Note: Descriptions are shown in the official language in which they were submitted.
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METHOD AND SYSTEM FOR HEALTH MONITORING
OF AIRCRAFT LANDING GEAR
FIELD OF THE INVENTION
The present invention relates to aircraft landing systems and more
particularly to a
method and system for deterrnining whether the landing gear is healthy or
whether it
requires maintenance, service and/or replacement. This invention will also
determine if the
risk of a catastrophic failure of the landing gear has changed as a result of
its in-service
operations.
BACKGROUND OF THE INVENTION
The goal of health monitoring technologies is to know at any time, for any
aircraft
in the fleet, the structural integrity of the landing gear, the amount of
remaining fatigue life
in the landing gear, the landing gear servicing information (such as shock
strut pressure and
fluid volume, tire pressure and temperature, and brake condition), and the
internal status of
all on-board electronics and systems related to the landing gear system.
Being able to measure and assess the safety and integrity of the landing gear
and
landing gear system is of vital interest to the public safety.
The current process for deciding that an airplane has had a "hard landing",
and thus
has compromised the safety and integrity of the landing gear, is based on a
subjective
assessment by the flight crew. Because of the lack of reliable quantitative
data, errors are
made in this assessment. As a result, an airplane may be grounded
unnecessarily, at a
considerable cost of time and money, or conversely, a damaged airplane can
continue in
service, thus compromising public safety.
In addition to this current practice, servicing and maintenance are scheduled
to take
place at pre-determined intervals. This results in some servicing and
inspections taking
place before it is required, thus resulting in considerable additional cost of
time and money.
Conversely, in some cases, the landing gear may be in need of servicing,
maintenance or
replacement before the next scheduled time. In the interests of the public
safety, it is better
to be safe than sorry and so maintenance and servicing schedules tend to be
very
conservative.
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Landing gear health monitoring systems involve several unique issues that
differentiate it from all other airplane systems and the airframe itself.
Airframes are made
from relatively ductile aluminum alloys that can withstand relatively long
cracks that grow
over time. These aluminum structures can sustain fairly significant corrosion
before the
airplane's fitness for service is compromised. In contrast, landing gears are
made from very
high strength (but relatively low toughness) steel, aluminum, and titanium
alloys with
critical defect sizes that are much smaller.
This significant difference is also reflected in the fact that aircraft design
and
approval methodologies are quite different between the airframe and the
landing gear. For
example, the airframe uses "damage tolerant" design methodologies, which allow
cracks of
known sizes to exist in structural members, applied to fatigue dominated zones
in the
airframe compared to "safe life" design methods, which do not permit cracks,
used in the
landing gear.
As a result, many of the technologies and articles related to health
monitoring of the
airframe, e.g. measuring the dynamic characteristics of the structure and then
inferring
whether certain joints have failed or cracks have grown, are of little
interest when
considering health monitoring of the landing gear. Similarly, the sensors and
technology
involved for airplane systems are not sensitive enough to resolve the very
small defects of
interest or displacements of interest for landing gear applications.
The present invention provides a system and method that utilizes extensive
destructive and non-destructive testing and analysis of full-scale landing
gear, extensive
engineering modeling of the landing gear design and modeling of the causes of
failure, and
extensive experience with analysis of landing gears in-service. This
integrated system and
method utilizes an arrangement of sensors and sub-systems and an extensive
database of
information such as the original manufactured condition of the landing gear,
amount and
type of maintenance, in-service history of similar landing gear, history of
the specific
landing gear of interest, prior in-service loads, and number and type of hard
landings; and
sophisticated analytical techniques in order to determine the safety of the
landing gear
and/or need for. service, maintenance or replacement. The present invention
can
disseminate and report the need for service, maintenance or replacement to a
spectrum of
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interested parties including: pilots and flight crews, maintenance personnel,
airline
operators, ground crew and regulatory authorities.
SUMMARY OF THE INVENTION
The present invention provides a system having a variety of sensors attached
to the
landing gear structure and equipment, a method and system to communicate the
data
measured by the sensors to the monitoring system, a method to analyze the data
to derive
relevant information about the health and safety of the landing gear, and a
method and
system to report the potential need for service, maintenance or replacement to
pilots and
flight crew, maintenance personnel, airline operators, ground crew and
regulatory
authorities.
