DAMPER

Abstract

An inter-blade vibration damper for a gas turbine engine has an elongate damper body. The damper body is formed as a truncated cone and has a longitudinal axis. The elongate damper body is formed as a truncated cone. A conic surface of the damper body contacts a portion of a first blade and a portion of an adjoining second blade.

Claims

1. An inter-blade vibration damper for a gas turbine engine, comprising: an elongate damper body having a longitudinal axis, wherein the elongate damper body is formed with a conic surface, the conic surface of the damper body contacts a portion of a first blade and a portion of an adjoining second blade.

2. The vibration damper as claimed in claim 1, wherein the conic surface is a truncated conic surface.

3. The vibration damper as claimed in claim 2, wherein the truncated conic surface is a truncated circular conic surface.

4. The vibration damper as claimed in claim 2, wherein the truncated conic surface has an included angle of between 4 and 60.

5. The vibration damper as claimed in claim 2, wherein the truncated conic surface is a truncated elliptical conic surface, the truncated elliptical conic suface being oriented such that a major axis of the elliptical base of the truncated conic surface extends along a mid-plane between the first and second blades.

6. The vibration damper as claimed in claim 1, wherein the conic surface extends partially around the longitudinal axis.

7. The vibration damper as claimed in claim 1, wherein the conic surface extends as a half cone from a mid-plane through the longitudinal axis.

8. The vibration damper as claimed in claim 1, wherein the conic surface is provided with a surface roughness Ra in the range of 0.1 to 50 m.

9. A rotor device for a gas turbine engine comprising: a disc wheel; at least two blades extending radially from the disc wheel; and at least one vibration damper as claimed in claim 1.

10. The rotor device as claimed in claim 9, wherein the longitudinal axis of the elongate damper body subtends an angle with an axis of rotation of the disc wheel.

11. A gas turbine engine for an aircraft comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein at least one of the turbine and the compressor comprises a rotor device as claimed in claim 10.

12. The gas turbine engine according to claim 11, wherein: the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0083] There now follows a description of an embodiment of the disclosure, by way of non-limiting example, with reference being made to the accompanying drawings in which:

[0084] FIG. 1 is a sectional side view of a gas turbine engine;

[0085] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0086] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0087] FIG. 4 shows a partial schematic view of a turbine blade damper according to the prior art;

[0088] FIG. 5 shows the damper arrangement of FIG. 4 in use with the vibration damper experiencing rotation;

[0089] FIG. 6 shows a schematic partial perspective view of a vibration damper arrangement according to a first embodiment of the disclosure;

[0090] FIG. 7 shows a schematic partial side view of the vibration damper of FIG. 6;

[0091] FIG. 8 illustrates the relationship between kinematic slip and cone included angle for the vibration damper of FIGS. 6 and 7;

[0092] FIG. 9 shows a schematic partial perspective view of a vibration damper arrangement according to a second embodiment of the disclosure;

[0093] FIG. 10 shows a schematic partial side view of the vibration damper of FIG. 9;

[0094] FIG. 11 shows a schematic partial perspective view of a vibration damper arrangement according to a third embodiment of the disclosure;

[0095] FIG. 12 shows a schematic partial side view of the vibration damper of FIG. 11;

[0096] FIG. 13 is a perspective view of the vibration damper of FIG. 6;

[0097] FIG. 14 is a perspective view of the vibration damper of FIG. 9; and

[0098] FIG. 15 is a perspective view of the vibration damper of FIG. 11;

[0099] It is noted that the drawings may not be to scale. The drawings are intended to depict only typical aspects of the disclosure, and therefore should not be considered as limiting the scope of the disclosure. In the drawings, like numbering represents like elements between the drawings.

DETAILED DESCRIPTION

[0100] 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.

[0101] 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.

[0102] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

[0103] 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.

[0104] The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in FIG. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0105] The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

[0106] It will be appreciated that the arrangement shown in FIGS. 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the FIG. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of FIG. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in FIG. 2.

[0107] Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

[0108] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0109] 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.

[0110] 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.

[0111] 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.

[0112] Referring to FIGS. 6, 7, and 13, a vibration damper according to a first embodiment of the disclosure is designated generally by the reference numeral 100, the holder

[0113] The vibration damper 100 has an elongate damper body 130. The damper body 130 is positioned circumferentially between two adjacent turbine blades 112A,112B that are attached to a turbine rotor 110. Each of the turbine blades 112A, 112B has a blade platform 114. The blade platform 114 extends laterally from the base of each of the blades in a circumferential direction. The radially outward surface of the blade platforms 114 co-operate to form a contiguous surface extending around a circumference of the rotor 110.

[0114] Each of the blade platforms 114 has an underside 116 in a radial sense. The underside 116 of each platform is angled in a radially outward sense extending from the respective blade 112 towards each adjoining blade 112. This blade platform angle 117 is shown in FIG. 6.

