Method and system for calculating accumulated damage on a structural component

Abstract

Methods for calculating the accumulated damage on a component of an aircraft, comprise either 1) determining the weight of the aircraft at a predetermined point of a flight cycle; 2) obtaining manoeuv

Description

FIELD OF THE INVENTION 5 The present invention relates to a method of calculating the accumulated damage of a component for an aircraft. BACKGROUND OF THE INVENTION 10 Structural components of aircraft subassemblies must function according to specification while they are in use to ensure the safety of the passengers. During every flight cycle the components may experience aging due to applied cyclic loading, i.e. applying and removing loads on the components. This aging may be 15 referred to as fatigue damage. The accumulation of fatigue damage may result in cracks or other damage on the components that can lead the structural component to fracture. Each component of an aircraft may have a known quantity of damage, after which failure may be expected. This may be referred to as a fatigue life. Once a component reaches the end of its fatigue life, there may be a requirement to 20 replace the component. The time left before the component needs to be replaced is known as Remaining Useful Life (RUL) and may be measured in flight hours or flight cycles. Conventionally, the RUL of a component is calculated based on a number of 25 previously performed flight cycles, and the fatigue life of the component as calculated during the design phase. The predefined fatigue life of the component may be calculated based on a worst case scenario for each flight cycle, such as a high number of manoeuvres and a maximum aircraft weight. 30 The conventional methods of calculating the RUL may result in shorter RUL than necessary, because for certain flight cycles the component may experience a 09 01 25 a spectrum than previously predicted. This may result in less fatigue damage than assumed in the calculation of the fatigue life during the design phase of the component. A shorter than necessary RUL may result in a premature replacement of the component, resulting in material waste and increased costs. 5 SUMMARY OF THE INVENTION A first aspect of the invention provides a method for calculating the accumulated damage on a component of an aircraft, the method comprising: determining the 10 weight of the aircraft at a predetermined point of a flight cycle; calculating, based at least partially on the determined weight of the aircraft, a calculated damage inflicted on the component during the flight cycle; obtaining historic data indicating the damage inflicted on the component before the flight cycle; and calculating the accumulated damage of the component, based on the calculated damage inflicted 15 on the component during the flight cycle and the historic data. Advantageously, the invention enables a more accurate way of assessing the accumulated damage or fatigue damage of a component by assessing the damage on the component for a given flight cycle not by using a theoretical weight of the 20 aircraft (which may be more or less than the actual weight) but by determining the actual weight and using the determined value in calculating the damage. This may result in a more accurate calculation of the damage inflicted on a component during the flight cycle and may in turn result in a more accurate calculation of the RUL, reducing waste and cost. 25 Determining the weight may be based on one or more of: the weight of the aircraft when empty; the weight of the cargo load at the predetermined point; the number of passengers at the predetermined point; and the quantity of fuel at the predetermined point. Any or all of the listed factors may be used in determining the 30 weight. The cargo load and quantity of fuel may be inferred from the number of passengers. The number of passengers may be the only retrieved data, as the weight of the aircraft may be known and the weight of cargo and fuel may be inferred from the number of passengers. By using such data, the weight at the 09 01 25 point may be calculated more accurately, improving the accuracy of the calculation of the calculated damage. The method may further comprise retrieving flight data from a flight computer 5 installed on the aircraft and determining the weight may be based at least partially on the flight data. By determining the weight based on flight data retrieved from a flight computer there is reduced need for further sensors or transceivers to be installed on the aircraft. 10 The method may further comprise retrieving system data from a remote system, and determining the weight may be based at least partially on the system data. By determining the weight based on system data retrieved from a remote system all the information needed to implement the method may be found externally to the aircraft. This may allow the method to be implemented externally to the aircraft 15 and may result in more efficient computation and a reduction in the use of aircraft resources used. The predetermined point of the flight cycle may be after loading and before taxiing for take-off. The predetermined point of the flight cycle may be after boarding is 20 complete and before pushback of the aircraft or immediately after the aircraft has been loaded according to flight specifications. This allows using the maximum actual weight of the aircraft for the relevant flight cycle, before any fuel has been used, and thus may result in more accurate calculation of the damage accrued on the component during the flight cycle. The maximum actual weight may be the most 25 relevant weight for assessing the fatigue damage to the component during taxiing. The method may further comprise obtaining manoeuvring data during the flight cycle and calculating the damage inflicted on the component during the flight cycle may be based at least partially on the manoeuvring data. Using manoeuvring data 30 in combination with the determined weight to calculate the damage inflicted on the component may provide a more accurate calculation allowing an operator to predict failure of a component more accurately. In this way, costs may be saved by avoiding premature replacement of a component and may improve safety by avoiding components being used beyond their expected life. 09 01 25 Calculating the damage inflicted on the component during the flight cycle may be based at least partially on the geometry of the component and / or the material of 5 the component. Taking the geometry and / or the material of the component into account when calculating the inflicted damage may result in a more accurate calculation as different materials may have different tolerances and different geometries may concentrate forces on specific points of the component. Thus the actual fatigue damage on the component may be calculated and the remaining 10 useful life of the component may be more accurately updated based on the actual fatigue damage accrued. The historic data may be based on historic manoeuvring data of at least one previously-completed flight cycle. This may result in a more accurate calculation of 15 damage, that may take into account the effect of manoeuvres performed on previous flight cycles on the structural health of the component. Additionally or alternatively, the historic data may be based on historic weight data of at least one previously-completed flight cycle. The weight as determined prior to the aircraft's departure for each flight may be logged and this log may be part of the historic 20 data. This may also result in a more accurate calculation of damage, that takes into account the weight of the aircraft during previous flight cycles and the damage on the structural health of the component due to said weight during previous flight cycles. 