Device for characterizing wear on a gas turbine engine

Abstract

Device (1) for characterizing wear on a gas turbine engine comprising: - an abradable material (8) including a wear end (80) adapted to be in contact with a flow of gaseous fluid (F) flowing from upstream to downstream of the gas turbine engine (2) when the gas turbine engine (2) is running; - an ultrasonic transceiver (7) adapted to emit an ultrasonic wave in the abradable material (8) and receive a corresponding ultrasonic wave reflected by an internal surface of the wear end (80); - an electronic circuit (10) connected to the ultrasonic transceiver (7) and configured to calculate information for characterizing the wear of the gas turbine engine (2). Figure for the abstract: Figure 2

Description

Title of the invention: Device for characterizing wear on a gas turbine engine FIELD OF INVENTION

[0001] The present invention relates to the field of gas turbine engines. More specifically, it relates to a device for characterizing wear in a gas turbine engine. STATE OF THE ART

[0002] Wear, i.e., erosion, is a mode of damage to gas turbine engines, particularly aircraft engines. It involves the removal of material by abrasion of particles carried along in the gaseous fluid flow circulating from the upstream to the downstream side of the gas turbine engine (air stream). This wear primarily affects the profiles of the blades and vanes of the fixed and moving parts of compressors and turbines.

[0003] Depending on the environments in which the aircraft operates, particularly for a helicopter, if the engine air intake is not equipped with a filtration system, in a desert environment, wear can be significant and reduce the interval between two major maintenance to a few dozen flight hours.

[0004] One common solution for ensuring that the wear level is below the certified limit guaranteeing the engine's operability throughout its entire flight envelope is to insert a videoscope equipped with a camera into the engine's airflow channel through endoscopic ports designed for this purpose and to inspect the interior via a control panel and its screen using the video feed provided by the videoscope. The acquired image allows the technician to determine the level of internal engine wear and thus determine whether maintenance is required.

[0005] Furthermore, as proposed in patent application no. FR3006013, it is possible to position an abradable material in the air stream that wears down as quickly as the engine part being monitored. This abradable material is visually inspected regularly to determine engine wear.

[0006] Most of the known methods require the use of a maintenance technician and specific tools. These methods are not automatic.

[0007] Automated wear measurement methods also exist, but these are implemented only by gas turbine engine manufacturers and are mainly used for break-in tests before delivery to customers. These tests are therefore carried out under very specific and controlled conditions.

[0008] However, the constraints to which aircraft operators are subjected require a technology that allows the wear of the various parts of gas turbine engines to be characterized in a more precise, robust, reliable and reproducible way.

[0009] There is therefore no satisfactory technology at present allowing for the measurement in real time, in a precise, robust, reliable and reproducible manner of wear in gas turbine engines. Description of the invention

[0010] One object of the invention is to characterize, in particular to measure, the wear of a gas turbine engine in a precise, robust, reliable and reproducible manner.

[0011] According to a first aspect, a device is proposed for characterizing the wear of a gas turbine engine comprising:

[0012] - an abradable material comprising a wear end adapted to be arranged in contact with a flow of gaseous fluid circulating from upstream to downstream of the gas turbine engine when the gas turbine engine is in operation;

[0013] - an ultrasonic transceiver adapted to emit an ultrasonic wave in the abradable material and receive a corresponding ultrasonic wave reflected by an internal surface of the wear end;

[0014] - an electronic circuit connected to the ultrasonic transceiver and configured for calculate information to characterize the wear of the gas turbine engine.

[0015] According to advantageous and non-limiting features, taken alone or in any combination: - the device includes a body adapted to carry the abradable material and the ultrasonic transceiver, the body being arranged between the abradable material and the ultrasonic transceiver; - The device includes a communication module connected to the electronic circuit and configured to transmit information characterizing the wear of the gas turbine engine to a data processing module configured to characterize the wear of the gas turbine engine based on the information characterizing the wear of the gas turbine engine.

[0016] According to a second aspect, a gas turbine engine is proposed comprising at least one device as described above, the device being arranged so that the wear end of the abradable material is in contact with a flow of gaseous fluid circulating from upstream to downstream of the gas turbine engine when the gas turbine engine is in operation.

[0017] According to advantageous and non-limiting features, the device is arranged in a hole drilled in the gas turbine engine, in particular the device is arranged in an endoscopic port of the gas turbine engine.

[0018] According to a third aspect, an aircraft comprising a gas turbine engine as previously presented is proposed.

[0019] According to a fourth aspect, a system is proposed for characterizing the wear of a gas turbine engine comprising: - a device for characterizing wear on a gas turbine engine as previously described; - a data processing module adapted to be connected to the electronic circuit, the data processing module being configured to characterize the wear of the gas turbine engine from the information enabling the characterization of the wear of the gas turbine engine.

[0020] According to a fifth aspect, a method is proposed for characterizing wear on a gas turbine engine by means of a system for characterizing wear on a gas turbine engine as previously described, the method comprising the steps of: - emission, by the ultrasonic transceiver, of an ultrasonic wave in the abradable material, - reception, by the ultrasonic transceiver, of a corresponding ultrasonic wave reflected by the internal surface of the wear end, - calculation, by the electronic circuit, of the information allowing to characterize the wear of the gas turbine engine.

