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Recent Advances in Inspection & Health Monitoring of Aerospace Materials & Structures for General Aviation Jay Amos (Cessna Aircraft) NDE Materials & Processes Engineering 1 Cessna Aircraft Cessna is part of Textron, a multi-industry company with strong brands such as Bell Helicopter, Kautex, Lycoming, AAI Corporation, E-Z-GO Jacobsen and Greenlee. Citation X Cessna's Citation X is the fastest business jet in the world (0.92 Mach), capable of flying from New York to London in just under 6 hours. 300th production Citation X rolled off the production line in 12/08; fleet has exceeded 1M flight hours Citation Sovereign Since the first delivery in late 2004, more than 250 Sovereigns have been delivered. Citation XLS+ FAA certification was achieved in 5/08 (EASA in 3/09); first retail XLS+ delivered in 12/08 With a global fleet of more than 700 and more than 1.4M flight hours (as of 8/08), the Excel/XLS is the best-selling business jet model Citation CJ4 CJ4 prototype completed first flight in 5/08; First flight of the production CJ4 in 8/08. Plan to begin customer deliveries in the first half of 2010. Mustang The first retail Citation Mustang entered service in April 2007, and Cessna delivered 45 in 2007 About three of every five Mustangs go to international customers, Certified in 60 countries through April 2009; Fleet reached 100 in August 2008 and 200 in May 2009 Corvalis and Corvalis TT The Corvalis TT received EASA certification February 2009 Corvalis TT – the fastest fixed-gear single-engine piston aircraft on the market Acquired certain assets of the former Columbia Aircraft Manufacturing Co. in December 2007 Skycatcher The Skycatcher Achieves ASTM Compliance in 7/09 Clearing the Way for Deliveries Skyhawk Cessna delivered the 40,000th Model 172 in 2008 - the most produced aircraft in history 2 Cessna History Cessna Aircraft Company is the world's largest manufacturer of general aviation airplanes in terms of units delivered. In 2008, Cessna delivered 1,301 aircraft, including 467 Citation business jets. Leading designer and manufacturer of light and mid-size business jets and utility turboprops, and a leading manufacturer of single-engine piston aircraft Since the company was originally established in 1927, some 192,000 Cessna airplanes have been delivered around the world, flying in nearly 80 countries including more than 24,000 twin-engine, 2,000 military jets and 6,000 Citations, making it the largest fleet of business jets in the world. Employment of 8019 as of 9/09. Most complete line of business jets in the industry - a Citation takes off or lands somewhere in the world every 20 seconds. 2009 approximate delivery rates (projected) 400 single-engine pistons 100 Caravans 275 Citation business jets 3 Cessna Facilities Wichita - Mid-Continent Headquarters and jet assembly, 3.5 million square feet Wichita - Pawnee Component production, 1.5 million square feet Wichita Citation Service Center 477,000 square feet – one of the largest buildings in Kansas Room for some 100 Citations for service, w/5,270 service orders in 2007 Independence, Kan. SEP and Mustang assembly, Opened in 1996, 528,000 square feet Columbus, Ga. Component production, Opened in 1996, 241,500 square feet Chihuahua, Mexico Component Production, Opened in 2006, 206,000 square feet Citation Service Centers Nine Cessna-owned Citation Service Centers or participating authorized facilities around the world. Customer Support for piston products Propeller-driven Cessnas are supported through a worldwide network of more than 350 independent service facilities operating with high standards of customer service. Cessna Pilot Centers CPC network consists of 264 domestic and 13 international affiliates. CPC has been the leader in flight training since 1973. Since 2000, CPCs have delivered private and instrument pilot training to more than 100,000 pilots around the world. CitationAir Offers the full range of transportation solutions, ranging from Jet Cards, Jet Shares (fractional), Jet Management (turn-key, cost- effective way to optimize aircraft ownership with guaranteed use of the aircraft to support or back-up the CitationAir fleet and revenue generation for the aircraft owners), Corporate Solutions (fleet access without whole aircraft ownership) 4 M&P Engineering - NDE ADVANCED TECHNOLOGY - AD, Experimental, Manufacturing and process sensing support Support process specifications Evaluate new methods of testing & inspection Support implementation of new techniques/technologies NDE SUSTAINING - Design review, DADT support, test article support Engineering drawing review Define NDI requirements for service-related inspections Inspection support for component test articles PRODUCTION OPERATIONS – Support for NDI methods at all facilities 5 Citation NDI Certification / Licensing Citation NDI Certification Program Established in 1985 to ensure independent facilities/personnel meet Cessna NDI requirements (reflects NDT industry standards). Provides Citation owners/operators a means to confirm a facilities NDI qualifications. 