The costs associated with the increasing maintenance and surveillance needs of aging structures are rising at an unexpected rate. 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. Aircraft maintenance and repairs represent about a quarter of a commercial fleet’s operating costs. The application of Structural Health Monitoring (SHM) systems using distributed sensor networks can reduce these costs by facilitating rapid and global assessments of structural integrity. The use of in-situ sensors for real-time health monitoring can 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. The ease of monitoring an entire on-board 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. SHM systems also allow for condition-based maintenance practices to be substituted for the current time-based or cycle-based maintenance approach thus optimizing maintenance labor. The Federal Aviation Administration has conducted a series of SHM validation and certification programs intended to comprehensively support the evolution and adoption of SHM practices into routine aircraft maintenance practices. This report presents one of those programs involving a Sandia Labs-aviation industry effort to move SHM into routine use for aircraft maintenance. The Airworthiness Assurance NDI Validation Center (AANC) at Sandia Labs, in conjunction with Sikorsky, Structural Monitoring Systems Ltd., Anodyne Electronics Manufacturing Corp., Acellent Technologies Inc., and the Federal Aviation Administration (FAA) carried out a trial validation and certification program to evaluate Comparative Vacuum Monitoring (CVM) and Piezoelectric Transducers (PZT) as a structural health monitoring solution to specific rotorcraft applications. 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 and PZT systems into rotorcraft maintenance programs and the document revisions necessary to allow for their routine use as an alternate means of performing periodic structural inspections. 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. Sandia Labs designed, implemented, and analyzed the results from a focused and statistically relevant experimental effort to quantify the reliability of a CVM system applied to Sikorsky S-92 fuselage frame application and a PZT system applied to an S-92 main gearbox mount beam application. The applications included both local and global damage detection assessments. 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 SHM sensors in a manner that agrees with current nondestructive inspection (NDI) validation requirements and is acceptable to both the aviation industry and regulatory bodies. The validation work completed in this program demonstrated the ability of both CVM and PZT SHM systems to detect cracks in rotorcraft components. It proved the ability to use final system response parameters to provide a Green Light/Red Light (“GO” – “NO GO”) decision on the presence of damage. In additional to quantifying the performance of each SHM system for the trial applications on the S-92 platform, this study also identified specific methods that can be used to optimize damage detection, guidance on deployment scenarios that can affect performance and considerations that must be made to properly apply CVM and PZT sensors. These results support the main goal of safely integrating SHM sensors into rotorcraft maintenance programs. Additional benefits from deploying rotorcraft Health and Usage Monitoring Systems (HUMS) may be realized when structural assessment data, collected by an SHM system, is also used to detect structural damage to compliment the operational environment monitoring. The use of in-situ sensors for health monitoring of rotorcraft structures can be a viable option for both flaw detection and maintenance planning activities. This formal SHM validation will allow aircraft manufacturers and airlines to confidently make informed decisions about the proper utilization of CVM and PZT technology. It will also streamline future regulatory actions and formal certification measures needed to assure the safe application of SHM solutions.
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.
Structural Health Monitoring 2019: Enabling Intelligent Life-Cycle Health Management for Industry Internet of Things (IIOT) - Proceedings of the 12th International Workshop on Structural Health Monitoring
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.
Wind energy is quickly becoming a significant contributor to the United States' overall energy portfolio. Wind turbine blades pose a unique set of inspection challenges that span from very thick and attenuative spar cap structures to porous bond lines, varying core material and a multitude of manufacturing defects of interest. The need for viable, accurate nondestructive inspection (NDI) technology becomes more important as the cost per blade, and lost revenue from downtime, grows. To address this growing need, Sandia and SkySpecs collaborated to evaluate NDI methods that are suitable for integration on an autonomous drone inspection platform. A trade study of these NDI methods was performed, and thermography was selected as the primary technique for further evaluation. Based on the capabilities of SkySpecs' custom inspection drone, a miniature microbolometer IR camera was successfully selected and tested in a benchtop setting. After identifying key operating parameters for inspecting wind blade materials, hardware and software integration of the IR camera was performed, and Sandia and SkySpecs conducted initial field testing. Finally, recommendations for a path forward for drone-deployed thermography inspections were provided.
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. This paper presents the results from successful SHM technology validation efforts that established the performance of sensor systems for aircraft fatigue crack detection. 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, and the potential inspector population. All factors that affect SHM sensitivity were included in this program including 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. This paper will also present the formal certification tasks including formal adoption of SHM systems into aircraft manuals and the release of an Alternate Means of Compliance and a modified Service Bulletin to allow for routine use of SHM sensors on commercial aircraft. This program also established a regulatory approval process that includes FAR Part 25 (Transport Category Aircraft) and shows compliance with 25.571 (fatigue) and 25.1529 (Instructions for Continued Airworthiness).