[0021] According to advantageous and non-limiting features, taken alone or in any combination:

[0022] - the process includes a step of determination by the processing module of data representing the wear level of the gas turbine engine, derived from information used to characterize the wear of the gas turbine engine; the wear level data being a wear thickness of the device calculated by the electronic circuit or data processing module from:

[0023] - the speed of ultrasonic waves, and

[0024] - a time elapsed between the emission of the ultrasonic wave in the abradable material and the reception of the corresponding ultrasonic wave reflected by the internal surface of the wear end,

[0025] - and optionally the amplitude of the reflected ultrasonic wave;

[0026] - the method includes a step of determining the engine wear level gas turbine, by the data processing module, based on a representative data of the wear level determined from the information allowing to characterize the wear of the gas turbine engine and a predetermined nomogram indicating a correspondence between the representative data of the wear level and a wear level of a gas turbine engine. DESCRIPTION OF THE FIGURES

[0027] Other features and advantages of the present invention will become apparent from the following description of a preferred embodiment. This description will be given with reference to the accompanying figures, including:

[0028] - Figure 1 schematically illustrates a gas turbine engine comprising at least one device to characterize wear on the gas turbine engine;

[0029] - Fig. Ibis schematically illustrates another embodiment of a gas turbine engine including at least one device for characterizing wear on the gas turbine engine;

[0030] - Figure 2 schematically illustrates a device for characterizing wear on a turbine engine gas arranged on an engine casing, the wear end of the abradable material of the device being unworn;

[0031] - Figure 2bis schematically illustrates, according to another embodiment, a device for characterize wear on the gas turbine engine arranged on an engine casing, the wear end of the abradable material of the device being unworn;

[0032] - Figure 3 schematically illustrates a device for characterizing wear on a turbine engine gas arranged on an engine casing, the wear end of the abradable material of the device being partially worn;

[0033] - Figure 4 schematically illustrates a device for characterizing wear on a turbine engine gas arranged on an engine casing, the wear end of the abradable material of the device being completely worn;

[0034] - [Fig.5] schematically illustrates the path of ultrasonic waves in the device;

[0035] - Figure 6 schematically illustrates a system for characterizing wear on a turbine engine gas;

[0036] - Figure 7 shows the steps of a process for characterizing wear on a gas turbine engine. DETAILED DESCRIPTION OF THE INVENTION Motor and device

[0037] With reference to [Fig.1], a gas turbine engine 2, or turbomachine, is proposed, which includes a gas generator 24 and optionally, particularly in the case of an aircraft engine, a fan 22, which may be shrouded or unshrouded.

[0038] In the example illustrated in [Fig.1], the gas generator 24 comprises, from upstream to downstream, the gases (gaseous fluid flow) flowing within the gas turbine engine 2 from upstream to downstream, a compressor 24A (or compressor section 24A), a combustion chamber 24B, a turbine 24C (or turbine section 24C) and an exhaust nozzle 26.

[0039] The blower 22 can be driven in rotation directly by a shaft 23 of the gas generator 24, for example a shaft of a low pressure body, or via a gearbox GB (“Gear Box”) or RGB (Anglo-Saxon acronym for “Reduction Gear Box”) mechanically connected to the compressor 24A.

[0040] The gas generator 24 can be of the twin-body type and comprise a low-pressure body and a high-pressure body.

[0041] The low-pressure body may include a low-pressure compressor 241A rotationally coupled with a low-pressure turbine 24IC via a low-pressure shaft, not shown.

[0042] The high-pressure body may include a high-pressure compressor 242A disposed downstream of the low-pressure compressor 241A and upstream of the combustion chamber 24B, and a high-pressure turbine 242C, disposed downstream of the combustion chamber 24B and upstream of the low-pressure turbine 241C, and coupled in rotation with the high-pressure compressor 242A by means of a high-pressure shaft, not shown.

[0043] The compressor 24A of the gas generator 24 can include the low and high pressure compressors 241A and 242A. The turbine 24C of the gas generator 24 can include the low and high pressure turbines 241C and 242C.

[0044] The [Fig. 1] is schematic, each compressor and each turbine being able to have one or more stages, each stage comprising a moving wheel (rotor) and a stator.

[0045] The exhaust nozzle 26 allows the exhaust gases that have circulated in the gas turbine engine 2 to exit the gas turbine engine 2.

[0046] The gas turbine engine 2 further comprises a casing 17 corresponding to its outer shell.

[0047] The casing 17 can be in contact, on its external surface, with the exterior of the gas turbine engine 2 and, on its internal surface, with the air stream. The casing 17 can be composed of all or some of the fixed parts (i.e., the stators or the housings of the compressor impellers) of the compressors and turbines.

[0048] The internal surfaces of the housing 17 may or may not be coated with an abradable layer 14. This abradable layer 14 ensures good contact behavior, particularly when a rotor, for example of a compressor, comes into contact with the housing of this compressor.

[0049] Figure 1bis illustrates another embodiment of a gas turbine engine 2. This gas turbine engine 2 is particularly suitable for helicopters. This engine does not include a fan. It comprises a gas generator section with a bi-centrifugal compressor 30 comprising two compression wheels 31, 32 respectively attached to a first coaxial turbine 33. The compression wheels 31, 32 are respectively designated as the first compressor 31 and Second compressor 32. The air stream 37 inside the casing 17 is annular and extends from an air inlet 37a, which guides the air to the axial inlet of the first compressor 31. The air inlet 37a can be axial or annular. The air compressed by the first compressor 31 is guided radially through a diffuser 37b. The air stream then forms a bend 37c to return the air towards the axis of the engine 2 to the axial inlet of the second compressor 32. The air is then guided to a combustion chamber 35, which supplies the first turbine 33 with hot gas. The expansion of the gases continues in a second turbine 34 of a second rotor attached to a power take-off shaft 23 for driving the load. The air stream is delimited by two coaxial walls, one of which is an internal wall 37i of the housing 17. In [Fig. Ibis], we see the bend 37c of the air stream, downstream of the diffuser 37b.The function of this elbow 37c is to divert the airflow from the diffuser towards the axis of the machine. The gases are then expelled through the exhaust nozzle 26.

[0050] The gas turbine engine 2 includes a device 1 for characterizing wear of the gas turbine engine 2.

[0051] Figure 1 schematically illustrates a gas turbine engine 2 particularly suitable for airplanes, and Figure 1bis schematically illustrates a gas turbine engine 2 particularly suitable for helicopters. However, device 1 is not limited to use in a particular type of gas turbine engine 2.