25 The component may be part of a landing gear. Landing gears are some of the subassemblies of an aircraft more affected by the weight of the aircraft or the manoeuvres performed by the aircraft while taxiing, and therefore it is especially beneficial to be able to calculate the fatigue damage of landing gears. 30 The method may further comprise calculating a remaining useful life of the component based on the calculated accumulated damage and at least one of certification data indicating the maximum allowable fatigue damage on the component and / or predicted values of one or more of the flight cycle attributes for 09 01 25 cles, wherein the predicted values have been determined based on previously completed flight cycles. A further aspect of the invention provides a system comprising a processor and a 5 memory for calculating the accumulated damage of a component for an aircraft. BRIEF DESCRIPTION OF THE DRAWINGS Embodiments of the invention will now be described with reference to the 10 accompanying drawings, in which: Figure 1 is a diagram of an aircraft. Figure 2 is a diagram of an aircraft landing gear. Figure 3 is a flowchart illustrating a known method of calculating the RUL of a component. 15 Figure 4 is a flowchart illustrating a method of calculating the accumulated damage to a component according to the invention. Figure 5 is a flowchart illustrating a second method of calculating the accumulated damage to a component according to the invention. Figure 6 is a flowchart illustrating a third method of calculating the accumulated 20 damage. Figure 7 is a flowchart illustrating a fourth method of calculating the accumulated damage to a component according to the invention. Figure 8 is a flowchart illustrating a method of calculating the accumulated damage to a component according to aspects of the invention. 25 Figure 9 is a flowchart illustrating a method according to the invention. Figure 10 is a diagram of a second aircraft landing gear. Figure 11 is a flowchart illustrating a method. DETAILED DESCRIPTION OF EMBODIMENT(S) 30 09 01 25 / ention relates to an improved method of calculating the remaining useful life of a component of an aircraft subassembly. The component may be a structural component, such as a lock link, a pin or a spring. Certain components of aircraft subassemblies may have a limited life, that is a limited number of flight 5 cycles during which the part performs according to specification. After this limited number of flight cycles, the part may be replaced to ensure that the subassembly continues performing as expected. An airline in possession of an aircraft benefits from an accurate estimate of the flight cycles remaining until replacement of a component is necessary because it allows accurately scheduling of the replacement. 10 Unnecessary replacement may also be avoided. The present invention enables a more accurate way of assessing the remaining flight cycles of a part or component before replacement is necessary, which may reduce wasted material and costs. Figure 1 is a diagram of an aircraft 10. The aircraft 10 includes subassemblies such 15 as a nose landing gear 12, main landing gear 14 and engines 16. Other aircraft subassemblies will be known to the skilled person. A subassembly can be a group of interconnected parts which are arranged to be fitted to the aircraft as a unit. Aircraft 10 may include also include a flight computer 13, communicatively coupled to various subassemblies fitted to the aircraft. 20 The flight computer 13 may be used by the pilot to determine characteristics of a flight. For example, the pilot may input to the flight computer 13 the number of passengers, the weight of the cargo load and the quantity of fuel in the tanks of the aircraft, to enable the flight computer to calculate the required thrust for take-25 off according to predefined specifications. These inputs may be described as flight data. Referring now to Figure 2, an aircraft assembly, namely an aircraft landing gear assembly, is shown generally at 14. The landing gear assembly 14 includes a 30 foldable stay 18 with stay links 18a and 18b, a lock link 20 with links 20a and 20b and a down lock spring assembly (not pictured) mounted to the stay 18 and arranged to urge the lock link 20 to assume a locked state. In addition, the landing gear assembly 14 also includes a shock absorber strut 24. The main strut 24, which may be hydraulic, pneumatic and / or sprung, comprises an upper portion 35 connectable to the underside 11 of the aircraft 10 at its upper end, and a lower 09 01 25 table to a ground contacting assembly 28 comprising wheels and a brake assembly. The components described above in relation to the aircraft landing gear assembly 5 14 are provided as examples only, and the skilled person would be well aware of other components of aircraft landing gear assemblies that may for example be part of nose landing gear 12 or the fuselage of the aircraft. The components or parts described above in relation to the aircraft landing gear 10 assembly 14 are designed and manufactured according to a set of predefined specifications that may differ based on the type of the component. For example, in the case of the foldable stay 18, link 18a link and 18b may be designed to be able to withstand a predefined maximum load or weight when the stay is in the locked position and the landing gear is extended. Once these components are 15 manufactured, they undergo a process of certification, to ensure that the components adhere to their respective specifications. Such certification processes are necessary to ensure the airworthiness of the aircraft that is fitted with these components. Once these certification processes are verified as producing reliable results, they are maintained as is, to enable reliably assessing the fatigue life of 20 components of the same type as described below in relation to Figure 3. Referring now to Figure 3, a known method 300 for calculating a remaining useful life (RUL) of a component is illustrated. The RUL is the remaining working life of a component, based on the total fatigue life and the accrued damage of the 25 component. The fatigue life of the component may be defined as the number of loading or stress cycles that the component may sustain before cracking occurs. In the context of an aircraft, a loading or stress cycle may be the same as a flight cycle. For a new component that has not sustained any flight cycles, the fatigue life may be equal to the RUL. 30 The known RUL calculation method 300 takes two inputs: Maximum ramp weight (MRW) 306 and component geometry 308. A certification tool 312 takes the inputs 306, 308 and calculates an inflicted damage to a component in an example or notional flight cycle. 