SHAFT FAILURE PROTECTION SYSTEM

Abstract

A shaft failure protection system includes an engine core comprising a turbine, a compressor, and a shaft connecting the turbine and compressor; a first braking element connected to a rotating part of the turbine; and a second braking element connected to a static part of the turbine. The first and second braking elements are arranged at an axial distance under normal operating conditions and configured to contact each other in case of a failure of the shaft and an associated axial displacement of the rotating part. The first braking element includes a first friction material and the second braking element comprises a second friction material, wherein the first and second friction materials each comprise a carbon-silica composite or a carbon-fibre-reinforced carbon. Upon shaft failure and associated axial displacement of the rotating part, the first and second friction materials contact each other to reduce speed of the rotating part.

Claims

1. A shaft failure protection system comprising: an engine core comprising a turbine, a compressor, and a shaft connecting the turbine and the compressor; a first braking element connected to a rotating part of the turbine; a second braking element connected to a static part of the turbine; wherein the first braking element and the second braking element are arranged at an axial distance under normal operating conditions and configured to contact each other in case of a failure of the shaft and an associated axial displacement of the rotating part of turbine, wherein the first braking element comprises a first friction material and the second braking element comprises a second friction material, wherein the first friction material and the second friction material each comprise a carbon-silica composite or a carbon-fibre-reinforced carbon; wherein, in case of a failure of the shaft and an associated axial displacement of the rotating part of the turbine, the first friction material and the second friction material contact each other to reduce rotational speed of the rotating part of the turbine by frictional forces.

2. The system of claim 1, wherein the first friction material and the second friction material each comprise a carbon-silica composite, wherein the carbon-silica composite is a carbon fibre reinforced silicon carbide, wherein carbon fibres are integrated in a silicon carbide (SiC) matrix.

3. The system of claim 1, wherein the first friction material and the second friction material are chosen such that the coefficient of kinetic friction between these materials is in the range between 0.15 and 0.8 at the operating temperature.

4. The system of claim 3, wherein the first friction material and the second friction material are chosen such that the coefficient of kinetic friction between these materials is in the range between 0.4 and 0.6 at the operating temperature.

5. The system of claim 3, wherein the operating temperature is 500° C.

6. The system of claim 3, wherein the operating temperature is 1300° C.

7. The system of claim 1, wherein the first friction material and the second friction material are identical.

8. The system of claim 1, wherein rotating part of the turbine is a rotor disc, wherein the first braking element is connected to a sealing element structure coupled to the rotor disc.

9. The system of claim 1, wherein the static part of the turbine to which the second braking element is connected is coupled to a bearing structure for the shaft.

10. The system of claim 1, wherein the first and second braking elements each comprise a surface, the surfaces contacting each other in case of a shaft failure.

11. The system of claim 10, wherein the surface is a flat surface.

12. The system of claim 10, wherein the surface of the first braking element and/or the surface of the second braking element which contact each other in case of a shaft failure have undergone a surface treatment that has increased the roughness of the surface.

13. The system of claim 1, wherein the first braking element and/or the second braking elements is in the form of a ring.

14. The system of claim 1, wherein the turbine is a high pressure turbine of the engine core.

15. A gas turbine engine for an aircraft comprising a system in accordance with claim 1.

Description

[0037] 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.

[0038] 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.

[0039] 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.

[0040] 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.

[0041] 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.

[0042] In both the high-pressure turbine 17 and the low pressure turbine 19 of the gas turbine engine 10, the turbine 17, 19 comprises at least one rotating part and at least one static part. The rotating part includes a rotating disc to which individual turbine blades are connected. The static part includes a stator that comprises turbine vanes.

[0043] In a gas turbine engine 10 as discussed with respect to FIG. 1, or in any other gas turbine engine, a shaft failure protection system may be implemented to limit the rotational speed of the rotating turbine disc by frictional forces in case of a shaft failure.

[0044] FIGS. 2 and 3 show an embodiment of such shaft failure protection system. The shaft failure protection system is implemented in a turbine of the gas turbine engine. In the embodiment depicted, the shaft failure protection system is implemented in the high-pressure turbine 17 of the gas turbine engine. FIG. 2 depicts a combustor 16 and nozzle guide vanes 6 located downstream of the combustor 16. The nozzle guide vanes 6 direct the gas flow from the combustor 16 onto turbine blades 171 which are connected to the outer rim of a rotor disc 170. The rotor disc 170 and the turbine blades 171 form a rotor of the high-pressure turbine 17. On passing through the nozzle guide vanes 16, gases from the combustor 16 are given a swirl in the direction of the rotation of the turbine rotor blades 171. The turbine rotor blades 171 receive a force from the gas flow which causes the turbine disc 170 to rotate at a high speed.

[0045] The turbine 17 further comprises a static part. The static part includes stator vanes 175 located in the gas path downstream of the rotor blades 171. The static part further includes structural components such as walls 177 which form the static part of a rear bearing arrangement 6 which includes two roller bearings 61, 62 that constrain movement of the shaft in the radial direction but do not constrain movement of the shaft in the axial direction. Static parts 177 may be coupled to a casing of the turbine 17.

