Preventive method for the thermal monitoring of a fixed internal structure inside a nacelle of an aircraft turbine engine

Abstract

The invention relates to a method for thermally monitoring an aircraft composite part, the method comprising the steps of: acquiring and storing, using a processor on-board the aircraft, temperature measurement data supplied by a plurality of temperature sensors (SNk) distributed over a surface of the composite part (1); defining a grid pattern on the surface of the composite part; determining, using the processor, a temperature (Ti,j) of each grid cell (Mi,j) according to the temperature measurements and respective positions of the temperature sensors; for each grid cell of which the temperature is above a corresponding threshold value, computing a damage index for the grid cell according to a difference between the temperature of the grid cell and the corresponding threshold value and according to a time during which the temperature of the grid cell remained above the corresponding threshold value.

Description

[0001] DESCRIPTION

[0002] TITLE: Preventive method for thermal monitoring of a fixed internal structure in an aircraft turbomachine nacelle

[0003] technical field

[0004] The present invention relates to a monitoring method aimed at defining preventively maintenance operations carried out in the nacelle of a turbomachine equipping an aircraft.

[0005] State of the art

[0006] A turbojet nacelle typically has a largely tubular structure comprising an air intake upstream of the turbojet, an intermediate assembly surrounding a turbojet fan, and a rear assembly, which may incorporate thrust reversing means, surrounding the combustion chamber and all or part of the turbojet's compressor and turbine stages. The nacelle is generally terminated by an exhaust nozzle whose outlet is located downstream of the turbojet.

[0007] Modern nacelles are designed to house a turbofan engine capable of generating, on the one hand, a hot airflow (also called the primary flow) from the engine's combustion chamber, circulating within a space defined by a roughly tubular compartment called the "core," and on the other hand, a cold airflow (called the "secondary flow") from the fan, circulating outside the engine through an annular passage, called the "filter," formed between an internal structure defining the engine's fairing and an internal wall of the nacelle. Both airflows are ejected from the engine through the nozzle at the rear of the nacelle.

[0008] The core compartment includes an external casing, known as the internal fixed structure (IFS) of the nacelle, which is typically made of composite materials. This internal fixed structure is subject to significant thermal stresses. To thermally protect this internal fixed structure (IFS), thermal protection panels are used to isolate the nacelle components from the engine environment, maintaining them at acceptable temperatures and thus maximizing their lifespan. These thermal protection panels also provide fire protection. They can be used in other areas of the nacelle where there is a risk of fire.

[0009] Thermal protection panels typically include at least one insulating layer, which can be made from silica fibers, ceramic, or a microporous material. This layer can be secured between metal strips, usually made of stainless steel.

[0010] During maintenance inspections on turbofan engines, numerous IFS components were found to have defects due to localized overheating of the composite structure. This overheating results from prior deterioration of the thermal protection, to the point where it can no longer protect the composite structure. Under these conditions, the composite structure can be subjected to a heat flux with an excessive temperature. This flux can significantly impact the mechanical integrity of a component within the structure, necessitating repair or even complete replacement. Late detection of such overheating therefore potentially leads to more complex and extensive replacements or repairs of the composite structure, and longer and more frequent aircraft grounding.However, at present, there is no solution to detect in flight heating of the composite structure of the 1'1 FS.

[0011] It is therefore desirable to propose a system that can more directly detect overheating of a composite structure such as that of the 1'1 FS during flight, analyze it, and define a maintenance operation within a timeframe compatible with the estimated damage. By optimizing maintenance operations in this way, it is possible to reduce the number of scrapped parts and the duration and frequency of aircraft grounding. The aircraft's operational reliability can thus be improved.

[0012] Summary

[0013] Embodiments relate to a method for thermal monitoring of a composite part in an aircraft, the method being implemented by a processor and comprising steps consisting of: acquiring and storing, by a processor on board the aircraft, temperature measurement data from a plurality of temperature sensors distributed over a surface of the composite part; defining a digital mesh of the surface of the composite part comprising a plurality of meshes; determining by the processor a temperature of each mesh as a function of the temperature measurement data and the respective positions of the temperature sensors on the surface of the composite part; for each mesh whose temperature is greater than a corresponding threshold value, calculating by the processor a damage index of the mesh as a function of a difference between the temperature of the mesh and the corresponding threshold value.

[0014] Thanks to these features, it is possible to determine in real time, with a spatial resolution dependent on the mesh used, the temperatures experienced locally by the composite part. Using the thermal properties of the composite part, it is possible to identify and anticipate potential damage to the composite part in the event of localized or more widespread overheating. Anticipating such damage allows for the determination of necessary maintenance operations while reducing the scope and therefore the duration of these operations. This results in improved operational reliability of the aircraft.

