Nondestructive inspection and structural health monitoring of aerospace materials
The high quality of materials used in aircraft structural components is essential for reliability and safety. Structural materials must be free from defects that would reduce the structural properties, failure strength and design life. However, despite careful control of the operations used in processing of aerospace materials, it is difficult to ensure that every structure is completely free from defects. For example, brittle intermetallic inclusions and gas holes form in the casting of metal components and voids and dry spots occur in the manufacture of fibre–polymer composite structures. Damage also occurs during in-service operation of aircraft from poor design, mechanical damage (e.g. impact, fatigue, battle damage) and environmental degradation (e.g. corrosion, moisture absorption, lightning strikes). Tables 23.1 and 23.2 list common manufacturing defects and in-service damage found in metal and fibre–polymer composite components.
|Defect||Description of damage|
|Porosity||Pocket of air or gas trapped in solid material during solidification|
|Intermetallic inclusion||Hard, brittle particle formed during casting or heat treatment|
|Shrinkage crack||Cracks formed by shrinkage of metal casting|
|Cracks, scratches||Defects caused by machining and drilling operations|
|Fatigue||Cracks formed owing to stress, thermal or acoustic fatigue|
|Corrosion||Material degradation (eg. pitting, cracking) owing to corrosion|
|Defect||Description of damage|
|Delamination||Separation of ply layers in laminate by cure stresses|
|Matrix crack||Crack within polymer matrix owing to cure stresses|
|Unbond||Area between prepreg layers that fail to bond together|
|Foreign object||Inclusion of a foreign object (e.g. peel ply)|
|Porosity||Pocket of air or gas trapped during lay-up or cure|
|Core crush||Crushing of core owing to excessive cure pressure|
|Impact damage||Delamination, matrix cracks and fibre breaks owing to impact loading|
|Moisture absorption||Softening and cracking owing to moisture absorption|
|Fatigue||Delamination, matrix cracks and fibre breaks owing to stress, thermal or acoustic fatigue|
|Core degradation||Corrosion of aluminium core and softening of Nomex core owing to moisture absorption|
Aviation safety authorities such as the Federal Aviation Administration apply strict regulations to the types and amount of damage allowed in structural materials without the replacement or repair of the damaged component. The regulations require that aircraft structures and engine components have a minimum level of damage tolerance, which means that the part must retain functionality and design structural properties when damage is present in the material. In commercial aviation, the maximum allowable defect (delamination) size for composites is typically 12.5 mm and the maximum porosity content is about 1.0% by volume. For military aircraft, which usually operate at higher stress levels than civil aircraft, the maximum allowable defect size is generally smaller, depending on the design and function of the structure. Maximum allowable defect sizes are also specified for metal aircraft components. Metal and composite structures containing damage greater than the allowable limits must be repaired or taken out of service.
Most damage is impossible to detect by eye because it is too small and buried below the component surface. The aerospace industry cannot rely solely on visual examination to determine material quality. Instead, the industry is reliant on nondestructive inspection (NDI), also called nondestructive testing (NDT) and nondestructive evaluation (NDE), to assess material integrity. As the name implies, NDI does not inflict damage on the component during the inspection process (as opposed to other methods which require destruction of the component to detect the damage).
Aviation safety regulations require the aerospace manufacturing companies to nondestructively inspect all primary structures before being built into the aircraft. Structures such as wing sections, fuselage panels, control surfaces and landing gear components are thoroughly inspected using NDI methods after fabrication to ensure they are free of defects. Airline companies and other aircraft operators (including the military) must also undertake regular NDI throughout the operating life of their aircraft. Aircraft and helicopters are taken out of service for routine NDI examinations to ensure the structures are damage-free.
Whether for the inspection of as-manufactured or in-service components, NDT is used to determine the type, size and location of damage which is essential information to assess the flight-worthiness and residual strength of the structure. There are various NDI methods, and the types most often used by the aerospace industry are listed in Table 23.3. The methods regularly used are: ultrasonics, dye penetrant, magnetic particle, radiography, thermography and eddy current. Quite often two or more methods must be used in combination to obtain a complete description of the type, size and location of internal damage. The types of damage that can be detected in metals and composites with the various methods are given in Table 23.3.
