Patent ID: 12258905

FIG.1illustrates a gas turbine engine10having a principal rotational axis9. The engine10comprises an air intake12and a propulsive fan23that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine10comprises a core11that receives the core airflow A. The engine core11comprises, in axial flow series, a low pressure compressor14, a high-pressure compressor15, combustion equipment16, a high-pressure turbine17, a low pressure turbine19and a core exhaust nozzle20. A nacelle21surrounds the gas turbine engine10and defines a bypass duct22and a bypass exhaust nozzle18. The bypass airflow B flows through the bypass duct22. The fan23is attached to and driven by the low pressure turbine19via a shaft26and an epicyclic gearbox30.

In use, the core airflow A is accelerated and compressed by the low pressure compressor14and directed into the high pressure compressor15where further compression takes place. The compressed air exhausted from the high pressure compressor15is directed into the combustion equipment16where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines17,19before being exhausted through the nozzle20to provide some propulsive thrust. The high pressure turbine17drives the high pressure compressor15by a suitable interconnecting shaft27. The fan23generally provides the majority of the propulsive thrust. The epicyclic gearbox30is a reduction gearbox.

Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e., not including the fan23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft26with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan23). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan23may be referred to as a first, or lowest pressure, compression stage.

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown inFIG.1has a split flow nozzle20,22meaning that the flow through the bypass duct22has its own nozzle that is separate to and radially outside the core engine nozzle20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct22and the flow through the core11are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine10may not comprise a gearbox30.

The geometry of the gas turbine engine10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis9), a radial direction (in the bottom-to-top direction inFIG.1), and a circumferential direction (perpendicular to the page in theFIG.1view). The axial, radial and circumferential directions are mutually perpendicular.

In the context of the present invention, a method and system of monitoring one or several components of the gas turbine engine are of relevance. The method and system are described in a general manner inFIGS.2to5.FIGS.6to14regard embodiments of monitoring and detecting mechanical failures of different components of a gas turbine engine.

FIG.2depicts a system in which a radar sensing element4is located in a chamber7of a gas turbine engine. The chamber7may be, e.g., a nacelle, a bypass duct or a core engine compartment of the gas turbine engine. Within chamber7is located as first component61a duct which passes through the chamber7. Within chamber7, there is further located a second component62.

The radar sensing element4emits radio waves41which are transmitted to the components61,62and reflected by the components61,62. Radar sensing element4may be a sensor of small size such as 8×10 mm size and with a power consumption in the range between 300 mW and 1000 mW. Components61,62located within the beam of the radar sensing element4reflect some portion of the beam energy back to the radar sensing element4, wherein an antenna (not shown) of the radar sensing element4detects the reflected radio waves. Of course, the depiction of two components61,62inFIG.2is to be understood as an example only. There may be located only one component in compartment7or more than two components in compartment7. Also, the monitored component may be a wall or a part of a wall of the chamber7.

Radar sensing element4is connected to a control and evaluation unit5which controls the radar sensing element4and receives information/data from the radar sensing element4. In particular, at least one property value of the reflected waves which are detected by the radar sensing element4is determined and provided to the control and evaluation unit5. Determination of the property value may be performed in the radar sensing element4or in the control and evaluation unit5. The property value that is determined from the detected reflected waves is, e.g., a time delay value, an energy value or one or several values of a frequency spectrum.

The control and evaluation unit5may be an Electronic Engine Control (EEC) unit of the gas turbine engine or a functional part of such EEC. The EEC is a digital control unit that combines engine sensor information with cockpit instructions to ensure that the engine performs both safely and at an optimal level. However, in principle, the control and evaluation unit5may be a unit separate from the EEC and interacting with the EEC. As all other components inFIG.1, the control and evaluation unit5is depicted only schematically.

More particularly, the control and evaluation unit5comprises a central processing unit51which receives data from the radar sensing element4. The control and evaluation unit5further comprises a power source54, a mass storage memory52in communication with the central processing unit51and in interface53for sending data, e.g., to an aircraft on-board communication unit. It is pointed out that only the components of the control and evaluation unit5relevant for the present invention are depicted inFIG.2.

Program instructions are stored in memory52which cause, when executed by central processing unit51, the performance of method steps as discussed with respect toFIGS.2to4.

FIG.3depicts a method of detecting a mechanical failure of a component which, in principle, may be any component in a gas turbine engine. In step301, a radar sensing element is provided that is configured to transmit and detect radio waves, such as radar sensing element4ofFIG.2. In step302, a state devoid of mechanical failure of the component is determined. Such determination is made by determining at least one property value of radio waves that have been transmitted from the radar sensing element to the component and have been reflected from the component, the component being in the state devoid of mechanical failure. Such a property value may be, e.g., a time delay value, an energy value or a value of a frequency spectrum. Such value is indicative of a characteristic of the component such as size, shape, orientation, material, distance, and velocity of the component in the state devoid of mechanical failure.

In step303, subsequently, a current state of the component is determined, wherein the at least one property value of the reflected radio waves is determined in the current state. Again, such at least one property value is determined using radar technology and the radar sensing element.

In step304, it is determined if the at least one property value has changed in a manner indicative of a mechanical failure. In such case, according to step305, a mechanical failure is reported, e.g., by sending a warning signal through interface53to an aircraft on-board communication unit.

FIGS.4and5discuss two different embodiments of how the determination of step304ofFIG.3that a mechanical failure is present is made.