In one embodiment the present invention provides an aircraft landing gear
health
monitoring system comprising at least one sensor coupled to at least one
component of the
aircraft landing gear system for measuring and recording real-time data
associated with the
status of at least one component. The system also includes at least one
processor connected
to the at least one sensor for receiving and processing the real-time data to
calculate the
condition of the at least one component of the landing gear system and
reporting means
operable to receive information from the at least one processor for reporting
at least one of
the condition of the landing gear system and the real-time data.
In another embodiment the present invention provides an aircraft landing gear
health monitoring system comprising a plurality of sub-systems and a reporting
device
connected to each of the sub-systems for receiving data from the sub-sytem.
Each sub-
system comprises a plurality of sensors each independently connected to
separate
components of a pre-determined sub-system of the aircraft landing gear and
operable to
measure and record real-time data associated with the status of each component
and a
processor connected to the plurality of sensors for receiving the real-time
data therefrom
and operable to analyse the real-time data to calculate the condition of the
sub-system. The
reporting device is operable to report at least one of real-time data and the
condition of
each of the sub-systems.
In a further embodiment, the present invention provides an aircraft landing
gear
health monitoring system comprising at least one sensor coupled to at least
one component
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of the aircraft landing gear system for recording real-time data associated
with the at least
one component, a processor operable to receive and compare the recorded real-
time data
with pre-determined health data associated with the at least one component and
to calculate
the condition of the at least one component, a communication device operable
to
communicate with the at least one sensor and the processor, and to receive
information
relating to the real-time data, the analysed data and the calculated condition
of the at least
one component and reporting means connected to the communication device and
operable
to receive and report information received therefrom relating to at least one
of the real-time
data, the analysed data and the condition of the at least one component.
In an alternative aspect the present invention provides a method for
monitoring the
health of an aircraft landing gear system comprising the steps of (i)
collecting real-time
data associated with the condition of at least one component of the aircraft
landing gear
system, (ii) analysing the real-time data to assess the condition of the at
least one
component and (iii) reporting the condition of the at least one component.
In an alternative embodiment of the present invention provides. a method for
monitoring and diagnosing the health of an aircraft landing gear system
comprising the
steps of (i) recording real-time data associated with the status of at least
one component of
the aircraft landing gear system, (ii) transmitting the real-time data to a
processor for
processing, (iii) processing the real-time data to calculate the current
condition of the at
least one component and to determine if any maintenance is required and (iv)
reporting at
least one of the real-time data, the calculated condition and any required
maintenance.
In addition, the method described above may also include repeating steps (i)
and (ii)
for additional sub-systems. The method may. occur while the aircraft is in
flight or while it
is on the ground. The methods described above may also include the additional
step of
transmitting at least one of the real-time data, the calculated condition and
any required
maintenance to a ground-based master landing gear database.
BRIEF DESCRIPTION OF THE DRAWINGS
The present invention will be described in further detail below with reference
to the
attached figures in which:
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Figure 1 is a schematic of one embodiment of the health monitoring system of
the
present invention;
Figures 2A-2C are a series of graphs showing the transformation of load data
into
load-damage data;
Figure 3 is a graph showing damage versus life remaining;
Figure 4 illustrates one embodiment as initial screen of a user interface for
the
present invention;
Figure 5 illustrates a query screen of a user interface for the present
invention;
Figure 6 illustrates a bulletin screen of a user interface for the present
invention;
Figure 7 illustrates a data accessing screen of a user interface for the
present
invention;
Figure 8 illustrates a multi-data display screen of a user interface for the
present
invention; and
Figure 9 illustrates a further query screen of a user interface of the present
invention.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
The invention relates to a new method and system for health monitoring of
aircraft
landing gear. The invention includes the use of a plurality of sensors that
are attached to
separate components of the landing gear structure and equipment (e.g., one or
more of
brakes, tires, hydraulics, electrical systems and switches) and analyzed to
report and alert
personnel such as pilots, maintenance personnel, airline operators, ground
crew and
regulatory authorities of the health of the landing gear and the potential
need for service,
maintenance or replacement.
