Patent Description:
An aircraft may experience problems or faults with one or more aircraft systems during a flight. An aircraft system may detect and record each of the faults that occurred during the flight. Then, upon landing, a maintenance crew may review the recorded faults and perform diagnostics procedures to identify the source of the faults and take any necessary corrective action.

Typically, a maintenance crew follows specific procedures related to specific detected faults. These procedures may involve performing one or more measurements related to the aircraft (e.g., voltages, temperatures). Different procedures may be followed and different measurements may be taken depending on the type of fault detected. The maintenance crew may then diagnose the fault based on the measurements and the procedures. Once the problem is diagnosed, one or more part may be replaced, or other corrective action may be taken.

However, this method of performing fault isolation and aircraft maintenance may be highly labor intensive. Furthermore, the information recorded about detected faults may be limited (e.g., information may include only a time and type of fault detected). Further, the measurements taken when the aircraft is on the ground may be different than when the fault occurred while the aircraft was in flight. This may limit the effectiveness of the fault isolation performed by the maintenance crew. Thus, there is a need for an improved method of performing fault isolation in an aircraft system and performing enhanced maintenance of aircraft.

<CIT> discloses apparatuses and methods that cooperate in assisting with the maintenance of mobile platforms by facilitating diagnosis of events detected onboard mobile platforms while such mobile platforms are in operation (e.g., transit, flight). <CIT> discloses a method for monitoring a controller for a vehicle, the method includes determining configuration information associated with the vehicle and determining vehicle operating states associated with a plurality of conditions. <CIT> discloses a maintenance system for generating maintenance instructions for at least one system, the maintenance system having a plurality of system components set up to receive operating data for a system from a central control unit of the system. <CIT> discloses a method and apparatus for chronologically combining fault data, repair data, and operational parameters for a railroad locomotive.

Claim <NUM> defines a computer implemented method of performing fault isolation for an aircraft, and claim <NUM> defines a corresponding electronic control unit. In the following, apparatus and/or methods referred to as embodiments that nevertheless do not fall within the scope of the claims should be understood as examples useful for understanding the invention.

The characteristics of the present technology, as well as the methods of operation and functions of the related elements of structure and the combination of parts and economies of manufacture, will become more apparent upon consideration of the following description and the appended claims with reference to the accompanying drawings, all of which form a part of this specification, wherein like reference numerals designate corresponding parts in the various figures. As used in the specification and in the claims, the singular form of 'a', 'an', and 'the' include plural referents unless the context clearly dictates otherwise.

The present disclosure generally relates to devices, systems, and methods for performing enhanced maintenance operations through situational knowledge. A system described herein may record fault information for an aircraft during a flight, including relevant data associated with faults at the time the fault occurs.

When the aircraft lands, a maintenance crew may attend to the aircraft to make any necessary repairs. The maintenance crew may access the information related to the faults that occurred during the flight. For each detected fault, the maintenance crew may have a checklist of procedures to perform to isolate and correct the fault. The checklist may come from a maintenance manual associated with the aircraft.

The checklist of procedures may involve gathering data associated with the aircraft and determining whether the data is within a normal or acceptable range. The system disclosed herein may provide this data to the maintenance crew, either by outputting current data of the aircraft, or by outputting the data recorded at the time the flight occurred. Based on this data, the maintenance crew may perform the procedures of the checklist. The maintenance crew may then identify and correct the fault.

Referring to <FIG>, an illustrative aircraft system <NUM> is schematically depicted. In the illustrated embodiment of <FIG>, the aircraft system <NUM> generally includes an aircraft <NUM>, which may include a fuselage <NUM>, wing assemblies <NUM>, and one or more engines <NUM>. While <FIG> depicts the aircraft <NUM> as being a fixed-wing craft having two wing assemblies <NUM> with one engine <NUM> mounted on each wing assembly <NUM> (two engines <NUM> total), other configurations are contemplated. For example, other configurations and/or aerial vehicles may include high speed compound rotary-wing aircraft with supplemental translational thrust systems, dual contra-rotating, coaxial rotor system aircraft, turboprops, tilt-rotors, tilt-wing aircraft, conventional take-off and landing aircraft and other turbine driven machines will also benefit from the present disclosure. Furthermore, other configurations may include more than two wing assemblies <NUM>, more than two engines <NUM> (e.g., trijets, quadjets, etc.), engines <NUM> that are not mounted to a wing assembly <NUM> (e.g., mounted to the fuselage <NUM>, mounted to the tail, mounted to the nose, etc.), non-fixed wings (e.g., rotary wing aircraft), and/or the like.

Turning back to the illustrated aircraft system depicted in <FIG>, as shown, a control mechanism <NUM> for controlling the aircraft <NUM> is included in the cockpit <NUM> and may be operated by a pilot located therein. It should be understood that the term "control mechanism" as used herein is a general term used to encompass all aircraft control components, particularly those typically found in the cockpit <NUM>.