Maintenance, Safety, Risk, Management and Life-Cycle Performance of Bridges - Proceedings of the 9th International Conference on Bridge Maintenance, Safety and Management, IABMAS 2018
Economic barriers to the replacement of bridges and other civil structures have created an aging infrastructure and placed greater demands on the deployment of effective and rapid health monitoring methods. To gain access for inspections, structure and sealant must be removed, disassembly processes must be completed and personnel must be transported to remote locations. Reliable Structural Health Monitoring (SHM) systems can automatically process data, assess structural condition, and signal the need for specific maintenance actions. They can reduce the costs associated with the increasing maintenance and surveillance needs of aging structures. 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 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 SHM systems were used to regularly, or even continuously, assess structural integrity. A research program was completed to develop and validate Comparative Vacuum Monitoring (CVM) sensors for crack detection. Sandia National Labs, in conjunction with private industry and the U.S. Department of Transportation, completed a series of CVM validation and certification programs aimed at establishing the overall viability of these sensors for monitoring bridge structures. Factors that affect SHM sensitivity include flaw size, shape, orientation and location relative to the sensors, along with operational environments. Statistical methods using one-sided tolerance intervals were employed to derive Probability of Flaw Detection (POD) levels for typical application scenarios. Complimentary, multi-year field tests were also conducted to study the deployment and long-term operation of CVM sensors on aircraft and bridges. This paper presents the quantitative crack detection capabilities of the CVM sensor, its performance in actual operating environments, and the prospects for structural health monitoring applications on a wide array of civil structures.
Wind turbine blades pose a unique set of inspection challenges that span from very thick and attentive spar cap structures to porous bond lines, varying core material and a multitude of manufacturing defects of interest. The need for viable, accurate nondestructive inspection (NDI) technology becomes more important as the cost per blade, and lost revenue from downtime, grows. NDI methods must not only be able to contend with the challenges associated with inspecting extremely thick composite laminates and subsurface bond lines, but must also address new inspection requirements stemming from the growing understanding of blade structural aging phenomena. Under its Blade Reliability Collaborative program, Sandia Labs quantitatively assessed the performance of a wide range of NDI methods that are candidates for wind blade inspections. Custom wind turbine blade test specimens, containing engineered defects, were used to determine critical aspects of NDI performance including sensitivity, accuracy, repeatability, speed of inspection coverage, and ease of equipment deployment. The detection of fabrication defects helps enhance plant reliability and increase blade life while improved inspection of operating blades can result in efficient blade maintenance, facilitate repairs before critical damage levels are reached and minimize turbine downtime. The Sandia Wind Blade Flaw Detection Experiment was completed to evaluate different NDI methods that have demonstrated promise for interrogating wind blades for manufacturing flaws or in-service damage. These tests provided the Probability of Detection information needed to generate industry-wide performance curves that quantify: 1) how well current inspection techniques are able to reliably find flaws in wind turbine blades (industry baseline) and 2) the degree of improvements possible through integrating more advanced NDI techniques and procedures. _____________ S a n d i a N a t i o n a l L a b o r a t o r i e s i s a m u l t i m i s s i o n l a b o r a t o r y m a n a g e d a n d o p e r a t e d b y N a t i o n a l T e c h n o l o g y a n d E n g i n e e r i n g S o l u t i o n s o f S a n d i a , L L C , a w h o l l y o w n e d s u b s i d i a r y o f H o n e y w e l l I n t e r n a t i o n a l , I n c . , f o r t h e U . S . D e p a r t m e n t o f E n e r g y ' s N a t i o n a l N u c l e a r S e c u r i t y A d m i n i s t r a t i o n u n d e r c o n t r a c t D E - N A 0 0 0 3 5 2 5 .
The aircraft industry continues to increase its use of composite materials, most noteworthy in the arena of principle structural elements. This expanded use, coupled with difficulties associated with damage tolerance analysis of composites, has placed greater emphasis on the application of accurate nondestructive inspection (NDI) methods. Traditionally, a few ultrasonic-based inspection methods have been used to inspect solid laminate structures. Recent developments in more advanced NDI techniques have produced a number of new inspection options. Many of these methods can be categorized as wide area techniques that produce two-dimensional flaw maps of the structure. An experiment has been developed to assess the ability of both conventional and advanced NDI techniques to detect voids, disbonds, delaminations, and impact damage in adhesively bonded composite aircraft structures. A series of solid laminate, carbon composite specimens with statistically relevant flaw profiles are being inspected using conventional, hand-held pulse echo UT and resonance, as well as, new NDI methods that have recently been introduced to improve sensitivity and repeatability of inspections. The primary factors affecting flaw detection in laminates are included in this study: material type, flaw profiles, presence of complex geometries like taper and substructure elements, presence of fasteners, secondarily bonded joints, and environmental conditions. One phase of this effort utilized airline personnel to study Probability of Detection (POD) in the field and to formulate improvements to existing inspection techniques. In addition, advanced NDI methods for laminate inspections — such as thermography, shearography, laser ultrasonics, microwave, and phased/linear array UT — were applied to quantify the improvements achievable through the use of more sophisticated NDI. This report presents the composite laminate experiment design and the POD results for advanced NDI with comparisons to results achieved by airline inspectors using conventional UT methods. A companion report provides the full set of results from the conventional NDI testing.