[0052] The gas turbine engine 2, more specifically its internal components such as the compressor(s) and turbine(s), can wear, i.e., erode, due to particles present in the gaseous fluid flow F circulating within the gas turbine engine 2. The gaseous fluid may include air and / or exhaust gases. The particles are volatile particles such as sand or gravel that are stirred up when the aircraft incorporating the gas turbine engine 2 lands or takes off on unprepared surfaces. This wear leads to malfunctions of the gas turbine engine. It is therefore necessary to characterize it, in particular to measure it, so as to be able to precisely identify when engine maintenance is required.

[0053] By "characterizing wear on the gas turbine engine 2," it is understood that one or more characteristics of the wear are to be determined. Preferably, the aim is to determine a level of wear on the gas turbine engine 2. This level of wear can be determined, for example, from the wear thickness of the engine. The wear thickness of the engine can refer to a thickness of material in the gas turbine engine 2 that has disappeared through wear, i.e., erosion, or a remaining thickness of material in the gas turbine engine 2.

[0054] Device 1 is thus adapted to calculate information enabling the characterization of the wear of the gas turbine engine 2. From this information, which will be detailed later, a level of wear of the gas turbine engine 2 can be determined.

[0055] Furthermore, by "characterizing wear on the gas turbine engine 2", it is understood that one seeks to determine wear on at least one part of the gas turbine engine 2, in particular the part which is located near the device 1. However, this wear may allow the wear of another part of the gas turbine engine 2 to be determined.

[0056] With reference to Figures 2 to 4, the device 1 for characterizing wear on the gas turbine engine 2 comprises an abradable material 8. The abradable material 8 includes a wear end 80 adapted to be arranged in contact with the flow of gaseous fluid F circulating from upstream to downstream of the gas turbine engine 2 when the gas turbine engine 2 is in operation. In Figures 2 to 4, the device 1 is shown arranged on a housing 17 of the engine 2.

[0057] By "in operation" is meant a state of the gas turbine engine 2 in which the moving parts (rotors) of the turbines and compressors are driven in rotation so that an air circulation from upstream to downstream of the gas turbine engine 2 takes place.

[0058] More specifically, in the case of the gas turbine engine 2 illustrated in [Fig. 1], the fan 22 is driven in rotation by a turbine, advantageously the low-pressure turbine 241C, resulting in air circulation from the upstream to the downstream side of the gas turbine engine 2. Part of the air passing through the fan 22 (air stream) passes successively through the low-pressure compressor 241A, the high-pressure compressor 242A, and is then injected into the combustion chamber 24B. In the combustion chamber 24B, the air is mixed with fuel. The combustion of the fuel generates exhaust gases which circulate successively through the high-pressure turbine 242C, then through the low-pressure turbine 241C, and are discharged via the exhaust nozzle 26.

[0059] In the case of the gas turbine engine 2 illustrated in [Fig. Ibis], air flows in the air stream 37, from the air inlet 37a, to the first compressor 31, then to the rectifier 37b, then through the elbow 37c, then through the second compressor 32 and is injected into the combustion chamber 35. In the combustion chamber 35, the air is mixed with fuel. The combustion of the fuel generates exhaust gases which flow successively through the first turbine 33, then through the second turbine, and are discharged via the exhaust nozzle 26.

[0060] By "abradable," it is understood that the abradable material 8 can be worn, i.e., eroded. Advantageously, the abradable material 8 is made of a material that wears, due to particles in the gaseous fluid flow F, at the same rate as the gas turbine engine element 2 for which wear is to be characterized, or at a rate proportional to the wear rate of the gas turbine engine element 2. which we wish to characterize the wear. Thus, it is easy to establish a correlation between the wear of the abradable material 8 and the wear of the element of the gas turbine engine 2 for which we wish to characterize the wear.

[0061] The abradable material 8 is advantageously made up of one or more materials allowing the circulation of ultrasound within it.

[0062] Preferably, the abradable material 8 comprises or is made of aluminium, an iron-nickel alloy (FeNi), nickel-coated graphite (Ni-graphite), a copper-silicon alloy (AISi) and / or a copper-silicon alloy (CuSi).

[0063] The abradable material 8 can take several forms. The shape is preferably adapted so as not to impair aerodynamics. For example, it can take the form of a parallelepiped or a cylinder. The cross-section of the abradable material 8 can be a polygon, as illustrated for example in Figures 2 and 4. Optionally, the polygon can include arcs of circles. In [Fig. 2], the cross-section of the abradable material 8 has a rectangular shape (this can, for example, be the case for an abradable material 8 in the shape of a cylinder).

[0064] The abradable material 8 includes a wear end 80 adapted to be in contact with the gaseous fluid stream F when the gas turbine engine 2 is operating. Advantageously, the wear end 80 is adapted to protrude from the abradable material 8 into the gaseous fluid stream F. The wear end 80 is therefore the part of the abradable material 8 adapted to wear away, potentially to the point of disappearance.

[0065] As illustrated in Figures 2 and 2bis, the abradable material 8 is complete and therefore unworn. Similarly, the abradable layer 14 coating the housing 17 is unworn.

[0066] Figure 3 illustrates the abradable material 8 being worn due to the flow of gaseous fluid F; it can be seen that the abradable material 8 has lost material. Similarly, the abradable layer 14 coating the housing 17 is worn and therefore thinner.

[0067] According to one embodiment, the abradable material 8 of [Fig.3] could correspond to a complete and therefore unworn abradable material 8.

[0068] Figure 4 illustrates the completely worn abradable material 8, i.e., the wear end 80 has disappeared. Similarly, the abradable layer 14 coating the housing 17 has disappeared.

[0069] The abradable layer 14, possibly present on the part of the motor 2 under study, is advantageously made of a material that erodes at the same rate as the abradable material 8. Consequently, the abradable layer 14 will be completely worn if the abradable material 8 is completely worn (i.e., the wear end 80 has disappeared), and vice versa. Thus, the wear of the abradable material 8 is directly representative of the wear of the part of the motor 2 under study. However, the abradable layer 14 may be made of a material that does not erode at the same rate as the abradable material 8. Therefore, the abradable material 8 will be able to erode faster or slower than the abradable layer 14. Preferably, if the abradable layer 14 is made of a material that does not erode at the same rate as the abradable material 8, the abradable layer 14 is made of a material that erodes at a rate proportional to the rate of wear of the abradable material 8.

