Gas turbine engine

Abstract

A gas turbine is provided for an aircraft comprising an engine core and a core flow path, a fan, a front drum cavity arranged radially inward of the core flow path, and a front bearing chamber. The front drum cavity comprises a front drum inlet, for providing air to the front drum cavity from the core air flow, located downstream of a stage of the compressor, and a front drum outlet, for ejecting air from the front drum cavity to the fan air flow, located axially between the fan and the compressor. The front drum inlet is through a seal, and the front drum outlet is through a spaced gap.

Claims

1. A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising at least one turbine and at least one compressor, and a core flow path for channelling a core air flow through the engine core; a fan located upstream of the engine core, the fan comprising a plurality of fan blades for producing a fan air flow; and a front drum cavity arranged radially inward of the core flow path; a front bearing chamber, comprising a front bearing, arranged radially inward of the core flow path and in fluid communication with the front drum cavity through one or more chamber seals; wherein the front drum cavity comprises a front drum inlet, for providing air to the front drum cavity from the core air flow, located downstream of a stage of the compressor, and a front drum outlet, for ejecting air from the front drum cavity to the fan air flow, located axially between the fan and the compressor; and wherein the front drum inlet is through a seal, and the front drum outlet is through a spaced gap.

2. The gas turbine of claim 1, wherein the flow resistance across the front drum inlet is higher than the flow resistance across the front drum outlet.

3. The gas turbine engine of claim 1, further comprising a bearing chamber vent line in fluid communication with the front bearing chamber, wherein the bearing chamber vent line comprises a vent pump for lowering the pressure in the front bearing chamber.

4. The gas turbine engine of claim 1, wherein the one or more chamber seals are contact carbon seals and/or air riding carbon seals.

5. The gas turbine engine of claim 1, wherein the front drum inlet is a labyrinth seal.

6. The gas turbine of claim 1, wherein the drum pressure ratio is less than 0.6 during operation of the gas turbine engine.

7. The gas turbine of claim 1, wherein the drum pressure ratio is less than 0.5 during operation of the gas turbine engine.

8. The gas turbine of claim 1, wherein the drum pressure ratio is less than 0.1 during operation of the gas turbine engine.

9. The gas turbine engine of claim 1, wherein the gas turbine engine further comprises a scavenge line in fluid communication with the front bearing chamber.

10. The gas turbine engine of claim 9, wherein the scavenge line comprises a scavenge pump.

11. The gas turbine engine of claim 1, further comprising a power gearbox, wherein optionally the gearbox is located within the front bearing chamber.

12. A method of designing and assembling a gas turbine engine of claim 1, the method comprising the steps of: defining the flow resistance of the front drum inlet and the front drum outlet such that the front drum cavity pressure is above the front bearing housing pressure; and installing the front drum inlet downstream of a compressor section and installing the front drum outlet axially between the fan and the compressor.

13. A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising at least one turbine and at least one compressor, and a core flow path for channelling a core air flow through the engine core; a fan located upstream of the engine core, the fan comprising a plurality of fan blades for producing a fan air flow; and a front drum cavity arranged radially inward of the core flow path; a front bearing chamber, comprising a front bearing, arranged radially inward of the core flow path and in fluid communication with the front drum cavity through one or more chamber seals; wherein the front drum cavity comprises a front drum inlet, for providing air to the front drum cavity from the core air flow, located downstream of a stage of the compressor, and a front drum outlet, for ejecting air from the front drum cavity to the fan air flow, located axially between the fan and the compressor; and wherein the drum pressure ratio is less than 0.6 during operation of the gas turbine engine.

14. The gas turbine of claim 13, wherein the drum pressure ratio is less than 0.5 during operation of the gas turbine engine.

15. The gas turbine of claim 13, wherein the drum pressure ratio is less than 0.1 during operation of the gas turbine engine.

Description

DESCRIPTION OF THE DRAWINGS

(1) Embodiments will now be described by way of example only, with reference to the Figures, in which:

(2) FIG. 1 is a sectional side view of an upstream portion of a prior art gas turbine engine;

(3) FIG. 2 is a sectional side view of a gas turbine engine;

(4) FIG. 3 is a close up sectional side view of an upstream portion of a gas turbine engine;

(5) FIG. 4 is a partially cut-away view of a gearbox for a gas turbine engine;

(6) FIG. 5a is a close up section side view of a front drum outlet;

(7) FIG. 5b is a close up section side view of a front drum inlet.