The NDI of aircraft is often an expensive, labour-intensive and slow process. The inspection of as-manufactured aircraft components adds considerably to the production time and cost (by as much as 50%). Inspection of in-service aircraft requires grounding and downtime, which affects the profitability of airlines and the operational capacity of airforces. For these reasons, the aerospace industry is assessing the use of structural health monitoring (SHM) to detect defects and damage. SHM systems use in situ sensor networks and intelligent data processing for the continuous inspection of aircraft structures. The sensor networks are embedded or surface-mounted on the aircraft, and can provide information on the presence of damage during manufacture or in-service operation with little or no human intervention. The concept of SHM is about determining the condition of a structure in real-time with structurally integrated sensing equipment. This way, corrective maintenance can be performed when required, rather than at intervals based on flight times. Many SHM technologies are currently being developed and evaluated by the aerospace industry, with the most promising including Bragg grating optical-fibre sensors, piezoelectric transducers, and comparative vacuum monitoring.
In this chapter, we examine the NDI techniques used to detect manufacturing defects and in-service damage in metals and composites. We study the operating principles of the techniques and learn about their capabilities and limitations. In addition, there is an introduction to SHM for aircraft and several SHM techniques that are emerging as damage-detection methods for aircraft are discussed.
Two of the simplest inspection methods are visual inspection and tap testing. Visual inspection involves the careful examination of the material surface with the eyes, often assisted with a magnifying glass. Although simple, visual inspection is the first step in any NDI method, and it can identify obvious signs of damage. However, visual inspection is only suitable for detecting surface damage such as large cracks or general or exfoliation corrosion in metals, and it is not suitable when the damage is buried below the surface of metal or composite components.
Tap testing involves the repeated tapping over the component surface using a coin, soft hammer or some other light object to produce a ringing noise (Fig. 23.1). Damage immediately below the surface may be detected using tap testing by a change in the pitch of the noise. For example, tapping the surface of damaged composite material containing large delamination cracks can produce a dull sound compared with the higher-pitch ringing noise of the damage-free material. However, tap testing is not always reliable and can easily fail to detect damage. Instrumented tap testing devices which measure and analyse the noise generated by the tapping are available to eliminate the need for human hearing, which is not always sensitive to small changes in pitch. Both visual inspection and tap testing can be used for the initial inspection of aircraft components, but more sophisticated NDI methods are needed for reliable inspections.
Ultrasonics is an NDI method used widely in the aerospace industry to inspect aircraft structures and engine components. Ultrasonics is used for the detection of both manufacturing defects and in-service damage. Although ultrasonics cannot detect every type of damage, it can determine the presence of common types of damage found in metals (e.g. voids, corrosion damage, fatigue cracks) and composites (e.g. delamination, porosity).
The operating principle of ultrasonics is shown schematically in Fig. 23.2. The method involves the transmission of ultrasonic pulses generated by a piezoelectric transducer through the material. The pulses are high frequency (typically 1 to 15 MHz) compressive or shear elastic waves. When the waves encounter a region with an acoustic impedance value different from the host material, such as cracks or voids, then they are reflected and scattered. The characteristic acoustic impedance (Z) of a medium, such as air, metal or composite, is a material property:
where ρ is the density of the medium and c is the longitudinal wave speed in the medium. The acoustic impedance values for the main types of aerospace materials and air (which is the typical value for a crack) are given in Table 23.4. A large difference in the acoustic impedance value results in the loss in acoustic intensity owing to reflection and scattering, which is called attenuation and is measured using a receiving transducer. The received signal is analysed to determine the location and size of defects and damage.
