METHOD AND SYSTEM OF MONITORING A COMPONENT OF AN AIRCRAFT

Abstract

A gas turbine engine for an aircraft that includes a nacelle, a fan, an engine core, a bypass duct extending between the engine core and the nacelle and guiding a bypass airflow through the bypass duct, and at least one non-structural strut extending in the radial direction within the bypass duct, wherein the non-structural strut includes an outside wall acting as a heat exchanger, and wherein the outside wall includes first transport means configured to transport in the outside wall at least one fluid to be cooled. It is provided that the non-structural strut further includes second transport means configured to transport a fluid to be heated, wherein the first transport means and the second transport means are configured such that the fluid to be heated is heated by the at least one fluid to be cooled and the at least one fluid to be cooled is cooled both by the bypass airflow and the fluid to be heated.

Claims

1. A gas turbine engine for an aircraft comprising: a nacelle; a fan; an engine core located downstream of the fan, the engine core comprising a primary duct guiding a core airflow through the engine core; a bypass duct located downstream of the fan, the bypass duct extending between the engine core and the nacelle and guiding a bypass airflow through the bypass duct; and at least one non-structural strut extending in the radial direction within the bypass duct; wherein the non-structural strut comprises an outside wall acting as a heat exchanger; wherein the outside wall includes first transport means configured to transport in the outside wall at least one fluid to be cooled, wherein the non-structural strut further comprises second transport means configured to transport a fluid to be heated, wherein the first transport means and the second transport means are configured such that the fluid to be heated is heated by the at least one fluid to be cooled and the at least one fluid to be cooled is cooled both by the bypass airflow and the fluid to be heated.

2. The gas turbine engine of claim 1, wherein as first transport means first transport channels configured to transport the at least one fluid to be cooled are integrated into the outside wall of the non-structural strut.

3. The gas turbine engine of claim 1, wherein the outside wall of the non-structural strut consists of aluminum or a metallic light alloy or a thermally conductive composite or polymer.

4. The gas turbine engine of claim 3, wherein the outside wall has a thickness in the range between 4 mm and 10 mm, in particular in the range between 6 mm and 8 mm.

5. The gas turbine engine of claim 2, wherein at least some the first transport channels are part of an oil system and configured to transport oil to be cooled.

6. The gas turbine engine of claim 1, wherein the second transport means configured to transport the fluid to be heated are also integrated as second transport channels into the outside wall.

7. The gas turbine engine of claim 6, wherein the second transport channels configured to transport the fluid to be heated are arranged in the outside wall to the inside of the first transport channels configured to transport the fluid to be cooled.

8. The gas turbine engine of claim 6, wherein the second transport channels configured to transport the fluid to be heated are arranged in the outside wall perpendicular to the first transport channels configured to transport the fluid to be cooled.

9. The gas turbine engine of claim 1, wherein the non-structural strut comprises an inside compartment surrounded by the outside wall, wherein the inside compartment comprises at least one of electrical lines, fluid ducts and drain pipes.

10. The gas turbine engine of claim 1, wherein the outside wall comprises two opposite walls spaced apart in the circumferential direction and extending in the axial direction, wherein at least the first transport channels extend in both walls.

11. The gas turbine engine of claim 10, wherein the outside wall forms a leading edge, wherein at least the first transport channels extend to the leading edge of the outside wall.

12. The gas turbine engine of claim 10, wherein the non-structural strut is configured as a cross flow heat exchanger, wherein at least the first transport channels extend radially in each of the opposite walls, wherein the fluid to be cooled is moved radially inward in the transport channels of one of the opposite walls and subsequently moved radially outward in the transport channels of the other opposite wall, or vice versa.

13. The gas turbine engine of claim 10, wherein the non-structural strut is configured as a partial counter flow and partial parallel flow heat exchanger, wherein at least the first transport channels extend axially in each of the opposite walls, wherein the fluid to be cooled is moved axially forward in the transport channels of one of the opposite walls and moved axially rearward in the transport channels of the other opposite wall.

14. The gas turbine engine of claim 10, wherein the non-structural strut is configured as a full counter flow heat exchanger, wherein at least the first transport channels extend axially in each of the opposite walls, and wherein the fluid to be cooled is moved axially forward in the transport channels of both opposite walls.

15. The gas turbine engine of claim 1, wherein at least some of the transport channels integrated into the outside wall each additionally comprise cooling fins extending towards the inside of the transport channel.

16. The gas turbine engine of claim 1, wherein a fluid inlet to the non-structural strut and a fluid outlet to the non-structural strut are both located at the radial outer side or the radial inner side of the non-structural strut.

17. The gas turbine engine of claim 1, wherein the at least one non-structural strut further comprises third transport means in the outside wall of the non-structural strut which are configured to transport a further fluid to be cooled.

