Method for determining wear

Abstract

The present invention relates to a method for determining (33) wear (43) on a component (2, 3, 4, 5) of an aircraft engine (1) of an aircraft, comprising determining (30, 31) flight route information (40) and / or environmental information (41) for at least one flight point on a journey by the aircraft from a starting point to a destination, determining (32) a thermodynamic state (42) of the component (2, 3, 4, 5) of the aircraft engine (1) at at least the flight point, and determining (33) the wear (43) on the component (2, 3, 4, 5) at at least the flight point on the basis of the flight route information (40) and / or the environmental information (41) and on the basis of the thermodynamic state (42), at least one particle speed (v) being taken into account when determining (33) the wear (43).

Description

[0001] METHOD FOR DETERMINING WEAR DESCRIPTION

[0002] Technical field

[0003] The present invention relates to a method for determining the wear of a component of an aircraft engine.

[0004] State of the art

[0005] Turbomachinery, such as aircraft engines or power units, is subjected to a multitude of stresses during operation. For example, certain phases of flight can involve high thermal and / or mechanical stresses, such as during takeoff or while flying through atmospheric regions with high particle concentrations. As a result of these stresses, turbomachinery components can wear, corrode, or develop cracks. For safety reasons, these components are therefore frequently inspected, which can be time-consuming and costly. For instance, if abnormalities are discovered during this inspection, the components are assessed to determine their continued operational capability.

[0006] Description of the invention

[0007] The present invention is based on the technical problem of providing a particularly advantageous method for determining the wear of a component of an aircraft engine.

[0008] This is achieved according to the invention using the method according to claim 1. In this method, for at least one flight point of a journey of the aircraft from a starting point to a destination, flight route information and / or environmental information is determined, a thermodynamic state of the component at the flight point is determined, and from this, the wear of the component of the aircraft engine at the flight point is determined based on the flight route and / or environmental information as well as on the thermodynamic state of the engine or its component.

[0009] In the context of this invention, wear can be understood as both direct or mechanical wear of the aircraft engine component and indirect wear of the respective component. Direct wear refers to a mechanical damage mechanism, such as a change in geometry caused by erosion (e.g., detectable as a change in blade length or blade thickness on the component) or by abrasion (e.g., detectable as a change in blade height and thus a blade tip gap). In contrast, indirect wear can be determined using a so-called "severity factor," which represents a measure of the degree of wear of a component or part. This "severity factor" can be dimensioned or dimensionless; it can be specified absolutely or relatively (e.g., in the form of a maximum permissible wear).For example, this value for an aircraft engine component may be 1.0 before maintenance is due and may still be 0.67 after 3000 flight cycles, resulting in a wear rate of 0.00023 per flight.

[0010] The inventors recognized that, in addition to flight route and / or environmental information, the thermodynamic state of the aircraft engine component is relevant for determining its wear. The thermodynamic state is defined here as the temporal profile of pressure, temperature, rotational speed, and air volume flow through the engine or component. This thermodynamic state is determined segment by segment, i.e., at the flight point(s) along the aircraft's journey from its starting point to its destination. From this, a particle velocity can be derived, specifically the velocity at which particles in the atmosphere at the respective flight point enter the gas duct and impact the component in question. The particle velocity, in turn, can significantly influence whether and to what extent wear actually occurs, as detailed below.In contrast to phenomenological modeling approaches, which also rely on flight path information, this method actually incorporates a factor relevant to damage from a physical or "microscopic" perspective. This allows the inventors to achieve accurate predictions of wear, enabling, for example, better planning of upcoming engine or component overhauls. Depending on the extent of wear, appropriate overhaul and refurbishment resources can be kept on hand and prepared.

[0011] Another factor influencing damage, which can also be derived from the thermodynamic state, is the component's temperature. For example, the introduction of sand particles into an aircraft engine component at a temperature of, say, 1100 °C may be less critical in terms of component damage than the same sand introduction at a temperature of 1500 °C. Compared to conventional methods, this approach therefore considers not only external information, such as flight path and / or environmental information, to determine the wear of the respective component within that segment, but also the thermodynamic state of the engine component within that segment. This can be done individually for each engine component, e.g., at the level of guide vanes and rotor blades, or cumulatively for several engine levels or modules.