[0046] In FIG. 2, there are further depicted flows of cooling air. For example, cooling air CA-1 is received from the high-pressure compressor and serves to cool the rotor disc 170 and the turbine blades 171. Cooling air CA-2 is received from the high-pressure compressor and/or the low-pressure compressor and serves to seal lubrication oil within the bearings 61, 62. To this end, cooling air CA-2 is led through a pipe 176 against the radial direction to the rear bearing arrangement 6. The cooling air is part of a secondary air system. Functions of the secondary air systems are, among others, cooling, sealing of oil cavities, sealing of the main gas path, and bearing load management.

[0047] A seal 7 is provided between the rotating part and the static part of the turbine 70. As shown in FIG. 3, the seal 7 comprises a static sealing element structure 71 connected to the static part of the turbine and a rotating sealing element structure 172 connected to the rotor disc 170.

[0048] The system further comprises two braking elements 4, 5. The first braking element 4 is connected to the rotating sealing element structure 172 by means of a connection 45 which is depicted schematically in FIG. 3. The second braking element 5 is connected to walls 177 of the static part which are coupled to the bearing structure 6. The connection of the second braking element 5 to walls 177 is provided by means of a connection 55 which is depicted schematically in FIG. 3.

[0049] FIG. 3 further depicts a flange connection 178 connecting static wall elements.

[0050] Under normal operation, as shown in FIGS. 2 and 3, the first braking element 4 and the second braking element 5 are arranged at an axial distance. However, in case of a shaft failure, the rotor disc 170 becomes axially displaced in the downstream direction such that the first braking element 4 and the second braking element 5 get into contact.

[0051] As shown in FIG. 3, in such case, the respective surfaces 41, 51 of the first and second braking elements 4, 5 form mating surfaces which get into contact, thereby creating frictional forces which reduce the rotational speed of the rotor disc 170, keeping the rotor disc 170 below the maximum permissible speed (terminal speed) and thereby preventing an otherwise possible braking of the rotor disc 170.

[0052] Both braking elements 4, 5 are in the form of a circumferential ring such that the surfaces 41, 51 which get into contact have a large surface area.

[0053] The surfaces 41, 51 are flat and arranged parallel to each other in the depicted embodiment. However, other corresponding forms of the surfaces 41, 51 may be implemented, such as a concave surface 41 of the first braking element 4 and a convex surface 51 of the second braking element or vice versa.

[0054] It is pointed out that the radial distance of the position of the braking elements 4, 5 from the main axis 9 (see FIG. 1) influences the resultant braking torque, as the braking torque is the force acting between the respective contact areas of the braking elements 4, 5 times the radial distance from the rotational axis.

[0055] This further means that the braking torque further depends on the size of the contact area between the braking elements 4, 5 as the size of this contact area determines the force acting between the braking elements 4, 5.

[0056] In view of this, a higher braking torque can be achieved when placing the braking elements at a larger distance from the main axis and having a large contact area. At the same time, larger contact areas lead to an increased weight of the braking elements. It is a design task to select the radius such that the braking power is sufficiently high while minimizing the weight of the braking elements.

[0057] The first braking element 4 consists of a first friction material and the second braking element 5 consists of a second friction material. Both friction materials consist of or comprise a carbon-silica composite such as carbon fibre reinforced silicon carbide (C/SiC) or a carbon-fibre-reinforced carbon (C/C). For example, the friction material of both braking elements 4, 5 is a carbon fiber reinforced silicon carbide (C/SiC). The first braking element 4 and the second braking element 5 may consist of the identical friction material.

[0058] Both carbon fibre reinforced silicon carbide (C/SiC) and carbon-fibre-reinforced carbon have a high coefficient of kinetic friction in the relevant temperature range between 500° C. and 1300° C., the coefficient of kinetic friction being in the range between 0.15 and 0.8. Carbon-fibre-reinforced carbon (C/C) is a composite material consisting of carbon fibre reinforcement in a matrix of graphite. Carbon fibre reinforced silicon carbide (C/SiC) is a composite made of a silicon carbide matrix with carbon fibre reinforcement. Both materials are well described in the scientific literature.

[0059] The friction material of the braking elements 4, 5 has material properties such that the coefficient of kinetic friction between the first braking element 4 and the second braking element 5 is higher than the coefficient of kinetic friction in a metal-to-metal contact (which would occur between the rotating part and static part of the turbine 17 without the braking elements 4, 5). In embodiments, the coefficient of kinetic friction lies in the range between 0.15 and 0.8, in particular in the range between 0.4 and 0.6. This coefficient of kinetic friction is present at the operating temperature of the turbine, which may be in the range between 500° C. and 1300° C.

[0060] To increase the frictional forces between the first braking element 4 and the second braking element 5, the surfaces 41, 51 of the braking elements 4, 5 may have experienced a surface treatment that increases the roughness of the surfaces 41, 51. In such case, the roughness of the surfaces 41, 51 of the braking elements may be higher than with other of the surfaces of the braking elements 4, 5.

[0061] The shaft failure protection system may comprise further components such as an automatic fuel shut off once a shaft failure occurs as known to the skilled person.

[0062] 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, the location of the first braking element 4 and the second braking element 5 within the turbine 17 may be different and the form of the first braking element 4 and of the second braking element 5 may be different than depicted in the embodiment of FIGS. 2 and 3.

[0063] 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.