[0015] According to one embodiment, the calculation of the damage index (Ejj) of each mesh depends on a time during which the temperature of the mesh remained above the corresponding threshold value.

[0016] According to one embodiment, the process includes a calculation of the remaining service life of the composite part based on the calculated damage indices.

[0017] Determining the remaining lifespan allows for the planning of maintenance operations.

[0018] According to one embodiment, the method includes the application by the processor of a spatial confidence test to the temperature measurement data, the spatial confidence test comprising steps consisting of: considering a set of sensor pairs, each associating two sensors from the plurality of temperature sensors, calculating for each sensor a spatial confidence index as a function of temperature deviations calculated on the basis of the measurement data from each sensor pair to which the sensor belongs at a given time and a distance between the sensors of the sensor pair; and combining the measurement data from each sensor with the spatial confidence index calculated for the sensor to decrease the influence of the measurement data from the sensor as a function of the corresponding confidence index in the determination of the temperature of each mesh.

[0019] This provision helps to eliminate false measurements and therefore increases the reliability of the measurements.

[0020] According to one embodiment, the method includes the application by the processor of a temporal confidence test to the temperature measurement data, the temporal confidence test comprising steps consisting of: calculating for each sensor a temporal confidence index as a function of a temperature deviation calculated on the basis of the measurement data from the sensor taken at different times; and combining the measurement data from each sensor with the temporal confidence index calculated for the sensor to reduce the influence of the measurement data from the sensor as a function of the corresponding confidence index in the determination of the temperature of each mesh.

[0021] This provision also helps to eliminate false measurements and thus increase the reliability of the measurements.

[0022] According to one embodiment, the method includes the calculation by the processor of a confidence coefficient for each sensor, combining the spatial confidence index and the temporal confidence index, the measurement data from each sensor being combined with the confidence coefficient calculated for the sensor to decrease the influence of the measurement data from the sensor in the determination of the temperature of each mesh.

[0023] By combining temporal and spatial comparisons, the detection of false measurements is more accurate and reliable.

[0024] According to one embodiment, the temperature of each mesh is determined as a function of curvilinear distances, following a curvature of the composite part, between a center of the mesh and the respective positions of at least a part of the temperature sensors.

[0025] This improves the accuracy of the temperatures in each mesh.

[0026] According to one embodiment, the process includes a step of refining the respective temperatures of the meshes by taking into account the evolution of the mesh temperatures during a time interval, in order to reduce an impact of noise present in the measurement data from the sensors.

[0027] According to one embodiment, the damage index of each mesh is determined as a function of a sum of differences between mesh temperatures exceeding the corresponding temperature threshold value and the temperature threshold value.

[0028] The mesh damage index thus calculated is more accurately representative of the condition of the mesh.

[0029] According to one embodiment, the method comprises steps consisting of: transmitting maintenance data from stored measurement data, via a transmission network, when the aircraft is on the ground; receiving and storing, via a ground monitoring system, the transmitted maintenance data; detecting, based on the maintenance data, potential damage to the composite part; and determining, via the maintenance system, maintenance operations to be carried out based on the potential damage detected to the composite part.

[0030] According to one embodiment, the sensors are distributed: according to the configuration of hot fluxes reaching the composite part, and / or according to potential thermal effects on the composite part in case of damage to the thermal protection layer, and / or on areas of the composite part likely to be subjected to the highest temperatures, and / or on areas of the composite part subjected to the greatest mechanical stresses.

[0031] Embodiments may also relate to a monitoring system for a composite part, carried in an aircraft, the system comprising: a set of temperature sensors distributed over a surface of a composite part of the nacelle, and a processor connected to the set of sensors, the monitoring system being configured to implement the process as previously defined.

[0032] According to one embodiment, the monitoring system includes a transmission interface connected to the processor to communicate with an external monitoring system.

[0033] According to one embodiment, the sensors include: temperature sensors of the type having an RFID communication interface, and / or Bragg gratings on optical fibers, and / or thermocouples, and / or one or more thermal image sensors.

[0034] Embodiments may also relate to a monitoring system for a composite part in an aircraft, comprising: an on-board monitoring system as previously defined, and a ground-based system configured to receive and store data emitted by the monitoring system, and to determine maintenance operations to be performed on the composite part based on the data received from the monitoring system.

[0035] Brief description of the figures

[0036] The present invention will be better understood with the aid of the following description of exemplary embodiments with reference to the accompanying figures, in which identical reference symbols correspond to structurally and / or functionally identical or similar elements.