A plurality of additional aircraft systems <NUM> that enable proper operation of the aircraft <NUM> may also be included in the aircraft <NUM> as well as an engine control system <NUM>, and a communication system having the aircraft wireless communications link <NUM>. The additional aircraft systems <NUM> may generally be any systems that effect control of one or more components of the aircraft <NUM>, such as, for example, cabin pressure controls, elevator controls, rudder controls, flap controls, spoiler controls, landing gear controls, heat exchanger controls, and/or the like. In some embodiments, the avionics of the aircraft <NUM> may be encompassed by one or more of the additional aircraft systems <NUM>. The aircraft wireless communications link <NUM> may generally be any air-to-ground communication system now known or later developed. Illustrative examples of the aircraft wireless communications link <NUM> include, but are not limited to, a transponder, a very high frequency (VHF) communication system, an aircraft communications addressing and reporting system (ACARS), a controller-pilot data link communications (CPDLC) system, a future air navigation system (FANS), and/or the like. The engine control system <NUM> may be communicatively coupled to the plurality of aircraft systems <NUM> and the engines <NUM>. In some embodiments, the engine control system <NUM> may be mounted on one or more of the engines <NUM> or mounted within the aircraft <NUM> and communicatively coupled to the engines <NUM>. While the embodiment depicted in <FIG> specifically refers to the engine control system <NUM>, it should be understood that other controllers may also be included within the aircraft <NUM> to control various other aircraft systems <NUM> that do not specifically relate to the engines <NUM>.

The engine control system <NUM> generally includes one or more components for controlling each of the engines <NUM>, such as, for example, a diagnostic computer, an engine-related digital electronic unit that is mounted on one or more of the engines <NUM> or the aircraft <NUM>, and/or the like. The engine control system <NUM> may also be referred to as a digital engine control system. Illustrative other components within the engine control system that may function with the engine control system <NUM> and may require software to operate include, but are not limited to, an electronic control unit (ECU) (<NUM>, <FIG>) and other controller devices. The software implemented in any one of these components may be software that is distributed as described herein.

The engine control system <NUM> may also be connected with other controllers of the aircraft <NUM>. In embodiments, the engine control system <NUM> may include a processor <NUM> and/or a non-transitory memory component <NUM>, including non-transitory memory. In some embodiments, the non-transitory memory component <NUM> may include random access memory (RAM), read-only memory (ROM), flash memory, or one or more different types of portable electronic memory, such as discs, DVDs, CD-ROMs, or the like, or any suitable combination of these types of memory. The processor <NUM> may carry out one or more programming instructions stored on the non-transitory memory component <NUM>, thereby causing operation of the engine control system <NUM>. That is, the processor <NUM> and the non-transitory memory component <NUM> within the engine control system <NUM> may be operable to carry out the various processes described herein with respect to the engine control system <NUM>, including operating various components of the aircraft <NUM> (such as the engine <NUM> and/or components thereof), monitoring the health of various components of the aircraft <NUM> (e.g., the engine <NUM> and/or components thereof), monitoring operation of the aircraft <NUM> and/or components thereof, installing software, installing software updates, modifying a record in a distributed ledger to indicate that software has been installed, and/or updated, carrying out processes according to installed and/or updated software, and/or the like.

In some embodiments, the engine control system <NUM> may include a full authority digital engine control (FADEC) system. Such a FADEC system can include various electronic components, one or more sensors, and/or one or more actuators that control each of the engines <NUM>. In particular embodiments, the FADEC system includes an ECU, as well as one or more additional components that are configured to control various aspects of performance of the engines <NUM>. The FADEC system generally has full authority over operating parameters of the engines <NUM> and cannot be manually overridden. A FADEC system generally functions by receiving a plurality of input variables of a current flight condition, including, but not limited to, air density, throttle lever position, engine temperature, engine pressure, and/or the like. The inputs are received, analyzed, and used to determine operating parameters such as, but not limited to, fuel flow, stator vane position, bleed valve position, and/or the like. The FADEC system may also control a start or a restart of the engines <NUM>. The operating parameters of the FADEC can be modified by installing and/or updating software, such as the software that is distributed by the aircraft system <NUM> described herein. As such, the FADEC can be programmatically controlled to determine engine limitations, receive engine health reports, receive engine maintenance reports and/or the like to undertake certain measures and/or actions in certain conditions.

The software run by the engine control system <NUM> (e.g., executed by the processor <NUM> and stored within the non-transitory memory component <NUM>) may include a computer program product that includes machine-readable media for carrying or having machine-executable instructions or data structures. Such machine-readable media may be any available media, which can be accessed by a general purpose or special purpose computer or other machine with a processor. Generally, such a computer program may include routines, programs, objects, components, data structures, algorithms, and/or the like that have the technical effect of performing particular tasks or implementing particular abstract data types. Machine-executable instructions, associated data structures, and programs represent examples of program code for executing the exchange of information as disclosed herein. Machine-executable instructions may include, for example, instructions and data, which cause a general purpose computer, special purpose computer, or special purpose processing machine to perform a certain function or group of functions. In some embodiments, the computer program product may be provided by a component external to the engine control system <NUM> and installed for use by the engine control system <NUM>. For example, the computer program product may be provided by the ground support equipment <NUM>, as described in greater detail herein. The computer program product may generally be updatable via a software update that is received from one or more components of the aircraft system <NUM>, such as, for example, the ground support equipment <NUM>, as described in greater detail herein. The software is generally updated by the engine control system <NUM> by installing the update such that the update supplements and/or overwrites one or more portions of the existing program code for the computer program product. The software update may allow the computer program product to more accurately diagnose and/or predict faults, provide additional functionality not originally offered, and/or the like.