DETAILED DESCRIPTION

(8) Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

(9) FIG. 2 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.

(10) 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.

(11) FIG. 3 shows an upstream portion of a geared gas turbine engine, for example the gas turbine engine 10. The low pressure turbine 19 (see FIG. 2) 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.

(12) 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.

(13) FIG. 3 shows a front drum cavity 54 comprising first chamber 60, second chamber 61 and third chamber 62. The first chamber 60, second chamber 61 and third chamber 62 are all linked to form a single cavity. The second chamber 61 and third chamber 64 are linked by drum passageway 76. The front drum cavity 54 may be at the same pressure across the cavity. The front drum cavity 54 may be at substantially the same pressure across the cavity, for example within +/−1 psi or +/−0.5 psi or +/−0.1 psi.

(14) FIG. 3 shows a front bearing chamber 56 comprising fourth chamber 64, fifth chamber 65 and sixth chamber 66. The fourth chamber 64, fifth chamber 65 and sixth chamber 66 are all linked to form a single cavity. The front bearing chamber 56 comprises a front bearing 68. The front bearing chamber 56 may be at the same pressure across the cavity. The front bearing chamber 56 may be at substantially the same pressure across the cavity, for example within +/−1 psi or +/−0.5 psi or +/−0.1 psi.

(15) In the FIG. 3 example the front bearing chamber comprises a power gearbox, however in other examples a power gearbox may not be present. For example the arrangement of the front drum cavity, for example the arrangement and/or flow resistances of the front drum inlet and outlet may be applied to other gas turbine arrangements, for example non-geared gas turbine engines.

(16) In the FIG. 3 example the front bearing chamber 56 is bounded by the shaft 26, linkages 36 (and seals between the shaft 26 and linkages 36), the second chamber 61 and third chamber 62 and a radially inner surface 59 of a core flow path 58. In other examples the front bearing chamber 56 may be bounded only by a shaft and the front drum cavity. In the FIG. 3 example chamber seals, for example chamber seal 70, are positioned between the front bearing chamber 56 and the front drum cavity 54. As shown in FIG. 3, the front drum cavity 54 is typically bounded by the radially inner surface 59, a structure of the fan 23, the shafts 26, 36, the front bearing housing 56 and a structure of a compressor section (for example the low pressure compressor or the high pressure compressor).

(17) FIG. 3 shows a front drum inlet 52 and a front drum outlet 50. In the FIG. 3 example the front drum inlet 52 is a labyrinth seal. In the FIG. 3 example the front drum outlet 50 is a spaced gap.

(18) The spaced gap is a gap between the radially inner surface 59 of the core flow path 58 and a gas washed surface that projects from the base of a fan blade 23. These two parts rotate with respect to each other. A spaced gap may comprise no features projecting off the parts that are spaced apart, for example fins. A spaced gap may resemble a labyrinth seal without the fins. In the FIG. 3 example the spaced gap is positioned at the gas washed surface, however in other examples the spaced gap may be recessed from the surface, for example within third chamber 62. In such arrangement there may be a gap at the gas washed surface greater than the separation at the spaced gap within the third chamber 62, and a structure between the gap at the gas washed surface and spaced gap within the third chamber 62.

(19) In the FIG. 3 example the front drum outlet 50 is adjacent the fan 23. For example the front drum outlet 50 is axially closer to the fan 23 than it is to the compressor 14 or the core splitter supporting structure 24.

(20) In the FIG. 3 example the front drum cavity 54 may be maintained at a pressure above the front bearing chamber 56 by, for example, a vent line comprising a vent line pump. A vent line pump (not shown) may be in fluid communication with the front bearing chamber 56 and reduces the pressure in the front bearing chamber 56, for example to below atmospheric pressure. Alternatively a scavenge pump, and advanced seals (for example chamber seal 70) between the front drum cavity 54 and the front bearing chamber 56, may maintain the overpressure (in addition to or instead of the vent line). Advanced seals may be, for example, contact carbon seals or air riding carbon seals.

(21) Indicative pressures in the arrangement of FIG. 3 are shown in table 1 below. Pin is the pressure at the entrance of the front drum inlet 52. Pdrum is the pressure in the front drum cavity 54. Pout is the pressure at the exit of the front drum outlet 50.