09 01 25 n tool 312 may calculate the inflicted damage to a component in a notional flight cycle or a number of notional flight cycles based on a predefined fatigue spectrum. The certification tool 312 may also calculate the inflicted damage to a component in a notional flight cycle based on expected maximum manoeuvring 5 data, that may indicate manoeuvres the aircraft would be expected to perform during a flight cycle. The certification tool 312 may be a computer-implemented tool or method and may be able to calculate the damage inflicted on the component at a notional flight cycle. 10 The damage inflicted on the component at a notional flight cycle can in turn be used to estimate the number of flight cycles that the component may sustain before breakage occurs. The number of flight cycles that the component may sustain before breakage occurs corresponds to the fatigue life 314 of the component. Additionally or alternatively, the fatigue life 314 of the component may be 15 measured in flight hours. The fatigue life as calculated and approved during the certification of the aircraft or the component may apply to all the components of the same manufacturing batch, or in other words for all components with the same part number, irrespective of the aircraft usage. 20 Once the manufacturer has calculated the fatigue life 314 of a component, the airline can schedule replacement of the component by keeping a record of the number of flight cycles and / or flight hours for which the component has been in use. As the number of flight cycles and / or flight hours approaches the fatigue life of the component, the component must be replaced. Replacing a component before 25 the end of the component's fatigue life is intended to ensure that the assembly comprising the component will perform according to specifications when in use, and in turn ensure the airworthiness and safety of the aircraft. Calculating the fatigue life of a component may be based on computing the fatigue 30 damage at key locations of the component for a flight cycle. Calculating the fatigue damage may be based at least partly on the geometry of the part and the load or forces inflicted on the part during a notional flight cycle. Calculating the fatigue damage may be based on a computer system simulating the part and the load or forces inflicted on the part during a simulated flight cycle. Although the geometry 35 of the part may be known to the manufacturer of the part, the load inflicted on the 09 01 25 nponent during an actual flight or a series of actual flights may be difficult to predict accurately. To ensure the safety of passengers, it is common during the certification of the part to calculate the load on the part as the load that would be inflicted on the part were the aircraft to be fully loaded and therefore have 5 its maximum weight, otherwise known as Max Ramp Weight (MRW) of the aircraft during various manoeuvres. The known method 300 therefore assumes a worst-case scenario for each flight as the maximum ramp weight is the maximum total weight an aircraft may have, when 10 it has a maximum load of passengers, cargo and fuel. The present inventors have realised that this worst-case scenario may be very conservative and may lead to premature replacement of parts. Figure 4 schematically illustrates an at least-partially computer-implemented 15 method 400 for calculating the accumulated damage of a component for an aircraft according to an embodiment of the present invention. The method of Figure 4, improves on the method of Figure 3 in various ways as will be appreciated based on the following. 20 The method of Figure 4 may be used to calculate the accumulated or accrued fatigue damage on a component having undergone a number of flight cycles. The accumulated fatigue damage can then be used in combination with the manufacturer's estimated or certified fatigue life to determine a RUL for a specific component. In some embodiments of the invention, calculating a RUL of the 25 component may be based at least partially on a combination of the calculated accumulated fatigue damage and specification or certification data that comprise the certified fatigue life . Additionally or alternatively, calculating the RUL of the component may be based at least partially on a comparison of the calculated accumulated fatigue damage and specification or certification data that comprise 30 the certified fatigue life. Advantageously, the method of Figure 4 enables calculating the RUL based on the actual use of the component, and therefore components that were part of the same manufacturing batch but may have been used with different e.g. frequency or in aircrafts with different loads may have a different lifespan. 09 01 25 previous methods that may rely on MRW to calculate the fatigue damage to the component, the method of Figure 4 is based on an estimation of the weight of the aircraft 402 for a given flight. Calculating the weight may be based on actual flight data, as opposed to a worst-case scenario. The flight data may be 5 one or more of: the weight of the aircraft when empty, the weight of the cargo load at a predetermined point, the number of passengers at the predetermined point and the quantity of fuel at a predetermined point. In some embodiments, calculating the weight may be based on the combination of all of the above. 10 The predetermined point may be a point of a flight cycle after the aircraft has been loaded and before taxiing for take-off. Alternatively the predetermined point may be a point of a flight cycle after the aircraft has landed, for example during touch down, or at a predefined time after touch down. Further alternatively, the predetermined point may be the point where the aircraft has stopped taxiing, or 15 once boarding has been completed and the aircraft is ready for departure. The weight of the aircraft after the aircraft has been loaded and before taxiing for takeoff may be known as in-service departure weight at the gate or ramp weight (RW). Calculating the weight of the aircraft may be based on information that may be 20 known to, for example, the airline or the aircraft's flight crew prior to the aircraft's departure. This information may be part of flight data provided to the aircraft's flight computer. Additionally or alternatively, this information may be system data provided by a remote system. For example, the weight of the aircraft when empty would be known once the certification process of the airworthiness of the plane has 25 been completed and may be provided to airlines by the airframer / aircraft manufacturer. The weight of the cargo load during a flight cycle may be known to the