103 Facilities Worldwide. 61 Independent Domestic Facilities 34 International Facilities 8 Service Centers NDI License Program Established May 2001 wherein Cessna owns all Citation NDI calibrations standards. Cessna provides control over the use and distribution of the standards. Preserves intent of the Certification Program. Ensures only approved NDI facilities are performing inspections. Ensures standards are controlled, certified and documented. Eliminates expensive acquisition costs for customers. Pay on a per use basis - easy to understand ROI. Mitigates Quality Control issues and lead times. 6 NDE Methods Liquid Penetrant Magnetic Particle Radiography Eddy Current (arrays) Ultrasonic (Squirter, Air-coupled, Manual, PAUT) Six 2.5D UT squirter gantries (up to 50’) Air-coupled UT – 8 ch high speed flat scanner (Prod) & 2 ch (Expr Shearography (6x6x20’ vacuum chamber w/heat excitation & robot scanner) Scanners GenScan free-scan system with Mimeo feedback 2-axis portable scanner for UT & ET MOI Thermography (PE & TTU) Acoustic Emission (16 ch) 7 Cessna Gantry UT Systems Gantries w/up to four TTU nozzle pairs and 28 ips 8 Cessna Air-coupled UT System Reduced inspection time Well-suited for composite or metal bond Inspected over 2M sq. ft. of bonded assemblies since installed in ’00. 9 Projects Slight Hidden Corrosion in thin metal bonded structures RFEC, ET array, phased array UT Multi-Layer Cracking – up to 4 layers 1.25” thick Pulsed Eddy Current, RFEC, MS & GMR Angle beam UT scanning w/portable scanner & instrument Physical Test Monitoring Acoustic Emission Metal Bond Process Monitoring PAA/Bond Primer Thickness Automated anodize trace analysis Semi-automated contamination detection system Residual stress after cold-working NDE Modeling Technique optimization 10 EC Modeling Model Part Probe Flaw Output (b) MEC03 Plate/Halfspace Air-core Complex (a) B,C Single crack, PLATE07 Multi-layers General A,B,C tight/open Single crack, BEM07 General General A,B,C tight/open (a) Flaw type: “Complex” = multi-ligament cracking, open or tight. (b) Output options: A = incident EC fields, B = part geometry signal, C = flaw signal Iowa State University CNDE 0.050” 0.050” 0.060” 0.100” 0.004” gap 0.250” 0.1 mm gap 0.250” 0.150” 11 UT Modeling Model can predict voltage of the defect echo, given info about the metal (density, velocity, surface), defect (size, shape, location) and system (waterpath, probe characteristics, reference echo, etc.). Diffraction, attenuation, transmission/ reflection coefficients, near/far fields, freq. dependencies, focused or flat probes, lenses & mode conversions SNR estimates can be made if noise properties of the microstructure are known (UT scatterer model for grain noise) 12 X-ray Modeling Simulation can predict radiographic density of various materials, flaw composition, flaw geometry POD estimation for 3D components with multiple shot orientations. 13 What is SHM Structural Health Monitoring is the continuous monitoring of structures/components using integrated or applied sensors. Aimed at assuring structural integrity of the aircraft, replacing on- event and periodic inspections to detect damages resulting from fatigue, corrosion, excessive loads, impact ... Monitoring of structures does not necessarily mean knowing the status of the structure in real-time. Structures are designed with acceptable margins such that, after normal or exceptional events, maintenance tasks can be planned at next appropriate inspection. Systems are available for aircraft condition monitoring - mostly for loads (accelerations, flight parameters, etc.) and enable decisions to be made based on actual flight load levels. Indirect surveillance of the structure is not comprehensive or reliable enough to avoid interval inspections. 