[0115] The damper body 130 has a longitudinal axis 132. The damper body 130 is formed as a truncated circular cone 140 having a conic surface 142. The conic surface 142 contacts the underside 116 of a platform 114 of a first blade 112A, and the underside of a platform 114 of an adjoining second blade 112B.

[0116] The longitudinal axis 132 of the damper body 130 is aligned with a plane extending along the joint between the first blade 112A and the second blade 112B.

[0117] Since the conic surface 142 of the damper body 130 contacts the underside 116 of each of the pair of adjoining platforms 114, the longitudinal axis 132 of the damper body 130 must be inclined relative to a rotational axis of the rotor 110. This inclination 134 is illustrated in FIG. 7.

[0118] In this arrangement, the damper body 130 has an included angle 144 of 30. In other arrangements, this included angle may be between approximately 4 and 60.

[0119] In use, as each of an adjoining pair of turbine blades 112 undergoes opposing radially directed motion (as illustrated in FIG. 5) the underside 116 of each turbine blade platform 114 will be forced to slide against the corresponding conic surface 142 of the damper body 130. This sliding motion will result in the frictional dissipation of vibrational energy.

[0120] FIG. 8 illustrates the relationship between the included angle of the conic surface of the damper body and the degree of kinematic slip between the conic surface and the abutting turbine blade platform for a range of damper platform angles. FIG. 8 shows that for a given included angle, an increase in platform angle will produce increased slip between the damper body and the blade platform.

[0121] Referring to FIGS. 9, 10, and 14, a vibration damper according to a second embodiment of the disclosure is designated generally by the reference numeral 200. Features of the vibration damper 200 which correspond to those of vibration damper 100 have been given corresponding reference numerals for ease of reference.

[0122] The vibration damper 200 is positioned between the undersides 116 of the blade platforms 114 of two adjoining turbine blades 112A,112B in the same way as outlined above for the first embodiment 100.

[0123] The vibration damper 200 comprises an elongate damper body 230 having a longitudinal axis 232. The elongate damper body 230 is formed as a half truncated circular cone 240. The half truncated circular cone 240 has a conic surface 242. The conic surface 242 has an included angle 244.

[0124] In the same way as outlined above for the first embodiment, the conic surface 242 abuts the underside 116 of the blade platform 114 of each adjoining pair of turbine blades 112A,112B. The longitudinal axis 232 is inclined relative to the axis of rotation 122 of the rotor 110.

[0125] In use, radially directed relative motion between adjoining turbine blades 112A,112B causes a sliding motion between the underside 116 of each blade platform 114 and the corresponding conic surface 242. This sliding motion results in the frictional dissipation of vibrational energy and hence damps this radial blade movement.

[0126] Referring to FIGS. 11, 12, and 15, a vibration damper according to a third embodiment of the disclosure is designated generally by the reference numeral 300. Features of the vibration damper 300 which correspond to those of vibration damper 100 have been given corresponding reference numerals for ease of reference.

[0127] The vibration damper 300 is positioned between the undersides 116 of the blade platforms 114 of two adjoining turbine blades 112A,112B in the same way as outlined above for the first embodiment 100.

[0128] The vibration damper 300 comprises an elongate damper body 330 having a longitudinal axis 332. The elongate damper body 330 is formed as an elliptical cone 340. The elliptical cone 340 has a conic surface 342. The conic surface 342 has a major included angle 344A, and a minor included angle 344B.

[0129] In the same way as outlined above for the first embodiment, the conic surface 342 abuts the underside 116 of the blade platform 114 of each adjoining pair of turbine blades 112A,112B. The longitudinal axis 332 is inclined relative to the axis of rotation 122 of the rotor 110.

[0130] In use, radially directed relative motion between adjoining turbine blades 112A,112B causes a sliding motion between the underside 116 of each blade platform 114 and the corresponding conic surface 342. This sliding motion results in the frictional dissipation of vibrational energy and hence damps this radial blade movement.

[0131] While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting.

[0132] Moreover, in determining extent of protection, due account shall be taken of any element which is equivalent to an element specified in the claims. Various changes to the described embodiments may be made without departing from the scope of the invention.

[0133] In addition, where a range of values is provided, it is understood that every intervening value, between the upper and lower limit of that range and any other stated or intervening value in that stated range, is encompassed within the invention.

[0134] Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

[0135] The foregoing description of various aspects of the disclosure has been presented for purposes of illustration and description. It is not intended to be exhaustive or to limit the disclosure to the precise form disclosed, and obviously, many modifications and variations are possible. Such modifications and variations that may be apparent to a person of skill in the art are included within the scope of the disclosure as defined by the accompanying claims.