In embodiments, each of the engines <NUM> may include a fan <NUM> and one or more sensors for sensing various characteristics of the fan <NUM> during operation of the engines <NUM>. Illustrative examples of the one or more sensors include, but are not limited to, a fan speed sensor <NUM>, a temperature sensor <NUM>, and a pressure sensor <NUM>. The fan speed sensor <NUM> is generally a sensor that measures a rotational speed of the fan <NUM> within the engine <NUM>. The temperature sensor <NUM> may be a sensor that measures a fluid temperature within the engine <NUM> (e.g., an engine air temperature), a temperature of fluid (e.g., air) at an engine intake location, a temperature of fluid (e.g., air) within a compressor, a temperature of fluid (e.g., air) within a turbine, a temperature of fluid (e.g., air) within a combustion chamber, a temperature of fluid (e.g., air) at an engine exhaust location, a temperature of cooling fluids and/or heating fluids used in heat exchangers in or around an engine, and/or the like. The pressure sensor <NUM> may be a sensor that measures a fluid pressure (e.g., air pressure) in various locations in and/or around the engine <NUM>, such as, for example, a fluid pressure (e.g., air pressure) at an engine intake, a fluid pressure (e.g., air pressure) within a compressor, a fluid pressure (e.g., air pressure) within a turbine, a fluid pressure (e.g., air pressure) within a combustion chamber, a fluid pressure (e.g., air pressure) at an engine exhaust location, and/or the like.

In some embodiments, each of the engines <NUM> may have a plurality of sensors associated therewith (including one or more fan speed sensors <NUM>, one or more temperature sensors <NUM>, and/or one or more pressure sensors <NUM>). That is, more than one of the same type of sensor may be used to sense characteristics of an engine <NUM> (e.g., a sensor for each of the different areas of the same engine <NUM>). In some embodiments, one or more of the sensors may be utilized to sense characteristics of more than one of the engines <NUM> (e.g., a single sensor may be used to sense characteristics of two engines <NUM>). The engines <NUM> may further include additional components not specifically described herein, and may include one or more additional sensors incorporated with or configured to sense such additional components in some embodiments.

In embodiments, each of the sensors (including, but not limited to, the fan speed sensors <NUM>, the temperature sensors <NUM>, and the pressure sensors <NUM>) may be communicatively coupled to one or more components of the aircraft <NUM> such that signals and/or data pertaining to one or more sensed characteristics are transmitted from the sensors for the purposes of determining, detecting, and/or predicting a fault, as well as completing one or more other actions in accordance with software that requires sensor information. As indicated by the dashed lines extending between the various sensors (e.g., the fan speed sensors <NUM>, the temperature sensors <NUM>, and the pressure sensors <NUM>) and the aircraft systems <NUM> and the engine control system <NUM> in the embodiment depicted in <FIG>, the various sensors may be communicatively coupled to the aircraft systems <NUM> and/or the engine control system <NUM> in some embodiments. As such, the various sensors may be communicatively coupled via wires or wirelessly to the aircraft systems <NUM> and/or the engine control system <NUM> to transmit signals and/or data to the aircraft systems <NUM> and/or the engine control system <NUM> via an aircraft bus.

An aircraft bus may enable an aircraft and/or one or more components of the aircraft to interface with one or more external system through wireless or wired means. An aircraft bus as used herein may be formed from any medium that is configured to transmit a signal. As non-limiting examples, the aircraft bus is formed of conductive wires, conductive traces, optical waveguides, or the like. The aircraft bus may also refer to the expanse in which electromagnetic radiation and their corresponding electromagnetic waves are propagated. Moreover, the aircraft bus may be formed from a combination of mediums configured to transmit signals. In one embodiment, the aircraft bus includes a combination of conductive traces, conductive wires, connectors, and buses that cooperate to permit the transmission of electrical data signals to and from the various components of the engine control system <NUM>. Additionally, it is noted that the term "signal" means a waveform (e.g., electrical, optical, magnetic, mechanical or electromagnetic) configured to travel through a medium, such as DC, AC, sinusoidal-wave, triangular-wave, square-wave, vibration, and the like.

For example, an interconnectivity of components coupled via a network, may include a wide area network, such as the internet, a local area network (LAN), a mobile communications network, a public service telephone network (PSTN) and/or other network and may be configured to electronically connect components. The illustrative components that may be connected via the network include, but are not limited to, a ground system <NUM> in communication with the aircraft <NUM> (e.g., via a ground wireless communications link <NUM> and an aircraft wireless communications link <NUM>), and/or a ground support equipment <NUM> via a wired or wireless system.