18. A heat management system for cooling oil in an oil system of a gas turbine engine having an engine core and a bypass duct, the heat management system comprising: a lube oil circuit of an oil system of the gas turbine engine, a fuel line providing fuel to a combustor of gas turbine engine, a non-structural strut extending in the radial direction within the bypass duct, the non-structural strut comprising an outside wall into which first transport channels and second transport channels are integrated, the outside wall forming a heat exchanger, wherein the first transport channels are configured to transport lube oil of the lube oil circuit; wherein the second transport channels are configured to transport fuel of the fuel line, wherein the first and second transport channels are configured such that the fuel is heated by the lube oil in the first transport channels and that the lube oil is cooled both by the bypass airflow and the fuel in the second transport channels.

19. The heat management system of claim 18, wherein the input to and the output from the first transport channels are both arranged at the radial inner side of the non-structural strut.

20. The heat management system of claim 18, wherein a further fluid circuit of an electrical generator or of the gearbox of a geared turbofan engine, wherein third transport channels are integrated into the outside wall and wherein the third transport channels are configured to transport the fluid of the further fluid circuit, wherein the fluid of the further fluid circuit is cooled by the bypass airflow and/or the fluid to be heated.

Description

[0048] The invention will be explained in more detail on the basis of exemplary embodiments with reference to the accompanying drawings in which:

[0049] FIG. 1 is a simplified schematic sectional view of a gas turbine engine in which the present invention can be realized;

[0050] FIG. 2 is a schematic depiction of a system of monitoring a component of a gas turbine engine, the system comprising a radar sensing element and a control and evaluation unit for determining a mechanical failure of the monitored component;

[0051] FIG. 3 is a flowchart of a method to detect a mechanical failure of a component of a gas turbine engine;

[0052] FIG. 4 is a flowchart of one method of determining if a property value is indicative of a mechanical failure in the flowchart of FIG. 3;

[0053] FIG. 5 is a flowchart of another method of determining if a property value is indicative of a mechanical failure in the flowchart of FIG. 3;

[0054] FIGS. 6, 7 is a gas turbine engine with a radar sensing element and an associated control and evaluation unit, wherein an air duct in a nacelle is monitored and a burst condition is detected;

[0055] FIGS. 8, 9 is a gas turbine engine with a radar sensing element and an associated control and evaluation unit, wherein an anti-ice air duct in a nacelle is monitored and a burst condition is detected;

[0056] FIGS. 10, 11 is a gas turbine engine with a radar sensing element and an associated control and evaluation unit, wherein an air duct for providing cooling air to a turbine is monitored and a burst condition is detected;

[0057] FIGS. 12, 13 is a gas turbine engine with a radar sensing element and an associated control and evaluation unit, wherein a combustor case of the core engine is monitored and a burn through of the combustor case is detected; and

[0058] FIG. 14 is a gas turbine engine with a radar sensing element and an associated control and evaluation unit, wherein a bleed valve is monitored and a bleed valve failure condition is detected.

[0059] FIG. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

[0060] In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where 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 turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

[0061] 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 fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e., not including the gearbox output shaft that drives the fan 23). 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 fan 23 may be referred to as a first, or lowest pressure, compression stage.

[0062] 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 in FIG. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are 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 engine 10 may not comprise a gearbox 30.

[0063] The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

[0064] 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 in FIGS. 2 to 5. FIGS. 6 to 14 regard embodiments of monitoring and detecting mechanical failures of different components of a gas turbine engine.

[0065] FIG. 2 depicts a system in which a radar sensing element 4 is located in a chamber 7 of a gas turbine engine. The chamber 7 may be, e.g., a nacelle, a bypass duct or a core engine compartment of the gas turbine engine. Within chamber 7 is located as first component 61 a duct which passes through the chamber 7. Within chamber 7, there is further located a second component 62.

[0066] The radar sensing element 4 emits radio waves 41 which are transmitted to the components 61, 62 and reflected by the components 61, 62. Radar sensing element 4 may 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. Components 61, 62 located within the beam of the radar sensing element 4 reflect some portion of the beam energy back to the radar sensing element 4, wherein an antenna (not shown) of the radar sensing element 4 detects the reflected radio waves. Of course, the depiction of two components 61, 62 in FIG. 2 is to be understood as an example only. There may be located only one component in compartment 7 or more than two components in compartment 7. Also, the monitored component may be a wall or a part of a wall of the chamber 7.

[0067] Radar sensing element 4 is connected to a control and evaluation unit 5 which controls the radar sensing element 4 and receives information/data from the radar sensing element 4. In particular, at least one property value of the reflected waves which are detected by the radar sensing element 4 is determined and provided to the control and evaluation unit 5. Determination of the property value may be performed in the radar sensing element 4 or in the control and evaluation unit 5. 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.

[0068] The control and evaluation unit 5 may 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 unit 5 may be a unit separate from the EEC and interacting with the EEC. As all other components in FIG. 1, the control and evaluation unit 5 is depicted only schematically.

[0069] More particularly, the control and evaluation unit 5 comprises a central processing unit 51 which receives data from the radar sensing element 4. The control and evaluation unit 5 further comprises a power source 54, a mass storage memory 52 in communication with the central processing unit 51 and in interface 53 for sending data, e.g., to an aircraft on-board communication unit. It is pointed out that only the components of the control and evaluation unit 5 relevant for the present invention are depicted in FIG. 2.