[0012] Overall, incorporating the thermodynamic state can improve wear determination for the thermodynamic component compared to previously known methods, as damage mechanisms dependent on pressure, temperature, rotational speed, and / or flow rate, such as erosion, oxidation, corrosion (e.g., due to hot gas flow or the salinity of the fluid flowing past), sulfidation, or adhesion, can be determined more reliably. For example, this can extend maintenance cycles, avoid unnecessary maintenance visits, reduce overall maintenance costs, and enable more reliable prediction of expected maintenance expenses. This can be advantageous in terms of reducing scrap rates per engine and, for example, increasing the number of flight cycles between maintenance visits.Furthermore, established damage limits and safety buffers can be validated during component design, and the latter can be reduced with regard to the damage mechanisms that actually occur during operation.

[0013] Preferred embodiments are found throughout the disclosure and particularly in the claims, although the description of features does not always differentiate between the various claim categories. For example, if a method for determining the wear of a component is described, this is to be understood simultaneously as a disclosure of a correspondingly designed computer program product and / or a corresponding use, and vice versa.

[0014] According to a preferred embodiment, particle velocity is taken into account when modeling wear, and this velocity is determined by referring to the aircraft speed and rotational speed of the component or engine. A high rotational speed can result in a high particle velocity, for example during takeoff, even if the aircraft itself is still traveling at a relatively low speed. The particle velocity, i.e., the speed at which ingested particles strike the components in the gas duct, especially moving components such as rotor blades or blade rings, can be a key factor in determining the extent of wear caused by erosion or abrasion.

[0015] Below is a formula for calculating Et, which is the amount of erosive wear per particle mass.

[0016] ■ (S) - fe)® fö) (W>” - («w - ■

[0017] 'PT 4" e-ompoHmr

[0018]

[0019] (Gig 1)

[0020]

[0021] 1 /

[0022] Here, denotes the particle velocity in the plane of incidence in m / s, where this velocity is quadratically involved. Other quantities, such as the Vickers hardness H, are also considered. vM The density (pt) depends on the material of the component or blade, while other values ​​depend on the material of the particle (such as the density pp). In addition to the particle velocity, an impact angle (β) can also be taken into account.

[0023] Abrasive wear (merosion per blade) per shovel can be calculated, for example, using...

[0024] > ^particle

[0025] erosion per blade — £ t '

[0026]

[0027] (Gig. 2)

[0028] with particle mass m particleand the number of shovel Nb loads-

[0029] In addition to particle velocity, a damage model focused on erosion can also consider mass flow, i.e., the engine's operating-dependent mass flow rate. This model thus represents, firstly, the velocity at which the ingested particles impact; secondly, it uses the mass flow rate to represent the number of ingested particles. The latter must also be considered in light of atmospheric data; only if particles are actually present will a high mass flow result in a high particle input.

[0030] In a preferred embodiment, at least one temperature-dependent damage model is considered when modeling wear. This model can, for example, focus on damage to the hot gas components, whereby the damage increases with rising temperature and / or may only begin above a certain threshold temperature; see also the preceding remark on sand ingress. In this case, the atmospheric data (are particles present, and if so, which ones?) and the mass flow (to what extent do any particles actually enter the engine?) can again serve as the basis. Sand drawn into the engine can, for example, cause erosive damage below a certain threshold temperature (provided the mass flow and particle velocity are sufficiently high). At very high temperatures, however, the sand particles can melt with molten deposits, figuratively speaking, resulting in vitrification. Another form of temperature-dependent damage would be, for example...Hot gas corrosion, which can be caused by salts or aerosols. Information on these particles is also available in the atmospheric data.