|Medium||Acoustic impedance (Pa s m−1)|
|Aluminium||17 × 106|
|Titanium||10 × 106|
|Nickel||54 × 106|
|Steel||45 × 106|
|Carbon–epoxy||9 × 106|
Ultrasonics is operated in two basic modes: pulse–echo (or back-reflection) and through-transmission (Fig. 23.3). The pulse–echo mode involves the use of a single transducer located at one side of the material to radiate and receive the acoustic waves. When a defect is blocking the wave path then part of the acoustic energy is reflected back to the transducer. The reflected acoustic wave is transformed into an electrical signal by the transducer and is displayed on an oscilloscope. Pulse–echo ultrasonics can accurately measure the size and depth of damage in metallic and composite materials. Through-transmission ultrasonics involves using one transducer to generate the waves and another transducer located on the other side of the material to receive the signal. The waves generated by the transmitting transducer propagate through the material. When the waves encounter a defect with an acoustic impedance value different from the host material they are scattered and back-reflected, which attenuates the transmitted wave signal. The receiving transducer records a weakened signal owing to blocking of the acoustic waves by the damage, which is used to indicate its presence.
Pulse–echo ultrasonics is the preferred method for the inspection of inservice aircraft because the equipment is portable for field use and only one-side access is required (no need to remove aircraft components to gain access to both sides). Through-transmission ultrasonics is used more for the inspection of as-manufactured components before they are assembled into the aircraft. This mode of inspection is faster and more easily automated than pulse–echo ultrasonics. Furthermore, ultrasonics can generate a two- or three-dimensional image (called a C-scan) of damage inside the material by measuring the time-of-flight of reflected acoustic waves. Figure 23.4 shows a C-scan image of an aerospace composite material, with the bright zone revealing internal delamination damage, caused by a low-energy impact event, that cannot be observed visually. Other ultrasonic methods are used occasionally by the aerospace industry, such as Lamb waves and laser ultrasonics, although their application is less common than pulse–echo and through-transmission ultrasonics. Ultrasonics is best suited for the detection of planar damage such as cracks aligned parallel with the surface. The technique is not well suited to detecting damage aligned parallel with the propagation direction of the acoustic waves, although angled probes can be used.
Radiography involves the use of radiation, such as x-rays, γ-rays or highspeed neutrons, to detect damage in solids (Fig. 23.5). Radiation is emitted from an energetic source, such as an x-ray tube, and directed to the test component. The radiation energy is absorbed during its passage through the material. However, the absorption rate changes when the radiation passes through a defective region having different absorption properties to the host material. The absorption value for an air gap is much lower than the aerospace materials and, therefore, less energy is absorbed during the passage of radiation through cracks and voids.
The radiation intensity is measured using x-ray film after passing out of the material. Regions of high-intensity radiation (owing to the presence of damage) and low-intensity radiation (pristine material) appear different on the x-ray image (as shown in Fig. 23.6). The size and shape of the defect is measured from the image. Radiography can detect defects such as voids, intermetallic inclusions, corrosion damage, and cracks larger than ~ 0.5–1.25 mm, which is below the critical damage size in aircraft structures. However, the shape and orientation of the defect affect how easily it is detected. Long cracks aligned parallel with the direction of radiation are more easily detected than cracks running perpendicular to the incident radiation. Therefore, it is necessary to inspect a component with different radiation angles to ensure cracks at different orientations are detected.
Thermography is used by the aerospace industry for the rapid, wide-area inspection of components. There are two main types of infrared thermography known as passive and active. Active thermography is the more widely applied of the two, and is shown in Fig. 23.7. The active method involves short duration heating (usually less than 1 s) of the component surface using flash tubes, hot-air guns or some other controllable heating device. The component surface is heated 10–20 °C above ambient temperature, with the heat being absorbed into the material. The method measures the heat dissipated from the heated surface as the component cools. The amount of heat dissipated depends on the thermal properties of the material together with the type, size and location of damage. Any defect which creates an air gap such as a delamination, void or corrosion cavity absorbs less heat than the parent material. Consequently, more heat is dissipated from the surface above the defective region. The damage is then observed using an infra-red (IR) camera as a ‘hot spot’ on the surface. For instance, Fig. 23.8 shows a hot spot in a thermographic image of a carbon–epoxy composite caused by delamination damage. Differences in surface temperature are recorded using an IR camera to reveal the damaged area, and temperature-data-processing methods are used to determine the damage depth. Thermography can detect delamination cracks, porous regions and foreign objects in composites and intermetallic inclusions and large voids in metals.