[0070] The device 1 further includes an ultrasonic transceiver 7. The ultrasonic transceiver 7 is adapted to emit an ultrasonic wave into the abradable material 8, in particular towards the wear end 80, and to receive a corresponding ultrasonic wave reflected by an internal surface 81 of the wear end 80.

[0071] By “corresponding reflected ultrasonic wave”, it is understood that a wave resulting from the reflection of the incident wave (i.e. wave emitted by the ultrasonic transceiver 7 in the abradable material 8 towards the wear end 80).

[0072] By "internal surface of the wear end 80", it is understood to mean the interface between the wear end 80 and the outside, the outside being able to correspond to the inside of the gas turbine engine 2 where the gaseous fluid flow F is intended to circulate when the device 1 is assembled to the gas turbine engine 2.

[0073] The ultrasonic transceiver 7 is configured to emit and receive, simultaneously or not, an ultrasonic wave. The emitted ultrasonic wave has a frequency advantageously suited to propagate through the abradable material 8.

[0074] The ultrasonic transceiver 7 is advantageously robust enough to be installed in the gas turbine engine 2.

[0075] The device 1 for characterizing wear of the gas turbine engine 2 advantageously comprises a body 9 adapted to carry the abradable material 8 and the ultrasonic transceiver 7. The body 9 is preferably arranged between the abradable material 8 and the ultrasonic transceiver 7.

[0076] The body 9 is advantageously adapted to carry the abradable material 8 and the ultrasonic transceiver 7 such that a wave emitted by the ultrasonic transceiver 7 can penetrate the abradable material 8 and a wave reflected by the internal surface 81 is received by the ultrasonic transceiver 7. In other words, the body 9 is advantageously adapted to carry the abradable material 8 and the ultrasonic transceiver 7 such that the ultrasonic transceiver 7 is oriented towards the abradable material 8.

[0077] Advantageously, the body 9 is made of a material that allows the propagation of ultrasonic waves. The body 9 can, for example, be made of stainless steel or aluminum.

[0078] The abradable material 8 is advantageously fixed to the body 9 so as to limit the waves reflected at the interface between the body 9 and the abradable material 8.

[0079] According to one embodiment, the abradable material 8 is inserted into a cavity formed in the body 9 and provided for this purpose. The abradable material 8 is in this embedded in the body 9. Embedded means that the abradable material 8 is partially embedded in the body 9, with at least the wear end 80 of the abradable material 8 not embedded in the body 9 so as to be exposed to the gas flow. This embodiment prevents damage to the abradable material 8 when the device 1 is installed on the gas turbine engine 2. Advantageously, there is no additional material between the body 9 and the abradable material 8 in this embodiment, which limits the reflected waves at the interface between the body 9 and the abradable material 8.

[0080] The abradable material 8 can be bonded, welded, or sprayed onto the body 9, as illustrated in [Fig. 2bis]. The embodiment of [Fig. 2bis] allows for improved sensitivity of the device 1. In other words, the device 1 will enable more precise wear measurements.

[0081] The abradable material 8 can be fixed to the body 9 both by insertion and by gluing and / or welding.

[0082] According to one embodiment, the body 9 and the abradable material 8 form a single piece and are therefore made of the same material. This ensures good circulation of ultrasound within the body 9.

[0083] Similarly, the ultrasonic transceiver 7 can be inserted into a cavity formed in the body and provided for this purpose and / or glued and / or welded to the body 9.

[0084] According to one embodiment, the ultrasonic transceiver 7 is reversibly attached to the body 9. In other words, the ultrasonic transceiver 7 is removable. For example, only the body 9 and the abradable material 8 can be attached to the gas turbine engine 2 so as to be permanently installed in the aircraft. The transceiver 7 can be used and connected to the body 9 only when a maintenance operation is carried out and the wear of the engine 2 is to be determined.

[0085] Preferably, the interfaces for connecting the body 9 and the abradable material 8 and / or the body 9 and the ultrasonic transceiver 7 are adapted to allow the flow of ultrasonic waves between the ultrasonic transceiver 7 and the abradable material 8, via the body 9.

[0086] Device 1 is advantageously arranged on a fixed part of the gas turbine engine 2. Preferably, device 1 is arranged on the engine 2 so as to be accessible from outside the engine 2 to facilitate maintenance and data acquisition.

[0087] Advantageously, the device 1 is adapted to be arranged in an endoscopic port of the motor 2 and is therefore adapted to replace an endoscopic plug of the motor 2. The body 9, and more generally the device 1, therefore advantageously has the shape of a plug.

[0088] The device 1 for characterizing wear of the gas turbine engine 2 comprises an electronic circuit 10 connected to the ultrasonic transceiver 7. The ultrasonic transceiver 7 is advantageously electrically connected to the electronic circuit 10.

[0089] The electronic circuit 10 is advantageously configured to supply the ultrasonic transceiver 7 with energy.

[0090] The electronic circuit 10 is configured to allow an exchange of information with the ultrasonic transceiver 7.

[0091] The electronic circuit 10 is advantageously configured to calculate information enabling the characterization of the wear of the gas turbine engine 2, from data received by the electronic circuit 10 from the ultrasonic transceiver 7.

[0092] The information used to characterize the wear of the gas turbine engine 2 can be a time elapsed between the emission of an ultrasonic wave in the abradable material 8 and the reception of the corresponding ultrasonic wave reflected by the internal surface 81 of the wear end 80. Indeed, it will be seen later that this time allows the determination of a wear thickness of the device 1 which makes it possible to characterize the wear of the device 1, in particular of the abradable material 8, which makes it possible to characterize the wear of the engine 2.