Composite structures are increasing in prevalence throughout the aerospace, wind, defense, and transportation industries, but the many advantages of these materials come with unique challenges, particularly in inspecting and repairing these structures. Because composites of- ten undergo sub-surface damage mechanisms which compromise the structure without a clear visual indication, inspection of these components is critical to safely deploying composite re- placements to traditionally metallic structures. Impact damage to composites presents one of the most signi fi cant challenges because the area which is vulnerable to impact damage is generally large and sometimes very dif fi cult to access. This work seeks to further evolve iden- ti fi cation technology by developing a system which can detect the impact load location and magnitude in real time, while giving an assessment of the con fi dence in that estimate. Fur- thermore, we identify ways by which impact damage could be more effectively identi fi ed by leveraging impact load identi fi cation information to better characterize damage. The impact load identi fi cation algorithm was applied to a commercial scale wind turbine blade, and results show the capability to detect impact magnitude and location using a single accelerometer, re- gardless of sensor location. A technique for better evaluating the uncertainty of the impact estimates was developed by quantifying how well the impact force estimate meets the assump- tions underlying the force estimation technique. This uncertainty quanti fi cation technique was found to reduce the 95% con fi dence interval by more than a factor of two for impact force estimates showing the least uncertainty, and widening the 95% con fi dence interval by a fac- tor of two for the most uncertain force estimates, avoiding the possibility of understating the uncertainty associated with these estimates. Linear vibration based damage detection tech- niques were investigated in the context of structural stiffness reductions and impact damage. A method by which the sensitivity to damage could be increased for simple structures was presented, and the challenges of applying that technique to a more complex structure were identi fi ed. The structural dynamic changes in a weak adhesive bond were investigated, and the results showed promise for identifying weak bonds that show little or no static reduction in stiffness. To address these challenges in identifying highly localized impact damage, the possi- bility of detecting damage through nonlinear dynamic characteristics was also identi fi ed, with a proposed technique which would leverage impact location estimates to enable the detection of impact damage. This nonlinear damage identi fi cation concept was evaluated on a composite panel with a substructure disbond, and the results showed that the nonlinear dynamics at the damage site could be observed without a baseline healthy reference. By further developing impact load identi fi cation technology and combining load and damage estimation techniques into an integrated solution, the challenges associated with impact detection in composite struc- tures can be effectively solved, thereby reducing costs, improving safety, and enhancing the operational readiness and availability of high value assets.
Recent structural failures such as the I-35W Mississippi River Bridge in Minnesota have underscored the urgent need for improved methods and procedures for evaluating our aging transportation infrastructure. This research seeks to develop a basis for a Structural Health Monitoring (SHM) system to provide quantitative information related to the structural integrity of metallic structures to make appropriate management decisions and ensuring public safety. This research employs advanced structural analysis and nondestructive testing (NDT) methods for an accurate fatigue analysis. Metal railroad bridges in New Mexico will be the focus since many of these structures are over 100 years old and classified as fracture-critical. The term fracture-critical indicates that failure of a single component may result in complete collapse of the structure such as the one experienced by the I-35W Bridge. Failure may originate from sources such as loss of section due to corrosion or cracking caused by fatigue loading. Because standard inspection practice is primarily visual, these types of defects can go undetected due to oversight, lack of access to critical areas, or, in riveted members, hidden defects that are beneath fasteners or connection angles. Another issue is that it is difficult to determine the fatigue damage that a structure has experienced and the rate at which damage is accumulating due to uncertain history and load distribution in supporting members. A SHM system has several advantages that can overcome these limitations. SHM allows critical areas of the structure to be monitored more quantitatively under actual loading. The research needed to apply SHM to metallic structures was performed and a case study was carried out to show the potential of SHM-driven fatigue evaluation to assess the condition of critical transportation infrastructure and to guide inspectors to potential problem areas. This project combines the expertise in transportation infrastructure at New Mexico State University with the expertise at Sandia National Laboratories in the emerging field of SHM.
Goal: to detect the subtle changes in laminate composite structures brought about by thermal, chemical, ultraviolet, and moisture exposure. Compare sensitivity of an array of NDI methods, including Fourier Transform Infrared Spectroscopy (FTIR), to detect subtle differences in composite materials due to deterioration. Inspection methods applied: ultrasonic pulse echo, through transmission ultrasonics, thermography, resonance testing, mechanical impedance analysis, eddy current, low frequency bond testing & FTIR. Comparisons between the NDI methods are being used to establish the potential of FTIR to provide the necessary sensitivity to non-visible, yet significant, damage in the resin and fiber matrix of composite structures. Comparison of NDI results with short beam shear tests are being used to relate NDI sensitivity to reduction in structural performance. Chemical analyses technique, which measures the infrared intensity versus wavelength of light reflected on the surface of a structure (chemical and physical information via this signature). Advances in instrumentation have resulted in hand-held portable devices that allow for field use (few seconds per scan). Shows promise for production quality assurance and in-service applications on composite aircraft structures (scarfed repairs). Statistical analysis on frequency spectrums produced by FTIR interrogations are being used to produce an NDI technique for assessing material integrity. Conclusions are: (1) Use of NDI to assess loss of composite laminate integrity brought about by thermal, chemical, ultraviolet, and moisture exposure. (2) Degradation trends between SBS strength and exposure levels (temperature and time) have been established for different materials. (3) Various NDI methods have been applied to evaluate damage and relate this to loss of integrity - PE UT shows greatest sensitivity. (4) FTIR shows promise for damage detection and calibration to predict structural integrity (short beam shear). (5) Detection of damage for medium exposure levels (possibly resin matrix degradation only) is more difficult and requires additional study. (6) These are initial results only - program is continuing with additional heat, UV, chemical and water exposure test specimens.
Airworthiness Assurance NDI Validation Center (AANC) objectives are: (1) Enhance aircraft safety and reliability; (2) Aid developing advanced aircraft designs and maintenance techniques; (3) Provide our customers with comprehensive, independent, and quantitative/qualitative evaluations of new and enhanced inspection, maintenance, and repair techniques; (4) Facilitate transferring effective technologies into the aviation industry; (5) Support FAA rulemaking process by providing guidance on content & necessary tools to meet requirements or recommendations of FARs, ADs, ACs, SBs, SSIDs, CPCP, and WFD; and (6) Coordinate with and respond to Airworthiness Assurance Working Group (AAWG) in support of FAA Aviation Rulemaking Advisory Committee (ARAC).