[0031] In a preferred embodiment, several damage models are considered when modeling wear. These can differ, for example, in at least their temperature dependence; alternatively or additionally, the dependence on particle velocity can also differ. One damage model can, for example, relate to erosion. If, for instance, sand particles below the melting point of sand are drawn in, this leads to erosive removal of the blade material, e.g., in the compressor, provided the mass flow and particle velocity are sufficiently high.

[0032] As mentioned previously, a damage model for CMAS corrosion can be considered alternatively or additionally. In this case, sand particles are also detected in the atmospheric data, which are drawn into the engine. However, these particles melt in the combustion chamber when the combustion temperature is high enough. They then deposit, for example, on porous thermal barrier layers in the combustion chamber or the high-pressure turbine, vitrifying them. Another damage model that can be considered alternatively or additionally concerns hot gas corrosion. A trigger for this, in addition to temperature, is the presence of salts / aerosols in the atmosphere or in the atmospheric data.

[0033] In a preferred embodiment, several analytical damage models are used for modeling the damage. This allows, for example, computationally intensive simulations to be reduced or avoided, which in turn can permit a higher-resolution evaluation of the flight route. This allows, for example, a larger number of flight points along the flight route to be considered. This can result in a better modeling of the wear. A "flight point" is a point on the flight route at which the modeling disclosed herein is performed, in particular calculations using several analytical damage models. In a preferred embodiment, the time interval between adjacent flight points is at most 10 minutes, and more preferably at most 5 minutes or 1 minute. Although the finest possible resolution may generally be desired, possible lower limits could be, for example, at least 1 second, 5 seconds, or 1 minute.The duration is 10 seconds. Preferably, the modeling or calculation is carried out over at least 70% of the flight duration, i.e., the journey, with a corresponding resolution, and more preferably over at least 80% or 90% or over the entire flight duration (100%).

[0034] In a preferred embodiment, the thermodynamic state encompasses a temporal profile of pressure and / or temperature states in the aircraft engine component during the flight segment. Generally, the "segment-wise" consideration of the thermodynamic state can also capture a continuous or quasi-continuous determination of the same throughout the flight. Alternatively, the process can be segmented, and an average value (e.g., of pressure, temperature, etc.) can be considered for each segment.

[0035] Alternatively or additionally, the thermodynamic state can include, for example, the rotational speed of the component, the flow rate of a fluid flowing through this component (e.g., ambient air containing particles), a bypass rate (i.e., the ratio of the amount of fluid bypassing the component to the amount flowing through it), a fuel quantity, a fuel composition, and / or the combustion residues of the fuel. Considering one or more of the aforementioned aspects can be relevant, for example, for assessing damage mechanisms dependent on the thermodynamic state, so that these can be taken into account when determining wear.

[0036] In a preferred embodiment, the thermodynamic state of the aircraft engine component is determined as a function of the component's rotational speed and / or the volumetric flow rate of a fluid flowing through the aircraft engine component. By including the aforementioned parameters individually or in combination, it becomes possible, for example, to model in detail the operating and load conditions occurring in the aircraft engine, namely the respective component, during each segment of the journey. A segment of the journey can, in particular, be a single phase, such as...a turbine-driven forward movement on the runway (English: taxi), a take-off phase, an initial climb, a climb to cruise altitude, a cruise, a descent, a holding, an approach and a landing.

[0037] In a preferred embodiment, the thermodynamic state in the segment is determined based on data points acquired during the flight within that segment. This is based on an analysis of so-called Continuous Engine Operational Data (CEOD) or Full-Flight Data (FFD). Individual data points are thus acquired—measured, for example, at a sampling frequency of 1 Hz—within the respective segment and stored for wear determination. Wear determination can be performed, for example, after the end of the flight, i.e., after the aircraft has arrived at its destination, which can be advantageous for a rapid assessment of component wear.

