single test. All relevant parameters must be identified and considered by properly merging results from multiple tests. Laboratory tests, for example, may have separate fatigue and environmental response components. Flight tests, which will likely not include statistically-relevant damage detection opportunities, will still play an important role in assessing overall SHM system performance under an aircraft operator’s control. One statistical method, the One-Sided Tolerance Interval (OSTI) approach, will be discussed along with the test methods used to acquire the data. Finally, prospects for streamlining the deployment of SHM solutions will be considered by comparing SHM data needs during what is now an introductory phase of SHM usage with future data needs after a substantial database of SHM data and usage history has been compiled.

There is much interest in the potential to use Structural Health Monitoring (SHM) technology to augment traditional Nondestructive Evaluation (NDE) methods to improve safety, increase asset availability, and reduce maintenance and inspection costs. SHM has the potential to be used in many areas of application including critical components in aircraft and pipelines. Probability of detection (POD) plays a critical role in aircraft structural integrity programs. As such, there has been a high interest in developing methods that can be used to assess POD in SHM applications. In contrast to traditional NDE laboratory experiments to assess POD that involve a set of specimens with cracks, SHM sensors are fixed and SHM data are acquired over time as cracks grow or otherwise evolve. Traditional statistical methods for assessing POD (e.g., as described in MIL-HDBK 1823A 2009) no longer apply to such repeated-measures data. This purpose of this paper is to review the basic statistical concepts of probability of detection (POD) and to show how these concepts can and should be applied to SHM POD studies by modifying and extending existing methods for estimating POD. The methods presented here are applicable when there is a scalar damage index or other response that will be used to make a detect decision. The paper compares a simple model based on length at detection and a random effects model to describe repeated measures data.

A typical aircraft can experience over 2,000 fatigue cycles (cabin pressurizations) and even greater flight hours in a single year. An unavoidable by-product of aircraft use is that crack, impact, and corrosion flaws develop throughout the aircraft's skin and substructure elements. Economic barriers to the purchase of new aircraft have placed even greater demands on efficient and safe repair methods. The use of bonded composite doublers offers the airframe manufacturers and aircraft maintenance facilities a cost effective method to safely extend the lives of their aircraft. Instead of riveting multiple steel or aluminum plates to facilitate an aircraft repair, it is now possible to bond a single Boron-Epoxy composite doubler to the damaged structure. The FAA's Airworthiness Assurance Center at Sandia National Labs (AANC), Boeing, and Federal Express completed a pilot program to validate and introduce composite doubler repair technology to the U.S. commercial aircraft industry. This project focused on repair of DC-10 fuselage structure and its primary goal was to demonstrate routine use of this repair technology using niche applications that streamline the design-to-installation process. As composite doubler repairs gradually appear in the commercial aircraft arena, successful flight operation data is being accumulated. These commercial aircraft repairs are not only demonstrating the engineering and economic advantages of composite doubler technology but they are also establishing the ability of commercial maintenance depots to safely adopt this repair technique. This report presents the array of engineering activities that were completed in order to make this technology available for widespread commercial aircraft use. Focused laboratory testing was conducted to compliment the field data and to address specific issues regarding damage tolerance and flaw growth in composite doubler repairs. Fatigue and strength tests were performed on a simulated wing repair using a substandard design and a flawed installation. In addition, the new Sol-Gel surface preparation technique was evaluated. Fatigue coupon tests produced Sol-Gel results that could be compared with a large performance database from conventional, riveted repairs. It was demonstrated that not only can composite doublers perform well in severe off-design conditions (low doubler stiffness and presence of defects in doubler installation) but that the Sol-Gel surface preparation technique is easier and quicker to carry out while still producing optimum bonding properties. Nondestructive inspection (NDI) methods were developed so that the potential for disbond and delamination growth could be monitored and crack growth mitigation could be quantified. The NDI methods were validated using full-scale test articles and the FedEx aircraft installations. It was demonstrated that specialized NDI techniques can detect flaws in composite doubler installations before they reach critical size. Probability of Detection studies were integrated into the FedEx training in order to quantify the ability of aircraft maintenance depots to properly monitor these repairs. In addition, Boeing Structural Repair and Nondestructive Testing Manuals were modified to include composite doubler repair and inspection procedures. This report presents the results from the FedEx Pilot Program that involved installation and surveillance of numerous repairs on operating aircraft. Results from critical NDI evaluations are reported in light of damage tolerance assessments for bonded composite doublers. This work has produced significant interest from airlines and aircraft manufacturers. The successful Pilot Program produced flight performance history to establish the durability of bonded composite patches as a permanent repair on commercial aircraft structures. This report discusses both the laboratory data and Pilot Program results from repair installations on operating aircraft to introduce composite doubler repairs into mainstream commercial aircraft use.