(22) The drum pressure ratio is defined by the following equation:

(23) TABLE-US-00001 TABLE 1 indicative pressures for the FIG. 3 arrangement. $\mathrm{Drum}\mathrm{pressure}\mathrm{ratio}=\frac{\mathrm{Pdrum}-\mathrm{Pout}}{\mathrm{Pin}-\mathrm{Pout}}$ Cruise (psi) Max-take off (psi) Pin 20 60 Pdrum 6 18 Pout 5 15 Drum pressure 0.07 0.07 ratio

(24) Indicative pressures for a prior art arrangement, for example the arrangement of FIG. 1, are shown below in table 2 for comparison.

(25) TABLE-US-00002 TABLE 2 indicative pressures for a prior art arrangement. Cruise (psi) Max-take off (psi) Pin 20 60 Pdrum 15.5 45 Pout 5 15 Drum pressure ratio 0.70 0.67

(26) It can be seen from tables 1 and 2 that the drum pressure ratio is reduced substantially compared to the prior art arrangement.

(27) The pressure loss across the front drum inlet 52 may be between 9 and 10 psi, or 9 and 14 psi at cruise condition. The pressure loss across the front drum inlet 52 may be between 28 and 30 psi, or 28 and 45 psi at mid take-off condition.

(28) The pressure loss across the front drum outlet 50 may be between 0.6 and 0.8, or 0.6 and 1.1 psi at cruise condition. The pressure loss across the front drum outlet 50 may be between 1.5 and 1.8, or 1.5 and 3.1 psi at mid take-off condition.

(29) Therefore the flow resistance of the front drum inlet 52 may be approximately 10 times the flow resistance of the front drum outlet 50.

(30) The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 4. 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. 4. 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.

(31) 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.

(32) 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. 3 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. 3. 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. 3.

(33) 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.

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

(35) 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. 2 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core exhaust 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.

(36) 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. 2), and a circumferential direction (perpendicular to the page in the FIG. 2 view). The axial, radial and circumferential directions are mutually perpendicular.

(37) FIG. 5a shows a close up sectional side view of a front drum outlet 82, which may be an example of the front drum outlet 50 of FIG. 3. FIG. 5a shows a fan projection 80 and a core projection 84. The fan projection 80 may be fixed to and rotate with the fan, for example the fan 23 of FIG. 3. The core projection 84 may not rotate and extend from the radially inner gas washed surface of the core flow path, for example the radially inner surface 59 of FIG. 3. The core projection 84 and fan projection 80 provide a spaced gap between a front drum cavity 86 and the gas flowpath 87.

(38) The fan projection 80 and core projection 84 overlap one another in the FIG. 5a example. The core projection 84 extends radially inward, or underneath, the fan projection 80. The core projection 84 and fan projection 80 extend parallel to one another over an axial length, for example a short axial length. A spaced gap is provided between the core projection 84 and the fan projection 80. The spaced gap is provided between opposing surfaces of the core projection 84 and the fan projection 80. The opposing surfaces have no flow restriction features, for example such as fins. The front drum outlet 82 does not provide a tortuous path to airflow passing through it. In the FIG. 5a example the spaced gap is a radial gap. However in other examples the core projection 84 and fan projection 80 may be at a different orientation, or located within the front drum cavity 86. In other examples the fan projection 80 and core projection 84 may not overlap. In other examples the front drum outlet 82 spaced gap may be a slot or hole.

(39) FIG. 5b shows a close up sectional side view of a front drum inlet 92, which may be an example of the front drum inlet 52 of FIG. 3. FIG. 5b shows a compressor section projection 88 and a core seal projection 94 that together form a labyrinth seal.

(40) The compressor section projection 88 extends from the compressor disc. The compressor disc may form part of, for example, the compressor 14 of FIG. 3. The core seal projection 94 extends from the radially inner surface of a core flow path, which may for example be the radially inner surface 59 shown in FIG. 3. A compressor support structure 90 is adjacent the front drum inlet 92. The support structure 90 may provide an improved seal, for example reducing the influence of vibration. The front drum inlet 92 provides a flow restriction between the core flow path 98 and a front drum cavity 96, which may be an example of the front drum cavity 54 of FIG. 3.

(41) In other examples the front drum inlet 92 may be a stepped labyrinth seal, a foil seal or a contact carbon seal. These seals may provide an equivalent flow restriction compared to the labyrinth seal shown in FIG. 5b.

(42) It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. 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.