The individual sensors measure and record data related to the component(s) to
which it is attached, e.g. a sensor attached to the shock strut may measure
oil pressure,
level, and/or temperature; a sensor attached to a tire may measure the tire
pressure. The
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sensors that may be used are known in the art and may be either releasably or
permanently
attached to the component after manufacturing or may be connected to the
component
during the original equipment manufacturing operations. The data collected
from the
sensor(s) is then either directly or indirectly, through analysis and/or
manipulation, used to
report a health issue which may be directly connected to the component being
monitored,
e.g. a flat tire, or indirectly related, e.g. low tire pressure that affects
the braking system.
The present system is operable to monitor and report critical health issues
associated with the landing gear such as the in-service loads due to landing
and taxiing, the
presence of structural defects such as cracks or pre-crack material damage,
tire pressure,
tire temperature, brake wear, hydraulic pressure, the status of on-board
electronics, the
status of the equipment and wiring, and the overall condition of the landing
gear and ability
to sustain another landing. This real-time information, also referred to
herein as real-time
data, can be analyzed in conjunction with an extensive database of
information, also
referred to herein as pre-determined health data, such as the original
manufactured
condition of the landing gear, amount and type of maintenance, in-service
history of similar
landing gear, history of the specific landing gear of interest, prior in-
service loads, and
number and type of hard landings in order to determine the safety of the
landing gear
and/or need for service, maintenance or replacement. The real-time information
and/or
information analyzed in conjunction with the database, can be used to alert
pilots using a
cockpit display screen and/or to alert the aircraft owners, operators,
maintenance staff,
ground crew and regulatory authorities via remote transmission providing such
personnel
with the option to take actions such as additional inspection, service,
maintenance and/or
replacement of the landing gear.
The present invention will now be discussed in further detail with reference
to the
attached Figures 1-9.
One embodiment of the health monitoring system 10 of the present invention is
illustrated in Figure 1 and includes a plurality of sensors 12 independently
connected to
separate components of the landing gear structure and equipment of an
aircraft, not shown.
The system 10 also includes a processor 16 to process and communicate the data
received
from the sensors 12. The processor 16 analyses the data received from the
sensors 12 to
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ascertain the status of the landing gear structure and may also compare the
analysed data to
pre-determined health data to determine the current condition of the landing
gear relative to
the historic condition. The processor 16 then reports the condition of the
landing gear using
a display 24 to at least one relevant personnel. The displaying of the
information may be in
the form of a query and reporting system that will be described in further
detail below.
The system 10 may also include at least one signal conditioning device 14
connected to the plurality of sensors 12 and the at least one processor 16. In
the illustrated
embodiment, the system 10 includes a plurality of signal conditioning devices
14
connected to the plurality of sensors 12. The plurality of signal conditioning
devices 14 are
in turn connected to the processor 16.
It will be understood that the processor 16 and the plurality of signal
conditioning
devices 14 may be separate components or may be one unitary component, i.e.
the
processor 16 may be operable to receive data directly from the sensors 12.
Further the
plurality of signal conditioning devices 14 may in fact be one unitary signal
conditioning
device 14 that is connected to each sensor 12 and to the processor 16.
Further, each sensor
12 may be connected to one signal conditioning device 14 or to several signal
conditioning
devices 14 or alternatively each sensor 12 may be directly connected to one
central
processor 16.
In one embodiment, the signal conditioning device 14 is operable to receive
data
from the sensor(s) 12 and convert it into a form that the processor is able to
understand and
analyse. As an example, the signal conditioning device may be an analog to
digital
converter to transfer the sensor information into a form that the processor
can understand.
In an alternative embodiment, the signal conditioning device 14 may include a
component for conditioning the data received which in turn is connected to a
network bus
or alternatively the data may be transferred from the sensor 12 to the network
bus and then
be converted into a form that the processor is able to understand and analyse.
It will
therefore be understood that the signal conditioning device 14 may be in the
form of a
distributed system where the sensor data is communicated over a network or
communications bus, such as ARINC-429, AFDX, CANbus or Time Triggered Protocol
-
or other similar devices known to a person skilled in the art.