[0038] In an alternative preferred configuration, the thermodynamic state at at least one flight point is determined using an engine model. The thermodynamic state can be determined, for example, over a flight phase or, more generally, over a time period of the flight mission or journey. The engine model here refers to a simulation model that represents the physical relationships within the aircraft engine, particularly those of its individual components. This engine model uses flight route information and / or environmental information to determine the thermodynamic state. For example, based on altitude, Mach number, outside temperature, or thrust lever position, the prevailing pressures, rotational speeds, and temperatures in the respective segment of the components are determined.This allows for a relatively reliable assessment of the thermodynamic state of the respective component during the journey in the segment.

[0039] In other words, the engine data, or rather the thermodynamic state, is modeled, e.g., the rotational speed, the pressure (in the engine or component), the temperature (in the engine or component), and / or the mass flow. This engine data is therefore not measured directly, but rather derived from flight route and engine information (e.g., engine type). The thrust requirement is determined from the flight route and weather data (e.g., wind). From the thrust requirement, together with the engine class or type, its thermodynamic state (rotational speed, pressure, temperature, mass flow) is derived.

[0040] The engine model can be a physical model, a data-based model, or a mathematical substitute model. A physical model, for example, is a representation of the physical states of the engine component during the aircraft's flight, determined using a thermodynamic performance simulation program. Put simply, in this case, the physical model is a representation of the aircraft engine or its components obtained through computer-based simulation methods. Alternatively, the engine model can also be a data-based model, such as a representation of the aircraft engine or its components obtained using a neural network. The neural network can, for example, approximate the physical states of the aircraft engine or its components, i.e., perform an approximate calculation.

[0041] Alternatively, the engine model can also be designed as a mathematical substitute model, i.e., based, for example, on an interpolation between individual data points. Examples of possible data points can be found in the following table, where the mass flow rate denotes a mass of the fluid taken into the aircraft engine, the abbreviation "BPR" stands for the bypass rate (see above), and the abbreviations T t 4 and To denote different temperatures in the aircraft engine. Flight phase mass flow BPR TU at I

[0042]

[0043] SA+15 at ISA-H5 Tmd 75 kg / s 5.2 1300 K 303.15 k lab •US 400 kg / s 4.5 1800 K 302.31 K Climb 200 kg A; 4.5 1790 K 283.17 K Cruise 150 kg A 4.9 1500 K 282.31 K Descent 50 g / s 5.9 1300 K 283, K Laßdiug 150 kg / s 5.9 1390 K 283.17 K Level 150 g / s 4.9 1500 K 202.31K

[0044] Preferably, the engine model is a physical model, although to reduce computing power or, for example, to simplify (preliminary) calculations, the data-based or substitute model can also be used.

[0045] In a preferred embodiment, the flight route information includes a flight level of the aircraft in the segment, an altitude of the aircraft in the segment, a geographical position (e.g., longitude and latitude) of the aircraft in the segment, a flight path of the aircraft in (or through) the segment, an orientation of the aircraft in the segment, a ground speed of the aircraft in the segment, an air speed of the aircraft in the segment, a flight distance in the segment, and / or a flight time in the segment, preferably each accompanied by a unique timestamp.Put simply, flight route information includes the essential aircraft-related parameters for tracking its position and route relative to the Earth's surface, preferably including the associated time information.

[0046] In a preferred embodiment, the environmental information includes air pressure, air temperature, wind speed, wind direction, humidity, the quantity and / or type of aerosols and / or aerosol particles, the composition of the atmosphere in the segment, and / or the state of the atmosphere in the segment. Put simply, the environmental information comprises the parameters of the ambient air in the respective segment. Possible parameters include, for example, a dust aerosol (0.03–0.55 μm) mixing ratio, a dust aerosol (0.55–0.9 μm) mixing ratio, and a dust aerosol (0.9–20 μm) mixing ratio.9 - 20 μm) mixing ratio), a mixing ratio of hydrophilic black carbon aerosol, a mixing ratio of hydrophobic black carbon aerosol, a mixing ratio of sea salt aerosol (0.03 - 0.5 μm), a mixing ratio of sea salt aerosol (0.5 - 5 μm), a mixing ratio of sea salt aerosol (5 - 20 μm), the specific humidity, a sulfate aerosol mixing ratio, a Sulphur dioxide content, temperature, U-component of wind and / or V-component of wind.Preferably, the environmental data is divided according to particle size and processed separately.