The goal is to create a systematic method and structure to compile, organize, and summarize SHM related data to identify the level of maturity and rate of evolution and have a quick and ongoing evaluation of the current state of SHM among research institutions and industry. Hundreds of technical publication and conference proceedings were read and analyzed to compile the database. Microsoft Excel was used to create a useable interface that could be filtered to compare any of the entered data fields.
The reinforced carbon-carbon (RCC) heat shield components on the Space Shuttle's wings must withstand harsh atmospheric reentry environments where the wing leading edge can reach temperatures of 3,000 F. Potential damage includes impact damage, micro cracks, oxidation in the silicon carbide-to-carbon-carbon layers, and interlaminar disbonds. Since accumulated damage in the thick, carbon-carbon and silicon-carbide layers of the heat shields can lead to catastrophic failure of the Shuttle's heat protection system, it was essential for NASA to institute an accurate health monitoring program. NASA's goal was to obtain turnkey inspection systems that could certify the integrity of the Shuttle heat shields prior to each mission. Because of the possibility of damaging the heat shields during removal, the NDI devices must be deployed without removing the leading edge panels from the wing. Recently, NASA selected a multi-method approach for inspecting the wing leading edge which includes eddy current, thermography, and ultrasonics. The complementary superposition of these three inspection techniques produces a rigorous Orbiter certification process that can reliably detect the array of flaws expected in the Shuttle's heat shields. Sandia Labs produced an in-situ ultrasonic inspection method while NASA Langley developed the eddy current and thermographic techniques. An extensive validation process, including blind inspections monitored by NASA officials, demonstrated the ability of these inspection systems to meet the accuracy, sensitivity, and reliability requirements. This report presents the ultrasonic NDI development process and the final hardware configuration. The work included the use of flight hardware and scrap heat shield panels to discover and overcome the obstacles associated with damage detection in the RCC material. Optimum combinations of custom ultrasonic probes and data analyses were merged with the inspection procedures needed to properly survey the heat shield panels. System features were introduced to minimize the potential for human factors errors in identifying and locating the flaws. The in-situ NDI team completed the transfer of this technology to NASA and USA employees so that they can complete 'Return-to-Flight' certification inspections on all Shuttle Orbiters prior to each launch.
An unavoidable by-product of a metallic structure's use is the appearance of crack and corrosion flaws. Economic barriers to the replacement of these structures have created an aging infrastructure and placed even greater demands on efficient and safe repair methods. In the past decade, an advanced composite repair technology has made great strides in commercial aviation use. Extensive testing and analysis, through joint programs between the Sandia Labs FAA Airworthiness Assurance Center and the aviation industry, have proven that composite materials can be used to repair damaged aluminum structure. Successful pilot programs have produced flight performance history to establish the durability of bonded composite patches as a permanent repair on commercial aircraft structures. With this foundation in place, this effort is adapting bonded composite repair technology to civil structures. 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 no rehabilitation options. It promises to be cost-effective with minimal disruption to the users of the structure. This report concludes a study into the application of composite patches on thick steel structures typically used in mining operations. Extreme fatigue, temperature, erosive, and corrosive environments induce an array of equipment damage. The current weld repair techniques for these structures provide a fatigue life that is inferior to that of the original plate. Subsequent cracking must be revisited on a regular basis. The use of composite doublers, which do not have brittle fracture problems such as those inherent in welds, can help extend the structure's fatigue life and reduce the equipment downtime. Two of the main issues for adapting aircraft composite repairs to civil applications are developing an installation technique for carbon steel and accommodating large repairs on extremely thick structures. This study developed and proved an optimum field installation process using specific mechanical and chemical surface preparation techniques coupled with unique, in-situ heating methods. In addition, a comprehensive performance assessment of composite doubler repairs was completed to establish the viability of this technology for large, steel structures. The factors influencing the durability of composite patches in severe field environments were evaluated along with related laminate design issues.
The FAA's Airworthiness Assurance NDI Validation Center, in conjunction with the Commercial Aircraft Composite Repair Committee, developed a set of composite reference standards to be used in NDT equipment calibration for accomplishment of damage assessment and post-repair inspection of all commercial aircraft composites. In this program, a series of NDI tests on a matrix of composite aircraft structures and prototype reference standards were completed in order to minimize the number of standards needed to carry out composite inspections on aircraft. Two tasks, related to composite laminates and non-metallic composite honeycomb configurations, were addressed. A suite of 64 honeycomb panels, representing the bounding conditions of honeycomb construction on aircraft, was inspected using a wide array of NDI techniques. An analysis of the resulting data determined the variables that play a key role in setting up NDT equipment. This has resulted in a set of minimum honeycomb NDI reference standards that include these key variables. A sequence of subsequent tests determined that this minimum honeycomb reference standard set is able to fully support inspections over the full range of honeycomb construction scenarios found on commercial aircraft. In the solid composite laminate arena, G11 Phenolic was identified as a good generic solid laminate reference standard material. Testing determined matches in key velocity and acoustic impedance properties, as well as, low attenuation relative to carbon laminates. Furthermore, comparisons of resonance testing response curves from the G11 Phenolic NDI reference standard was very similar to the resonance response curves measured on the existing carbon and fiberglass laminates. NDI data shows that this material should work for both pulse-echo (velocity-based) and resonance (acoustic impedance-based) inspections.