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It will also be understood by a person skilled in the art that each different
type of
sensor will require suitable interface circuitry to adapt it to be read by the
processor. For
instance, strain gauged based sensors need excitation, amplification,
filtering, and then
conversion from the analog domain to the digital domain. A capacitive fluid
level sensor
may require conversion either using a direct to digital converter, or by using
a capacitive
bridge circuit, excitation circuitry, an demodulation circuitry. Preferably,
the sensors will
be of an analog nature, with the exception of sensors that behave like a
switch, e.g.
proximity sensors.
An appropriate number and type of sensors 12 are attached to the components of
the
landing gear (not shown) in appropriate locations for each component to be
monitored
which will be described in further detail below. It will be understood that a
person skilled
in the art will know where particular sensors 12 should be located and how to
attach them
to a component of the landing gear. It will be understood that the choice of
number and
type of sensor 12 may vary depending on the type and number of components to
be
monitored. The minimum number of sensors 12 may depend on both the geometry of
the
landing gear, and the information desired. For example, the system may include
only one
sensor if a small amount of specific data is required whereas other systems
will require
additional sensors. The state of some components may be assessed using data
from one
particular sensor or from a combination of sensors.
Examples of the types of sensors 12 and components to which the sensors 12 may
be attached include, but are not limited to, tires, brakes, hydraulics,
electronics, landing
gear doors, oil pressure, oil temperature, oil level, shock strut position,
loads, strain gauges,
structural integrity, magnetic permeability, brake pressure, and aircraft bus
data including
airplane velocity, position, attitude and altitude
As an example, in a system for measuring the servicing state of the shock
strut, a
measurement of one of the following could be used alone or in combination with
the other
measurements to assess the servicing state: internal oil pressure, oil level,
shock strut
extension, and oil temperature. Alternatively, one could not measure the oil
level and
instead use a stored measurement that was made with the aircraft in flight -
two sets of data
at known conditions - strut position, temperature, and pressure can be used to
determine
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the service state. For the determination of the loads through the landing
gear, the landing
gear geometry directly affects how this can be achieved - there are three
loading directions
of interest - vertical, side, and drag - these could define a minimum.
This real-time data, i.e. the data recorded by each sensor 12, may be
conditioned or
transformed, by either the sensor itself or the signal conditioning device 14
or the processor
16, into information that is more directly relevant to a condition to be
monitored and
reported (e.g. converting voltage to pressure). For example, if a
communications bus is
used the sensors may be of the 'smart sensor' variety. Therefore, the sensors
may employ a
local microprocessor, signal conditioning circuitry, and data conversion
circuitry to convert
the measured signal to a digital signal, then relay that signal over the
communications bus
to the processor 16.
This information may then be analysed in conjunction with an on-board database
18, containing the pre-determined health data associated with each component,
by the
processor 16 based on pre-determined algorithms, heuristics, or alternative
methodology
such as neural networks or fuzzy logic. This set of analytical techniques may
be referred to
as the "Analysis Method Library" and is indicated generally by numeral 20. It
will be
understood that the analytical techniques may be stored in the processor 16 or
may be
stored within a system with which the processor 16 is operable to communicate
to retrieve
the required information. Likewise the on-board database 18 may be part of the
processor
16 oir may be a separate component in communication with the processor 16.
The resulting analysis determines the landing gear condition and may be
reported,
along with any alerts deemed necessary as a result of the analysis through a
display 24
accessible by, for example, the on-board crew including the pilot(s), co-
pilot(s) and flight
crew. Alternatively, the display 24 may be part of a ground-based system that
is in
communication with the processor 16 and is accessible by ground personnel.
In an alternative embodiment the system 10 includes a communications sub-
system
22 that receives the information relating to the condition of the landing gear
components
from the processor 16 and in turn communicates the information to the display
24. The
communications sub-system 22 may be part of the processor 16 or may be a co-
processor.
The sub-system 22 is operable to collect all communications into and out from
the
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processor 16, and optionally all other components in the system 10. The sub-
system 22
provides a separate communication from the processor 16 if desired or
required.
Examples of the types of alerts that may be transmitted as a result of the
analysis
include, but are not limited to: the remaining life of the landing gear; the
need for
servicing; the need for maintenance; the need for inspection and a calculated
risk of failure
of the landing gear upon next landing.