[0047] In a preferred embodiment, the journey comprises a plurality of segments, preferably at least a first segment, a second segment, and a third segment. A segment here is understood to be, for example, a respective travel section ("phase"), as described above.

[0048] In a preferred implementation, the wear and tear during the journey is summed across the individual segments of the journey. Put simply, the wear and tear for each segment is determined based on flight route information, environmental information, and the thermodynamic state, and a total wear and tear is calculated as the sum of the wear and tear for each segment.

[0049] In a preferred embodiment, the aircraft engine component is a guide vane grid or a rotor blade grid. In simplified terms, the wear is determined for at least one guide vane grid or rotor blade grid, or the corresponding blade assembly. However, several guide vane and rotor blade grids arranged along a central axis of the turbomachine can also jointly form the aircraft engine component, for example, a guide vane grid and a rotor blade grid together. In general, the present method can be applied to aircraft engine components of any size, for example, to an entire engine module (i.e., an entire compressor module or an entire turbine module). Alternatively, the aircraft engine component can also be a static component, such as a combustion chamber of the aircraft engine.

[0050] The invention further relates to a method for predicting the remaining service life and / or the overhaul effort of a component of an aircraft engine, comprising determining the wear of the component in a segment of an aircraft's journey from a starting point to a destination based on flight route and / or environmental information as well as on the thermodynamic state in the segment, preferably using a method according to one of the above aspects, comparing the wear of the aircraft engine component with a wear limit value and predicting the remaining service life of the component based on the state comparison.

[0051] In simplified terms, the wear determined using the method described above is compared with wear limits to calculate the remaining service life of each component. This wear limit can be a fixed value, or it can depend on flight route information, environmental data, and / or the component's thermodynamic state during the flight. In other words, the remaining service life prediction can be based on a predefined safety limit or determined specifically for one or more remaining flight segments. This makes it possible, in particular, to determine a safe remaining service life for the aircraft depending on its usage.

[0052] In a preferred implementation, the forecast involves summing the wear and tear of the component over a multiple or numerous flight journeys. These, in turn, each comprise a multitude of flight points, which can be considered, for example, segmented into different categories.

[0053] The invention further relates to a method for overhauling at least one component of an aircraft engine. In this method, the wear of the component(s) is determined, preferably accumulated over a large number of flights. In a method described herein, the overhaul effort can, for example, be predicted. Depending on the determined wear and tear or the overhaul effort, the overhaul can be scheduled, for example, with regard to the required duration. Alternatively or additionally, the resources required for the overhaul can also be provided; that is, preparatory and peripheral work can be initiated before the actual overhaul is carried out. This is possible because, in addition to phenomenological modeling, the factors relevant to wear are taken into account.

[0054] The invention further relates to a computer program product comprising instructions which, when the program is executed by a computer, cause it to execute the method for determining the wear of a component and / or the method for predicting the remaining service life of a component of an aircraft engine according to one of the aforementioned aspects.

[0055] The invention further includes a computer-readable medium comprising instructions which, when executed by a computer, cause it to execute the method for determining the wear of a component and / or the method for predicting the remaining service life of a component of an aircraft engine according to one of the aforementioned aspects.

[0056] Brief description of the drawings

[0057] The invention is explained in more detail below using an exemplary embodiment, whereby the individual features within the scope of the dependent claims may also be essential to the invention in other combinations, and no distinction is made in detail between the different claim categories. In detail,

[0058] Figure 1 shows a schematic axial section of a turbofan engine;

[0059] Figure 2a shows a schematic overview of segments of an aircraft's journey from a starting point to a destination;

[0060] Figure 2b shows a schematic representation of flight route information;

[0061] Figure 2c is a schematic representation of environmental information;

[0062] Figure 3 shows a schematic representation of the process flow.