In the past decade, an advanced composite repair technology has made great strides in commercial aviation use. Extensive testing and analysis, through joint programs between the Sandia Labs FAA Airworthiness Assurance Center and the aviation industry, have proven that composite materials can be used to repair damaged aluminum structure. Successful pilot programs have produced flight performance history to establish the viability and durability of bonded composite patches as a permanent repair on commercial aircraft structures. With this foundation in place, efforts are underway to adapt bonded composite repair technology to civil structures. This paper presents a study in the application of composite patches on large trucks and hydraulic shovels typically used in mining operations. Extreme fatigue, temperature, erosive, and corrosive environments induce an array of equipment damage. The current weld repair techniques for these structures provide a fatigue life that is inferior to that of the original plate. Subsequent cracking must be revisited on a regular basis. It is believed that the use of composite doublers, which do not have brittle fracture problems such as those inherent in welds, will help extend the structure's fatigue life and reduce the equipment downtime. Two of the main issues for adapting aircraft composite repairs to civil applications are developing an installation technique for carbon steel structure and accommodating large repairs on extremely thick structures. This paper will focus on the first phase of this study which evaluated the performance of different mechanical and chemical surface preparation techniques. The factors influencing the durability of composite patches in severe field environments will be discussed along with related laminate design and installation issues.
One of the concerns surrounding composite doubler technology pertains to long-term survivability, especially in the presence of non-optimum installations. This test program demonstrated the damage-tolerance capabilities of bonded composite doublers. The fatigue and strength tests quantified the structural response and crack-abatement capabilities of boron-epoxy doublers in the presence of worst-case flaw scenarios. The engineered flaws included cracks in the parent material, disbonds in the adhesive layer, and impact damage to the composite laminate. Environmental conditions representing temperature and humidity exposure were also included in the coupon tests. Large strains immediately adjacent to the doubler flaws emphasize the fact that relatively large disbond or delamination flaws (up to 100 diameter) in the composite doubler have only localized effects on strain and minimal effect on the overall doubler performance (i.e., undesirable strain relief over disbond but favorable load transfer immediately next to disbond). This statement is made relative to the inspection requirement that result in the detection of disbonds/delaminations of 0.5 '' diameter or greater. The point at which disbonds become detrimental depends upon the size and location of the disbond and the strain field around the doubler. This study did not attempt to determine a "flaw size vs. effect" relation. Rather, it used flaws that were twice as large as the detectable limit to demonstrate the ability of composite doublers to tolerate potential damage.
The number of commercial airframes exceeding twenty years of service continues to grow. 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 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. The use of bonded composite doublers offers the airframe manufacturers and aircraft maintenance facilities a cost effective method to safety 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) is conducting a program with Boeing and Federal Express to validate and introduce composite doubler repair technology to the US commercial aircraft industry. This project focuses on repair of DC-10 structure and builds on the foundation of the successful L-1011 door corner repair that was completed by the AANC, Lockheed-Martin, and Delta Air Lines. The L-1011 composite doubler repair was installed in 1997 and has not developed any flaws in over three years of service, As a follow-on effort, this DC-1O repair program investigated design, analysis, performance (durability, flaw containment, reliability), installation, and nondestructive inspection issues. Current activities are demonstrating regular use of composite doubler repairs on commercial aircraft. The primary goal of this program is to move the technology into niche applications and to streamline the design-to-installation process. Using the data accumulated to date, the team has designed, analyzed, and developed inspection techniques for an array of composite doubler repairs with high-use fuselage skin applications. The general DC-10 repair areas which provide a high payoff to FedEx and which minimize design and installation complexities have been identified as follows: (1) gouges, dents, lightning strike, and impact skin damage, and (2) corrosion grind outs in surface skin. This paper presents the engineering activities that have been completed in order to make this technology available for widespread commercial aircraft use.
Rotorcraft structures do not readily lend themselves to quantifiable inspection methods due to airframe construction techniques. Periodic visual inspections are a common practice for detecting corrosion. Unfortunately, when the telltale signs of corrosion appear visually, extensive repair or refurbishment is required. There is a need to nondestructively evaluate airframe structures in order to recognize and quantify corrosion before visual indications are present. Nondestructive evaluations of rotorcraft airframes face inherent problems different from those of the fixed wing industry. Most rotorcraft lap joints are very narrow, contain raised fastener heads, may possess distortion, and consist of thinner gage materials ({approximately}0.012--0.125 inches). In addition the structures involve stack-ups of two and three layers of thin gage skins that are separated by sealant of varying thickness. Industry lacks the necessary data techniques, and experience to adequately perform routine corrosion inspection of rotorcraft. In order to address these problems, a program is currently underway to validate the use of eddy current inspection on specific rotorcraft lap joints. Probability of detection (POD) specimens have been produced that simulate two lap joint configurations on a model TH-57/206 helicopter. The FAA's Airworthiness Assurance Center (AANC) at Sandia Labs and Bell Helicopter have applied single and dual frequency eddy current (EC) techniques to these test specimens. The test results showed enough promise to justify beta site testing of the eddy current methods evolved in this study. The technique allows users to distinguish between corrosion signals and those caused by varying gaps between the assembly of skins. Specific structural joints were defined as prime corrosion areas and a series of corrosion specimens were produced with 5--20% corrosion distributed among the layers of each joint. Complete helicopter test beds were used to validate the laboratory findings. This paper will present the laboratory and field results that quantify the EC technique's corrosion detection performance. Plans for beta site testing, adoption of the new inspection procedure into routine rotorcraft maintenance, and NDI training issues will also be discussed.