The communications sub-system 22 and/or the processor 16 may also be operable
to communicate with a ground-based master landing gear database 26. Therefore,
all
relevant information may be transmitted between the communications sub-system
22/processor 16, including the on-board database 18 and the ground-based
master landing
gear database 26 which will allow for any updates of information from the
master landing
gear database 26, for example to the algorithms used to be made. The master
landing gear
database 26 may include information such as landing gear system information,
i.e. built in
test results for each piece of avionics, reported anomalies, brake system
information, i.e.
brake temperatures, pressures, wear information, tire information, i.e. tire
pressure, tire
wear information, tire temperatures and landing gear information, i.e. landing
gear usage
including loads, forces, time histories, individual part fatigue information
and life
consumed. Preferably the master landing gear database 26 includes at least the
landing gear
information.
When the information is communicated to the ground-based master landing gear
database 26, the status of the landing gear and/or alerts can be sent through
a reporting sub-
system 28 to the aircraft owners and operators, maintenance staff, ground
crew, and
regulatory authorities, indicated generally at 30, who may decide to take
actions such as
additional inspection, service, maintenance and/or replacement of the landing
gear. Any
actions taken on the landing gear, such as servicing, maintenance or
inspection, indicated
generally at 32, can then be entered and uploaded back to the master landing
gear database
26, which can in turn update the on-board database 18 in preparation for the
next takeoff.
In an alternative embodiment of the invention, there are multiple sub-systems
that
are included in the overall system. The sub-systems each focus on one
component of the
aircraft, including structural, tires, brakes, hydraulics, electronics,
position,
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communications, database, analysis, and reporting. Each modular sub-system is
dedicated
to obtaining, conditioning and analyzing information of interest regarding the
component.
However, the information recorded by each sub-system can be shared between sub-
systems
and used to assess the condition of other components in the overall system.
In one example, a structural integrity sub-system is provided in which the
sensors
12 are attached directly onto the structural portions of the landing gear,
either during a
retrofit operation or during the original manufacture. These, sensors 12 are
used to measure
all relevant loads experienced by the structure during .taxiing, take-off and
landing,
including, but not limited to, for example torsional, axial, fatigue and shock
loads. The data
recorded by the sensors 12 of the system may then be processed to calculate
information
related to the presence of defects, discontinuities and/or pre-crack damage to
the structure.
This information can then be used to calculate the current health of the
landing gear
structure which in turn can be compared to the original manufactured condition
of the
landing gear.
In another example, in order to monitor the tire pressure a pressure sensor is
attached to the tire in such a way as to obtain and communicate the pressure
information to
the monitoring system or brake monitoring sub-system. In order to monitor the
loads on the
structure, load sensors are attached to the structure in the appropriate
locations so as to
obtain and communicate the load information to the monitoring system or
structural
integrity monitoring sub-system.
The sensor information is analyzed in a variety of ways, depending on the
specific
sub-system. For example, the tire pressure sensor can measure the pressure
directly. By
knowing the change in tire pressure over time, an assessment can be made
whether the tire
is leaking air. Depending on rate of pressure decrease, ambient temperature,
prior service
history, e.g. if a valve has just been replaced, and correlation with tires in
the rest of the
fleet, an assessment may be made to replace the tire, fix a valve stem, simply
re-inflate the
tire, or leave the tire alone because the pressure drop was caused by a drop
in ambient
temperature.
Each sub-system will have its own method for analyzing the raw sensor
information, conditioning or converting the raw information into more directly
relevant
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information as appropriate, e.g. converting voltage to pressure or converting
the time rate
of voltage change to magnetic permeability to the presence of pre-crack
damage, analyzing
the information in conjunction with a database of information and,reporting
the need for
service, maintenance or replacement.
The real-time information from each sub-system can be analyzed in conjunction
with an extensive database of information such as the original manufactured
condition of
the landing gear, amount and type of maintenance, in-service history of
similar landing
gear, history of the specific landing gear of interest, prior in-service
loads, and number and
type of hard landings in order to determine the safety.of the landing gear
and/or need for
service, maintenance or replacement.
The real-time information and/or information analyzed in conjunction with the
database, can be used to alert pilots using a cockpit display screen, and/or
remotely
transmitted to the aircraft owners and operators, maintenance staff, ground
crew, and
regulatory authorities who may decide to take actions such as additional
inspection,
service, maintenance and/or replacement of the landing gear.