14 Conventional vs. Condition Based Management Currently, NDT is applied starting with visual inspections followed by for more subtle or hidden flaws, procedures are defined based on eddy currents, ultrasonics, x-rays, etc… Inspection intervals are usually based on knowledge of the structure residual strength, operating environment, applied loads, damage growth rate and failure consequences. Of course, inspections result in downtime and inaccessible areas of structure often require significant effort to remove equipment or strip protective coatings for access, which then must be restored after the inspection. Monitoring activity comes at a considerable cost and accounts for an average of 44% of all on-aircraft maintenance man-hours for commercial aircraft (Andresen, 2006). In terms of life cycle cost, a US DoD study attributed 27% of the total cost of an aircraft being maintenance related with structural inspection being a significant driver of this cost (Kudva, et.al. 1999) suggesting that SHM could save up to 44% of current inspection time on modern fighter aircraft. Ultimate concept imitates the human nervous system, though SHM will better since structures are monitoring directly, measuring the effect of damage. Compared to conventional NDT, SHM has many advantages: No access to the inspection area necessary – fewer access panels & component removal requirements No physical operation in the area - safe inspection of hazardous areas No use of scanners necessary – eliminating time consuming setup Sensors used in the inspection are integral to the structure Automated process - no human factors influence on inspection POD Interrogating many locations or wide field at once - significant time saving H. Speckmann, Materials & Processes - Testing Technology, Airbus 15 CBM Approach – SHM Potential Time spent inspecting the structure to assure continued airworthiness increases as aircraft age. To allow for statistical variation of the real life of the structure, a safety factor is applied to the demonstrated lives of components. To reduce the inspection burden, some industries have introduced automated on-line structural health monitoring systems with maintenance only being carried out when the health of the structure indicates a need for it. CBM approach to maintenance if applied to aerospace structures has the potential to not only reduce the time spent inspecting these structures, but also improve airworthiness by detecting damage at an earlier stage than possible during discrete periodic inspections. For safe life components it would also be possible to detect early failures and withdraw them from service ahead of their expected life, or for healthy components continue to use them beyond the design life and only withdraw them when their health indicates a requirement to do so. Potential to both improve airworthiness and gain economic benefits was originally conceived within the context of a rotorcraft Health and Usage Monitoring System (HUMS) but is applicable to any health or usage monitoring system. T. Ewbank, Cranfield University, Application Of Condition Based Maintenance On Aerospace Structures 16 SHM Applications • Difficult to access inspections • Hot spot monitoring 17 Desirable SHM Attributes • Impact damage in composite • Wireless/passive sensing • Appropriate for in-situ embedded or attached robust sensors during component manufacture • Cost effective, lightweight sensors (optical, acoustic, electromagnetic, etc.) Acoustic Emission Acellent Technologies 18 SHM Outcomes Physics-based material properties measurements to determine material state throughout life cycle – allow design conservativeness to be minimized Intelligent structures (self-diagnostic, self-healing, health monitoring & diagnostics, manufacturability w/sensors, etc.) Leading to design optimization, weight saving, less fuel consumption & environmental impact 19 Challenges to SHM Develop and demonstrate SHM technologies that can be used to monitor structural integrity in service conditions with high reliability & durability. As in conventional NDT, a single technology will not be suited for the entire range of applications, based on different materials, component geometries and damage scenarios. Diagnosis must have high reliability over the aircraft lifetime, since un-justified maintenance actions are quite costly to the operator and spurious warnings degrades confidence in the system. Accuracy and reliability may even be more stringent, since further optimization of structural design will rely on SHM with better knowledge of actual flight loads & condition. 20 System Qualification Components shall qualified, as part of aircraft certification, meaning they shall perform the specified function while withstanding the specified environmental conditions. Include large variations of temperature, vibrations, impacts, Electro- Magnetic Hazards, chemical fluids, etc… as per RTCA DO 160 and aircraft integrator directives. Components qualification shall demonstrate that the system performs reliably in the specified environmental conditions, in all the aircraft operational conditions over it’s lifetime. Specific issues need to be considered particularly: Easy installation and application on surfaces Accuracy and reliability when used on painted surfaces No corrosive damage to surfaces where applied No delaminating between sensor and monitored structure Suitable for metal, composite, sandwich structures Suitable for various damage: cracks, corrosion, delamination, de-bonding… Clearly different requirements will apply onto the system components: sensors, processors, computers, wiring, power supply, … depending on the technologies, architecture and installation location retained. Micro-Nano-Technologies have the potential for supporting qualification requirements and the SHM business case. H. Speckmann, Materials & Processes - Testing Technology, Airbus 21 Implementing SHM SHM is not a new concept - it is already implemented on military aircraft, with a different rationale but some converging features. Still the constraints of airworthiness certification and the existing cost/benefit have limited its introduction in commercial aviation There is currently no specific FAA or EASA policy on CBM for civil aircraft. However, some guidance is provided in FAR-29 (FAA, 2003) on achieving maintenance or airworthiness credits with HUMS that could be developed. US DoD stated the requirement to transition to a CBM program by the end of 2015. Confidence needs to be built that SHM will bring the expected benefits, while maintaining or improving the safety and efficiency of modern aircraft – by progressive introduction and proving the reliability & benefits. 1st generation of SHM shall target maintenance cost reduction and increased aircraft availability - technology will allow saving cost and time in regulatory inspections. 2nd and 3rd generations of SHM shall integrate a new certifiable design philosophy and will permit weight reduction. Sensors and their local processors would be more integrated with microelectronics allowing more decentralized architecture where local processors perform & record the first level of SHM processing until transmission to the upper level processor. H. Speckmann, Materials & Processes - Testing Technology, Airbus 22 USAF Experiences LAHMP Health Monitoring System F-15 Flight Tests In 2003, the Army awarded an SBIR Phase II contract to TRI/Austin to develop a diagnostic/prognostic system that could monitor aircraft and rotorcraft structural components in flight. focused on ruggedizing the system, optimizing performance, reducing power draw, refining the prognostic/diagnostic algorithms and building a system for test. successfully conducted third-party independent testing included acceleration testing of up to 6G on 6 axes as well as RFI/EMI testing. In-house thermal testing showed the system to be operational in the specified range of -40 to 85C, including thermal shock. effort culminated in a successful flight test of the LAHMP system acquiring data from three areas on the F-15. developed patented algorithms to determine structural health from on board sensor readings. In addition, we designed the health management platform to be fully customizable for a wide variety of aircraft." JSF program has CBM features within the aircraft’s design with rudimentary corrosion sensors installed and strain gauge monitoring of loads on a limited number of aircraft loads are then coupled with flight data monitoring to allow parametric usage monitoring as a tool for CBM across the entire fleet using maneuver recognition algorithms to determine the loads on aircraft that are not monitored (Reed, JSF CBM Features, 2007). 23 Current Technology State Of various fatigue damage detection technologies being researched to enable CBM on aerospace structures, Comparative Vacuum Monitoring & acoustic emission are most mature and are currently marketed as commercial structural health monitoring solutions – however both have application limitations CVM has the capability to detect the presence, location and extent of damage, which when combined with a usage monitoring and a prognostic system could provide a full CBM capability. However, at present CVM falls into a grey area between on-line structural health monitoring and NDT as the sensors are permanently installed but the vacuum and flow detector are only connected on the ground for off-line damage assessment. Given the simplicity of this process it still offers significant advantages, especially for inaccessible structure. However, it is limited to areas of structure where the damage mechanism is well understood and predictable (localized damage detection) Nevertheless, it is a elegantly simple concept that is gaining mainstream acceptance from aircraft OEMs and operators. Corrosion sensing technology is generally crude - moisture detectors that could be placed in areas of corrosion prone structure are under development. However, apart from using UT to measure the reduction in plate thickness, this and all the present techniques give an indication of the probability of corrosion on the structure that must be verified by visual inspection and to quantify its extent. Of the currently developing damage detection technologies, guided waves and electrical impedance measurement appear to have promise but need to be tested in realistic structures under environmental conditions. Fiber Bragg Gratings are also promising for increased structural coverage with minimal calibration requirements. 