It should be understood that the aircraft <NUM> merely represents one illustrative embodiment that may be configured to implement embodiments or portions of embodiments of the devices, systems, and methods described herein. During operation, by way of non-limiting example, the control mechanism <NUM> may be utilized to operate one or more of the aircraft systems <NUM>. Various sensors, including, but not limited to, the fan speed sensors <NUM>, the temperature sensors <NUM>, and/or the pressure sensors <NUM> may output data relevant to various characteristics of the engine <NUM> and/or the other aircraft systems <NUM>. The engine control system <NUM> may utilize inputs from the control mechanism <NUM>, the fan speed sensors <NUM>, the temperature sensors <NUM>, the pressure sensors <NUM>, the various aircraft systems <NUM>, one or more database, and/or information from airline control, flight operations, or the like to diagnose, detect, and/or predict faults that airline maintenance crew may be unaware of. Among other things, the engine control system <NUM> may analyze the data output by the various sensors (e.g., the fan speed sensors <NUM>, the temperature sensors <NUM>, the pressure sensors <NUM>, etc.), over a period of time to determine drifts, trends, steps, or spikes in the operation of the engines <NUM> and/or the various other aircraft systems <NUM>. The engine control system <NUM> may also analyze the system data to determine historic pressures, historic temperatures, pressure differences between the plurality of engines <NUM> on the aircraft <NUM>, temperature differences between the plurality of engines <NUM> on the aircraft <NUM>, and/or the like, and to diagnose, detect, and/or predict faults in the engines <NUM> and/or the various other aircraft systems <NUM> based thereon. The aircraft wireless communications link <NUM> and the ground wireless communications link <NUM> may transmit data such that data and/or information pertaining to the fault may be transmitted off the aircraft <NUM>.

While the embodiment of <FIG> specifically relates to components within an aircraft <NUM>, the present disclosure is not limited to such. That is, the various components depicted with respect to the aircraft <NUM> may be incorporated within various other types of craft and may function in a similar manner to deliver and install new software and/or updated software to the engine control system <NUM> as described herein. For example, the various components described herein with respect to the aircraft <NUM> may be present in watercraft, spacecraft, and/or the like without departing from the scope of the present disclosure.

Furthermore, it should be appreciated that, although a particular aerial vehicle has been illustrated and described, other configurations and/or aerial vehicles, such as high speed compound rotary-wing aircraft with supplemental translational thrust systems, dual contra-rotating, coaxial rotor system aircraft, turboprops, tilt-rotors, tilt-wing aircraft, conventional take-off and landing aircraft and other turbine driven machines will also benefit from the present disclosure.

Still referring to <FIG>, the ground system <NUM> is generally a transmission system located on the ground that is capable of transmitting and/or receiving signals to/from the aircraft <NUM>. That is, the ground system <NUM> may include a ground wireless communications link <NUM> that is communicatively coupled to the aircraft wireless communications link <NUM> wirelessly to transmit and/or receive signals and/or data. In some embodiments, the ground system <NUM> may be an air traffic control (ATC) tower and/or one or more components or systems thereof. Accordingly, the ground wireless communications link <NUM> may be a VHF communication system, an ACARS unit, a CPDLC system, FANS, and/or the like. Using the ground system <NUM> and the ground wireless communications link <NUM>, the various non-aircraft components depicted in the embodiment of <FIG> may be communicatively coupled to the aircraft <NUM>, even in instances where the aircraft <NUM> is airborne and in flight, thereby allowing for on-demand transmission of software and/or software updates whenever such software and/or software updates may be needed. However, it should be understood that the embodiment depicted in <FIG> is merely illustrative. In other embodiments, the aircraft <NUM> may be communicatively coupled to the various other components of the aircraft system <NUM> when on the ground and physically coupled to one of the components of the aircraft system <NUM>, such as, for example, the ground support equipment <NUM>.

The ground support equipment (GSE) <NUM> is an external equipment used to support and test the engine control system <NUM> and/or other components of the aircraft <NUM>. The ground support equipment <NUM> is configured to provide software updates to the engine control system <NUM> and download data obtained by the engine control system <NUM> during a flight. As another non-limiting example, the GSE <NUM> may include production support equipment for restricted data monitoring, test support equipment for comprehensive data monitoring and changing adjustable parameters, and integration test rigs for system and software testing. In embodiments, the GSE <NUM> may be connected to the engine control system <NUM> via wired local area network, or Ethernet. The GSE <NUM> may communicate with the engine control system <NUM> according to Ethernet protocols. The GSE <NUM> may be a portable maintenance access terminal. The GSE <NUM> may test a ballistic mode of the aircraft by directly communicating with the ECU <NUM> of the engine control system <NUM>, which is described in more detail herein.

<FIG> will now describe illustrative embodiments of a fault isolation system <NUM>. Turning to <FIG>, an illustrative system diagram of the electronic control unit (ECU) <NUM> is depicted. The ECU <NUM> may include a computing device having the ability to operate and interface with components of the aircraft system <NUM>, for example, the components described herein with respect to <FIG>. The ECU <NUM> may include a processor <NUM>, input/output hardware <NUM>, network interface hardware <NUM>, a data storage component <NUM>, and a memory component <NUM>. In the illustrated example, the memory component <NUM> comprises the fault isolation system <NUM>. In other examples, the fault isolation system <NUM> may reside in other components of the aircraft system <NUM> such as an aircraft health management unit.