airline or the flight crew because the cargo may be weighed before being loaded on the plane. The number of passengers during a flight cycle may be known to the airline that schedules the flight. The quantity of fuel loaded on the fuel tanks of the 30 aircraft may be known to the ground crew and / or the pilot before the aircraft starts taxiing. The skilled person will readily recognize multiple other ways to obtain this information. Based on the determined ramp weight of the aircraft 402 for a flight cycle, the 35 loads or strains on each subassembly for a flight cycle may be assessed. Based on 09 01 25 :he subassembly and the physical characteristics 408 of each component of the subassembly, a load or load spectrum on a component 410 of the subassembly for the flight cycle may be calculated. As will be discussed later, once the actual load on a component 410 is calculated, the fatigue damage for the 5 flight cycle for that component may be calculated. The physical characteristics 408 of a component remain the same for all flight cycles for the duration of the fatigue life of the component. The physical characteristics 408 of each component may be e.g., the geometry of the component and / or the 10 material of the component. The physical characteristics 408 of each component are known in advance and may be provided by the manufacturer of the component after the component has been certified. As such, calculating the damage inflicted on the component during a flight cycle may be based at least partially on the geometry of the component and / or the material of the component. The skilled 15 person would be well aware of various methods of simulating a mechanical component under load or strain and of corresponding methods of measuring the effects of such loads on a component. For example, scale models may be used or computational finite element analysis may be performed. 20 Having calculated the actual load on the component based on the RW for a flight cycle, tool / algorithm 412 can calculate the fatigue damage inflicted on the component for said flight cycle based on the actual load on the component Calculating the fatigue damage may also be based on other measurements obtained by tool 412, for example pressure experienced by the component and / or 25 temperature of the component. The pressure and the temperature may be obtained from transducers installed on the aircraft or on the subassembly comprising the component. For example, in the case of a component of a landing gear, for example parts 18a and 18b of the foldable stay 18 of Figure 2, transducers installed on the shock absorber strut 24 may provide information relating to the temperature and 30 the pressure. The skilled person will appreciate that different assemblies may comprise different sensors that provide information to the flight computer of the aircraft that may be provided to tool 412 to calculate the actual fatigue damage inflicted on the component. 09 01 25 :ompute an Average Damage per Flight cycle (ADF). ADF represents the average fatigue damage sustained by a component or part of a specific section or subassembly of the aircraft, during one flight cycle, for a specified RW. The ADF may be a function of RW. Calculating the ADF may be based on calculating the 5 fatigue damage on a component over a number of flight cycles. Tool 412 may use the same methods or algorithms used by certification tool 312. Advantageously, having tool 412 using the same methods as certification tool 312 for calculating fatigue damage ensures that the fatigue damage will be calculated 10 as reliably as during the certification process of the component. Advantageously, using the determined weight, RW, instead of the MRW may improve the accuracy of the calculation of the fatigue damage, and may result in extending the RUL of the component. 15 Tool 412 may be configured to store the calculated actual fatigue damage or the ADF in a register during a flight cycle, or after a flight cycle is completed. The register may contain previously calculated values of fatigue damage or ADF inflicted on the component before the current flight cycle. The values on the register may be considered to be historic data indicating the damage inflicted on the component 20 before the flight cycle. The historic data may be part of the flight data retrieved from the flight computer installed on the aircraft or from the system data retrieved from a remote system. In some embodiments, the historic data may be based on one or both of historic manoeuvring data of at least one previously-completed flight cycle and historic weight or load data for at least one previously-completed flight 25 cycle. The historic data may comprise a number of previous flight cycles to which the component has been subjected. Tool 412 may calculate the total accumulated fatigue damage 414 based on the calculated actual damage or ADF inflicted on the component during a flight cycle 30 and the historic data indicating the damage inflicted on the component before the flight cycle. Thus, the method of Figure 4 enables calculating an adapted or actual cumulative fatigue damage (or more simply an accumulated damage) on demand, based on 09 01 25 aft weight history for one or more of the components that constitute a subassembly of an aircraft (for example, the landing gear installed on an aircraft). In some embodiments, a further tool 418 may be capable of obtaining or may be 5 provided with specification or certification data and / or historical damage data 416 indicating the maximum allowable damage of the component and / or previously inflicted damage to the component. In such embodiments, further tool 418 may be capable of calculating a RUL 420 of the component based at least partially on a combination of the calculated accumulated fatigue damage and specification or 10 certification data that comprise the certified fatigue life . Additionally or alternatively, calculating the RUL of the component may be based at least partially on a comparison of the calculated accumulated damage and the specification or certification data. 15 In some embodiments, it may be possible to calculate an adapted fatigue life of a component, based on the calculated accumulated damage thus far and an expectation of future fatigue damage per flight cycle that may be inflicted on the component. The future fatigue damage per flight cycle may be based on the average fatigue damage experienced by the component on a predefined number of 20 previously completed flight cycles. In some cases, the fatigue damage may be first calculated per section or per subassembly, and then the fatigue damage for a section of the component of the subassembly may be calculated. For example, for a section y, the ADF may be 25 calculated based on equation 1: Total fatigue damage at section y, at a RW of X tons ADFtsectiony.RWx) = Total number of fUghts Based on equation 1, the accumulated fatigue damage for a section y at a given number of flight cycles z may be calculated based on equation 2: z FC 30 Accumulated fatigue damage^sec^-iony zpc^ {section y, rwx at flight z) (eg. 0 09 01 25 umulated fatigue damage per subassembly is calculated, the accumulated fatigue damage per section of component may be calculated based on the physical characteristics 408 of each component. The accumulated fatigue damage per section of the component may be used to determine the accumulated 5 fatigue damage of the component in total. For example, if several section of a component are used in fatigue calculation, the fatigue damage of a section of the component that has sustained the highest damage may be used as the fatigue damage of the total of the component.. 