24 Comparative Vacuum Monitoring (CVM) CVM sensors work on the basis of differential pressure - pressure changes in a system of small capillaries provide an indication of structural defects (cracks, corrosion and loss of bonding contact). Each sensor, which is ~ 125 mm thick, is perforated with fine galleries alternately containing air and a vacuum. The presence of a crack or other defect in the monitored material creates a connection between the two types of gallery, altering the distribution of pressure inside the sensor at this point. Used by Airbus in acceptance testing of GLARE (GLAss-fibre REinforced aluminum) composites - a laminate consisting of three layers of Al held together by intermediate layers of glass-fiber reinforced epoxy resin (A380 upper skins) Since CVM sensors were permanently attached, inspection was done in a fraction of the time required by conventional testing methods each measurement could be reproduced under exactly the same conditions onerous task of installing and removing sensors only had to be carried out once, which saved a great deal of time, especially in the less easily accessible parts of the airframe To monitor cracking from the fastener holes, CVM sensors were positioned inside the lap joint before riveting began able to detect cracks of a magnitude of 1-2 mm – cracks that most other test methods were incapable of detecting H. Speckmann, Materials & Processes - Testing Technology, Airbus 25 CVM Sensitivity & Durability Sandia led project to mount a series of 26 sensors on structure in four different DC-9, 757, and 767 aircraft in NWA and Delta fleet Periodic testing is being used to study the long-term operation of the sensors in actual operating environments compliments lab flaw detection as part of an overall CVM certification effort. In conjunction with Boeing, Northwest Airlines, Delta Airlines, Structural Monitoring Systems, the University of Arizona, and the FAA, validation testing conducted on the CVM system in an effort to adopt Comparative Vacuum Monitoring as a standard NDI practice. Fatigue tests conducted on simulated aircraft panels to grow cracks in riveted specimens while the vacuum pressure within the various sensor galleries are simultaneously recorded. Crack is propagated until it engages, and fractures, one of the vacuum galleries such that crack detection is achieved (sensor indicates the presence of a crack by its inability to maintain a vacuum). In order to properly consider the effects of crack closure in an unloaded condition (i.e. during sensor monitoring), a crack was deemed to be detected when a permanent alarm was produced and the CVM sensor did not maintain a vacuum even if the stress was reduced to zero. Unpainted 0.040" Skin 0.040" Skin w/Primer 0.002-0.030” long cracks 0.002-0.010” long cracks 26 Acoustic Emission An arbitrary mechanical excitation applied to a plate will generate a multiplicity of Lamb waves carrying energy across a range of frequencies - such is the case for the AE wave. The challenge is to recognize the multiple Lamb wave components in the received waveform and to interpret them in terms of source motion. This contrasts with the situation in UT, where the first challenge is to generate a single, well-controlled Lamb wave mode at a single frequency. But even in UT, mode conversion takes place when the generated Lamb wave interacts with flaws, so the interpretation of reflected signals compounded from multiple modes becomes a means of flaw characterization AE’s fundamental limitation is the ability to only detect the growth of damage and not reliably give a measure of its extent, which makes the assessment of current and future load carrying capability impossible to reliably determine. Consequently, once damage is detected a second method is required to confirm and assess the extent of the damage, which with current available technologies will require manual NDT. this may still yield benefits but will need to be combined with another damage detection technology such as guided waves to be a truly on-line CBM enabling technology. Many applications have struggled because it is difficult to determine exactly the position of an indication. Also the false call rate and POD can be problematic 27 AE Applications Airbus A320 fatigue test certification of inner wing (Staszewski, et al., 2003), and the monitoring of the A340 Landing Gear Support Structure during the full scale fatigue test of the A340-600 (Lloyd, et al., 2003). In both cases the AE system was used to identify