[0093] Advantageously, the electronic circuit 10 is configured to identify the correct reflected wave, i.e., the one reflected by the internal surface 81 of the wear end 80, in order to calculate said duration. Indeed, when an ultrasonic wave is emitted towards the abradable material 8, the wave is reflected by the interface between the ultrasonic transceiver 7 and the abradable material 8. The wave is also reflected by the internal surface 81 of the wear end 80 (i.e., by the interface between the abradable material 8, more precisely the wear end 80, and the outside) due to the change in material (change from the abradable material 8 to the air of the vein). There are at least two reflected waves arising from the incident wave.

[0094] As illustrated in [Fig. 5], in the case where the device 1 comprises a body 9, there are at least three reflected waves: a wave reflected by the interface between the ultrasonic transceiver 7 and the body 9 (not shown), a wave reflected OR1 by the interface between the body 9 and the abradable material 8, and a wave reflected OR2 by the internal surface 81 of the wear end 80. The wave OI in [Fig. 5] corresponds to the incident wave, i.e., the wave emitted by the ultrasonic transceiver 7. [Fig. 5] is schematic, and it is evident that several incident waves OI will be emitted at the same instant by the ultrasonic transceiver 7, the incident waves OI each arriving at a different point on the interface between the body 9 and the abradable material 8 and on the internal surface 81 of the wear end. 80. Thus, several reflected waves OR1 and several reflected waves OR2 will be received by the ultrasonic transceiver 7.

[0095] Consequently, the electronic circuit 10 is advantageously configured so as to know in advance an expected time between the emission of an ultrasonic wave in the abradable material 8 and the reception of a corresponding ultrasonic wave reflected by the internal surface 81 of the wear end 80. The electronic circuit 10 is therefore advantageously configured so as not to take into account the other reflected waves which are received before the wave reflected by the internal surface 81 of the wear end 80.

[0096] It is understood that the electronic circuit 10 is advantageously calibrated to know in advance an approximate length between the transmitter-receiver 7 and the wear end 80 of the abradable material 8.

[0097] The information enabling the characterization of the wear of the gas turbine engine 2 can be the wear thickness of the device 1.

[0098] The expression "wear thickness" can refer to several quantities. The wear thickness can be the worn thickness, i.e. the thickness that has disappeared, of the device 1.

[0099] The wear thickness can be the worn thickness of the abradable material 8. The worn thickness of the device 1 corresponds in practice to the worn thickness of the abradable material 8.

[0100] The wear thickness can be the remaining thickness of the device 1. In other words, the wear thickness can be the instantaneous thickness of the abradable material 8 or the instantaneous thickness of the assembly formed by the body 9 and the abradable material 8. In other words, the wear thickness can be the instantaneous distance between the ultrasonic transceiver 7 and the internal surface 81 of the wear end 80, referred to as the instantaneous distance which will be detailed later.

[0101] By "instantaneous," it is understood to mean "actual." In other words, an instantaneous value is a point value measured at a precise instant. The instantaneous thickness and the instantaneous distance correspond respectively to an effective thickness and distance at the instant when an ultrasonic wave transmission and reception are implemented by the ultrasonic transceiver 7.

[0102] Thus, according to one embodiment, the electronic circuit 10 can itself be configured to calculate the instantaneous distance between the ultrasonic transceiver 7 and the wear end 80, called the instantaneous distance.

[0103] As illustrated for example in [Fig.3], the instantaneous distance DA advantageously corresponds to the distance between, on the one hand, the interface between the transmitter-receiver 7 and the abradable material 8 or the body 9 (depending on whether the device 1 includes a body 9), and, on the other hand, the interface between the abradable material 8 and the outside.

[0104] In other words, the instantaneous distance DA advantageously corresponds to the distance between, on the one hand, the interface between the transmitter-receiver 7 and the material abradable 8 or the body 9 (depending on whether the device 1 includes a body 9), and, on the other hand, a point of the wear end 80.

[0105] Advantageously, the instantaneous distance DA corresponds to an average distance between, on the one hand, the interface between the transceiver 7 and the abradable material 8 or the body 9 (depending on whether the device 1 includes a body 9), and, on the other hand, a point on the wear end 80. Indeed, as explained previously, at the same instant, the ultrasonic transceiver 7 emits several incident waves which arrive at different points on the internal surface 81 of the wear end 80. Several reflected waves result from this, and an averaging of these reflected waves is implemented by the transceiver 7 or by the electronic circuit 10. Consequently, if the abradable material 8 is worn non-uniformly along its wear end, an average instantaneous distance DA will be obtained.

[0106] The instantaneous distance therefore corresponds to the thickness of the abradable material 8, if the device 1 does not include a body 9. The instantaneous distance corresponds to the thickness of the assembly formed by the abradable material 8 and the body 9, if the device 1 includes a body 9.

[0107] The instantaneous distance can be calculated simply from the speed of the ultrasonic waves and the time elapsed between the emission of an ultrasonic wave in the abradable material 8 (and in the body 9, if applicable) and the reception of the corresponding ultrasonic wave reflected by the inner surface 81 of the wear end 80. The following formula can be applied: time * speed. The speed depends on the material traversed. The instantaneous distance can also be calculated by taking into account the amplitude of the emitted signal and the amplitude of the reflected signal. Indeed, the ultrasonic signal loses power after reflection, i.e., its amplitude decreases. This makes it possible, in particular, to identify the wave reflected by the inner surface 81 of the wear end 80.

[0108] The propagation speed of ultrasonic waves depends on the material in which the wave propagates. The electronic circuit 10 is therefore preferably configured to have access to the propagation speed of ultrasonic waves in the abradable material 8 and in the body 9, if applicable.

[0109] According to one embodiment, the information enabling the characterization of the wear of the gas turbine engine 2 is the instantaneous thickness of the abradable material 8.

[0110] The instantaneous thickness of the abradable material 8 can be calculated from the instantaneous distance DA, from which the thickness of the body 9 is subtracted, if applicable. The thickness of the body 9 is in this case accessible to the electronic circuit 10.

[0111] According to one embodiment, the information enabling the characterization of the wear of the gas turbine engine 2 is the worn thickness of the abradable material 8 (which is also in practice the worn thickness of the device 1).