Determining the number, location and type of these sensors 12 requires
engineering
modeling and testing of the landing gear in order to optimize performance and
sensitivity.
It should be noted that the number, location, and type, e.g. number of
windings, of sensors
12 will be identical from one set of landing gear to another within a given
type of landing
gear, but will vary from one type of landing gear to another depending on the
engineering
design analysis and full scale destructive testing results. The raw data taken
from the load
sensor must be conditioned in order to determine the actual load.
The following provides an example of the type of analysis that may be
performed.
Analysis of the in-service loads measured using the structural integrity sub-
system can be
used to determine the weight and balance of the airplane, presence of hard
landings, and
other loads that may contribute to a reduction in the remaining useful life of
the structure.
Generally, the most common way for engineers to measure the life of a
structure is to
construct a S-N Curve showing the number of cycles to failure (N) for a given
applied load
(S). Statistical analysis is used to predict the probability of failure of a
landing gear that has
experienced a certain number of cycles of a given load. However, if the
landing gear is
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subjected to several significantly higher loads than S, the effect on the
remaining number
of cycles to failure (N) is far more difficult to calculate. Furthermore,
during the lifetime of
a landing gear there will be large fluctuations in the range of loads (S),
requiring a more
sophisticated technique for assessing the total amount of load-damage (e.g. a
modification
to the calculation of S which we will name "D") and thus remaining useful life
(a
modification to the calculation of N which we will name "L"). The structural
integrity sub-
system must continually update the actual D that the particular landing gear
has been
subjected to, so as to continuously be calculating the remaining number of
cycles to failure
(L) and probability of failure for any potential future landing (e.g. the
probability of failure
during a hard landing will be different than the probability of failure during
a soft landing).
As an example, Figure 2 is a graphical illustration of how to transform the
load data
(S) into more relevant information (D). As can be seen, certain small loads
may not affect
the overall damage, however, a large load may cause a significant increase in
the damage.
In the example shown in Figure 2, the first set of loads (S) cause a slight
increase in D.
However, this slightly increased D has almost no effect on the Life Remaining
(L), as
shown in the lower arrow that points from the second graph to the third graph.
In contrast,
a higher load, such as shown taking place at a later time (t), can increase
the damage
further resulting in a significant effect on the Life Remaining (L), as shown
in the top
arrow that points from the second graph to the third graph.
As shown in Figure 3, there may be significant scatter in the experimental
data
when performing actual destructive tests on full-scale landing gear. For
example,
seemingly identical specimens may have a significant difference in the number
of cycles to
failure or remaining life. This uncertainty is addressed using statistical
analyses as well as
conservative projections. As can be seen from Figure 3, using a conservative
remaining life
projection (the lower edge of the curve) may result in almost half of the
anticipated lifetime
compared to using the projection from the centre or top of the curve.
Therefore, using only the directly measured in-service loads to predict
remaining
life will result in very conservative predictions. However, these predictions
can be
significantly improved, often resulting in a much higher calculated remaining
life, if the D-
L curve can be modified based on additional sources of information such as
direct
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measurement of the damage, knowledge of the original manufactured condition,
and
knowledge of the service history of the landing gear. A system, as provided by
the present
invention, that includes continual monitoring of the landing gear components
and a
comparison of the monitored data with pre-determined structural health
measurements
provides more accuracy to such calculations.
In an alternative embodiment, the sensors 12 in the structural integrity sub-
system
will also directly measure the material properties in a way that will provide
direct evidence
of the presence of defects and/or pre-crack damage. In one embodiment of the
invention,
meandering winding magnetometer sensors are used to measure the magnetic
permeability
of the material in such a way as to be highly correlated with the presence of
defects and/or
pre-crack damage. An example of the sensors that may be used are the sensors
manufactured b'y Jentek Sensors Inc.
Each landing gear has its own continuously updated D-L curve, that calculates
the
remaining life (L) as a function of the total spectrum of applied loads, which
are used to
calculate (D). However, once a defect, crack or pre-crack defect has been
found, for
example by measuring the magnetic permeability of the material, the D-L curve
will be
shifted to a new D-L curve which shows a lower remaining life, which will be
named "Ld"
for L in the presence of a known defect.