According to the method ofFIG.4, in step401, at least one property value of the reflected waves in the current state is determined. In step402, this property value of the current state is compared with a stored property value in the original state, i.e., the state devoid of mechanical failure. Such original state property value may be stored in memory52ofFIG.2. In step403, a comparison is made if the property value between the original state and the current state differs at least by a specified amount. Accordingly, it is determined if the difference between the two values is larger than a predefined threshold. In such case, according to step404, a mechanical failure is identified. If the difference is below the threshold, a mechanical failure is not identified and the method continues in step401with determination and evaluation of further current property values.

In the method according toFIG.5, artificial intelligence is implemented to determine if a mechanical failure is present. In step501, and artificial intelligence engine is trained with property values of the state devoid of mechanical failure of the component. For example, a plurality of properties of the waves reflected from the component in the state devoid of mechanical failure are determined and stored, e.g., in memory52of control and evaluation unit5ofFIG.2.

Further, in step502, the artificial intelligence engine is trained with property values of states of mechanical failure of the component. For example, several possible failure scenarios are implemented such as a burst duct or a burned through surface having a hole. For these failure scenarios, a plurality of properties of the waves reflected from the component are determined and also stored, e.g., in memory52of control and evaluation unit5ofFIG.2.

In one embodiment, only step501or step502is implemented. However, to increase the artificial intelligence of the artificial intelligence engine and its ability to discriminate between the state devoid of mechanical failure and states of mechanical failure, it is preferable to train the artificial intelligence engine both on the state devoid of mechanical failure and states of mechanical failure, thus implementing both steps501and502.

In step503, at least one property value of the reflected waves for the current state of the component is determined. For example, an actual time delay value is determined. This value is fed in step504in the artificial intelligence engine. The artificial intelligence engine determines in step505if the change is indicative of a mechanical failure. Such determination is a direct result of the artificial intelligence engine. If so, in step506, a mechanical failure is reported. If not, the method continues with step503.

The artificial intelligence engine may be implemented in central processing unit51or may be implemented as a separate component of the control and evaluation unit5.

There exist multiple failure scenarios of why damage can be created to a gas turbine engine, such failure scenarios including a burst duct of a pressurized air pipe, a combustor burn through, a failed open bleed valve, a liquid (oil, flued, hydraulic fluid) pipe leak and a cooling system airflow failure.

FIGS.6to10show embodiments in which the radar monitoring technology of the present invention allows the early detection of a nacelle damage due to hot air impingement. InFIGS.6and7, a radar sensing element4is located in a rear part of a nacelle21. An EEC5which serves as a control and evaluation unit in accordance with the present invention is also located in nacelle21. Air ducts71provide bleed air from two compressor stages, wherein the bleed air is provided in a duct63to an environmental control system of an aircraft. The radar sensing element4is able to detect changes in the geometry in case of a leak630in duct63, wherein the leak630itself and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element4. Such changes in geometry may be a delamination of parts and/or cracks.

In this respect, it is pointed out that, in an embodiment, the radar sensing element4may be configured to steer the beam of radio waves into different directions such that a high intensity beam can be directed to different areas and to different components.

InFIGS.8and9, compared toFIGS.6and7, additionally a radar sensing element4′ is located in a front part of nacelle21. Air ducts71provide bleed air from two compressor stages, wherein the bleed air is provided in a duct64towards the inlet lip210of the nacelle21in order to prevent the formation of ice at the inlet lip210. The radar sensing element4′ is able to detect changes in the geometry in case of a leak640in duct64, wherein the leak640itself and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element4′. Such changes in geometry may be a delamination of parts and/or cracks.

InFIGS.10and11, a radar sensing element4is located in a compartment of the core engine11. An EEC5which serves as a control and evaluation unit in accordance with the present invention is located in nacelle21. An air duct65of a secondary air system provides cooling bleed air from a compressor stage to a turbine stage of the core engine11. The radar sensing element4is able to detect changes in the geometry in case of a leak650in duct65, wherein the leak650itself and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element4. Such changes in geometry may be a delamination of parts and/or cracks.

InFIGS.12and13, similar as inFIGS.10and11, a radar sensing element4is located in a compartment of the core engine11. An EEC5which serves as a control and evaluation unit in accordance with the present invention is located in nacelle21. A combustor16comprises an outer case66. The radar sensing element4is able to detect changes in the geometry of the combustor case66in case of a leak660, wherein the leak660itself and/or other changes in the geometry of the combustor case66and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element4. Such changes in geometry may be a delamination of parts and/or cracks.

InFIG.14, a radar sensing element4is located in a rear part of a nacelle21. An EEC5which serves as a control and evaluation unit in accordance with the present invention is also located in nacelle21. A compressor bleed valve67is provided at compressor15. If the bleed valve67has a failure in that it opens at high power for unlimited time, such opening will result in damage in the outer wall68of bypass duct22, which will result in damage in the outer wall62, including perforation and leaking of bypass air into the fan compartment. The changes in geometry associated there with can be detected by the radar sensing element4. In this case, the radar sensing element4emits radio wave towards the wall68which represents the monitored component. In addition, a radar sensing element may be provided in proximity to bleed valve67to also monitor bleed valve67.

It should be understood that the above description is intended for illustrative purposes only and is not intended to limit the scope of the present disclosure in any way. For example, it is pointed out that the present invention is not limited in its application to a propulsion system but may be implemented at the whole aircraft level. Other embodiments regard, among others, an anti-ice system in an aircraft wing or the cargo bay of the aircraft.

Also, those skilled in the art will appreciate that other aspects of the disclosure can be obtained from a study of the drawings, the disclosure, and the appended claims. All methods described herein can be performed in any suitable order unless otherwise indicated herein or otherwise clearly contradicted by context. Various features of the various embodiments disclosed herein can be combined in different combinations to create new embodiments within the scope of the present disclosure. In particular, the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein. Any ranges given herein include any and all specific values within the range and any and all sub-ranges within the given range.