A project to develop non-intrusive active sensors that can be applied on existing aging aerospace structures for monitoring the onset and progress of structural damage (fatigue cracks and corrosion) is presented. The state of the art in active sensors structural health monitoring and damage detection is reviewed. Methods based on (a) elastic wave propagation and (b) electro-mechanical (NM) impedance technique are sighted and briefly discussed. The instrumentation of these specimens with piezoelectric active sensors is illustrated. The main detection strategies (E/M impedance for local area detection and wave propagation for wide area interrogation) are discussed. The signal processing and damage interpretation algorithms are tuned to the specific structural interrogation method used. In the high-frequency EIM impedance approach, pattern recognition methods are used to compare impedance signatures taken at various time intervals and to identify damage presence and progression from the change in these signatures. In the wave propagation approach, the acoustic-ultrasonic methods identifying additional reflection generated from the damage site and changes in transmission velocity and phase are used. Both approaches benefit from the use of artificial intelligence neural networks algorithms that can extract damage features based on a learning process. Design and fabrication of a set of structural specimens representative of aging aerospace structures is presented. Three built-up specimens, (pristine, with cracks, and with corrosion damage) are used. The specimen instrumentation with active sensors fabricated at the University of South Carolina is illustrated. Preliminary results obtained with the E/M impedance method on pristine and cracked specimens are presented.
A project to develop non-intrusive active sensors that can be applied on existing aging aerospace structures for monitoring the onset and progress of structural damage (fatigue cracks and corrosion) is presented. The state of the art in active sensors structural health monitoring and damage detection is reviewed. Methods based on (a) elastic wave propagation and (b) electro-mechanical (E/M) impedance technique are cited and briefly discussed. The instrumentation of these specimens with piezoelectric active sensors is illustrated. The main detection strategies (E/M impedance for local area detection and wave propagation for wide area interrogation) are discussed. The signal processing and damage interpretation algorithms are tuned to the specific structural interrogation method used. In the high-frequency E/M impedance approach, pattern recognition methods are used to compare impedance signatures taken at various time intervals and to identify damage presence and progression from the change in these signatures. In the wave propagation approach, the acousto-ultrasonic methods identifying additional reflection generated from the damage site and changes in transmission velocity and phase are used. Both approaches benefit from the use of artificial intelligence neural networks algorithms that can extract damage features based on a learning process. Design and fabrication of a set of structural specimens representative of aging aerospace structures is presented. Three built-up specimens (pristine, with cracks, and with corrosion damage) are used. The specimen instrumentation with active sensors fabricated at the University of South Carolina is illustrated. Preliminary results obtained with the E/M impedance method on pristine and cracked specimens are presented.
The rapidly increasing use of composites on commercial airplanes coupled with the potential for economic savings associated with their use in aircraft structures means that the demand for composite materials technology will continue to increase. Inspecting these composite structures is a critical element in assuring their continued airworthiness. The FAA's Airworthiness Assurance NDI Validation Center, in conjunction with the Commercial Aircraft Composite Repair Committee (CACRC), is developing a set of composite reference standards to be used in NDT equipment calibration for accomplishment of damage assessment and post-repair inspection of all commercial aircraft composites. In this program, a series of NDI tests on a matrix of composite aircraft structures and prototype reference standards were completed in order to minimize the number of standards needed to carry out composite inspections on aircraft. Two tasks, related to composite laminates and non-metallic composite honeycomb configurations, were addressed. A suite of 64 honeycomb panels, representing the bounding conditions of honeycomb construction on aircraft, were inspected using a wide array of NDI techniques. An analysis of the resulting data determined the variables that play a key role in setting up NDT equipment. This has resulted in a prototype set of minimum honeycomb reference standards that include these key variables. A sequence of subsequent tests determined that this minimum honeycomb reference standard set is able to fully support inspections over the fill range of honeycomb construction scenarios. Current tasks are aimed at optimizing the methods used to engineer realistic flaws into the specimens. In the solid composite laminate arena, we have identified what appears to be an excellent candidate, G11 Phenolic, as a generic solid laminate reference standard material. Testing to date has determined matches in key velocity and acoustic impedance properties, as well as, low attenuation relative to carbon laminates. Furthermore, comparisons of resonance testing response curves from the G11 Phenolic prototype standard was very similar to the resonance response curves measured on the existing carbon and fiberglass laminates. NDI data shows that this material should work for both pulse-echo (velocity-based) and resonance (acoustic impedance-based) inspections. Additional testing and industry review activities are underway to complete the validation of this material.
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 possible to bond a single Boron-Epoxy composite doubler to the damaged structure. Most of the concerns surrounding composite doubler technology pertain to long-term survivability, especially in the presence of non-optimum installations, and the validation of appropriate inspection procedures. This report focuses on a series of full-scale structural and nondestructive inspection (NDI) tests that were conducted to investigate the performance of Boron-Epoxy composite doublers. Full-scale tests were conducted on fuselage panels cut from retired aircraft. These full-scale tests studied stress reductions, crack mitigation, and load transfer capabilities of composite doublers using simulated flight conditions of cabin pressure and axial stress. Also, structures which modeled key aspects of aircraft structure repairs were subjected to extreme tension, shear and bending loads to examine the composite laminate's resistance to disbond and delamination flaws. Several of the structures were loaded to failure in order to determine doubler design margins. Nondestructive inspections were conducted throughout the test series in order to validate appropriate techniques on actual aircraft structure. The test results showed that a properly designed and installed composite doubler is able to enhance fatigue life, transfer load away from damaged structure, and avoid the introduction of new stress risers (i.e. eliminate global reduction in the fatigue life of the structure). Comparisons with test data obtained prior to the doubler installation revealed that stresses in the parent material can be reduced 30%--60% through the use of the composite doubler. Tests to failure demonstrated that the bondline is able to transfer plastic strains into the doubler and that the parent aluminum skin must experience significant yield strains before any damage to the doubler will occur.