The passive thermography method uses internally generated heat, often from damage growth, rather than externally applied heat. Heat generated by damage growth is measured as a hot spot on the component surface. This technique is not as popular as active thermography because the material must be damaged to generate the internal heat and, therefore, it is rarely used to inspect aircraft components.
The eddy current method is widely used to inspect metallic aircraft components for surface cracks and corrosion damage. Figure 23.9 shows the eddy current process of inspection. Eddy current testing equipment contains a conductive metal coil which is electrified with an alternating current. The current generates a magnetic field around the coil. This magnetic field expands as the alternating current rises and collapses as the current drops. When the coil is placed in close proximity to another conducting material, such as a metal component, then an alternating current, called an eddy current, is induced in this material by the magnetic field. The eddy currents are induced by electrical currents that flow in circular paths. Surface and near-surface cracks interrupt the eddy currents, and this is detected by changes in the coil’s impedance. By passing the eddy current equipment, which is a hand-held device, above the component surface it is possible to detect cracks, including shallow and tight surface cracks that may be less than 50 μm deep and 5 μm wide. Cracks that cut across the eddy current path are easily detected, although cracks parallel to the current flow can be missed because they do not disturb the eddy currents.
Eddy current equipment is lightweight and portable, thus allowing for inspection of grounded aircraft. However, access to both sides of the component is necessary for complete inspection. Eddy current testing only works on conductive materials such as aerospace metal alloys. It cannot be used to inspect insulating materials with low electrical conductivity, such as fibreglass composites, and difficulties are experienced with ferromagnetic materials.
Magnetic particle inspection is a simple NDI method used to detect cracks at the surface of ferromagnetic materials such as steels and nickel-based alloys. The inspection process begins with the magnetisation of the component. The surface is then coated with small magnetic particles, which is usually a dry or wet suspension of iron filings. Surface cracks or corrosion pits create a flux leakage field in the magnetised component, as shown in Fig. 23.10. The magnetic particles are attracted to the flux leakage and thereby cluster at the crack. This cluster of particles is easier to see than the actual crack, and this is the basis for magnetic particle inspection.
Liquid dye penetrant is used to locate surface cracks in materials, but cannot detect subsurface damage. The method is shown schematically in Fig. 23.11. The component surface is cleaned before a visible or fluorescent liquid dye is applied using a spray, brush or bath. The dye seeps into surface cracks by capillary action. Excess dye retained on the surface is wiped off leaving only the dye that has seeped into the cracks. Chemical developer is then applied and it reacts with the dye, drawing it from the crack on to the surface. The dye can then be observed, either because it changes the colour of the developer or because it fluoresces under ultraviolet light.
Liquid dye penetrant is a popular inspection method because it is simple, inexpensive and can detect cracks to a depth of about 2 mm. The main drawback is that the method can only detect surface breaking cracks. Despite this problem, it is often used to inspect aircraft components, particularly engine parts, which are susceptible to surface cracking.
Acoustic emission involves the detection of defects using sounds generated by the defects themselves. Figure 23.12 shows the operating principles of the acoustic emission test method. The component is subjected to an applied elastic stress; usually just below the design load limit. When cracks and voids exist in materials, the stress levels immediately ahead of the defect are several times higher than the surrounding material (as explained in chapter 18). This is because cracks act as a stress raiser. Any plastic yielding and microcracking that occurs ahead of the defect owing to the stress concentration effect can generate acoustic stress waves before any significant damage growth. The waves are generated by the transient release of strain energy owing to microcracking. The waves are detected using sensitive acoustic transducers located at the surface. The transducers are passive, that is, they only ‘listen’ for sounds and do not generate the acoustic waves that the transducers used in ultrasonics do. Several transducers are placed over the test surface to determine the damage location. Certain defects have a characteristic sound frequency value and this is used to determine the type of damage present in the material. For example, delaminations in carbon–epoxy composites have a characteristic frequency of about 100 kHz whereas damaged fibres emit sound at around 400 kHz.