[0037] Figures 1A, 1B, and 1C schematically represent in cross-section an internal composite structural element of a turbojet nacelle associated with thermal protection, according to different embodiments. Figure 2 schematically represents, in a projected view onto a plane, the composite structural element, according to one embodiment.

[0038] Figures 3A, 3B, 3C and 3D schematically represent an embedded part of a monitoring system for a fixed composite structure according to various embodiments,

[0039] Figure 4 represents steps in a process executed by the monitoring system, according to one embodiment,

[0040] Figure 5 is a projected view onto a plane of the internal composite structural element, illustrating one step of the process, according to one embodiment.

[0041] Figure 6 is a projected view onto a plane of the internal composite structural element, illustrating one step of the process, according to one embodiment.

[0042] Figure 7 represents a temperature variation curve in a zone of the internal composite structure, illustrating one step of the process, according to one embodiment,

[0043] Figure 8 represents curves showing the variation of the structural deflection of parts of the structural element as a function of time and temperature experienced by the part of the structural element, illustrating one step of the process, according to one embodiment,

[0044] Figure 9 represents an example of a corrected service life variation curve of a part of the structural element as a function of time, illustrating a step in the process, according to one embodiment.

[0045] Figure 10 schematically represents modules of the monitoring system, according to one embodiment,

[0046] Figure 11 schematically represents modules of the monitoring system, according to another embodiment,

[0047] Detailed description

[0048] Figures 1A, 1B, 1C, and 2 depict a thermal protection layer 2 covering an aircraft composite structure 1, such as a fixed composite component of a turbofan engine nacelle. Layer 2 can be formed from several juxtaposed panels. In one embodiment, the thermal protection layer 2 is associated with temperature sensors SN distributed over the external surface of layer 2 (Figure 1A). The sensors can also be integrated within layer 2 (Figure 1B) or distributed across the surface of the composite structure 1, covered by layer 2, as illustrated in Figure 1C. The sensors SN can include temperature sensors and optionally, pressure sensors.

[0049] Figures 3A, 3B, 3C, and 3D depict an embedded component of a monitoring system adapted for monitoring a composite part. The embedded system includes a PRC processor and a TXI transmission interface connected to the PRC processor for communication with an external monitoring system, for example, a ground-based system when the aircraft is on the ground. The PRC processor is connected to the SN temperature sensors directly or indirectly via wired or wireless links. The PRC processor is configured to process the signals from the SN sensors.

[0050] In the example shown in Figure 3A, the TS temperature sensors are of the type featuring an RFID (Radio-Frequency Identification) communication interface communicating with an RFID reader connected to the PRC processor and one or more antennas, for example, of the UHF type. In the example shown in Figure 3B, the temperature sensors comprise Bragg gratings on OF optical fibers distributed across the composite structure 1, with the PRC processor connected to the optical fibers via an optoelectronic interface circuit (OPI). Each optical fiber transmits light pulses, with a portion of the incident light being reflected by each Bragg grating at the Bragg wavelength, while the remainder of the incident light is transmitted through the optical fiber. When the optical fiber undergoes deformation or a temperature change, the Bragg wavelength shifts.By detecting and measuring such a shift, it is possible to obtain localized temperature measurements along the fiber. In the example in Figure 3C, the sensors include thermocouples (TC) distributed across the composite structure 1 and connected to the PRC processor. In the example in Figure 3D, the sensors include one or more thermal imaging sensors (TIS) distributed across the composite structure, with the PRC processor able to directly receive and process the thermal images from the sensors. If necessary, an interface circuit can be provided to preprocess the thermal images. The TIS sensors can be positioned taking into account the configuration of heat fluxes likely to reach the composite structure 1.

[0051] The distribution of the SN sensors (TS, TC, OF, TIS) on the composite part 1 is determined based on the potential thermal effects in the event of damage to the thermal protection layer 2. In one embodiment, the temperature sensors are positioned to detect heat flow at an excessive temperature on the composite structural part, such an excessive temperature being able to reveal a breach in the thermal protection. For this purpose, a measured temperature can be considered excessive when the corresponding heat flow has sufficient energy to damage the composite structural part.

[0052] According to one embodiment, the temperature sensors are preferably positioned on areas of the composite structure 1 subjected to the greatest mechanical stresses. Indeed, these areas are predominant in the material health of the composite structure 1 as a whole.

[0053] To reduce the number of sensors required, they can be positioned on areas of the composite structure most likely to be subjected to the highest temperatures. These areas can be identified beforehand, for example, using a thermal imaging camera. As a result, the distribution of the SN sensors on the structure is not necessarily uniform, as illustrated in Figure 2.

