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 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
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 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).
USAF Experiences
24
 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.
Current Technology State
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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
　
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