[0063] Preferred embodiment of the invention

[0064] Figure 1 shows an example of an aircraft engine 1, namely an exemplary turbofan engine, in an axial section along a central axis X. It is functionally divided into a fan 5, a compressor 2 (comprising a low-pressure compressor 2.1 and a high-pressure compressor 2.2), a combustion chamber 3, and a turbine 4 (comprising a high-pressure turbine 4.1 and a low-pressure turbine 4.2), which are arranged along two shafts 7 parallel to the central axis X, namely a high-pressure shaft connecting the high-pressure compressor 2.2 with the high-pressure turbine 4.1 and a low-pressure shaft connecting the fan 5 with the low-pressure compressor 2.1 and the low-pressure turbine 4.2.

[0065] In the low- and high-pressure compressors 2.1 and 2.2, intake air is compressed. In the combustion chamber 3, fuel, e.g., kerosene, is added and this mixture is burned. The hot gas is expanded in the high- and low-pressure turbines 4.1 and 4.2, whereby energy is extracted from the hot gas. This energy is used to drive the rotors of the compressors 2.1 and 2.2 as well as the fan 5 (via shaft 7). The compressor 2.1, 2.2 and the turbine 4.1, 4.2 each have at least one stage with at least one guide vane (also referred to as a "guide vane row") and one rotor blade (also referred to as a rotor blade row). Figure 1 also shows the airflow of the outside ambient air through the aircraft engine, namely that a certain proportion of the outside ambient air is guided past a core engine formed from compressor 2, combustion chamber 3 and turbine 4 (see arrows in Fig. 1).Overall, only a certain proportion of the outside ambient air required for the propulsion / thrust of the aircraft is passed through the combustion chamber (usually less than 20%), while a larger proportion is routed past the core engine (so-called "bypass").

[0066] The ambient air carries a quantity and / or type of aerosols and / or aerosol particles which (depending on a thermodynamic state) can lead to wear and tear and consequently damage to one or more components 2, 3, 4, 5 of the aircraft engine 1.

[0067] Figure 2a shows a schematic overview of the segments of an aircraft's journey from a starting point to a destination. An example of altitude (ALT), engine exhaust gas temperature (EGT), and engine speed (NH, high-pressure shaft speed) are also sketched over these segments.

[0068] Figure 2b shows a schematic representation of flight route information in relation to the Earth's surface, i.e., the distances covered by the respective aircraft during its operation.

[0069] Figure 2c, on the other hand, shows an exemplary environmental information using the example of air pressure maps.

[0070] Figure 3 summarizes some process steps in a flowchart. The process can include, for each segment of the flight, determining 30 flight route information 40 and determining 31 environmental information 41. Furthermore, a thermodynamic state 42 is determined using a model for determining the thermodynamic state of the aircraft engine in the segment 32, and subsequently, based on this data, preferably using a model for determining damage or wear of a component, the wear 43 of the component is determined 33.

[0071] For example, depending on the component under consideration (Fan 5, Low-pressure compressor 2.1, High-pressure compressor 2.2, Combustion chamber 3, High-pressure turbine 4.1, Low-pressure turbine 4.2) and the relevant damage mechanism (e.g., erosion, abrasion, corrosion, sulfidation, etc.), different damage models can be used to determine the respective damage (e.g., erosion in the high-pressure compressor 2.1 (stage 1) or sulfidation in the low-pressure turbine 4.2 (stage 3)). [REFERENCE SYMBOL LIST]

[0072] Aircraft engine (turbocharged engine) 1 Compressor 2 Low-pressure compressor of compressor 2 2.1 High-pressure compressor of compressor 2 2.2 Combustion chamber 3 High-pressure turbine of turbine 4 4.1 Low-pressure turbine of turbine 4 4.2 Fan 5 Shaft of aircraft engine 1 7 Central axis of aircraft engine 1 X Determine (flight route information) 30 Determine (environmental information) 31 Determine (thermodynamic state) 32 Determine (component wear) 33 Flight route information 40 Environmental information 41 Thermodynamic state 42 Component wear 43