[0054] Figure 4 illustrates steps S1 to S8 of a process executed by the monitoring system, and in particular the PRC processor, according to one embodiment. In step S1, the processor receives temperature measurement signals Tk (k = 1, ..., m) from m temperature sensors SNi to SN1. mThese signals are sampled into Tk,t measurements per unit of time t, according to a sampling frequency. In step S2, the PRC processor constructs a temperature map at each unit of time t, considering a mesh dividing the surface of the composite structure 1 into cells Mu, ..., Mjj, ... for example, as shown in Figure 5. The PRC processor calculates, for each unit of time t, a temperature Tjj.t at the center of each cell Mj based on the temperature samples Tk,t from the SNk sensors. To this end, it uses a distance Djj.k between each sensor k and the center of each cell Mj. The distances Djj.k, which can be provided by a database DB, are calculated in a calculation step S10 by a DC distance calculation function, based on the position Pk of each SNk sensor and the position of the center MCj of each cell Mj.The temperature Ti,j,t provided for each cell Mj can be obtained, for example, by a weighted sum of the temperatures Tk,t. In this weighted sum, each temperature Tk,t is multiplied by a coefficient FD that varies according to the distance Djj.k and a parameter representing the temperature propagation in the composite structure. Each of the distances Djj.k can be a curvilinear distance, i.e., following the curvature of the composite structure between the sensor SNk and the center of the cell Mjj. In calculating the temperature of each cell Mjj, only a subset of the sensors SNk can be considered, for example, only those located at a distance Djj.k from the cell that is less than a threshold distance value. Figure 6 illustrates the thermal map TM(t) obtained at the end of step S2.The TM(t) map extends over the entire surface of the monitored composite structure 1 and presents the temperature Ti,j,t calculated for each mesh Mjj for the temperature samples Tk,t taken at time t.

[0055] The thermal map TM(t) can be refined by taking into account its evolution during the time interval from t - e to t between time t and the time corresponding to the previous temperature sample, in order, for example, to reduce the impact of noise in the measurements. For this purpose, the PRC processor can use the following equation: in which e defines the duration of the time interval considered and v is a function determining a weighting according to the duration of the interval e.

[0056] In step S3, the PRC processor compares the temperature Tjj.t of each unit cell Mj,j at each time t to a maximum threshold temperature TTHjj defined for that unit cell Mjj. The threshold values ​​TTHjj can be read from the database DB. Below the threshold temperature TTHj, the unit cell is considered to be free of degradation due to excessive temperature. Above this threshold temperature, the unit cell is considered to be in a state of overheating. If none of the threshold values ​​TTHjj are exceeded, the PRC processor can assess the "health" state (SOH) of the composite structure 1 in step S4. If one or more of the threshold values ​​TTHj are reached or exceeded (at least one unit cell Mjj is considered to be in a state of overheating), the PRC processor executes step S5.

[0057] In step S5, the PRC processor calculates a damage index that depends on the amplitude of the overheating and the time each cell remained in a superheated state. The calculation of the damage index for a cell is illustrated in Figure 7, which shows a curve C1 of the temperature variation of a cell as a function of time t. In the example in Figure 7, the temperature Tjj of cell Mjj is approximately 55°C at time t0, reaches the threshold value TTHjj of 120°C at time t1, and continues to increase until it reaches 150°C at the current time t2. The damage index Ejj for cell Mjj can be estimated as a function of the area (t2 - t1) - (Tjj(t2) - Tjj(t1)) of rectangle R1. This area is also equal to (t2 - t1 )-(Tjj(t2) - TTHjj(t1)).

[0058] According to another embodiment, the damage index Ejj for the unit cell Mjj can be estimated using a DF function applied to the area OTSjj between the temperature variation curve C1 and the line corresponding to the threshold value TTHjj. The area OTSjj can be estimated for the unit cell Mjj by the following summation:

[0059] OTSj = Sf2 t1(TijW-TTHj) (2) In the next step S6, the PRC processor evaluates a structural degradation SAB(t) of the composite structure (1) at time t by applying a function FA to the indices Ej(t) calculated for each of the unit cells Mj of the composite structure. The structural degradation SAB(t) quantifies the extent and impact of overheating damage on the properties of the composite structure 1 relative to a nominal state. The structural degradation can be expressed as a percentage equal to (100% - residual mechanical strength), the residual mechanical strength representing the degradation of the mechanical properties of a part relative to an initial state, for example, at the end of a production line. The calculation of the structural degradation SAB(t) can use models of the behavior of the material forming the composite structure to identify and estimate the impact of overheating on its properties.These models, which are stored in the DB database, may include charts obtained by subjecting the materials of the composite structure and the entire composite structure to mechanical strength tests (e.g., compression and shear strength tests).