Multi-site fatigue damage, hidden cracks in hard-to-reach locations, disbonded joints, erosion, impact, and corrosion are among the major flaws encountered in today’s extensive fleet of aging aircraft. The use of in-situ sensors for real-time health monitoring of aircraft structures, coupled with remote interrogation, provides a viable option to overcome inspection impediments stemming from accessibility limitations, complex geometries, and the location and depth of hidden damage. Reliable, Structural Health Monitoring (SHM) systems can automatically process data, assess structural condition, and signal the need for human intervention. Prevention of unexpected flaw growth and structural failure can be improved if on-board health monitoring systems are used to continuously assess structural integrity. Such systems can detect incipient damage before catastrophic failures occurs. Other advantages of on-board distributed sensor systems are that they can eliminate costly and potentially damaging disassembly, improve sensitivity by producing optimum placement of sensors and decrease maintenance costs by eliminating more time-consuming manual inspections.

Structural Health Monitoring (SHM) has the potential to revolutionize aircraft maintenance. Less complex and less time-consuming procedures – when compared to current non-destructive inspection (NDI) technologies – allowed by SHM can reduce the amount of time and burden of the inspection tasks and, consequently, reduce costs. Reliable SHM systems will be able to automatically assess structural condition and inform maintenance personnel the presence of a structural flaw. In the long term, when fully automated and integrated, SHM will be able to inform the need for an intervention and to provide targeted information for planning maintenance actions and logistics. This will also allow the improvement of structural performance, and may eliminate certain scheduled inspection tasks. Over the years Embraer has investigated different SHM technologies, and two of them were selected for a more in-depth investigation. After demonstrating strong results on ground tests and in an Embraer- 190 flight test aircraft, Comparative Vacuum Monitoring (CVM) and Lamb Waves (LW) technologies were included in an effort to move SHM, in particular the concept of Scheduled Structural Health Monitoring (S-SHM), into routine use for aircraft maintenance procedures. Embraer teamed with the Airworthiness Assurance NDI Validation Center (AANC) at Sandia Labs and Azul Airlines in a joint effort with Agência Nacional de Aviação Civil (ANAC, the Brazilian civil aviation regulator) for the qualification of SHM technologies and to validate the performance of such systems in real-life operational environment. The work aimed to develop and carry out a qualification process for SHM damage detection systems, which includes laboratory tests for the assessment of detection capabilities in terms of Probability of Detection (POD) and tests with systems installed on a number of aircraft from Azul Airlines. Demonstration of equipment configuration, sensor interrogation procedures, systems’ resolutions, conditions during the execution of inspections, and other aspects, were addressed. A formal process for SHM implementation was reached with the close consultation of the regulatory agency, and the feasibility of the use of SHM systems was also demonstrated to support the validation.

The use of in-situ sensors for real-time health monitoring of a wide array of civil structures can be a viable option to overcome inspection impediments stemming from accessibility limitations, complex geometries, and the location and depth of hidden damage. The maturity of Structural Health Monitoring (SHM) sensors has evolved to the point where many networks have demonstrated sensitivities that meet or exceed current damage detection requirements. As a result, there is a growing need for well-defined methods to statistically quantify the performance of sensors and sensor networks. Statistical methods can be applied to laboratory and flight test data to derive Probability of Detection (POD) values for SHM sensors in a fashion that agrees with current nondestructive inspection (NDI) validation requirements. However, while there are many agreed-upon procedures for quantifying the performance of NDI techniques, there are no guidelines for assessing SHM systems. While the intended function of the SHM and NDI systems may be very similar, there are distinct differences in the parameters that affect their performance and differences in their implementation that require special consideration. Factors that affect SHM sensitivity include flaw size, shape, orientation and location relative to the sensors, operational and environmental variables and issues related to the presence of multiple flaws within a sensor network. The FAA Airworthiness Assurance NDI Validation Center (AANC) at Sandia Labs, in conjunction with the FAA WJH Technical Center, has conducted a series of SHM validation and certification programs aimed at establishing the overall viability of SHM systems and producing appropriate precedents and guidelines for the safe adoption of SHM solutions for aircraft maintenance. This paper will present the use of several different statistical methods, some of them adapted from NDI performance assessments and some proposed to address the unique nature of damage detection via SHM systems, and discuss how they can converge to produce a confident quantification of SHM performance. Comparisons of hit-miss, a versus ȃ, and One Sided Tolerance Intervals will provide valuable insights into how the characteristics of the collected SHM data affect the formulation of that system’s POD curve. Similarities between NDI and SHM assessments will be highlighted in order to provide a foundation in traditional flaw detection performance measures. In addition, considerations of the controlling factors to be considered when collecting SHM response data will be discussed