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 possible to bond a single boron-epoxy composite doubler to the damaged structure. In order for the use of composite doublers to achieve widespread use in the civil aviation industry, it is imperative that methods be developed which can quickly and reliably assess the integrity of the doubler. In this study, a specific composite application was chosen on an L-1011 aircraft in order to focus the tasks on application and operation issues. Primary among inspection requirements for these doublers is the identification of disbonds, between the composite laminate and aluminum parent material, and delaminations in the composite laminate. Surveillance of cracks or corrosion in the parent aluminum material beneath the doubler is also a concern. No single nondestructive inspection (NDI) method can inspect for every flaw type, therefore it is important to be aware of available NDI techniques and to properly address their capabilities and limitations. A series of NDI tests were conducted on laboratory test structures and on full-scale aircraft fuselage sections. Specific challenges, unique to bonded composite doubler applications, were highlighted. An array of conventional and advanced NDI techniques were evaluated. Flaw detection sensitivity studies were conducted on applicable eddy current, ultrasonic, X-ray and thermography based devices. The application of these NDI techniques to composite doublers and the results from test specimens, which were loaded to provide a changing flaw profile, are presented in this report. It was found that a team of these techniques can identify flaws in composite doubler installations well before they reach critical size.
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. This paper focuses on a specialized structural test facility which was developed to evaluate the performance of composite doublers on actual aircraft structure. The facility can subject an aircraft fuselage section to a combined load environment of pressure (hoop stress) and axial, or longitudinal, stress. The tests simulate maximum cabin pressure loads and use a computerized feedback system to maintain the proper ratio between hoop and axial loads. Through the use of this full-scale test facility it was possible to: (1) assess general composite doubler response in representative flight load scenarios, and (2) verify the design and analysis approaches as applied to an L-1011 door corner repair.
The number of commercial airframes exceeding twenty years of service continues to grow. In addition, Service Life Extension Programs are attempting to extend the {open_quotes}economic{close_quotes} service life of commercial airframes to thirty years. The use of bonded composites may offer the airframe manufacturers and aircraft maintenance facilities a cost effective method to extend the lives of their aircraft. The Federal Aviation Administration Assurance NDI Validation Center (AANC) to validate the use of bonded composite doublers on commercial aircraft.
Commercial airframes exceeding 20 service years often develop crack and corrosion flaws. Bonded composite doublers offer a cost effective method to safely extend aircraft lives. The Federal Aircraft Authority (FAA) has completed a project to introduce composite doubler repair technology to the commercial aircraft industry. Instead of riveting steel or aluminum plates for repair, a single composite doubler may be bonded to the damaged structure. Adhesive bonding eliminates stress concentrations caused by fastener holes. Composites are readily formed into complex shapes for repairing irregular components. Also, composite doublers can be tailored to meet specific anisotropy needs, eliminating structural stiffening in directions other than those required. Other advantages include corrosion resistance, a high strength-to-weight ratio, and potential time savings in installation. One phase of this study developed general methodologies and test programs to ensure proper performance of the technique. A second phase focused on reinforcement of an L-1011 door frame, and encompassed all lifetime tasks such as design, analysis, installation, and nondestructive inspection. This paper overviews the project and details the activities conducted to gain FAA approval for composite doubler use. Structural tests evaluated the damage tolerance and fatigue performance of composite doublers while finite element models were generated to study doubler design issues. Nondestructive inspection procedures were developed and validated using full-scale test articles. Installation dry-runs demonstrated the viability of applying composite doublers in hangar environments. The project`s documentation package was used to support installation of a Boron-Epoxy composite repair on a Delta Air Lines L-1011 aircraft. A second product of the results is a Lockheed Service Bulletin which allows the door corner composite doubler to be installed on all L-1011 aircraft. 9 refs., 10 figs., 2 tabs.
One of the major thrusts established under the FAA's National Aging Aircraft Research Program is to foster new technologies associated with civil aircraft maintenance. Recent DOD and other government developments in the use of bonded composite doublers on metal structures has supported the need for research and validation of such doubler applications on US certificated airplanes. Composite doubler technology is rapidly maturing and shows promise of cost savings on aging aircraft. While there have been numerous studies and military aircraft installations of composite doublers, the technology has not been certified for use on commercial aircraft. Before the use of composite doublers can be accepted by the civil aviation industry, it is imperative that methods be developed which can quickly and reliably assess the integrity of the doubler. In this study, a specific composite application was chosen on an L-1011 aircraft in order to focus the tasks on application and operation issues. Primary among inspection requirements for these doublers is the identification of disbonds, between the composite laminate and aluminum parent material, and delaminations in the composite laminate. Surveillance of cracks or corrosion in the inspection (NDI) method can inspect for every flaw type, therefore it is important to be aware of available NDI techniques and to properly address their capabilities and limitations. This paper reports on a series of NDI tests which have been conducted on laboratory test structures and on a fuselage section cut from a retired L-1011 aircraft. Specific challenges, unique to bonded composite doubler applications, will be highlighted. In order to quickly integrate this technology into existing aircraft maintenance depots, the use of conventional NDI, ultrasonics, x-ray, and eddy current, is stressed. The application of these NDI technique to composite doublers and the results from test specimens, which were loaded to provide a changing flaw profile, are presented in this paper. The development of appropriate inspection calibration standards will also be discussed.