the presence and source of damage with conventional NDT being used to both confirm and quantify the extent of the damage. The A340 test was conducted with Ultra Electronics BALRUE system, which used 24 narrow bandwidth ceramic AE sensors at 300 kHz. During the one year trial, all damage detected by conventional NDT was also detected with the AE system. Furthermore, several damage sites were detected by AE before being found by conventional NDT techniques. This system has now been modified and qualified for airborne use and is now marketed by Ultra Electronics as the AAIMS as an additional tool for SHM. Apparently only one operator implemented, on one P-3 fire fighting aircraft (Aero Union) 12 similar sensors were installed on the front spar to monitor the spar structure between the fuselage and inboard engine. Data collected by these sensors was then stored together with 28 several other flight parameters such as spar cap strain, indicated airspeed, tank volume and vertical acceleration to a Data Acquisition Unit (DAU). The equipment was installed on the P-3 aircraft during depot level maintenance and after just 47 flying hours the analysis showed emissions. Further analysis of these results showed consistent crack growth in several areas, which were then confirmed by conventional NDT. A 12 mm crack on the lower spar cap, almost certainly present during maintenance and not detected by the NDT carried out, was found to correlate with the damage. It was located ~ 20 cm away from the true site of the damage and identified the need to use an alternative technique for damage location in realistic structures. Approach requires the characterization of an installation by inserting 300 kHz signals into the structure in known locations during the system’s installation, and then using signature recognition and a 3D model of the structure in the analysis, which has improved the positional accuracy to within 2 cm. 28 AE Application on MLG Fitting For the Tornado GR4 retraction jack fitting, the only technically viable damage detection method to provide an on-line condition monitoring capability was considered to be acoustic emission. With CVM, flanges on the bushings may preclude damage detection and the movement of the retraction jack was likely to damage the sensors. Given the limitations of CVM for wide area unpredictable damage mechanisms fatigue damage may occur in virtually any location of the retraction jack fitting and take a number of directions. Over a 10 year period the labor savings of introducing an AE on-line condition monitoring capability for the Tornado GR4 retraction jack fitting are estimated to be approximately £800k in NPV terms. estimated to increase aircraft availability by 61 aircraft days per year across the fleet. When acquisition and design incorporation costs of an AE system are taken into account to provide an on-line condition monitoring it is unlikely that the system would be cost effective in terms of labor savings alone. However, the increased availability may yield sufficient savings to make the system viable and was recommended to be further investigated. T. Ewbank, Cranfield University, Application Of Condition Based Maintenance On Aerospace Structures 29 PZT Network System using Lamb Waves Acellent Technologies uses built-in network of piezoelectric transducers embedded in a thin dielectric carrier film. system includes the PZT network connected to portable, diagnostic hardware and software. Performs in-situ monitoring, data collection, signal processing, and real-time data interpretation to produce a two-dimensional image of the structure being interrogated. Software controls the actuators to generate pre-selected diagnostic signals and transmit them to neighboring sensors. wave types including 3, 5, and 10-peak narrow band frequency waveforms, chirp, random, and user defined excitations Software links each sensor with its neighbors to form a web, or network, covering the area of interest and collects responses from each of the sensor sets as each PZT is activated. Changes in Lamb waves generated within the structure are used with triangulation methods to detect the presence of anomalies and to determine size & location. 