The Federal Aviation Administration (FAA) started a research program in structural health monitoring (SHM) in 2011. The program’s goal was to understand the technical gaps of implementing SHM on commercial aircraft and the potential effects on FAA regulations and guidance. The program evolved into a demonstration program consisting of a team from Sandia National Labs Airworthiness Assurance NDI Center (AANC), the Boeing Corporation, Delta Air Lines, Structural Monitoring Systems (SMS), Anodyne Electronics Manufacturing Corp (AEM) and the FAA. This paper will discuss the program from the selection of the inspection problem, the SHM system (Comparative Vacuum Monitoring-CVM) that was selected as the inspection solution and the testing completed to provide sufficient data to gain the first approved use of an SHM system for routine maintenance on commercial US aircraft.

A self-healing, toughened epoxy adhesive is demonstrated based on a commercial structural adhesive film. Self-healing is achieved via embedded microcapsules containing dicyclopentadiene monomer and solid particles of bis(tricyclohexylphosphine)-benzylidine ruthenium (IV) dichloride (Grubbs') catalyst. Recovery of fracture toughness is assessed through fracture testing of width tapered double cantilever beam (WTDCB) specimens. Healing efficiencies as high as 58% were achieved for 6.6wt% DCPD microcapsules and 10mg Grubbs’ catalyst. However, virgin fracture toughness is reduced with the addition of ca. 117μm diameter microcapsules as a result of suppression of the damage zone as revealed by transmission optical micrographs. The uniform dispersal of microcapsules throughout a rubber toughened epoxy adhesive formulated using EPON 828, piperidine and CTBN alleviated the suppression effect and demonstrated retention of virgin fracture toughness of adhesives.

Transverse impact from hail ice can create internal damage to composite structures that is not visually detectable and is therefore a damage tolerance concern. This paper focuses on the experimental characterization of the damage resistance of laminates made from T800/3900-2 carbon/epoxy tape material to impact by high velocity ice spheres, i.e., simulated hail ice (SHI). The failure threshold energy (FTE) defining the onset of damage was found for three panel thicknesses (1.59, 3.11, and 4.66mm), each impacted by three ice diameters (38.1, 50.8, and 61.0mm). Non-destructive investigation techniques were used to detect, map, and characterize the delaminated area. A regression analysis was used to quantitatively determine the FTE of the tape laminates, which was found to closely match previous woven carbon/epoxy FTE data. Both data sets were found to exhibit a linear and common relationship to the ratio of panel thickness to ice diameter (H/D). The resulting delamination patterns of the current and previous panels were found to be similar at damage initiation, but to differ for higher damage states.

To address a critical aspect of the fast growing use of composites in aircraft and aerospace industry, the influence of the porosity on the shear strength of composites property was investigated as a mean for determining the critical values of porosity. Acid digestion techniques were applied to determine the void volume ratio of two families of carbon epoxy laminates (8 and 16 plies). Ultrasonic inspections revealed the corresponding attenuation coefficients. The void morphology was investigated by optical microscopy. Results from Interlaminar shear and Iosipescu shear tests were correlated with the attenuation coefficient to determine critical values of porosity using a modified Mar-Lin fracture criteria. It has been shown that the shear strength decrease with the increase of void volume ratio and the effects are more significant in thicker laminates. This work showed that by using the Mar-Lin criteria the singularity order, which is an indicative of the sensibility to voids in composites, is dependent of type of loading and void distribution.