In another embodiment, the original manufactured condition of the landing gear
is
known. This is achieved by creating a "birth certificate" or "fingerprint",
indicated
generally at 40 in the schematic of Figure 1, for each newly manufactured
landing gear
composed of a three dimensional geometric inspection and complemented with
enhanced
non-destructive inspection results, such as magnetic permeability or
Barkhausen noise
inspection. The "birth certificate" or "fingerprint" establishes the part's
expected fatigue
and strength performance. The "birth certificate" includes such data as non-
destructive
inspection results, surface discontinuities, coating thickness, tube wall
thickness, heat
treatment history, repair and rework occurring during manufacturing. By
comparing the in-
service material properties, e.g. magnetic permeability or Barkhausen noise,
the "new
fingerprint"- to the original fingerprint, a better determination can be made
as to the
presence of a defect and/or pre-crack damage. Thus a better determination can
be made as
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WO 2006/053433 PCT/CA2005/001750
to the remaining useful life (Ld) of the landing gear. This information may be
stored in the
on-board database 18 along with the pre-determined health data for each
component.
Alternatively, the "birth certificate" or "fingerprint" for each component may
be stored
separately.
Using an array of information based on real-time data (e.g. load and magnetic
permeability), provided by the sensors 12 and/or signal conditioning device 14
in
conjunction with the on-board database 18 (e.g. prior birth certificate
fingerprint of
magnetic permeability) and sophisticated algorithms (e.g. the method for
transforming S
into D to calculate L or Ld) or heuristics, neural networks or fuzzy logic
(e.g. the Analysis
Method Library 20), the processor 16 can determine the need for service,
maintenance or
replacement.
The structural integrity sub-system can. provide a plethora of useful
information
including weight and balance information (which is of immediate interest and
concern to
the pilot and flight crew who may wish to move passengers or decline to take-
off until the
weather has changed), hard landing indication (which may be used for
regulatory authority
notification), and notification that the remaining useful life has been
compromised so that
the landing gear can be removed, inspected, or serviced.
If a hard landing has taken place, the structural integrity sub-system can
calculate
the immediate effect of this on the remaining life of the landing gear and the
cost of this
reduction in life can be charged to the operator.
In a further embodiment, the sensors 12 in the health monitoring system are
connected to a plurality of measurement and analysis units (not shown) (one
per landing
gear assembly) that are in close proximity to the sensors, and that contain
internal,
rechargeable power supplies. When the aircraft avionics are on (such as when
the aircraft
is flying or taxiing) the remote measurement and analysis unit(s), which are
connected by
electrical cabling to the aircraft avionics, are recharged by the aircraft
electrical system and
the measured data contained in the units are transferred to the aircraft
avionics. This
system permits the measurement of landing gear structural integrity when the
aircraft
power is not on (during towing, parking, and storage or maintenance
activities). This
facility provides the structural integrity sub-system with capability to
detect damage during
CA 02587750 2007-05-03
WO 2006/053433 PCT/CA2005/001750
these times when conventional systems would not be operational. Considerable
landing
gear damage can occur when the aircraft is not powered on.
Alternatively, each sensor 12 may include its own source of power separate
from
the aircraft power system, which allows the sensor to continue to monitor data
even when
the aircraft power is switched off. In this embodiment, each sensor 12 for
which continual
monitoring is required, is equipped with, or connected to, a separate power
supply/source.
In a preferred embodiment of the system, subsets of the Master Landing Gear
Database including accrued Damage (D) information and available life (Ld)
information
will be stored (along with fingerprint information) in an electronic memory
that is attached
to components of the landing gear. As critical components of a landing gear
may be
removed for maintenance and replaced with other components from a rotatable
pool of
parts, a means is required to track the current composition of the landing
gears on the
airplane. By storing pertinent excerpts from the Master Landing Gear Database
on the
actual landing gear components, and by being able to retrieve them
electronically (e.g. by
using RFID tags and scanners) the processor 16 will always be aware of the
exact damage
status of the components on, the aircraft. For example, if a component in the
landing gear is
changed when the aircraft is powered off, once the power is returned the
processor 16 will
automatically read and download the information about the component contained
on it.