10 Figure 5 is a flow chart of a method 500 for calculating the accumulated damage of a component for an aircraft according to the present invention. The method of Figure 5 may be computer-implemented. At step 501, flight data is retrieved from the flight computer installed on the aircraft 15 and / or wherein system data is retrieved from a remote system. However, in some cases step 501 may be missed or data may be input by alternative means. At step 502, the weight of the aircraft at a predetermined point of a flight cycle is determined. The predetermined point of the flight cycle may be after loading and 20 before taxiing for take-of, otherwise known as ramp weight, RW. In some embodiments the estimation may be made by the flight computer of the aircraft once the flight computer is provided with one or more of the weight of the aircraft when empty, the weight of the cargo load at a predetermined point, the number of passengers at a predetermined point and the quantity of fuel at a predetermined 25 point. In other embodiments the estimation may be made by a computer external to the aircraft that is provided with one or more of the above values. In other embodiments of the invention, the weight may be determined by the pilot, the flight crew or the ground crew. 30 At step 504, the damage inflicted on the component during the flight cycle is calculated based at least partially on the determined weight of the aircraft or ramp weight, RW as described above with reference to Figure 4. In some embodiments the weight is determined based at least partially on the flight data. In some embodiments the weight is determined based at least partially on the system data. 09 01 25 of the damage inflicted on the component may require the use of a computer system. The computer system may comprise the flight computer of the aircraft. Additionally or alternatively, the computer system may comprise computing units external to the aircraft. 5 At step 506, historic data indicating the damage inflicted on the component before the flight cycle may be obtained. The historic data may be based on historic weight data from at least one previously-completed flight cycle. The historic data may be stored in an electronic register. 10 At step 508, the accumulated damage of the component is calculated based on the calculated fatigue damage or ADF inflicted on the component during the flight cycle and the historic data. The computer system implementing the method may output a signal indicating the accumulated damage. 15 At step 510, specification or certification data indicating the maximum allowable damage according to a specification of the component may be obtained. At step 512, the RUL of the component may be calculated based at least partially 20 on a comparison of the calculated accumulated damage and the specification or certification data. The RUL may be calculated based on the remaining damage, RD. RD may be calculated based on equation 3: RD = Maximum Allowable damage — Accumulated fatigue damage (eq. 3) 25 wherein the Maximum Allowable damage may be part of the specification or certification data. If an ADF for a component has been calculated the RUL may be measured in terms of flight cycles based on equation 4: RD ™l = -(^.4) 30 Alternatively, the RUL may be determined in terms of flight hours based on equation 5, wherein AFT is the Average Flight Time: RD AFT RUL = ^X^(e<?'5) The system implementing the method may output a signal indicating the RUL of the component. The remaining useful life may be measured in flight hours, flight cycles, or may be a percentage value indicating the remaining useful life in comparison to 5 an estimated updated fatigue life. 09 01 25 Figure 6 schematically illustrates an at least-partially computer-implemented method for calculating the accumulated damage of a component for an aircraft. 10 The method of Figure 6 may be used to calculate the accumulated or accrued fatigue damage on a component having undergone a number of flight cycles. The accumulated fatigue damage can then be used in combination with the manufacturer estimated or certified fatigue life to determine the remaining useful life for that specific component. Advantageously, the method of Figure 6 enables 15 calculating a remaining useful life based on the actual use of the component and therefore components that were part of the same manufacturing batch but may have been used with different frequency or in aircrafts with different loads may have different lifespans. 20 The method of Figure 6 is based on determining the forces inflicted on a subassembly or a structural part of the subassembly as a result of the aircraft performing various manoeuvres while taxiing. As the aircraft taxies prior to take off and / or after landing, forces may be developed on the subassemblies as a result of lateral of axial accelerations. As such, manoeuvring events performed by the 25 aircraft while taxiing may result in strain and fatigue damage at various subassemblies. For example, a right turn may result in additional load inflicted on a landing gear on the right side of the aircraft. As another example, a braking event may result in additional load inflicted on the front landing gear. 30 The method 600 of Figure 6 has as an input manoeuvring data 604. The manoeuvring data 604 may indicate a number and, optionally, an order of turning cases at specific speeds, of braking cases and of other specific landing and / or taxiing events. The manoeuvring data may indicate the speed, lateral acceleration, on, braking and steering during various points and times in which the aircraft is taxiing. Using manoeuvring data 604 may enable identifying actual manoeuvre case pairs along with their load magnitudes. Thus, the actual component damage accrued on the component due to manoeuvres may be 5 calculated. Thus the need to use the enveloping design certification loads and certification fatigue spectra to calculate a conservative theoretical enveloping generic maximum damage may be reduced. 