[0060] To illustrate the influence of temperature on structural degradation, Figure 8 shows curves of the variation of structural degradation of parts of the composite structure as a function of time and the temperature experienced by each part of the composite structure. The structural degradation values ​​shown in Figure 8 can be stored as tables in the database DB, providing the structural degradation as a function of the damage index Ejj of each part of the composite structure. Figure 8 shows that the structural degradation degrades more or less rapidly with increasing temperature, depending on the part of the composite structure considered.

[0061] In the next step, S7, the PRC processor evaluates the structural health status SOH(t) of the material constituting the composite structure 1 at the current time t. To this end, the PRC processor determines a structural margin MS. In the case of a composite material, the structural margin MS can be obtained by the ratio between an allowable deformation value of the material and a corresponding deformation value actually observed in the composite structure under a given loading situation (stress + thermal load), the deformation value of the composite structure being multiplied by the structural deflection SAB(t). These values ​​are defined during the design of the composite structure and provided by the database DB.

[0062] The allowable deformation of a material is defined using charts provided by the database. In the case of a composite material, the allowable deformation depends on several parameters, including the number of carbon plies, the orientation of the carbon fibers, and the characteristics of the honeycomb structure. This allowable deformation can be expressed in pStrain. In the case of a composite structure equipping an aircraft, this structure is designed to have a structural margin greater than 1 for a given load. This ensures that the part will have a structural capacity compatible with the aircraft's lifespan. Thus, the calculated structural margin allows for a precise evaluation of the composite structure's ability to remain operational. If this margin is insufficient to ensure the structural integrity of the composite structure, the PRC processor calculates a remaining service life (RLT) for the composite structure in step S8.

[0063] The remaining service life (RLT) determines the maximum period before the next maintenance operation on the composite structure. Data estimated in steps S7 and S8, including the remaining service life (RLT) and maintenance recommendations, can be stored in the database (DB), for example, at the end of each flight mission. The remaining service life (RLT) varies depending on the structural margin (MS). For example, the remaining service life (RLT) can be related to the structural margin (MS) by a linear function.

[0064] Figure 9 illustrates an example of the variation of the remaining service life (RSL) over time. In the example shown in Figure 9, the composite structure underwent two temperature failures: an initial overheating of 200°C for 200 hours after approximately 15,000 hours of service, followed by a second overheating of 140°C for 3,600 hours after approximately 30,000 hours. The first overheating reduced the remaining service life of the composite structure by approximately 60,000 hours out of its initial 120,000 hours. The second overheating reduced the remaining service life of the composite structure by approximately 25,000 hours, bringing the remaining service life to almost zero.

[0065] At the end of a mission (a flight for an aircraft), the DB database can be updated with the structural margin MS and the remaining life RLT which have just been calculated and maintenance recommendations determined in particular based on the remaining life and a schedule of planned maintenance operations.

[0066] In one embodiment, the PRC processor implements a learning loop configured to update the thermal reference data in the DB database, based on thermal readings from sensors, thus closely reflecting the operational conditions of the composite structure equipping aircraft. The DB database can also be used to enrich a general database that aggregates the databases of an aircraft fleet, and the learning loop can be applied to the general database.

[0067] Furthermore, damage assessment during maintenance operations can be used to adjust the TTHjj threshold values. This provision allows the composite structure monitoring parameters to be adapted to actual operating conditions, thus improving this monitoring.

[0068] In one embodiment, spatial confidence tests TC1 and temporal confidence tests TC2 are applied to temperature measurement signal samples Tk(t) to determine a confidence coefficient CCk(t) for each sample Tk(t). The confidence test TC1 comprises steps S11 and S12. In step S11, the PRC processor calculates, for each existing sensor pair (SNk, SNki), confidence indices Csk,ki based on the temperature measurement samples Tk(t) and Tki(t) recorded at time t and the distances D k,ki between the SNk and SNki sensors of the sensor pair. Thus, each confidence index CSk.ki can be calculated by applying a function FD1 to the temperature difference (T k - T k i) and at a distance D k ,ki between the SNk and SNki sensors. The FD1 function can be a polynomial type comparison function, for example in 1 / x.

[0069] Here too, each of the distances Dk,ki can be a curvilinear distance, that is to say, following the curvature of the composite structure between the sensors SNk and SNkij.