This information will provide the processor 16 with all relevant damage
information and
the information will be updated by the processor 16 during and after flight.
The central
database can also be updated with the relevant information relating to the
parts in use and
ariy damage applied thereto.
In a preferred embodiment of the system, there will also be sub-systems for
the
following: tires (pressure, temperature, wear and remaining life), brakes
(temperature,
integrity, wear and remaining life), hydraulics (pressure, temperature and
viscosity),
electronics (power, integrity and status), position (of the landing gear doors
and landing
gear), communications (between the sensors, on-board systems, pilot cockpit
display, on-
board database, and ground-based systems), Master Landing Gear Database (of
the
maintenance history, in-service load history, similar landing gear systems,
and maintenance
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WO 2006/053433 PCT/CA2005/001750
history), and analysis and reporting (to show alerts, recommendations for
servicing or
maintenance, and provision of information).
Each of the above named sub-systems can be implemented using a similar
methodology as described for the structural integrity sub-system: taking key
data from
sensors attached in the appropriate location, analyzing the data to determine
critical
information of interest (e.g. =the condition of the brakes), and analyzing
this information in
conjunction with the on-board database by the central processing unit to
determine the need
to take actions such as alerting the pilot, performing additional inspection,
removing the
landing gear, and/or performing servicing or maintenance.
In another embodiment of the system, data collected from each of the sub-
systems
is returned electronically from the aircraft to an analysis center. Each
report is accepted
into a database system (such as Teamcenter from UGS) that attaches the data
report to the
data records for that part number and serial number of part or assembly. In
the case of data
returned from the structural integrity sub-system, data is electronically
attached to a top
level landing gear assembly. The software aligns all data records with
original design
specifications and as-built records. For structural data on an assembly, the
data is
automatically routed to individual data processing and analysis routines that
generate the
damage and life information for each sub component. This information is then
automatically appended to the appropriate part numbers and serial numbers
within the
database. Upon completion, electronic messages are dispatched to the aircraft
avionics to
update the onboard databases, and to operators and customer service personnel.
Figures 4 through 9 demonstrate one embodiment of a user interface to the
Querying and Reporting Sub-System for the present invention (in this example,
using the
trade name "SmartStrut - your landing gear health monitoring system"). As
shown in
Figure 4, the system can be accessed using a standard web browser such as
Microsoft
Explorer or Netscape Navigator and can be password protected to restrict
access to conduct
queries and permission to modify the database.
Figure 5 demonstrates the ability to conduct a query for a specific customer,
aircraft
and landing gear. Figure 6 demonstrates the ability to tie into other
databases such as
service bulletins and news bulletins. Figure 7 demonstrates the ability to
directly access the
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WO 2006/053433 PCT/CA2005/001750
real-time data and/or most recently updated information provided to the Master
Landing
Gear Database from the on-board systems. Figure 8 demonstrates the ability to
access
information associated with several sub-systems at once, including a down-lock
sensor
component fault, the oil level and nitrogen pressure. As can be seen, simple
heuristics can
be used to determine the potential need for servicing. In this case, the
heuristic for oil level
is that the oil level is critically high when above one number and critically
low if the oil
level is below another number.
In one embodiment of the invention, the oil level, rate of change of oil
level,
nitrogen pressure, and rate of change of nitrogen pressure (and/or other
information) are
used to report a single value - "need to perform service", "no need to perform
service", or
"service needed soon".
Figure 9 demonstrates the ability to log the access to the system, email
alerts and
conduct further queries. Additional searching, querying, analysis and
reporting functions
are available through this user interface.
While this invention has been described with reference to illustrative
embodiments
and examples, the description is not intended to be construed in a limiting
sense. Thus,
various modifications of the illustrative embodiments, as well as other
embodiments of the
invention, will be apparent to persons skilled in the art upon reference to
this description.
It is therefore contemplated that the appended claims will cover any such
modifications or
embodiments. Further, all of the claims are hereby incorporated 'by reference
into the
description of the preferred embodiments.
All publications, patents and patent applications referred to herein are
incorporated
by reference in their entirety to the same extent as if each individual
publication, patent or
patent application was specifically and individually iridicated to be
incorporated by
reference in its entirety. _
18