09 01 25 In some embodiments, the manoeuvring data may be based on inputs from a 10 steering control in the aircraft and / or an accelerometer in the aircraft. Manoeuvring events and / or manoeuvring case pairings may be grouped in categories based on the amount of damage they inflict on components. Manoeuvring events may be grouped in categories based on the amount of acceleration they effect on components. The acceleration may be measure in g or m / s2. For example a right 15 turn under a specific velocity may result in 0.2g acceleration for a landing gear and may be grouped together with a left turn that results in 0.2g on the landing gear or a braking event that results in 0.2g on the landing gear. In some embodiments, the manoeuvring data may be stored in the aircraft's flight 20 computer as they are performed. Additionally or alternatively, the manoeuvring data may be stored in a computer external to the aircraft and combined with previous historic data to keep a record of the ground manoeuvres performed by the aircraft. 25 In some embodiments, each manoeuvring event and / or manoeuvring case pairings may be associated with a respective individual damage value. The individual damage value may represent the damage inflicted on the component due to the manoeuvring event and / or manoeuvring case pairings. In such embodiments calculating the damage inflicted on the component during the flight cycle may be 30 based on the individual damage values associated with manoeuvring events performed by the aircraft during the flight cycle. In some embodiments, calculating the damage inflicted on the component may be based on summing the damage value for each manoeuvring event performed by the aircraft while taxiing. 09 01 25 lanoeuvring data 604 of the aircraft for a flight cycle and the physical characteristics of the component 608, the load or load spectrum 610 on each subassembly for a flight cycle may be assessed. In some embodiments assessing the load or load spectrum 610 on each subassembly may comprise using the MRW 5 as the weight of the aircraft. In other embodiments, other predefined values may be used as the weight of the aircraft. Based on the manoeuvring data 604 for the flight cycle and the physical characteristics 608 of each component of the subassembly, the actual load or load 10 spectrum on a component 610 for the flight cycle may be calculated. As discussed above, once the actual load or load spectrum on a component 610 is calculated, the fatigue damage per flight cycle 614 for that component may be calculated using a tool 612, which may be similar to or the same as the tool 412 described above with reference to Figure 4. 15 Figure 7 illustrates a flow chart of a method 700 for calculating the accumulated damage of a component for an aircraft according to an embodiment of the present invention. 20 At step 702, manoeuvring data based on manoeuvres performed during a flight cycle are obtained. The manoeuvring data may be based on inputs from a steering control in the aircraft and / or an accelerometer in the aircraft. At step 704, the damage inflicted on the component during the flight cycle is 25 calculated based at least partially on the obtained manoeuvring data as described above with reference to Figure 6. At step 706, historic data indicating the damage inflicted on the component before the flight cycle may be obtained. The historic data may be based on historic 30 manoeuvring data of at least one previously-completed flight cycle. The historic data may be stored in an electronic register. 09 01 25 e accumulated damage on the component is calculated based on the calculated fatigue damage or ADF and the historic data. The computer system implementing the method may output a signal indicating the accumulated damage. 5 At step 710, specification or certification data indicating the maximum allowable damage according to a specification of the component may be obtained. At step 712, a remaining useful life of the component based at least partially on a comparison of the calculated accumulated damage and the specification or 10 certification data similarly to the process described previously with respect to step 512 of figure 5. Figure 8 schematically illustrates an at least-partially computer-implemented method 800 for calculating the accumulated damage of a component for an aircraft 15 according to an embodiment of the present invention. The method of Figure 8 can be considered as a combination of the methods illustrated in Figures 4 and 6. The method of Figure 8 may be used to calculate the accumulated or accrued fatigue damage on a component having undergone a 20 number of flight cycles. The method of Figure 8 is based on determining the load or forces inflicted on a subassembly or a structural part of the subassembly as a result of the aircraft performing various manoeuvres while taxiing whilst also using a determined weight, or ramp weight to assess the actual load inflicted on a subassembly or component. 25 As such, the method of Figure 8 is based at least partly on an estimation of the weight of the aircraft 802. Estimation of the weight of the aircraft 802 can be performed as described above in relation to the method of Figure 4. 30 The method of Figure 8 is also based at least partly on determining the load or forces inflicted on a subassembly or a structural part of the subassembly as a result of the aircraft performing various manoeuvres while taxiing. Thus the method of Figure 6 has as an input manoeuvring data 804. Manoeuvring data 804 may be the 09 01 25 • to manoeuvring data 604 as described above with reference to the method illustrated in Figure 6. Thus, the method of Figure 8 can determine a load spectrum or ground load 806, 5 i.e. a load based on the combination of the RW of the aircraft and of manoeuvres performed by the aircraft while taxiing. Based on the load spectrum or ground load 806 for the flight cycle and the physical characteristics 808 of each component of the subassembly, the actual load or actual 10 load spectrum on a component 810 for the flight cycle may be calculated. As discussed above, once the actual load or actual load spectrum on a component 810 is calculated, the fatigue damage per flight cycle 814 for that component may be calculated in a similar manner using tool 812, which may be the same as tool 412 or tool 612. 15 Figure 9 illustrates a flow chart of a method 900 for calculating the accumulated damage of a component for an aircraft according to an embodiment of the present invention. 20 At step 902, the weight of the aircraft at a predetermined point of a flight cycle is determined. The predetermined point of the flight cycle may be after loading and before taxiing for take-off. In some embodiments the estimation may be made by the flight computer of the aircraft once the flight computer is provided with one or more of the weight of the aircraft when empty, the weight of the cargo load at a 25 predetermined point, the number of passengers at a predetermined point and the quantity of fuel at a predetermined point. In other embodiments the estimation may be made by a computer external to the aircraft that is provided with one or more of the above values. Yet in other embodiments of the invention, determining the weight may be made by the pilot, the flight crew or the ground crew. In some 30 embodiments, step 902 is preceded by step 901, wherein flight data is retrieved from a flight computer installed on the aircraft and / or wherein system data is retrieved from a remote system. 