[0112] The worn thickness of the abradable material 8 corresponds to the thickness of the abradable material 8 which has disappeared by wear, i.e. which has eroded.

[0113] In this case, the electronic circuit 10 is advantageously configured to have access to the initial distance between the ultrasonic transceiver 7 and the wear end 80, referred to as the initial distance. As illustrated in [Fig. 2], the initial distance DI is the distance between the ultrasonic transceiver 7 and the wear end 80 before the abradable material 8 is worn.

[0114] To calculate the worn thickness, the electronic circuit 10 is advantageously configured to apply the following formula: initial distance - instantaneous distance.

[0115] In this case, it is understood that the electronic circuit 10 is advantageously configured to determine the time elapsed between the emission of an ultrasonic wave in the abradable material 8 and the reception of the corresponding ultrasonic wave reflected by the internal surface 81 of the wear end 80, to deduce the instantaneous distance and thus to deduce the worn thickness of the abradable material 8.

[0116] The device 1 preferably includes a communication module 11 connected to the electronic circuit 10 and configured to transmit to an electronic data collection system 3 the information enabling the characterization of the wear of the gas turbine engine 2.

[0117] The communication module 11 can transmit information in a wired manner, for example with a wired CAN (Analog-to-Digital Converter) link or wirelessly, for example according to an RFID (Radio Frequency Identification) process.

[0118] The communication module 11 is configured to allow the transmission of information enabling the characterization of the wear of the gas turbine engine 2 so that it can be analyzed in order to determine a level of wear of the engine 2 and thus to determine the need or not for maintenance of the engine 2.

[0119] As illustrated in Figures 1 and 1bis, the device 1 is advantageously arranged on fixed parts of the gas turbine engine 2, for example on the casing 17 of the gas turbine engine 2.

[0120] Here, it is understood that the device 1 is arranged on the motor 2 at least in its operating configuration, i.e., when a wear measurement is to be performed. Certain elements such as the transceiver 7, the electronic circuit 10, and / or the communication module 11 can be deactivated or removed (and thus physically detached from the body 9 and the abradable material 8) when no wear measurement is being performed. Thus, the transceiver 7, the electronic circuit 10, and / or the communication module 11 can be attached to the body 9 only when a wear measurement is required, so that these elements are not permanently attached to the body 9, and therefore not permanently installed on the motor 2. even when the motor 2 is running. On the other hand, it is understood that the abradable material 8 and, where applicable, the body 9 remain in the same position on the motor 2 even if the transceiver 7, the electronic circuit 10 and / or the communication module 11 are removed.

[0121] Advantageously, the device 1 is arranged on the gas turbine engine 2 so that the wear end 80 of the abradable material 8 is in contact with a flow of gaseous fluid F circulating from upstream to downstream of the gas turbine engine 2 when the gas turbine engine 2 is in operation.

[0122] Preferably, the device 1 is arranged so that the transceiver 7, the electronic circuit 10, and the communication module 11 are not in contact with a gaseous fluid flow F when the gas turbine engine 2 is running. This protects these components. For example, the transceiver 7, the electronic circuit 10, and the communication module 11 can be arranged outside the gas turbine engine 2.

[0123] Thus, preferably, the device 1 is fixed to a housing so that a part of the device 1 is outside the engine and another part is inside the engine.

[0124] According to a preferred embodiment, if the transceiver 7, the electronic circuit 10 and / or the communication module 11 are permanently fixed to the body 9, they are protected by another housing and are therefore not in contact with the outside. This prevents wear on these components.

[0125] According to a preferred embodiment, the device 1 is arranged in an endoscopic port 21 of the gas turbine engine 2. Indeed, endoscopic ports are generally provided in the gas turbine engine 2 to allow inspection of the inside of the gas turbine engine 2 by means of an endoscope inserted into the engine 2 from the outside of the engine 2. These ports are generally, when not in use for inspection, plugged with endoscopic plugs. The device 1 can therefore replace an endoscopic plug. Thus, it is not necessary to modify a gas turbine engine 2 to allow its inspection by means of the device 1.

[0126] Alternatively, the device 1 can be arranged in a hole drilled in a casing of the gas turbine engine 2. The drilled hole gives access to the air stream, i.e. to the gas flow.

[0127] The device 1 can be arranged in different locations of the gas turbine engine 2, depending on the area of ​​the engine that one wishes to monitor.

[0128] For example, as illustrated in Figures 1 and 1bis, the device 1 can be arranged at the stator of the low-pressure compressor or at the rotor of the low-pressure compressor.

[0129] The device 1 can also be arranged at the leading edge of a compressor rectifier or at the leading edge of a compressor diffuser. In this case, the device 1 is better exposed to the particles of the flow of gas (angle of attack at 90°) without generating any additional aerodynamic wake. Therefore, device 1 has no impact on the aerodynamic flow and thus no impact on the performance of the gas turbine engine 2.

[0130] Device 1 can be arranged at an air inlet, for example air inlet 37a shown in [Fig. Ibis]. This position is advantageous because it is easy to arrange device 1 there.

[0131] Figures 1 and 1bis schematically illustrate different possible positions of device 1 in the gas turbine engine 2.

[0132] The gas turbine engine 2 may include several devices 1. System

[0133] With reference to [Fig.6], a system 100 is proposed for characterizing wear of a gas turbine engine 2.

[0134] The system 100 comprises at least one device 1 as previously described.

[0135] The system 100 advantageously includes an electronic data collection system 3 adapted to be connected to the device 1. The electronic data collection system 3 can be connected to the device 1, more specifically to the communication module 11, by wired or wireless means.

[0136] For example, the system 100 may include a wireless reading system 4 adapted to be connected to the electronic data collection system 3 and configured to receive data from the communication module 11 using an RFID method. The wireless reading system 4 advantageously comprises an antenna and a reader connected to each other. In another embodiment, the electronic data collection system 3 includes the wireless reading system 4, i.e., the functions of the electronic data collection system 3 and the wireless reading system 4 are implemented by the same physical system.