Current maintenance operations and integrity checks on a wide array of structures require personnel entry into normally-inaccessible or hazardous areas to perform necessary nondestructive inspections. To gain access for these inspections, structure must be disassembled and removed or personnel must be transported to remote locations. The use of in-situ sensors, coupled with remote interrogation, can be employed to overcome a myriad of inspection impediments stemming from accessibility limitations, complex geometries, the location and depth of hidden damage, and the isolated location of the structure. Furthermore, prevention of unexpected flaw growth and structural failure could be improved if on-board health monitoring systems were used to more regularly assess structural integrity. A research program has been completed to develop and validate Comparative Vacuum Monitoring (CVM) Sensors for surface crack detection. Statistical methods using one-sided tolerance intervals were employed to derive Probability of Detection (POD) levels for a wide array of application scenarios. Multi-year field tests were also conducted to study the deployment and long-term operation of CVM sensors on aircraft. This paper presents the quantitative crack detection capabilities of the CVM sensor, its performance in actual flight environments, and the prospects for structural health monitoring applications on aircraft and other civil structures.

Background The number of commercial airframes exceeding twenty years of service continues to grow. An unavoidable by-product of aircraft use is that crack and corrosion flaws develop throughout the aircraft's skin and substructure elements. Economic barriers to the purchase of new aircraft have created an aging aircraft fleet and placed even greater demands on efficient and safe repair methods. Composite doublers, or repair patches, provide an innovative repair technique which can enhance the way aircraft are maintained. Instead of riveting multiple steel or aluminum plates to facilitate an aircraft repair, it is now possible to bond a single Boron-Epoxy composite doubler to the damaged structure. The composite doubler repair process produces both engineering and economic benefits. The FAA's Airworthiness Assurance Center at Sandia National Labs completed a project to introduce composite doubler repair technology to the commercial aircraft industry [I].

Health monitoring of structures is a growing concern of the engineering community. This need is even more intense in the case of aging aerospace and civil structures many of which are operating well beyond their initial design lives. An unavoidable by-product of a metallic structure's use is the development of fatigue cracks that degrade integrity. Significant cost barriers to replace these structures have created an aging infrastructure. These economic realities, coupled with an explosion in infrastructure development and unexpected phenomena that have accelerated decay, have produced even greater demands for in-situ structural health monitoring (SHM). Towards that end, innovative sensor designs and deployment methods must be employed to overcome a myriad of inspection impediments stemming from accessibility limitations, complex geometries, and the location and depth of hidden damage. Prevention of unexpected flaw growth and structural failure could be improved if on-board health monitoring systems are used to continuously assess structural integrity. Eddy current (EC) sensing is an effective, low-cost, low-complexity method to detect flaws in structures. However, existing commercial EC equipment is not able to support extensive, remote, field operation for SHM. This paper presents an EC impedance analysis sensor that has been designed to work with low-profile, planar coils to detect cracks in metallic or other conductive structures. By incorporating the planar coils into flexible circuit boards, it was possible to develop a sensor system that is capable of inspecting a large lateral area with sufficient power to detect surface or deeply embedded cracks. The use of a rectangular planar coil, collocated with miniaturized signal processing, conditioning, and wireless communication electronics provides a stand-alone system to perform high-quality, remote, in-situ, real-time, structural health monitoring. Automated inspections of bridges, ships, railcars, and pipelines can be performed over a period of many months without any direct human interaction. Use of this sensor and data acquisition system can yield significant inspection labor savings without sacrificing fidelity, while providing relevant data to determine a structure's integrity. Furthermore, the remote operation and data logging capabilities allow for more frequent structural interrogation. This sensor overcomes a myriad of typical inspection impediments including "lift-off" from the structure being inspected, the presence of non-conductive or conductive barriers on the surface, and the need to conform to uneven surfaces or other complex geometry. A comprehensive performance assessment program will be presented to demonstrate the potential use of this EC sensor system in a wide array of SHM applications. Test results will clearly quantify its ability to detect small fatigue cracks via in-situ, remote operation and highlight its use in guiding condition-based maintenance activities.

