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 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.
0.002-0.030” long cracks 0.002-0.010” long cracks
Unpainted 0.040" Skin 0.040" Skin w/Primer
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 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
Acoustic Emission
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 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
　
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