Acoustic emission has several advantages, including rapid inspection of large components and the capability to determine the location and type of damage. The method can be used to continuously monitor components while operating in-service, and in this respect it is a SHM technology. The downside is that the component must be further damaged to generate the acoustic emission signal.
Conventional NDI poses several problems for the aerospace industry that are difficult to resolve using the standard test methods. NDI methods such as ultrasonics, radiography and magnetic particle inspection require point-by-point inspections which are slow and labour intensive. Wide-area NDI techniques such as thermography allow for rapid inspection, although it is difficult to detect damage in nonplanar and complex structures such as joints. Most NDI methods require the grounding and stand-down of aircraft, often for lengthy periods because the inspections are slow. Further problems with many NDI tests are difficulties in the detection of damage in physically inaccessible areas of the aircraft and the inability to continuously monitor the formation and growth of damage over the aircraft life. For these reasons, there is growing awareness in the aviation industry that the real-time, continuous monitoring of in-service aircraft requires the use of SHM.
SHM involves the continuous measurement and assessment of in-service structures with little or no human intervention. The information provided by SHM systems about the physical condition of an aircraft structure is used by airline companies and aerospace engineers to make continuous lifecycle management decisions. Continuous assessment can significantly reduce aircraft downtime because it minimises the need for routine ground-based inspections for damage that may not be present. The early detection of damage using SHM improves the structural reliability and safety of aircraft. Early damage detection also offers the possibility of reducing design safety factors applied for damage tolerance, which translates into more lightweight aircraft structures, as well as minimising the repair process.
SHM systems are classified as passive or active. A passive system relies on taking measurements during normal in-service operation or detecting individual damage events, such as bird strike. In contrast, active systems stimulate the structure with an input and measure the output response of the structure. Data collected from the sensors is analysed using intelligent software processors for the determination of structural health. The processors are stored on the aircraft, acting much like a black-box flight recorder, or located at a ground-based facility with the signal transmitted direct from the aircraft.
SHM uses surface-mounted or embedded sensors for the real-time structural monitoring of damage initiation and growth. Various sensor types are available, including the following: fibre-optic sensors, piezoelectric transducers, dielectric sensors, and comparative vacuum galleries.
There are several similarities between the SHM of aircraft and the nervous system of humans. The idea of a SHM ‘nervous’ system installed in a commercial airliner is shown in Fig. 23.13. SHM sensors are distributed throughout an aircraft, in much the same way as nerves are spread throughout the body. The sensors detect changes to the ‘health condition’ of the structure in real-time, again like the nervous system, and this information is sent to a central processor. The processor provides immediate information about the damage during aircraft operations. Although it is possible to distribute sensors at many locations throughout an aircraft to assess the overall structural condition, this is usually unnecessary. It is only necessary to locate sensors at critical and highly loaded structures, such as the wing box, wing–fuselage connections and landing gear, and at structures prone to environmental or impact damage, such as leading edges susceptible to bird strike.
SHM systems are designed for either local or wide-area monitoring. A local system is concerned only with pre-determined structural ‘hot spots’ such as joints, leading edges and door frames. SHM systems for local monitoring include Bragg grating optical fibre systems and comparative vacuum monitoring. Global systems are concerned with damage detection and identification over a much larger area and generally aim to detect larger major structural damage, and examples include acousto-ultrasonics and random decrement analysis based on vibrational responses.
Unlike conventional NDI, SHM for aircraft damage is not a fully mature technology, and SHM is currently not widely used in aircraft or helicopters. There are several military aircraft and, to a lesser extent, civil airliners fitted with SHM sensors. However, SHM is not a mainstream technology for damage monitoring, and the aviation industry is still heavily reliant on NDI technology. As the industry gains a better understanding of SHM then it is likely that sensors will be used increasingly in aircraft. Short descriptions of several promising SHM technologies follow – i.e. optical fibres, piezoelectric sensors and comparative vacuum monitoring – to demonstrate their potential aerospace applications.