[0070] Thus, when the temperature difference is small compared to a nearby sensor, the temperature measurement has a relatively high confidence level. When the temperature difference is large compared to a distant sensor, the temperature measurement has a slightly lower confidence level. When the temperature difference is large compared to a nearby sensor, the temperature measurement has a significantly lower confidence level.

[0071] In step S12, the PRC processor calculates, for each temperature value Tk(t) at time t, a spatial confidence coefficient C1 k(t) by combining, using a function FC1, all the confidence indices CSk,ki obtained in step S11 for the value Tk(t). The FC1 function calculates, for example, the product of all the error coefficients CSk.ki determined for the measurement sample Tk(t), where k1 varies from 1 to the number m of sensors SNk, according to the following equation:

[0072] C1 k= n!?i=i CS k , k1 (3)

[0073] The FD1 function and more generally the calculation of the confidence coefficient C1 k (t) can be the subject of a learning loop so that the behavior of the confidence coefficient C1 k (t) adapts to the actual conditions encountered. According to one embodiment, the SNki sensors considered in equation (3) are restricted to those whose distance Dk,ki is less than a potentially variable threshold value.

[0074] The TC2 confidence test includes a step S13, during which the PRC processor calculates, for each value Tk(t), a temporal confidence coefficient C2k(t) based on the current temperature measurement sample Tk(t) and previous measurement samples Tk(t-1), Tk(te), recorded at times t-1, te by the SNk sensor. Thus, each confidence coefficient C2k can be calculated by applying a comparison function FT to the samples Tk(t), ..., Tk(te).

[0075] At step S14, the PRC processor combines the spatial confidence coefficient C1k(t) and the temporal confidence coefficient C2k(t) for each temperature sample value Tk(t) at time t using a combination function FC, to produce a confidence coefficient CCk(t) for each sample value Tk(t). The confidence coefficients CCk(t) thus obtained can be used in the calculations performed at step S2 by combining (e.g., multiplying) them by the respective sample values ​​Tk(t), in order to reduce the influence of each sample according to its corresponding confidence coefficient. In this case, the confidence coefficient CCk(t) can take values ​​between 1 if the corresponding sample value Tk(t) is completely reliable, and 0 if this value is determined to be completely erroneous.

[0076] Figure 10 illustrates the composite structure monitoring system. The system comprises an aircraft-mounted ACT control unit and a ground-based GCT monitoring control unit, GND, communicating with the ACT control unit via an NT network. The ACT control unit is implemented by the PRC processor, which is connected to the SN sensors and runs a monitoring MNTM module, a data storage DSM module, and a data transmission DTTX ​​module. The MNTM module receives measurements from the SN sensors at a rate determined by a polling cycle during periods when the turbofan engine is running. The MNTM module transmits the acquired measurements to the storage DSM module, which stores the data locally. The DTTX ​​module is activated when the aircraft is on the ground, and the control unit is connected to a ground-based GCT monitoring control unit.When activated, the DTTX ​​module transmits the data stored by the DSM storage module. For this purpose, the DTTX ​​module can access the NT network, both for transmission and reception, via an NT-compatible ATXI communication interface module.

[0077] The ACT control unit may also include a BITM test module that receives measurements from the SN sensors. The BITM module can be activated before aircraft takeoff to verify that the entire ACT unit and SN sensors are operational. If the test results do not meet expectations, the monitoring system can take action to avoid impacting aircraft availability.

[0078] In one embodiment, the ACT control unit may include a self-contained power supply unit, independent of the avionics network. This independent power supply unit optimizes the aircraft's power consumption. Activation of this independent unit can be performed according to the aircraft's use case, such as when using BITM and DTTX ​​modules.

[0079] In one embodiment, the ACT control unit is configured to acquire measurement data from the SN sensors at a sampling rate adapted to the flight phase. Thus, the sampling rate can be reduced when the current flight phase does not require high-frequency measurements. This reduces the volume of measurement data to be transmitted, thereby also reducing the power consumption of the onboard monitoring system.

[0080] The GCT ground surveillance control unit comprises a processor that implements an AMT access management module, a DTMT downloaded data management module, and a DTPR data processing module. The AMT module manages access to the ACT control unit for aircraft located on the ground nearby. To this end, the AMT module can communicate with the NT network via a GTXI communication interface module compatible with the NT network. The DTMT module stores the data downloaded by the AMT module in a ground database that stores downloaded data from a fleet of aircraft. The DTPR module is configured to access the downloaded database for an aircraft and to perform a maintenance diagnosis of composite part 1 in an aircraft nacelle.