[0070] Program instructions are stored in memory 52 which cause, when executed by central processing unit 51, the performance of method steps as discussed with respect to FIGS. 2 to 4.

[0071] FIG. 3 depicts a method of detecting a mechanical failure of a component which, in principle, may be any component in a gas turbine engine. In step 301, a radar sensing element is provided that is configured to transmit and detect radio waves, such as radar sensing element 4 of FIG. 2. In step 302, 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.

[0072] In step 303, 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.

[0073] In step 304, 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 step 305, a mechanical failure is reported, e.g., by sending a warning signal through interface 53 to an aircraft on-board communication unit.

[0074] FIGS. 4 and 5 discuss two different embodiments of how the determination of step 304 of FIG. 3 that a mechanical failure is present is made.

[0075] According to the method of FIG. 4, in step 401, at least one property value of the reflected waves in the current state is determined. In step 402, 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 memory 52 of FIG. 2. In step 403, 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 step 404, a mechanical failure is identified. If the difference is below the threshold, a mechanical failure is not identified and the method continues in step 401 with determination and evaluation of further current property values.

[0076] In the method according to FIG. 5, artificial intelligence is implemented to determine if a mechanical failure is present. In step 501, 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 memory 52 of control and evaluation unit 5 of FIG. 2.

[0077] Further, in step 502, 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 memory 52 of control and evaluation unit 5 of FIG. 2.

[0078] In one embodiment, only step 501 or step 502 is 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 steps 501 and 502.

[0079] In step 503, 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 step 504 in the artificial intelligence engine. The artificial intelligence engine determines in step 505 if the change is indicative of a mechanical failure. Such determination is a direct result of the artificial intelligence engine. If so, in step 506, a mechanical failure is reported. If not, the method continues with step 503.

[0080] The artificial intelligence engine may be implemented in central processing unit 51 or may be implemented as a separate component of the control and evaluation unit 5.

[0081] 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.

[0082] FIGS. 6 to 10 show embodiments in which the radar monitoring technology of the present invention allows the early detection of a nacelle damage due to hot air impingement. In FIGS. 6 and 7, a radar sensing element 4 is located in a rear part of a nacelle 21. An EEC 5 which serves as a control and evaluation unit in accordance with the present invention is also located in nacelle 21. Air ducts 71 provide bleed air from two compressor stages, wherein the bleed air is provided in a duct 63 to an environmental control system of an aircraft. The radar sensing element 4 is able to detect changes in the geometry in case of a leak 630 in duct 63, wherein the leak 630 itself and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element 4. Such changes in geometry may be a delamination of parts and/or cracks.

[0083] In this respect, it is pointed out that, in an embodiment, the radar sensing element 4 may 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.

[0084] In FIGS. 8 and 9, compared to FIGS. 6 and 7, additionally a radar sensing element 4′ is located in a front part of nacelle 21. Air ducts 71 provide bleed air from two compressor stages, wherein the bleed air is provided in a duct 64 towards the inlet lip 210 of the nacelle 21 in order to prevent the formation of ice at the inlet lip 210. The radar sensing element 4′ is able to detect changes in the geometry in case of a leak 640 in duct 64, wherein the leak 640 itself and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element 4′. Such changes in geometry may be a delamination of parts and/or cracks.

[0085] In FIGS. 10 and 11, a radar sensing element 4 is located in a compartment of the core engine 11. An EEC 5 which serves as a control and evaluation unit in accordance with the present invention is located in nacelle 21. An air duct 65 of a secondary air system provides cooling bleed air from a compressor stage to a turbine stage of the core engine 11. The radar sensing element 4 is able to detect changes in the geometry in case of a leak 650 in duct 65, wherein the leak 650 itself and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element 4. Such changes in geometry may be a delamination of parts and/or cracks.

[0086] In FIGS. 12 and 13, similar as in FIGS. 10 and 11, a radar sensing element 4 is located in a compartment of the core engine 11. An EEC 5 which serves as a control and evaluation unit in accordance with the present invention is located in nacelle 21. A combustor 16 comprises an outer case 66. The radar sensing element 4 is able to detect changes in the geometry of the combustor case 66 in case of a leak 660, wherein the leak 660 itself and/or other changes in the geometry of the combustor case 66 and/or changes in the geometry of the hardware close to the leak can be detected by the radar sensing element 4. Such changes in geometry may be a delamination of parts and/or cracks.

[0087] In FIG. 14, a radar sensing element 4 is located in a rear part of a nacelle 21. An EEC 5 which serves as a control and evaluation unit in accordance with the present invention is also located in nacelle 21. A compressor bleed valve 67 is provided at compressor 15. If the bleed valve 67 has a failure in that it opens at high power for unlimited time, such opening will result in damage in the outer wall 68 of bypass duct 22, which will result in damage in the outer wall 62, 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 element 4. In this case, the radar sensing element 4 emits radio wave towards the wall 68 which represents the monitored component. In addition, a radar sensing element may be provided in proximity to bleed valve 67 to also monitor bleed valve 67.

[0088] 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.

[0089] 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.