Optical fibre sensor systems are one of the more mature SHM techniques. The optical fibre sensor consists of a central silica core surrounded by an annular silica cladding with a protective coating (Fig. 23.14). Sensors are also made of translucent plastic materials. The fibres are long and thin (50–250 μm in diameter) and are surface mounted on metal and composite structures. Sensors can be embedded within fibre–polymer composites and along the bond-line of structural joints.
The method of damage detection involves shining monochromatic light along the fibre. The fibre core has a higher refractive index than the cladding, which allows the light to be confined within the core with minimal loss over long distances. The core is inscribed with Bragg gratings, which are lengths in which the grating lines lead to changed reflection. The light is partly reflected back at the Bragg gratings. The spacing distance between neighbouring gratings is measured by the wavelength of the reflected light. The spacing between Bragg gratings changes when the fibre is strained; the spacing increases under tension and contracts under compression. The strain level applied to the sensor (and therefore the structure to which it is attached) is measured from the change in reflected wavelength caused by the change in spacing of the Bragg gratings. Certain types of damage, such as delaminations and broken fibres in composites and corrosion in metals, can reduce the local stiffness of structures. Damage is detected using optical fibres by an unexpected change in the measured strain (or stiffness) compared with the strain of the defect-free structure. Optical fibres are small and robust which makes them suitable for installation on aircraft structures. However, they can only detect damage in the vicinity of the Bragg gratings and, therefore, a large number of sensors are required for wide-area inspection of large aircraft structures.
The basic operating concept of a SHM system called acousto-ultrasonics that uses piezoelectric sensors is shown in Fig. 23.15. Piezoelectric sensors produce an electric charge upon the application of strain and, conversely, they can expand when subject to an electric field (Fig. 23.16). This piezoelectric effect is caused by the disturbance of electric dipoles from their equilibrium state within the sensor material.
Piezoelectric properties are established by applying a high electric field in a direction known as the polling direction, at an elevated temperature, in order to align all the electric dipoles within the material. Common piezoelectric materials include quartz, barium titanate (BaTiO3), lead zirconate titanate [Pb(ZrTi)O3] and polyvinylidene fluoride. Piezoelectric materials come in various shapes and forms, including wafers, plates, strips or fibres, which are attached to the structural component. The sensors are often thin and small (less than ~ 15 mm), and are bonded directly to the structure.
The system shown in Fig. 23.15 has three piezoelectric devices: an actuator that releases elastic stress waves when activated by an alternating electric charge and two receiving sensors that measure the strength and frequency of the waves by changes to their electrical properties. When an aircraft structure fitted with a piezoelectric SHM system is damage-free, the signal recorded by the receiving sensor has a characteristic wave intensity and frequency. Damage that changes the elastic properties of the structure, such as corrosion, fatigue or impact cracking, is detected by a change in the signal wave properties. It is feasible to use many piezoelectric actuators and sensors distributed over a large structure to perform wide-area, real-time and continuous inspections. The potential application of piezoelectric devices for the structural monitoring of aircraft has been demonstrated.
Comparative vacuum monitoring (CVM) is an emerging SHM system for use in large metallic and composite aircraft structures. The operating principle of CVM is simple; the system consists of a thin polymer sensor containing a series of thin holes that are connected to sensing and recording equipment (Fig. 23.17). The sensor is bonded to the aircraft structure, usually where damage is expected. The arrangement of the narrow holes, called ‘galleries’, inside the sensor is shown in Fig. 23.18. Air within every second gallery in a series is removed by a vacuum pump to place them in a state of low pressure (called vacuum galleries). Air is retained in the other galleries, and this creates a sensor system consisting of a series of parallel galleries that alternate between low pressure and ambient pressure.
CVM is based on the principle that a crack growing under the sensor links a low-pressure gallery and a neighbouring gallery at ambient pressure. Air flows through the crack into the neighbouring low-pressure gallery, and the pressure change is detected by sensing equipment. Using this approach, it is possible to determine the location and size of surface cracks caused by corrosion, fatigue or impact, provided of course the sensor is located at the defective location.