[0081] In one embodiment, the DSM data storage module applies compression processing to the data before storing it. This reduces the volume of data to be stored and transmitted to the ground. The DTMT or DTPR module is then configured to apply a corresponding decompression process to the data received from the aircraft.

[0082] In one embodiment, the ATXI module is an onboard transmission system used by other aircraft equipment. An alternative solution involves using the DTTX ​​transmission module of the ACT control unit to communicate data collected during flight to the ground monitoring system GND via the NT network.

[0083] The DTTX ​​module is configured to detect the communication protocol implemented in the aircraft and activate a corresponding communication interface. Thus, the DTTX ​​module can implement, for example, the ARINC, AFDX ("Avionics Full DupleX"), or CAN ("Controller Area Network") protocols, and select the one used in the aircraft where the nacelle composite structure monitoring system is installed. The GTXI ground interface can be configured to implement the same communication protocol as the ATXI interface. The NT network can be LoRa, 5G, or WiFi™.

[0084] In another embodiment, the NT network is separate from the network used by other aircraft equipment to communicate with ground systems and is dedicated to transmitting measurement data from the SN sensors. The ATXI and GTXI communication interface modules, both onboard and on the ground, are specific interfaces. This gives the monitoring system greater autonomy, facilitating the integration of the ACT unit onboard the aircraft and expanding its application possibilities to a wider range of aircraft.

[0085] In one embodiment, the ATXI and GTXI communication interface modules onboard the aircraft and on the ground implement an authentication protocol to allow the AMT access management module to approve or deny the download of data from the ACT control unit, based on the authentication result. The implemented authentication protocol can be configured to allow the onboard ACT unit to send a request to the GND ground control unit, which then performs an analysis to approve or deny the download of data from the ACT unit to the DTMT data management module. This measure aims to prevent potential cybersecurity risks.

[0086] In one embodiment, the raw, unprocessed data from the SN sensors is stored by the DSM module and transmitted to the ground by the DTTX ​​module. Data processing is handled entirely by the DTPR module of the GCT unit on the ground. The DTPR module can be configured to (1) validate the data measured by the SN sensors by applying confidence tests such as TC1 and TC2, (2) construct a temperature distribution map of the composite structure (step S2, Figure 6), and (3) assess the material health of the composite structure (steps S3 to S7), for example, by calculating margins or applying thresholds. The material health of the composite structure allows for the definition of any necessary maintenance operations. In another embodiment illustrated in Figure 11, the ACT unit includes a DPPR module for preprocessing the data from the SN sensors.The DPPR module processes the data received from the sensors and stored by the DSM module in real time. The DTTX ​​module can then transmit only the results of the data processing performed by the DPPR module. This reduces the volume of data transmitted over the NT network, thus allowing the NT network used by the system to have a lower data rate.

[0087] In one example, the DPPR module is configured to perform all or part of the data validation processes for measurement data from sensors, to construct a temperature distribution map, and to assess the material health of the composite structure. The DPPR module can thus be configured to perform maintenance diagnostics and generate alerts that are transmitted to the ground-based GCT unit. In this case, the GCT unit is responsible for supplementing the processing performed by the DPPR module, particularly for generating diagnostic reports and recommendations for aircraft operators.

[0088] Thanks to these provisions, the nacelle's IFS environment can be secured by enabling damage detection as soon as it appears, or even potential damage detection before it appears, and the determination of appropriate maintenance operation recommendations.

[0089] It will be readily apparent to those skilled in the art that the present invention is susceptible to various embodiments and applications. In particular, not all the steps for calculating the remaining lifespan of the composite structure are necessarily performed by the onboard PRO processor; some of the final steps of the process can be executed by a ground-based computer using the thermal map of the composite structure, formed from the respective temperatures of the individual cells.

[0090] Furthermore, the damage index Ejj of each mesh Mj can be evaluated without taking into account the time during which the temperature of the mesh remained above the threshold value associated with the mesh.

Claims

DEMANDS 1. Method for thermal monitoring of a composite aircraft part, the method being implemented by a processor (PRC) and comprising steps consisting of: acquiring and storing, by a processor (PRC) on board the aircraft, temperature measurement data from a plurality of temperature sensors (SN) distributed over a surface of the composite part (1); defining a digital mesh of the surface of the composite part comprising a plurality of meshes (Mj); determining by the processor a temperature (Tj) of each mesh as a function of the temperature measurement data and the respective positions of the temperature sensors on the surface of the composite part; for each mesh whose temperature is greater than a corresponding threshold value (TTHij), calculating by the processor a damage index (Ejj) of the mesh as a function of a difference between the temperature of the mesh and the corresponding threshold value.