[0137] The electronic data collection system 3 can be an electronic data collection box (Data Collector Box, DCB) or simply a memory, in general.

[0138] The system 100 includes a data processing module 12 adapted to be connected to the communication module 11 of the device 1, optionally via the electronic data collection system 3. The data processing module 12 can be integrated into the electronic data collection system 3.

[0139] The data processing module 12 is configured to determine a wear level of the engine 2 from the information enabling the characterization of the wear of the gas turbine engine 2.

[0140] The data processing module 12 is therefore configured to analyze the information used to characterize the wear of the gas turbine engine 2 and to deduce the wear level of the gas turbine engine 2. For this purpose, the module Data processing 12 is advantageously configured to determine, from the information, a representative data of the wear level allowing the wear level to be determined from a nomogram as will be described later.

[0141] If the information enabling the characterization of the wear of the gas turbine engine 2 is the instantaneous thickness of the abradable material 8, the data processing module 12 can be configured to calculate the worn thickness of the abradable material 8 as a function of its instantaneous thickness and its initial thickness.

[0142] If the information enabling the characterization of the wear of the gas turbine engine 2 is the instantaneous distance between the ultrasonic transceiver 7 and the wear end 80 (said instantaneous distance being able to correspond in one embodiment to the instantaneous thickness of the abradable material 8), the data processing module 12 can be configured to calculate the worn thickness of the abradable material 8 from the instantaneous distance and the initial distance between the ultrasonic transceiver 7 and the wear end 80, the initial distance being accessible to the data processing module 12. As explained previously, the worn thickness can be calculated according to the following formula: initial distance - instantaneous distance.

[0143] Alternatively or in addition, the data processing module 12 can be configured to determine the instantaneous thickness of the abradable material 8 from the instantaneous distance and the thickness of the body 9 to which the data processing module 12 can have access.

[0144] If the information is the time elapsed between the emission of an ultrasonic wave in the abradable material 8 and the reception of the corresponding ultrasonic wave reflected by the internal surface 81 of the wear end 80, the data processing module 12 can be configured to calculate, from this time (and the speed of the ultrasonic waves, which would be known to the data processing module 12), the instantaneous distance. Then, from the instantaneous distance (and the initial distance, which would be known to the data processing module 12), the data processing module 12 can be configured to calculate the worn thickness and / or the instantaneous thickness of the abradable material 8.

[0145] Thus, the data representing the wear level can be the information itself or be calculated from the information. The data representing the wear level is advantageously the wear thickness of the device 1, which can be, for example, the worn thickness of the abradable material 8, the instantaneous thickness of the abradable material 8, or the instantaneous distance between the ultrasonic transceiver 7 and the internal surface 81 of the wear end 80.

[0146] Advantageously, the data processing module 12 is configured to determine a wear level of the gas turbine engine 2 from the data representative of the level of wear and therefore, directly or indirectly, from the information allowing to characterize the wear of the gas turbine engine 2.

[0147] The wear level of the gas turbine engine 2 can, for example, be expressed as a percentage, with 100% corresponding to an unworn engine so that the engine reaches 100% of its maximum performance and 0% corresponding to a worn engine so that the engine is out of service or can no longer be operated over its entire range of use with the same performance guarantees.

[0148] The wear level of the gas turbine engine 2 can also be expressed as the remaining service life of the gas turbine engine 2. For example, the wear level could be the number of days remaining before maintenance. Maintenance could be, for example, a simple inspection of the engine 2 or the replacement of a component of the engine 2, possibly following an inspection.

[0149] The level of wear can be expressed according to a class belonging to a group of classes such as: not worn, slightly worn, worn, very worn. The worn and very worn classes could correspond to wear requiring engine maintenance.

[0150] The wear level is used to indicate to an operator whether or not engine maintenance is required. Alternatively, or equally, the operator may be informed of the remaining time before maintenance and possibly the nature of the maintenance.

[0151] Advantageously, the data processing module 12 is configured to determine the wear level of the engine from a predetermined nomogram indicating a correspondence between the data representing the wear level obtained from the information enabling the characterization of the wear of the gas turbine engine 2 and a wear level of a gas turbine engine.

[0152] Advantageously, the nomogram indicates a correspondence between, on the one hand, the wear thickness of the device 1, i.e. the worn thickness of the abradable material 8, the instantaneous thickness of the abradable material 8 or the instantaneous distance between the ultrasonic transmitter-receiver 7 and the internal surface 81 of the wear end 80, and, on the other hand, a wear level of the gas turbine engine 2. Advantageously, the nomogram indicates a correspondence between, on the one hand, the wear thickness of the device 1, i.e. the worn thickness of the abradable material 8, and, on the other hand, a wear level of the gas turbine engine 2.

[0153] The nomogram considered has been advantageously constructed on the basis of a gas turbine engine 2 having similar properties, or even being identical, to the gas turbine engine 2 that one seeks to monitor.

[0154] The nomogram considered was advantageously constructed using a gas turbine engine 2 in an environment similar to that in which the gas turbine engine 2 that one seeks to monitor is used. The nomogram is therefore advantageously associated with a type of environment. An environment can be characterized by different Parameters such as soil type and the type of volatile particles present are taken into account. The environment can, for example, be categorized as "low erosion," "erosive," or "high erosion." Therefore, there may be several nomograms for the same engine, with a different nomogram used depending on the environment. The environment can be determined, for example, based on the aircraft's GPS position.

[0155] The nomogram(s) can be stored by the electronic data collection system 3 or by the data processing module 12. The nomogram(s) can therefore be carried on board the aircraft or not, depending on whether the electronic data collection system 3 or the module 12 is carried on board or not.

[0156] The data processing module 12 is therefore preferably configured to determine the wear level of the gas turbine engine 2 corresponding to the representative wear level data from the nomogram.

[0157] The system 100 can be configured to determine the wear level automatically at regular intervals. Alternatively or in addition, the wear level determination can be implemented following an operator command.