The number of commercial aircraft operating at or beyond their initial design lives continues to grow. Multisite fatigue damage, hidden cracks in hard‐to‐reach locations, disbonded joints, erosion, impact, and corrosion are among the major flaws encountered in today's extensive fleet of aging aircraft and space vehicles. Furthermore, the extreme damage tolerance and high strength‐to‐weight ratio of composites have motivated designers to expand the role of advanced materials in aircraft structures. These developments, coupled with new and unexpected phenomena, have placed greater demands on the application of advanced nondestructive inspection (NDI) and health monitoring techniques. In addition, innovative deployment methods must be employed to overcome a myriad of inspection impediments stemming from accessibility limitations, complex geometries, and the location and depth of hidden damage. The use of in situ sensors for real‐time health monitoring of aircraft structures appears to be a viable option to address these concerns. Recent requests for real‐time monitoring of structures have produced a niche for active sensor systems using remote frequency eddy currents, fiber‐optics, piezoelectric materials, and comparative vacuum monitoring. Reliable, structural health monitoring systems can automatically process data, assess structural condition, and signal the need for human intervention. Prevention of unexpected flaw growth and structural failure could be improved if onboard health monitoring systems are used to continuously assess structural integrity. Such systems would be able to detect incipient damage before catastrophic failure occurs. Condition‐based maintenance practices could be substituted for the current time‐based maintenance approach. Other advantages of onboard distributed sensor systems are that they can eliminate costly, and potentially damaging, disassembly, improve sensitivity by producing optimum placement of sensors with minimized human factor concerns in deployment, and decrease maintenance costs by eliminating more time‐consuming manual inspections. In addition to mature MEMS (microelectromechanical systems) devices such as acoustic emission sensors, accelerometers, strain gauges, and pressure sensors, recent advances in microsensors have produced miniature eddy‐current, ultrasonic, piezoelectric, fiber‐optic, and other devices that lend themselves more directly to damage detection. Technology exists to co‐locate the processing electronics with in situ sensor networks to produce real‐time transmission of data and real‐time diagnostics of structural health. When combined in a systems approach that includes sensors to monitor electronics, hydraulics, and avionics, it is possible to produce an aircraft prognostic health management architecture that can assist in maintenance scheduling and tracking. This article focuses on developments and testing of mountable sensors and how they can be integrated into such a health management system. Specific example applications are discussed along with issues that must be addressed to realistically deploy leave‐in‐place sensors. Successful field testing is presented to quantify the performance of real‐time, health monitoring systems and to highlight their use in guiding condition‐based maintenance activities.

This paper discusses ongoing work to develop structural health monitoring techniques for composite aerospace structures such as air& control surfaces, fuselage sections or repairs, and reusable launch vehicle fuel tanks. The overall project is divided into four tasks: operational evaluation, diagnostic measurements, information condensation, and damage detection. Five composite plates were constructed to study delaminations, disbonds, and fluid retention issues as the initial step in creating an operational system. These four square feet plates were graphite-epoxy with nomex honeycomb cores. The diagnostic measurements are composed of modal tests with a scanning laser vibrometer at over 500 scan points per plate covering the frequency range up to 2000 Hz. This data has been reduced into experimental dynamics matrices using a generic software package developed at the University of Colorado at Boulder. The continuing effort will entail performing a series of damage identification studies to detect, localize, and determine the extent of the damage. This work is providing understanding and algorithm development for a global NDE technique for composite aerospace structures.

An unavoidable by-product of a metallic structure's use is the appearance of crack, corrosion, erosion and other flaws. Economic barriers to the replacement of these structures have created an aging civil and military infrastructure and placed even greater demands on efficient and safe repair and inspection methods. As a result of Homeland Security issues and these aging infrastructure concerns, increased attention has been focused on the rapid repair and preemptive reinforcement of structures such as buildings and bridges. This Laboratory Directed Research and Development (LDRD) program established the viability of using bonded composite patches to repair metallic structures. High modulus fiber-reinforced polymer (FRP) material may be used in lieu of mechanically fastened metallic patches or welds to reinforce or repair damaged structures. Their use produces a wide array of engineering and economic advantages. Current techniques for strengthening steel structures have several drawbacks including requiring heavy equipment for installation, poor fatigue performance, and the need for ongoing maintenance due to continued corrosion attack or crack growth. The use of bonded composite doublers has the potential to correct the difficulties associated with current repair techniques and the ability to be applied where there are currently no rehabilitation options. Applications include such diverse structures as: buildings, bridges, railroad cars, trucks and other heavy machinery, steel power and communication towers, pipelines, factories, mining equipment, ships, tanks and other military vehicles. This LDRD also proved the concept of a living infrastructure by developing custom sensors and self-healing chemistry and linking this technology with the application of advanced composite materials. Structural Health Monitoring (SHM) systems and mountable, miniature sensors were designed to continuously or periodically assess structural integrity. Such systems are able to detect incipient damage before catastrophic failure occurs. The ease of monitoring an entire network of distributed sensors means that structural health assessments can occur more often, allowing operators to be even more vigilant with respect to flaw onset. In addition, the realization of smart structures, through the use of in-situ sensors, allows condition-based maintenance to be substituted for conventional time-based maintenance practices. The sensitivity and reliability of a series of sensor systems was quantified in laboratory and real-world environments. Finally, self healing methods for composite materials were evolved--using resin modules that are released in response to the onset of delaminations--so that these components can provide a living infrastructure with minimal need for human intervention. This program consisted of four related research elements: (1) design, installation, and performance assessment of composite repairs, (2) in-situ sensors for real-time health monitoring, (3) self healing of in-service damage in a repair, and (4) numerical modeling. Deployment of FRP materials and bonded joints requires proper design, suitable surface preparation methods, and adequate surveillance to ensure structural integrity. By encompassing all 'cradle-to-grave' tasks --including design, analysis, installation, durability, flaw containment, and inspection--this program is designed to firmly establish the capabilities of composite doubler repairs and introduce technology to incorporate self-monitoring and self-healing (living structures) methodologies. A proof-of-concept repair was completed on a steel highway bridge in order to demonstrate the potential of composite doubler technology for critical infrastructure use.