One of the major thrusts established under the FAA`s National Aging Aircraft Research Program is to foster new technologies associated with civil aircraft maintenance. Recent DOD and other government developments in the use of bonded composite patches on metal structures has supported the need for research and validation of such doubler applications on U.S. certificated airplanes. Composite patching is a rapidly maturing technology which shows promise of cost savings on aging aircraft. Sandia Labs is conducting a proof-of-concept project with Delta Air Lines, Lockheed Martin, Textron, and the FAA which seeks to remove any remaining obstacles to the approved use of composite doublers. By focusing on a specific commercial aircraft application - reinforcement of the L-1011 door frame - and encompassing all {open_quotes}cradle-to-grave{close_quotes} tasks such as design, analysis, installation, and inspection, this program is designed to prove the capabilities of composite doublers. This paper reports on a series of structural tests which have been conducted on coupons and subsize test articles. Tension-tension fatigue and residual strength tests attempted to grow engineered flaws in coupons with composite doublers bonded to aluminum skin. Also, structures which modeled key aspects of the door corner installation were subjected to extreme tension, shear, and bending loads. In this manner it was possible to study strain fields in and around the Lockheed-designed composite doubler using realistic aircraft load scenarios and to assess the potential for interply delaminations and disbonds between the aluminum and the laminate. The data acquired was also used to validate finite element models (FEM) and associated Damage Tolerance Analyses.
The Aging Aircraft NDI Validation Center (AANC) was established by the Federal Aviation Administration Technical Center (FAATC) at Sandia National Laboratories in August of 1991. The goal of the AANC is to provide independent validation of technologies intended to enhance the structural inspection of aging commuter and transport aircraft. The deliverables from the AANC`s validation activities are assessments of the reliability of existing and emerging inspection technologies as well as analyses of the cost benefits to be derived from their implementation. This paper describes the methodology developed by the AANC to assess the performance of NDI techniques. In particular, an experiment being developed to evaluate corrosion detection devices will be presented. The experiment uses engineered test specimens, as well as complete aircraft test beds to provide metrics for NDI validation.
Researchers contend that composite repairs (or structural reinforcement doublers) offer numerous advantages over metallic patches including corrosion resistance, light weight, high strength, elimination of rivets, and time savings in installation. Their use in commercial aviation has been stifled by uncertainties surrounding their application, subsequent inspection and long-term endurance. The process of repairing or reinforcing airplane structures is time consuming and the design is dependent upon an accompanying stress and fatigue analysis. A repair that is too stiff may result in a loss of fatigue life, continued growth of the crack being repaired, and the initiation of a new flaw in the undesirable high stress field around the patch. Uncertainties in load spectrums used to design repairs exacerbates these problems as does the use of rivets to apply conventional doublers. Many of these repair or structural reinforcement difficulties can be addressed through the use of composite doublers. Primary among unknown entities are the effects of non-optimum installations and the certification of adequate inspection procedures. This paper presents on overview of a program intended to introduce composite doubler technology to the US commercial aircraft fleet. In this project, a specific composite application has been chosen on an L-1011 aircraft in order to focus the tasks on application and operation issues. Through the use of laboratory test structures and flight demonstrations on an in-service L-1011 airplane, this study is investigating composite doubler design, fabrication, installation, structural integrity, and non-destructive evaluation. In addition to providing an overview of the L-1011 project, this paper focuses on a series of fatigue and strength tests which have been conducted in order to study the damage tolerance of composite doublers. Test results to-date are presented.
This report identifies and describes emerging nondestructive inspection (NDI) methods that can potentially be used to inspect commercial transport and commuter aircraft for structural damage. The nine categories of emerging NDI techniques are: acoustic emission, x-ray computed tomography, backscatter radiation, reverse geometry x-ray, advanced electromagnetics, including magnetooptic imaging and advanced eddy current techniques, coherent optics, advanced ultrasonics, advanced visual, and infrared thermography. The physical principles, generalized performance characteristics, and typical applications associated with each method are described. In addition, aircraft inspection applications are discussed along with the associated technical considerations. Finally, the status of each technique is presented, with a discussion on when it may be available for use in actual aircraft maintenance programs. It should be noted that this is a companion document to DOT/FAA/CT-91/5, Current Nondestructive Inspection Methods for Aging Aircraft.
In response to a request from the US Government, Sandia National Laboratories has developed an instrumentation system to monitor the conditions along an underwater, rotating drive shaft. It was desired to study the structural integrity and signal acquisition capabilities of the Shaft Instrumentation System (SIS) in an environment which closely simulates the actual deployment conditions. In this manner, the SIS response to ill-defined conditions, such as flow field turbulence or temperature fluctuations, could be determined. An Underwater Spin Facility was developed in order to verify the operation of the instrumentation and telemetric data acquisition system in a combined environment of external pressure, transient axial loads and centrifugal force. The main components of the Underwater Spin Facility are a large, five foot diameter pressure vessel, a dynamically sealed shaft, a drive train assembly and a shaker table interface which is used to apply the axial loads. This paper presents a detailed description of the design of the Underwater Spin Facility. It also discusses the SIS certification test program in order to demonstrate the successful performance of the Underwater Spin Facility. 8 refs., 10 figs.