[0158] For example, the system 100 may include a maintenance console 13 adapted to be connected to the data processing module 12 and adapted to be connected to the electronic data collection system 3 which would allow the operator to control and monitor the wear level.

[0159] The operator can be the aircraft pilot himself. Indeed, the entire system 100 can be carried on board the aircraft.

[0160] Alternatively, part of the system 100, for example the data processing module 12 and the maintenance console 13, might not be carried on board the aircraft and might only be present at maintenance sites. The electronic data collection system 3, the wireless reading system 4, and even the communication module 11, the electronic circuit 10, and / or the transceiver 7 might not be permanently carried on board the aircraft. Method

[0161] With reference to [Fig.7], a method is further proposed for characterizing wear of a gas turbine engine 2 by means of a system 100 as previously presented.

[0162] The method can be implemented when the gas turbine engine 2 is running or stopped.

[0163] The process includes a step a) of emitting, by the ultrasonic transceiver 7, an ultrasonic wave into the abradable material 8, advantageously towards the wear end 80 of the abradable material 8.

[0164] The ultrasonic wave travels through the abradable material 8 to the wear end 80 and in particular to the internal surface 81 of the wear end 80. If the device 1 of the system 100 includes a body 9, the ultrasonic wave travels through the body 9 and then through the abradable material 8.

[0165] The method includes a step b) of receiving, by the ultrasonic transceiver 7, a corresponding ultrasonic wave reflected by the internal surface 81 of the wear end 80.

[0166] The method includes a step c) of calculating, by the electronic circuit 10 of the device, information enabling the wear of the gas turbine engine 2 to be characterized.

[0167] As explained previously, this information can be, for example, the worn thickness of the abradable material 8, the instantaneous distance between the ultrasonic transmitter-receiver 7 and the wear end 80, the instantaneous thickness of the abradable material 8 or the time elapsed between the emission of an ultrasonic wave in the abradable material 8 and the reception of the corresponding ultrasonic wave reflected by the internal surface 81 of the wear end 80.

[0168] The method advantageously includes a step d) of calculating a data representative of the wear level by the data processing module 12 from the information.

[0169] Preferably, the method includes a step e) of determining a wear level of the gas turbine engine 2 based on the representative wear level data and the predetermined nomogram indicating a correspondence between the representative wear level data and a wear level of a gas turbine engine. Preferably, the determination step e) is implemented by the data processing module 12.

[0170] Thus, it is possible to monitor in real time and above all in a very simple and efficient way the state of wear of a gas turbine engine 2.

Claims

Demands

1. Device (1) for characterizing wear on a gas turbine engine (2) comprising: - an abradable material (8) including a wear end (80) adapted to be arranged in contact with a flow of gaseous fluid (F) flowing from upstream to downstream of the gas turbine engine (2) when the gas turbine engine (2) is in operation; - an ultrasonic transceiver (7) adapted to emit an ultrasonic wave in the abradable material (8) and receive a corresponding ultrasonic wave reflected by an internal surface (81) of the wear end (80); - an electronic circuit (10) connected to the ultrasonic transceiver (7) and configured to calculate information enabling the characterization of the wear on the gas turbine engine (2).

2. Device according to claim 1, comprising a body (9) adapted to carry the abradable material (8) and the ultrasonic transceiver (7), the body (9) being arranged between the abradable material (8) and the ultrasonic transceiver (7).

3. Device according to any one of claims 1 and 2, comprising a communication module (11) connected to the electronic circuit (10) and configured to transmit information enabling the characterization of the wear of the gas turbine engine (2) to a data processing module (12) configured to characterize the wear of the gas turbine engine (2) from the information enabling the characterization of the wear of the gas turbine engine (2).

4. Gas turbine engine (2) comprising at least one device (1) according to any one of claims 1 to 3, the device (1) being arranged so that the wear end (80) of the abradable material (8) is in contact with a flow of gaseous fluid (F) flowing from upstream to downstream of the gas turbine engine (2) when the gas turbine engine (2) is in operation.

5. Motor according to claim 4, wherein the device (1) is arranged in a hole drilled in the gas turbine motor (2), in particular the device (1) is arranged in an endoscopic port (21) of the gas turbine motor (2).

6. Aircraft comprising a gas turbine engine (2) according to any one of claims 4 and 5.

7. System (100) for characterizing wear on a gas turbine engine (2) comprising: - a device (1) according to any one of claims 1 to 3; - a data processing module (12) adapted to be connected to the electronic circuit (10), the data processing module (12) being configured to characterize the wear of the gas turbine engine (2) from the information enabling the characterization of the wear of the gas turbine engine (2).

8. A method for characterizing wear on a gas turbine engine (2) by means of a system (100) according to claim 7, the method comprising the steps of: - emission (a), by the ultrasonic transceiver (7), of an ultrasonic wave in the abradable material (8), - reception (b), by the ultrasonic transceiver (7), of a corresponding ultrasonic wave reflected by the internal surface (81) of the wear end (80), - calculation (c), by the electronic circuit (10), of the information enabling the characterization of the wear of the gas turbine engine (2).

9. A method according to claim 8, comprising a step of: - determination (d) by the data processing module (12) of a data point representative of the wear level of the gas turbine engine (2) from the information used to characterize the wear of the gas turbine engine (2), the data point representative of the wear level being a wear thickness of the device (1) calculated by the electronic circuit (10) or the data processing module (12) from: • the speed of ultrasonic waves, and • a time elapsed between the emission of the wave ultrasonic in the abradable material (8) and the reception of the corresponding ultrasonic wave reflected by the internal surface (81) of the wear end (80).

10. A method according to any one of claims 8 and 9, comprising a step of determining (e) a wear level of the gas turbine engine (2), by the data processing module (12), as a function of a data representative of the wear level determined from the information enabling the characterization of the wear of the gas turbine engine (2) and a predetermined nomogram indicating a correspondence between the data representative of the wear level and a wear level of a gas turbine engine.

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