Ultrasonic and Advanced Methods for Nondestructive Testing and Material Characterization, pp. 403-437 (2007) No AccessSMART STRUCTURES APPROACH: ACHIEVING GLOBAL STRUCTURAL HEALTH MONITORING THROUGH THE APPLICATION OF DISTRIBUTED, EMBEDDED MICRO-SENSOR SYSTEMSDennis RoachDennis RoachSandia National Laboratories, Albuquerque, NM 87185, USAhttps://doi.org/10.1142/9789812770943_0018Cited by:0 PreviousNext AboutSectionsPDF/EPUB ToolsAdd to favoritesDownload CitationsTrack CitationsRecommend to Library ShareShare onFacebookTwitterLinked InRedditEmail Abstract: Multi-site fatigue damage, hidden cracks in hard-to-reach locations, disbonded joints, erosion, impact, and corrosion are among the major flaws encountered in today's extensive array of aerospace vehicles and civil structures. These damage scenarios, coupled with new and unexpected phenomena, have placed greater demands on the application of advanced nondestructive inspection (NDI) and health monitoring techniques. Reliable, structural health monitoring systems can automatically process real-time data, assess structural condition, and signal the need for human intervention. Prevention of unexpected flaw growth and structural failure could be improved if on-board health monitoring systems are used to continuously assess structural integrity. Such systems would be able to detect incipient damage before catastrophic failures occurs. Condition-based maintenance practices could be substituted for the current time-based maintenance approach. Other advantages of on-board distributed sensor systems are that they can eliminate costly, and potentially damaging, disassembly, improve sensitivity by producing optimum placement of sensors with minimized human factors concerns in deployment, and decrease maintenance costs by eliminating more time-consuming manual inspections. This chapter focuses on developments in mountable sensors and how they can be integrated into such a Structural Health Monitoring (SHM) system to guide condition-based maintenance activities. FiguresReferencesRelatedDetails Ultrasonic and Advanced Methods for Nondestructive Testing and Material CharacterizationMetrics History PDF download

The reinforced carbon-carbon (RCC) heat shield components on the Space Shuttle's wings must withstand harsh atmospheric re-entry environments where the wing leading edge can reach temperatures of 3000°F (about 1650°C). Potential damage includes impact damage, micro cracks, oxidation in the silicon carbide-to-carbon-carbon layers, and interlaminar disbands. The thick, carbon-carbon and silicon-carbide layers in the heat shield panels are difficult to inspect and existing techniques were deemed to be inadequate. Since accumulated damage in these structures can lead to catastrophic failure of the Shuttle's heat protection system, it became imperative for NASA to develop quickly an acceptable health monitoring programme. NASA selected a multi-method approach for inspecting the wing leading edge which includes eddy current, thermography, and ultrasonics. Sandia Labs produced the in-situ ultrasonic pitch-catch inspection method. Optimum combinations of custom ultrasonic probes and data analysis were merged into the overall inspection method needed to properly survey the heat shield panels. Comprehensive validation tests revealed that the ultrasonic pitch-catch inspection system is capable of reliably finding flaws of 0.25" diameter or less in the heat shields. The inspection system is now in use at NASA to complete 'Return-to-Flight' certification inspections prior to each Shuttle launch.