30 PZT tests on Boron Epoxy patch Similar to conventional UT, PZT data analysis can include one or more of the following measurements: Time of wave transit (or delay), path length, frequency, phase, amplitude and angle of wave deflection (reflection & refraction) A series of excitation frequencies were used to optimize detection: 50 kHz, 200 kHz, 350 kHz, and 500 kHz. Results revealed that disbond flaws were most strongly detected with 50 kHz, while the crack growth was monitored best with the highest 500 kHz excitation Signal attenuation, corresponding to disbonds between the patch and metal skin were apparent Both flaws from one complete disbonded due to a Teflon insert, and a weak bond produced by a mold release agent D. Roach, Sandia Labs, HEALTH MONITORING OF AIRCRAFT STRUCTURES USING DISTRIBUTED SENSOR SYSTEMS 31 Lamb Wave Applications A wingbox was tested (Grondel, Assaad, Delebarre, & Moulin, 2004) with delamination of the plate sections of the composite structure and disbonds, with stringers being readily detected using amplitude analysis. However, one key conclusion drawn was the need to identify the Lamb wave propagation modes possible in the structure at the frequency used In this case there were 4 modes at the 400 kHz transducer frequency with wavelengths ranging from 3.75 mm to 15 mm. Higher frequency modes were very sensitive to damage, whereas the modes with the longer wavelengths were relatively insensitive. 1. Use a frequency region where only fundamental propagation modes exist. 2. Chose a propagation mode and frequency where dispersion (wave velocity is a function of frequency and thickness of the plate) is kept to a minimum in order to simplify signal analysis. Damage Localization in a Stiffened Composite Panel (D. Chetwynd, et.al. University of Sheffield) work conducted as part of the Aircraft Reliability Through Intelligent Materials Application (ARTIMA) EU project. Case study of damage detection in a curved carbon-fiber reinforced panel with two omega stiffeners investigated using UT Lamb waves. Outlier statistical analysis was used as a way of pre-processing data prior to damage classification. Multilayer perceptron neural networks were used for classification and regression problems of damage detection. It was then investigated whether using wavelet analysis to perform prior wavelet decompositions of experimental data could facilitate damage classification. 32 Fiber Bragg Gratings (FBG) Fiber-optic sensors with elastic properties similar to those of the tested material Can be used to monitor temperature, thermal and mechanical stress, damage caused by collision or impact, and delamination. Fiber-optic sensors operate in similar manner as strain gauges. As the material under test expands due to the effect of temperature or mechanical forces, properties of the sensor fibers vary in an easily measured way. In a strain gauge, the electrical resistance varies in proportion to it’s distension In fiber Bragg gratings, the characteristics of the reflected light change based on the position of tiny mirrors that make up the Bragg grating and with which the optical fiber is doped using a laser technique. Up to 25 measurement points can be integrated in a small single fiber To achieve the same number of measurement points using a strain gauge, it would be necessary to lay 25 thick multi-wire cables – a hardware density that is already too high to be of practical use in a test configuration. As a result, the network of measurement points used in conventional testing is correspondingly widely spaced; FBGs would allow more closely spaced data resolution. Quality of data would also be enhanced, using far less elaborate means Unlike more conventional types of sensors, fiber-optic sensors are not subject to interference by EM fields, Cross-section of embedded fiber so do not require elaborate shielding. 33 FBG Applications Betz, Staszewski, Thursby, & Culshaw; Structural Damage Identification Using Multifunctional Bragg Grating Sensors: Damage Detection Results and Analysis, 2006). Work attempted to determine sensitivity to temperature variation and, most significantly, give an indication of damage detection sensitivity. Included the use of two driving frequencies at 260 kHz and 460 kHz to explore the effect of frequency on damage detection. Also involved the use of several analytical tools and the use of both piezoelectric and fiber optic Bragg Grating (FBG) sensors for the ultrasonics. Results suggest that analysis of amplitude and the propagation period of the first two Lamb wave packets received gave the best damage detection correlation with the ability to discriminate between damage sizes of 0.8, 1.4 and 2 mm but for damage greater than 15 mm in size, a saturation effect was observed. When using the analysis of amplitude and the propagation period of the first two Lamb wave packets both the piezoelectric and FBG sensors worked equally well, and it was possible to correlate damage size with both of the driving frequencies used. Although analysis did show some sensitivity to changes in temperature, these effects were very minor and therefore well suited to further processing. 34 Acousto-Ultrasonics (AU) Technique that sends acoustic waves into the structure and intercepts them when they emerge on the other side. Deviations from the expected wave pattern indicate the presence of cracks or delamination. 