The processor <NUM> may include any processing component(s) configured to receive and execute instructions (such as from the data storage component <NUM> and/or the memory component <NUM>). The instructions may be in the form of a machine-readable instruction set stored in the data storage component <NUM> and/or the memory component <NUM>. The input/output hardware <NUM> may include a display, keyboard, mouse, printer, camera, microphone, speaker, and/or other device for receiving, sending, and/or presenting data. The network interface hardware <NUM> may include any wired or wireless networking hardware, such as a modem, LAN port, Wi-Fi card, WiMax card, mobile communications hardware, and/or other hardware for communicating with other networks and/or devices. For example, the network interface hardware <NUM> may include a transceiver.

The memory component <NUM> may be machine-readable memory (which may also be referred to as a non-transitory processor readable memory). The memory component <NUM> may be configured as volatile and/or nonvolatile memory and, as such, may include random access memory (including SRAM, DRAM, and/or other types of random access memory), flash memory, registers, compact discs (CD), digital versatile discs (DVD), and/or other types of storage components. Additionally, the memory component <NUM> may be configured to store a fault detection module <NUM>, a fault parameter identification module <NUM>, a parameter measurement module <NUM>, a fault output module <NUM>, a parameter output module <NUM>, a checklist output module <NUM>, and/or other modules that may be necessary for enabling operation of the fault isolation system <NUM> (each of which may be embodied as a computer program, firmware, or hardware, as an example).

A local interface <NUM> is also included in <FIG> and may be implemented as a bus or other interface to facilitate communication among the components of the ECU <NUM>.

The data storage component <NUM> may reside local to and/or remote from the ECU <NUM> and may be configured to store one or more pieces of data for access by the ECU <NUM> and/or other components. As illustrated in <FIG>, the data storage component <NUM> may store fault parameters 238a, fault data 238b, checklist data 238c, and/or other data sets, as disclosed in further detail below.

As explained above, the aircraft system <NUM> may include a variety of sensors that may detect faults during a flight. In known systems, the faults may be identified and recorded by the ECU <NUM> or other components of the aircraft system <NUM> based on the sensor data. Then, when the aircraft <NUM> lands, a maintenance crew may view the faults that were recorded and perform the procedures outlined in a maintenance manual for each of the identified faults. A maintenance manual may include a set of procedures for a maintenance crew to follow in order to perform fault isolation and identify a source of a fault on the aircraft. Different types of faults may require different procedures to be performed. These procedures often involve measuring values of various parameters of the aircraft <NUM> (e.g., voltages, currents, temperatures, etc.). Based on these measured values, the maintenance manual may identify what the source of a fault is and may identify a particular part that should be changed and/or what other corrective action should be taken. For example, if a particular voltage is too high, the maintenance manual may recommend replacing a particular part.

However, after the aircraft <NUM> lands, the conditions are not the same as when the aircraft <NUM> was in flight. For example, the temperature on the ground may be significantly higher than the temperature when the aircraft <NUM> is in flight at <NUM>,<NUM> feet. In addition, other systems of the aircraft <NUM> may no longer be in use or may be in a different state than when the aircraft <NUM> was in flight. As such, it may be difficult to diagnose the source of faults detected during a flight when the conditions that occurred when the fault was detected cannot be recreated on the ground.

As such, the aircraft system <NUM> of the present disclosure measures important values of parameters associated with the aircraft <NUM> when a fault is detected, as disclosed herein. Then, when the aircraft <NUM> lands, a maintenance crew may review the measured values recorded when a fault occurred in order to perform the procedures of a maintenance manual. Thus, a more accurate fault isolation method may be performed based on conditions of the aircraft <NUM> when a fault occurred during flight rather than conditions of the aircraft <NUM> on the ground.

Referring still to <FIG>, the fault parameters 238a may comprise a list of aircraft parameters associated with a variety of faults. The fault parameters 238a may be taken from a maintenance manual associated with the aircraft <NUM>. For example, for a particular fault, associated aircraft parameters may include a voltage of one component, and a temperature of another component. Any number of parameters may be associated with each fault. In embodiments, the parameters associated with a particular fault include the parameters that a maintenance manual indicates should be measured when the particular fault occurs. By storing the fault parameters 238a in the data storage component <NUM>, the fault isolation system <NUM> is able to measure appropriate values when a particular fault occurs for later use by a maintenance crew, as explained in further detail below.

The fault data 238b comprises measured values of parameters of the aircraft <NUM> when a fault is detected. The fault data 238b is stored in the data storage component <NUM> such that a maintenance crew can review the data after the aircraft <NUM> lands. The fault data 238b is discussed in further detail below.

The checklist data 238c comprises data associated with a checklist to be performed when a fault is detected. As explained above, when a fault is detected, a maintenance manual includes a set of procedures to be performed to identify the source of the fault. This set of procedures may be in the form of a checklist. In some embodiments, the fault isolation system <NUM> may implement a smart checklist procedure based on the checklist data 238c. The smart checklist procedure is discussed in further detail below.