2. Method according to claim 1, wherein the calculation of the damage index (Ejj) of each mesh depends on a time during which the temperature of the mesh remained above the corresponding threshold value.

3. Method according to claim 1 or 2, comprising a calculation of a remaining life (RLT) of the composite part (1) as a function of the calculated damage indices (Ejj).

4. A method according to any one of claims 1 to 3, comprising the application by the processor (PRC) of a spatial confidence test (TC1) to the temperature measurement data, the spatial confidence test comprising the steps of: considering a set of sensor pairs, each pair associating two sensors from the plurality of temperature sensors (SN), calculating for each sensor a spatial confidence index (C1) as a function of temperature deviations calculated on the basis of the measurement data from each sensor pair to which the sensor belongs at a given time and a distance between the sensors in the sensor pair; and combining the measurement data from each sensor with the spatial confidence index (C1) calculated for the sensor to reduce the influence of the data. measurement from the sensor according to the corresponding confidence index in determining the temperature (Tjj) of each mesh (Mj).

5. A method according to any one of claims 1 to 4, comprising the application by the processor (PRC) of a temporal confidence test (TC2) to the temperature measurement data, the temporal confidence test comprising steps of: calculating for each sensor (SN) a temporal confidence index (C2) as a function of a temperature deviation calculated on the basis of the measurement data from the sensor taken at different times; and combining the measurement data from each sensor with the temporal confidence index (C2) calculated for the sensor to decrease the influence of the measurement data from the sensor as a function of the corresponding confidence index in the determination of the temperature (Tj) of each mesh (Mj).

6. Method according to claims 4 and 5, comprising the calculation by the processor of a confidence coefficient (CC) for each sensor (SN), combining the spatial confidence index (C1) and the temporal confidence index (C2), the measurement data from each sensor being combined with the confidence coefficient (CC) calculated for the sensor to decrease the influence of the measurement data from the sensor in the determination of the temperature (Tj) of each mesh (Mj).

7. Method according to any one of claims 1 to 6, wherein the temperature (Tj) of each mesh (Mj) is determined as a function of curvilinear distances, following a curvature of the composite part (1), between a center of the mesh and the respective positions of at least a part of the temperature sensors (SN).

8. A method according to any one of claims 1 to 7, comprising a step of refining the respective temperatures (Tj) of the meshes (Mj) taking into account the evolution of the temperatures of the meshes during a time interval, in order to reduce a noise impact present in the measurement data from the sensors (SN).

9. A method according to any one of claims 1 to 8, wherein the damage index (Ejj) of each mesh (Mj) is determined as a function of a sum of differences between mesh temperatures (Tj) exceeding the corresponding temperature threshold value (TTHj) and the temperature threshold value.

10. A method according to any one of claims 1 to 9, comprising steps of: transmitting maintenance data from stored measurement data, via a transmission network (TN), when the aircraft is on the ground; receiving and storing, via a ground-based monitoring system (GCT), the transmitted maintenance data; detecting, based on the maintenance data, potential damage to the composite part; and determining, by the maintenance system, maintenance operations to be carried out based on the potential damage detected to the composite part.

11. A method according to any one of claims 1 to 10, wherein the sensors (SN, TS, TC, OF, TIS) are distributed: according to the configuration of hot fluxes reaching the composite part (1), and / or according to potential thermal effects on the composite part in case of damage to the thermal protection layer (2), and / or on areas of the composite part likely to be subjected to the highest temperatures, and / or on areas of the composite part subjected to the greatest mechanical stresses.

12. A composite part monitoring system, carried in an aircraft, the system comprising: a set of temperature sensors (SN) distributed over a surface of a composite part (1) of the aircraft, and a processor (PRC) connected to the set of sensors, the monitoring system being configured to implement the method according to any one of claims 1 to 11.

13. Monitoring system according to claim 12, comprising a transmission interface (TXI) connected to the processor (PRC) for communicating with an external monitoring system.

14. Monitoring system according to claim 12 or 13, wherein the sensors (SN, TC, OF, TIS) comprise: temperature sensors (SN) of the type having an RFID communication interface, and / or Bragg gratings on optical fibers (OF), and / or thermocouples (TC), and / or one or more thermal imaging sensors (TIS).

15. A composite part monitoring system in an aircraft, comprising: an on-board monitoring system, according to any one of claims 12 to 14, and a ground-based system configured to receive and store data emitted by the monitoring system, and to determine maintenance operations to be performed on the composite part based on the data received from the monitoring system.

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