35 Eddy Current Eddy Current Testing Foil Sensors (ETFS) are suitable for use on metallic structures. Cracks and corrosion alter the electromagnetic field induced by the eddy current generated by the sensor (flexible). Micro Eddy Current Sensor - Sandia Labs is developing a customized eddy current sensor for crack detection in thick steel structure. The probe must be able to detect deep, second-layer cracks as much as 0.5” below the surface. Impedance bridge and other differential circuits were explored to maximize the magnetic flux density and corresponding eddy current strength. Successful crack detection was achieved with a dual coil configuration that combines a pancake excitation inductor with a co-located pickup coil to produce a transducer that requires very little drive current (75 mA) and operates in the desired 10 kHz range. Excellent crack detection was achieved even when inspecting through composite repair doublers approaching 0.5” thickness. More sophisticated & rugged electronics package – including digital signal processing to filter and detect phase shifts - to further improve probe sensitivity were being developed. GE Inspection Technologies 36 In-situ Sensors & Durability Advantage/disadvantage compared to established NDT techniques is that future sensors remain permanently attached / embedded in place, required to withstand many decades of aircraft service life mechanical structural stresses and ensure sensor performance & substrate bonding hot/cold and wet conditions Embedded sensors can integrate well with composite materials - piezoelectric fibers or fiber-optic sensors can be fabricated with CFRP or GFRP (mitigating risks of sensor debonding) however two difficulties arise if component is replaced due of wear or damage, the embedded sensor is also maintaining the sensor is difficult, virtually impossible to repair, and not replaceable Once installed, sensors provide a simple means of monitoring even areas that are difficult or dangerous for inspectors to access such as fuel tanks, wing spars, engine beams, etc. – areas that are difficult to detect microscopic cracks or corrosion. Detection might be in online (measured continuously in flight) or offline (data downloaded at next inspection or maintenance) modes. In-situ SHM sensors would be capable of spotting defects much faster, leading to considerably shorter inspection times. Still, SHM should not entirely replace conventional NDT inspection practices Conventional maintenance checks are rarely limited to the precise area specified in the maintenance manual - maintenance technicians typically take a careful look at surrounding area outside the actual inspection range. 37 Signal Analysis & Power Mgt Processor capabilities now are not limiting damage detection signal processing, but weight / power are always concerns. Data must be managed in a robust system to ensure value and translated into useful structural health and usage information. DO-178B provide some structured development guidelines, which is recognized by FAA & EASA Many published works recognize patterns from known damage rather than identify unknown damage. Parametric recognition systems are beginning to appear (UK published policy in ’07 to guide development for future military aircraft). Parallel processing appears to be mature for some limited applications such as maneuver recognition in usage monitoring, but not prevalent in damage detection. Energy harvesting methods have been studied that can power sensors systems by converting structural stresses (strain energy harvesting) into electrical power via piezoelectric transducers (Sandia Labs, Kansas State, etc.) 38 Directions Aircraft structural design optimization is the ultimate benefit of SHM Structural maintenance inspections are a significant factor in operator’s Direct Operating Costs. Minimize conventional NDI for periodic interval inspections for airworthiness and due to unusual events (hard landing, impact, lightning strike, etc.) since structural health data is available. Inspection intervals are calculated conservatively based on fatigue and corrosion growth models. SHM will allow optimizing these assumptions with actual aircraft flight data. SHM sensors can greatly simplify inspections since affixed permanently, can be activated quickly and reliably (without operator variability) Immediate benefit in faster and less costly inspection No difficult or dirty access to critical inspection zones 39 Conclusions Some technologies are well ahead of others in terms of development maturity – reliability, robustness, durability and data analysis are still key issues. Future innovative approaches are being developed in microelectronics, nanomaterials, MEMS, etc. which will improve effectiveness and costs. 40