Referring still to <FIG>, the memory component <NUM> may include a fault detection module <NUM>, a fault parameter identification module <NUM>, a parameter measurement module <NUM>, a fault output module <NUM>, a parameter output module <NUM>, and a checklist output module <NUM>. As explained above, the ECU <NUM> may utilize data from one or more sensors of the aircraft system <NUM> to identify faults that occur during a flight. In embodiments, this fault detection may be performed by the fault detection module <NUM> using known techniques. The fault detection module <NUM> may store information regarding detected faults as fault data 238b in the data storage component <NUM>.

The fault parameter identification module <NUM> may identify one or more parameters associated with a fault detected by the fault detection module <NUM>. Specifically, the fault parameter identification module <NUM> may access the fault parameters 238a stored in the data storage component <NUM> to identify the parameters associated with the fault detected by the fault detection module <NUM>. As discussed above, the parameters associated with a particular fault comprise values associated with one or more components of the aircraft <NUM> to be measured to assist in a later diagnosis of the fault. For example, the parameters may comprise voltages, currents, pressures, temperatures, positions, or other values associated with components of the aircraft <NUM>.

The parameter measurement module <NUM> measures values of the parameters identified by the fault parameter identification module <NUM>. The parameter measurement module <NUM> may utilize one or more sensors of the aircraft system <NUM> to perform the measurement of the parameters. The parameter measurement module <NUM> may measure values of the parameters at the same time or shortly after the fault detection module <NUM> detects a fault. Thus, the measured values may comprise values for the parameters when a fault occurs. As such, when a maintenance crew performs maintenance on the aircraft <NUM> after it lands, the maintenance crew may diagnose a fault based on data at the time a fault occurred, rather than based on data measured while the aircraft <NUM> is on the ground in different conditions than when the fault occurred. The parameter measurement module <NUM> may store the measured values of parameters as fault data 238b in the data storage component <NUM>.

After the aircraft <NUM> completes a flight and lands, and a maintenance crew performs maintenance on the aircraft <NUM>, the fault output module <NUM> may access the fault data 238b in the data storage component <NUM> and output an indication of each fault that occurred during the flight. In one example, the fault information may be output to the GSE <NUM>. In another example, the fault information may be output to the input/output hardware <NUM>. In still other examples, the fault information may be output to other devices using the network interface hardware <NUM>. The fault output module <NUM> outputs an indication of each fault that occurred during the flight such that a maintenance crew can view the information and perform appropriate maintenance.

The parameter output module <NUM> may access the fault data 238b in the data storage component <NUM> and output the values of the parameters associated with faults that were measured when those faults occurred. Similar to the fault output module <NUM>, the parameter output module <NUM> may output the measured parameter values to the GSE <NUM>, or the input/output hardware <NUM>, or the network interface hardware <NUM>, or otherwise such that they may be viewed by a maintenance crew. The parameters output by the parameter output module <NUM> and the faults output by the fault output module <NUM> may allow a maintenance crew to perform maintenance on the aircraft <NUM> to identify the source of the faults. In some examples, the maintenance crew may use a checklist in a maintenance manual to perform maintenance. In other examples, the maintenance crew may utilize a smart checklist procedure, as discussed below.

Referring still to <FIG>, the checklist output module <NUM> may access the checklist data 238c from the data storage component <NUM> and output checklist information such that it may be viewed by a maintenance crew. The checklist output module <NUM> may output checklist information associated with the faults output by the fault output module <NUM>. Thus, the maintenance crew may perform the procedures included in the checklist information output by the checklist output module <NUM>. This may include procedures to be performed by the maintenance crew to determine the source of the identified faults. As explained above, these procedures may be based on a maintenance module for the aircraft <NUM>.

According to the invention, the checklist output module <NUM> implements a smart checklist procedure. In this example, the checklist output module <NUM> may access the checklist data 238c and identifies one or more checklists associated with the faults output by the fault output module <NUM>. These checklists include procedures to be performed by a maintenance crew These procedures may involve checking the values of one or more parameters and determining whether the values are within acceptable ranges. The acceptable ranges may be defined by the checklists. Thus, for each checklist output by the checklist output module <NUM>, the checklist output module <NUM> may automatically perform certain steps of the checklist. Specifically, for steps of the checklist that comprise checking whether parameter values are within acceptable ranges, the checklist output module <NUM> may access the fault data 238b in the data storage component <NUM> and determine whether parameter values identified in the checklist are within acceptable ranges. Then, the checklist output module <NUM> outputs an indication of the parameter values that are within acceptable ranges and a corresponding indication that those values do not need to be checked by the maintenance crew. As such, the number of steps of the checklist that need to be performed by the maintenance crew is reduced, thereby increasing the efficiency of the maintenance crew.

Turning to <FIG>, an illustrative flow diagram depicting a method of performing fault isolation for an aircraft is shown. The method of <FIG> may be performed by the ECU <NUM>.