09 01 25 lanoeuvring data based on manoeuvres performed during the flight cycle are obtained. The manoeuvring data may be based on inputs from a steering control in the aircraft and / or an accelerometer in the aircraft. 5 At step 906, the damage inflicted on the component during the flight cycle is calculated based at least partially on the determined weight of the aircraft or ramp weight, RW as described above with reference to Figure 4, and at least partially on the obtained manoeuvring data as described above with reference to Figure 6. 10 At step 908, historic data indicating the damage inflicted on the component before the flight cycle may be obtained. The historic data may be based on historic manoeuvring data of at least one previously-completed flight cycle and / or historic weight data. The historic data may be stored in, and may be retrieved from, an electronic register. 15 At step 910, the accumulated damage of the component is calculated based on the calculated fatigue damage or ADF and the historic data. The computer system implementing the method may output a signal indicating the accumulated damage. 20 At step 912, specification or certification data indicating the maximum allowable damage according to a specification of the component may be obtained. At step 914, a remaining useful life of the component based at least partially on a comparison of the calculated accumulated damage and the specification or 25 certification data similarly to the process described previously with respect to steps 512 and 712. Figure 10 shows a second aircraft landing gear 1014. The aircraft landing gear is substantially similar to the gear 14 shown in Figure 2 and the same parts are 30 labelled with the same references. For brevity, the structural parts of the landing gear will not be described again here. 09 01 25 ar 1014 has a load sensor 1001, which is fixed to the upper stay link 18a. The load sensor may be a strain gauge, an ultrasonic transducer, a deflection sensor or any other means of measuring a force applied to the upper stay link 18a. 5 Based on the output of the load sensor 1001, a load in the upper stay link 18a may be determined. This may be used to determine fatigue in the upper stay link 18a. Alternatively, or additionally, a further load, such as the load in the lower stay link 18b may be inferred from the measurement of the load sensor 1001. Additionally or alternatively, the load sensor 1001 may comprise a system of temperature and 10 pressure sensors that can be used to infer the actual load spectrum in components of the subassembly. For example, in the case where the subassembly is a landing gear, sensors measuring the temperature and pressure in the shock absorber (24) and the stroke of the shock absorber (24) may provide outputs that can be combined to determine the actual load or actual load spectrum on components of 15 the shock absorber as discussed below Figure 11 is a flow chart 1100 demonstrating how data from the measurement sensor 1001 may be used in the determination of fatigue damage to a component. At step 1110, data is obtained from the load measurement unit. Obtaining the data 20 may involve transferring the data by wired communication to a data storage device onboard the aircraft. Alternatively, data may be stored locally to the sensor 1001 to be downloaded at a later time. As a further alternative, the sensor 1001 may be arranged to transmit data wirelessly. 25 At step 1112, the data is processed using a tool, such as a tool to determine local maxima and local minima of the load, and in this way loading cycles, which may also be referred to as case pairings, may be determined. Subsequently, the fatigue damage due to each loading cycle can be calculated as a loading cycle damage value. The tool may therefore calculate a plurality of loading cycle damage values 30 for a single flight cycle. The loading cycle damage values may be summed together to determine a combined fatigue damage for a flight cycle. At step 1114, a fatigue damage output may be provided, indicating a level of damage accrued by the component during a flight cycle. The fatigue damage output 09 01 25 leasured loads may also be compared to outputs 614, 814 calculated based on the weight and / or manoeuvre data. A checking function may compare the outputs and may indicate a significant disparity to a user. Alternatively, the greater of the compared values may be used in subsequent calculations in order to avoid 5 unexpected crack propagation in a component. In subsequent steps, a remaining useful life of the component may be determined substantially similarly to the methods described previously. 10 In some embodiments of the invention, the methods described above may comprise further steps. For example, in some embodiments once the ADF is calculated, the method may further include calculating an adapted value for the fatigue life of a component or a subassembly that is based on the expected or predicted fatigue damage. The expected fatigue damage may be inferred based on the potential 15 future usage of the aircraft. The potential future usage of the aircraft may be extrapolated from the historic data described above. Usage of the aircraft may be defined by a set of parameters including average length of flight, average distance travelled, average load of the aircraft, average quantity of fuel loaded on the aircraft before departure etc. The predicted values for any of these set of parameters may 20 be determined based on previously completed flight cycles. In some embodiments of the invention, the historic data may comprise flight information derived from previous flight cycles that may include: length of flight, distance travelled, type of flight, flight manoeuvres, experienced temperatures, 25 altitude readings etc. The methods described above may be implemented using a system. The system may comprise at least a processor and a memory. The at least one processor may be configured to perform any or all of the steps of the above methods. The register 30 discussed with relation to tool 412 may be stored on the memory. The system may comprise the flight computer of the aircraft or may be external to the flight computer. LO CXI escribed above may be used to calculate the RUL of a subassembly or an assembly of the aircraft based on the calculated accumulated damage of the plurality of components used in the subassembly. As such, although certain components may not have reached the end their fatigue life, if one of the 5 components has reached the end of its fatigue life but cannot be replaced, the subassembly or assembly may be replaced. As such, in some embodiments of the invention the remaining useful life of an aircraft subassembly may be the remaining useful life of a component of the subassembly with the least useful life left. 10 Although the invention has been described above with reference to one or more preferred embodiments, it will be appreciated that various changes or modifications may be made without departing from the scope of the invention as defined in the appended claims. 03 07 25