Aircraft structures and engine components must be nondestructively inspected after manufacturing and throughout their operational life for the presence of defects and damage. Most inspections are currently performed using NDT methods such as ultrasonics, radiography and thermography. Structural health monitoring (SHM) is emerging as an alternative to conventional NDI, in which sensor systems are used with little or no human invention to monitor aircraft for damage.
NDI methods have the capability to detect certain (but not all) types of damage in metals and composites. Ultrasonics, thermography and eddy current inspections are capable of detecting damage and cracks aligned parallel with the material surface whereas radiography is better suited to detecting cracks normal to the surface. It is often necessary to use two or more inspection methods to obtain a complete description of the type, amount and location of the damage.
Some NDI techniques can be used to inspect metals but not fibre–polymer composites. Of the NDI methods described in this chapter, damage in composites is difficult to detect using eddy current and magnetic particle owing to their low electromagnetic properties, and using liquid dye penetrant because most damage is internal (e.g. delaminations) and does not break the surface.
SHM has the potential to reduce aircraft downtime for routine inspections and reduce design safety factors for damage tolerance because of the early detection of damage. It is often only necessary to locate SHM sensors in components prone to damage (e.g. heavily-loaded parts, parts susceptible to impact damage), rather than covering the entire aircraft with a complex, integrated sensor network system.
SHM techniques are classified as local or global (wide-area). Examples of local health monitoring include Bragg grating optical fibre sensors and comparative vacuum monitoring, whereas wide-area monitoring techniques are acoustic emission and acousto-ultrasonics.
Bragg grating: Periodically spaced zones in an optical fiber core with refractive indexes that are slightly higher than the core. The gratings selectively reflect a very narrow range of wavelengths while transmitting others. Used as an SHM sensor for the measurement of strain.
C-scan: Data presentation method (usually as a two-dimensional image) applied to ultrasonic techniques. Image shows the size and shape of the defective region, and by using time-of-flight data can reveal defect depth.
Comparative vacuum monitoring: SHM technique that uses a sensor to measure the differential pressure between fine galleries at a low vacuum alternating with galleries at atmosphere. If no flaw is present, the vacuum remains at a stable pressure. However, if a flaw develops, air flows through the passage created from the atmosphere to the vacuum galleries.
Eddy current: An NDE technique that uses an induced electric current formed within conductive materials which are exposed to a time varying magnetic field. Damage is recorded as a disturbance to the current.
Lamb waves: Elastic waves whose motion is along the plane of the plate. Lamb waves have lower frequencies than conventional ultrasonic waves, which allows them to travel longer distances along plates with a consistent wave pattern.
Liquid dye penetrant: An NDE technique for detecting surface porosity or cracks in metals. The part to be inspected is cleaned and coated with a dye that penetrates any flaws that may be present. The surface is wiped clean and coated with a chemical to absorb the dye retained in the surface defects indicating their location.
Magnetic particle: An NDE technique for determining defects in ferromagnetic materials. Finely divided magnetic particles, applied to the magnetised material, are attracted to and outline the pattern of the magnetic leakage fields created by the damage.
Piezoelectricity: The generation of electricity or electrical polarity in dielectric crystals subjected to mechanical stress, or the generation of stress in such crystals subjected to an applied voltage.
Radiography: An NDE technique that relies on the transmission of radiation (usually x-rays) through a solid to produce an image produced on a radiosensitive surface, such as photographic film. Defective regions absorb the radiation at a different rate to the pristine material, and damage appears as a darker or brighter spot in the image.
Structural health monitoring (SHM): Process involving the observation of a structure over time using periodically sampled dynamic response measurements from an array of sensors, the extraction of damage-sensitive features from these measurements, and the statistical analysis of these features to determine the current state of system health.
Through-transmission ultrasonics: Ultrasonic technique that relies on the attenuation of the acoustic beam when passing through damage within a material to determine the location and size of the defective region.
Ultrasonics: NDE technique which relies on an ultrasonic beam passing through the material to detect the presence of damage by either back- reflection (pulse–echo) or attenuation (through–transmission) of the acoustic waves.