At step <NUM>, the fault detection module <NUM> identifies a fault that occurred during a flight of the aircraft <NUM>. The fault detection module <NUM> may identify the fault based on data captured by one or more sensors of the aircraft system <NUM>. The fault detection module <NUM> may record the detected fault as fault data 238b in the data storage component <NUM>. In addition, the fault output module <NUM> may outputs an indication of the fault. The fault output module <NUM> may output the indication of the fault to a maintenance crew when the aircraft <NUM> is on the ground after it has landed. The fault output module <NUM> may use the fault data 238b stored in the data storage component <NUM> to output the indication of the fault. The indication of the fault output by the fault output module <NUM> may include a time that the fault occurred, a type of fault that occurred, and other information related to the fault.

At step <NUM>, the fault parameter identification module <NUM> identifies a first set of parameters associated with the aircraft <NUM> based on the identification of the fault. The fault parameter identification module <NUM> may identify the set of parameters based on the fault parameters 238a stored in the data storage component <NUM>.

At step <NUM>, the fault isolation system <NUM> automatically determines values of the first set of parameters to obtain a first set of measured values. In a traditional mode of operation, the parameter measurement module <NUM> measures current values of the first set of parameters. As such, in the traditional mode of operation, the first set of measured values comprise values of parameters (e.g., voltages, temperatures) measured while the aircraft <NUM> is being serviced by a maintenance crew. This is similar to traditional maintenance where the maintenance crew would obtain measurements of parameters while the aircraft is on the ground. However, in the traditional mode of operation of the fault isolation system <NUM>, the values of the parameters are measured and output to the maintenance crew automatically. Thus, the maintenance crew does not need to manually obtain the measurements. As such, in the traditional mode of operation, the second set of measured values of the set of parameters comprises current values of the parameters. Specifically, the parameter measurement module <NUM> may record current values of the parameters associated with the fault and the parameter output module <NUM> may output the measured value of the parameters to the maintenance crew.

Alternatively, in a historic mode of operation, the parameter output module <NUM> outputs historic values of the parameters such that the first set of measured values comprise values of the first set of parameters measured at a time that the fault occurred. Specifically, the parameter output module <NUM> outputs values of the parameters recorded at the time the fault occurred. As such, the maintenance crew may be able to obtain a more accurate picture of the conditions that existed at the time the fault occurred. This may allow the maintenance crew to better diagnose and correct the fault. Thus, in the historic mode of operation, the second set of the measured values of the set of parameters comprise the recorded first set of measured values. In either the traditional or historic mode of operation, the first set of measured values may be recorded as fault data 238b in the data storage component <NUM>.

At step <NUM>, the checklist output module <NUM> determines whether the first set of measured values are within acceptable ranges. The acceptable ranges for the first set of measured values may be based on a maintenance manual associated with the aircraft. Then, at step <NUM>, the checklist output module identifies a source of the fault based on the determination of whether the first set of measured values are within the acceptable ranges. The source of the fault may comprise a component of the aircraft.

Turning to <FIG>, an illustrative flow diagram depicting another method of performing fault isolation for an aircraft is shown. The method of <FIG> may be performed by the ECU <NUM>.

At step <NUM>, the fault detection module <NUM> identifies a fault that occurred during a flight of the aircraft <NUM>. The fault detection module <NUM> may identify the fault based on data captured by one or more sensors of the aircraft system <NUM>.

At step <NUM>. the fault isolation system <NUM> automatically determines values of the first set of parameters to obtain a first set of measured values. As disclosed above, the fault isolation system <NUM> may operate in either traditional or historic mode. In the traditional mode of operation, the first measured values comprise current values of the first set of parameters. In the historic mode of operation, the first set of measured values comprise values of the first set of parameters measured at a time that the fault occurred.

At step <NUM>, the checklist output module <NUM> determines whether the first set of measured values are within acceptable ranges. Then at step <NUM>, the checklist output module <NUM> determines one or more maintenance procedures to be performed based on the first set of measured values and the determination of whether the first set of measured values are within the acceptable ranges. The one or more maintenance procedures comprise maintenance procedures from a checklist based on a maintenance manual associated with the aircraft.

The checklist output module <NUM> also outputs the checklist, which comprises a set of procedures to be performed to identify a source of the fault, wherein the checklist is based on a type of the fault. The checklist comprises checklist data 238c stored in the data storage component <NUM>. In some examples, the checklist is based on a maintenance manual associated with the aircraft <NUM>. The checklist data 238c comprises a set of procedures to be performed by a maintenance crew based on the identified fault. One or more of these procedures may comprise determining whether one or more values of parameters are within an acceptable range.

In some examples, the fault isolation system <NUM> may automatically perform certain steps of a checklist associated with the fault by determining whether one or more parameter values are within an acceptable range. The fault isolation system <NUM> then outputs to the maintenance crew a subset of the one or more maintenance procedures from the checklist that do not need to be performed, wherein the subset is based on the first set of measured values and the determination of whether the first set of measured values are within the acceptable ranges. When the fault isolation system <NUM> is operating in the traditional mode of operation, it may determine that certain procedures of a checklist do not need to be performed based on current parameter values. Alternatively, when the fault isolation system <NUM> is operating in the historic mode of operation, it may determine that certain procedures of a checklist do not need to be performed based on historic parameter values (e.g., parameter values measured at the time of the fault).