Claims

1. A method for calculating the accumulated damage on a component of an aircraft, the method comprising:determining the weight of the aircraft at a predetermined point of a flight cycle, wherein determining the weight is based on one or more of:a weight of the aircraft when empty;a weight of the cargo load at the predetermined point;a number of passengers at the predetermined point; anda quantity of fuel at the predetermined point;calculating, based at least partially on the determined weight of the aircraft, a calculated damage inflicted on the component during the flight cycle;obtaining historic data indicating the damage inflicted on the component before the flight cycle; andcalculating the accumulated damage of the component, based on the calculated damage inflicted on the component during the flight cycle and the historic data.

2. A method according to any one of the preceding claims, wherein the method comprises retrieving flight data from a flight computer installed on the aircraft, and wherein the determining the weight is based at least partially on the flight data.

3. A method according to any one of the preceding claims, wherein the method comprises retrieving system data from a remote system, and wherein determining the weight is based at least partially on the system data.

4. A method according to any of the preceding claims, wherein the predetermined point of the flight cycle is after loading and before taxiing for ta ke-off.03 07 255. A method according to any of the preceding claims, wherein the method comprises obtaining manoeuvring data during the flight cycle and wherein calculating the damage inflicted on the component during the flight cycle is based at least partially on the manoeuvring data.

6. A method according to any one of the preceding claims, wherein calculating the damage inflicted on the component during the flight cycle is based at least partially on the geometry of the component and / or the material of the component.

7. A method according to any of the preceding claims, wherein the historic data is based on one or both of:historic manoeuvring data of at least one previously-completed flight cycle; andhistoric weight data of at least one previously-completed flight cycle.

8. A method according to any of the preceding claims, further comprising calculating a remaining useful life of the component based on the calculated accumulated damage and at least one of:certification data indicating the maximum allowable fatigue damage on the component; and / orpredicted values of one or more of the flight cycle attributes for future flight cycles, wherein the predicted values have been determined based on previously completed flight cycles.

9. A method according to any of the preceding claims, wherein the component is part of a landing gear of the aircraft.

10. A system comprising at least a processor and a memory, the processor configured to perform a method according to any of the preceding claims.

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