Claims

REQUIREMENTS 1. Method for determining (33) wear of at least one component (2, 3, 4, 5) of an aircraft engine (1) of an aircraft, comprising: Determine (30, 31) flight route information (40) and / or environmental information (41) for at least one flight point of a journey of the aircraft from a starting point to a destination; - Determining (32) a thermodynamic state (42) of the component (2, 3, 4, 5) of the aircraft engine (1) at least at the flight point; and determining (33) the wear (43) of the component (2, 3, 4, 5) at least at the flight point based on the flight route and / or environmental information (40, 41) as well as on the thermodynamic state (42), taking into account at least one particle velocity (v) when determining (33) the wear (43).

2. Method according to claim 1, wherein the particle velocity (v) is determined based on a flight speed of the aircraft and a rotational speed of the component (2, 4, 5) of the aircraft engine (1).

3. Method according to claim 1 or 2, wherein the thermodynamic state (42) of the component (2, 3, 4, 5) of the aircraft engine (1) comprises a time course of temperature states in the component (2, 3, 4, 5) of the aircraft engine (1), wherein at least one temperature-dependent damage model is taken into account when determining (33) the wear (43).

4. Method according to one of the preceding claims, wherein several damage models are taken into account when determining (33) the wear (43), preferably analytical damage models.

5. Method according to one of the preceding claims, in which the determination (33) of the wear (43) at a plurality of flight points of the journey is carried out and a wear summed over the plurality of flight points is determined.

6. The method according to claim 5, wherein the time interval between nearest adjacent flight points is at most 10 minutes.

7. Method according to one of the preceding claims, wherein the wear of component (2, 3, 4, 5) of the aircraft engine (1) is determined as a function of a mass flow through the aircraft engine (1).

8. Method according to one of the preceding claims, wherein the thermodynamic state (42) at least at the flight point is determined with an engine model, in particular at least one of a rotational speed, a pressure, a temperature and a mass flow is derived from the flight route information (40).

9. Method according to any of the preceding claims, wherein the flight route information (40) comprises a flight level of the aircraft, a flight altitude of the aircraft, a geographical position of the aircraft, a flight path of the aircraft, an orientation of the aircraft, a speed of the aircraft, a flight distance and / or a flight time, preferably each provided with a unique timestamp.

10. A method according to any one of the preceding claims, wherein the environmental information (41) comprises air pressure, air temperature, wind speed, wind direction, humidity, quantity and / or type of aerosols and / or aerosol particles, atmospheric composition at least at the flight point, and / or the state of the atmosphere at least at the flight point.

11. A method for predicting the remaining service life and / or overhaul effort of at least one component (2, 3, 4, 5) of an aircraft engine (1), comprising: Determining and summing up the wear (43) of component (2, 3, 4, 5) for a multitude of flights of the aircraft in a method according to one of the preceding claims; Comparing the wear (43) of component (2, 3, 4, 5) of the aircraft engine (1) with a wear limit; and Forecast of the remaining service life and / or the reconditioning effort of component (2, 3, 4, 5) based on the condition comparison.

12. Method for overhauling at least one component (2, 3, 4, 5) of an aircraft engine (1), comprising: Determining (33) wear of component (2, 3, 4, 5) of the aircraft engine (1) in a method according to any one of claims 1 to 10, in particular for predicting overhaul costs in a method according to claim 11, Depending on this: Scheduling and / or preparing the overhaul, overhauling component (2, 3, 4, 5) of the aircraft engine (1).

13. Device for data processing, comprising means for carrying out the method according to any one of claims 1 to 10 or the forecasting method according to claim 11.

14. Computer program product comprising instructions which, when executed by a computer, cause the computer to execute the method according to any one of claims 1 to 10 or the prediction method according to claim 11.

15. Computer-readable medium comprising instructions which, when executed by a computer, cause the computer to execute the method according to any one of claims 1 to 10 or the prediction method according to claim 11.

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