In the traditional mode of operation, the parameter measurement module <NUM> may measure current values of the set of parameters. That is, the parameter measurement module <NUM> may measure parameter values while the aircraft <NUM> is on the ground. The checklist output module <NUM> then determine one or more procedures of the set of procedures of the checklist that do not need to be performed based on the measured current values of the set of parameters. Specifically, the checklist output module <NUM> may determine that the current value of one or more parameters are within an acceptable range and, thus, procedures comprising checking whether these values are within an acceptable range need not be performed. The checklist output module <NUM> may then output the one or more procedures that do not need to be performed.

In the historic mode of operation, the checklist output module <NUM> determine one or more procedures of the set of procedures of the checklist that do not need to be performed based on the first set of measured values. That is, the checklist output module <NUM> may determine procedures that do not need to be performed based on the historic values of parameters measured at the time the fault occurred. Specifically, the checklist output module <NUM> may determine that historic value of one or more parameters are within an acceptable range and, thus, procedures comprising checking whether these values are within an acceptable range need not be performed. The checklist output module <NUM> may then output the one or more procedures that do not need to be performed.

Turning to <FIG>, an illustrative flow diagram depicting a method of performing aircraft maintenance for an aircraft that may be performed by an aircraft maintenance crew is shown.

At step <NUM>, the maintenance crew receives an indication of a fault that occurred during a flight of the aircraft. When the fault isolation system <NUM> is operating in the traditional mode of operation, the measured values of the set of parameters comprise current values of the parameters. That is, values of the parameters at the time that maintenance is being performed. When the fault isolation system <NUM> is operating in the historic mode of operation, the measured values of the set of parameters comprise historic values of the first set of parameters recorded at a time when the fault occurred.

At step <NUM>, the maintenance crew receives a checklist comprising a set of procedures to be performed to identify a source of the fault, wherein the checklist is based on a type of fault that occurred. The checklist may be based on a maintenance manual associated with the aircraft <NUM>. In some examples, the maintenance crew receives a subset of the procedures of the checklist that do not need to be performed, wherein the subset is based on the measured values of the one or more parameters of the set of parameters. In the traditional mode of operation, the subset is based on current values of the one or more parameters of the set of parameters. In the historic mode of operation, the subset is based on historic values of the first set of parameters recorded at a time when the fault occurred.

At step <NUM>, the maintenance crew receives measured values of a set of parameters associated with the aircraft, wherein the set of parameters is based on the checklist. In some examples, the maintenance crew may receive parameters associated with one procedure on the checklist.

At step <NUM>, the maintenance crew determines whether the measured values of the set of parameters are within acceptable ranges. The acceptable ranges for the measured values of the set of parameters may be identified in the checklist and/or in a maintenance manual associated with the aircraft <NUM>.

At step <NUM>, the maintenance crew determines whether additional procedures remain on the checklist. If the maintenance crew determines that additional procedures remain on the checklist (yes at step <NUM>), then control returns to step <NUM> and additional parameter values are received associated with another procedure on the checklist. If the maintenance crew determines that no additional procedures remain on the checklist (no at step <NUM>), then control passes to step <NUM>.

At step <NUM>, the maintenance crew identifies a source of the fault based on the determination of whether the measured values of the set of parameters are within acceptable ranges. In some examples, the source of the fault comprises a component of the aircraft. In these examples, the maintenance crew may replace the faulty component.

The functional blocks and/or flowchart elements described herein may be translated onto machine-readable instructions. As non-limiting examples, the machine-readable instructions may be written using any programming protocol, such as: descriptive text to be parsed (e.g., such as hypertext markup language, extensible markup language, etc.), (ii) assembly language, (iii) object code generated from source code by a compiler, (iv) source code written using syntax from any suitable programming language for execution by an interpreter, (v) source code for compilation and execution by a just-in-time compiler, etc. Alternatively, the machine-readable instructions may be written in a hardware description language (HDL), such as logic implemented via either a field programmable gate array (FPGA) configuration or an application-specific integrated circuit (ASIC), or their equivalents. Accordingly, the functionality described herein may be implemented in any conventional computer programming language, as pre-programmed hardware elements, or as a combination of hardware and software components.

Claim 1:
A computer implemented method of performing fault isolation for an aircraft (<NUM>) comprising:
identifying (<NUM>, <NUM>, <NUM>) a fault that occurred during a flight of the aircraft;
identifying (<NUM>) a first set of parameters associated with the aircraft based on the identification of the fault;
automatically determining (<NUM>, <NUM>, <NUM>) values of the first set of parameters to obtain a first set of measured values, wherein the first set of measured values comprise values of the first set of parameters measured at a time that the fault occurred;
determining (<NUM>) whether the first set of measured values are within acceptable ranges;
wherein the method is characterized by using a checklist output module (<NUM>) to:
determine a checklist comprising a first set of maintenance procedures to be performed by a maintenance crew to identify a source of a fault, wherein the checklist is based on a type of the fault and a maintenance manual associated with the aircraft; and
determine a subset of the first set of maintenance procedures that does not need to be performed based on the determination of whether the first set of measured values are within the acceptable ranges.