GAS TURBINE ENGINE

Abstract

A gas turbine engine, and arrangements of turbine blades around an exhaust nozzle of an engine core. Example embodiments include 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, the engine core comprising an inlet upstream of the compressor and an exhaust nozzle at a downstream outlet of the turbine; a fan located upstream of the engine core inlet; and a set of exhaust nozzle vanes spanning the exhaust nozzle, the turbine comprising a first row of turbine blades upstream of the exhaust nozzle vanes and a second row of turbine blades downstream of the exhaust nozzle guide vanes, one or more of the exhaust nozzle guide vanes comprising a passage configured to direct airflow downstream from the first row of turbine blades towards the second row of turbine blades.

Claims

1. 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, the engine core comprising an inlet upstream of the compressor and an exhaust nozzle at a downstream outlet of the turbine; a fan located upstream of the engine core inlet; and a set of exhaust nozzle vanes spanning the exhaust nozzle, the turbine comprising a first row of turbine blades upstream of the exhaust nozzle vanes and a second row of turbine blades downstream of the exhaust nozzle guide vanes, each of the exhaust nozzle guide vanes comprising a passage configured to direct airflow downstream from the first row of turbine blades towards the second row of turbine blades.

2. The gas turbine engine of claim 1, wherein an inlet of the passage is positioned to entrain airflow at a maximum pressure exiting from the first row of turbine blades.

3. The gas turbine engine of claim 1, wherein the inlet of the passage extends between around 20% and 80% of the span of the exhaust nozzle guide vanes.

4. The gas turbine engine of claim 1, wherein an inlet of the passage extends between around 30% and 70% of a span of the exhaust nozzle guide vanes.

5. The gas turbine engine of claim 1, wherein the first row of turbine blades is a final downstream row of turbine blades immediately upstream of the exhaust nozzle guide vanes, the first row of turbine blades have a first blade tip diameter, the second row of turbine blades have a second blade tip diameter, and the second blade tip diameter is smaller than the first blade tip diameter.

6. The gas turbine engine of claim 5, wherein the second blade tip diameter is less than 70% of the first blade tip diameter.

7. The gas turbine engine of claim 6, wherein the second blade tip diameter is greater than around 30% of the first blade tip diameter.

8. The gas turbine engine of claim 5, wherein a span of each blade in the second row of turbine blades is between around 20% and around 50% of a span of each blade in the first row of turbine blades.

9. The gas turbine engine of claim 1, wherein the passage extends to a duct surrounding the second row of turbine blades.

10. The gas turbine engine of claim 1, wherein the passage comprises an adjustable valve arranged to control an amount of airflow through the passage.

11. The gas turbine engine of claim 1, further comprising a rotating machine connected to the second row of turbine blades.

12. The gas turbine engine of claim 11, wherein the rotating machine is an electrical generator or a hydraulic pump.

13. The gas turbine engine of claim 11, further comprising a gearbox connected between the rotating machine and the second row of turbine blades.

14. The gas turbine engine of claim 1, wherein the first and second rows of turbine blades are mounted for rotation on a common shaft.

15. The gas turbine engine according to claim 1, 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

DESCRIPTION OF THE DRAWINGS

[0050] Embodiments will now be described by way of example only, with reference to the accompanying drawings, in which:

[0051] FIG. 1 is a sectional side view of a gas turbine engine;

[0052] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0053] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0054] FIG. 4 is a schematic diagram of an aft section of an example gas turbine engine;

[0055] FIG. 5 is a schematic diagram of an aft section of an example gas turbine engine with an additional row or turbine blades; and

[0056] FIG. 6 is a schematic partial sectional view of the gas turbine engine of FIG. 5.

DETAILED DESCRIPTION

[0057] 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.

[0058] 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.

[0059] 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 core exhaust 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.

[0060] 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.

[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] 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.

[0063] 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.

[0064] 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.

[0065] 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.

[0066] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0067] 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 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 engine 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.

[0068] 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.

[0069] FIG. 4 illustrates in simplified schematic sectional form the aft section of an example gas turbine engine 400, illustrating a gas flow path from a high pressure turbine stage 401 through to a low pressure turbine 402 and exhaust 403. The low pressure turbine 402 comprises a plurality of rows 404 of vanes alternating with a plurality of rows 405 of turbine blades. The rows 405 of turbine blades are mounted to respective discs 406 and rotate together about a rotational axis 407 on a turbine shaft 408, which connects through to the front of the engine for driving a compressor and/or fan of the engine. The shaft 408 is mounted for rotation to a bearing 410. A row of exit vanes 411 span the exhaust 403, supporting the engine 400 to a rear engine mount static structure 409. In operation, airflow from the high pressure turbine 401 expands as it flows through to the low pressure turbine 402 and exits the engine 400 between the exit vanes 411 over an engine plug 412, along with airflow from the bypass duct 22 (FIG. 1).

[0070] FIG. 5 illustrates in simplified schematic sectional form the aft section of an example gas turbine engine 500 similar to that of FIG. 4, but with a second row 501 of turbine blades downstream of the exhaust nozzle guide vanes 511 in addition to a first row 505 of turbine blades upstream of the vanes 511, the first row 505 being part of the low pressure turbine 402 of the engine 500. A passage 502 is provided in each guide vane 511 that is configured to direct airflow downstream from the first row 505 of turbine blades towards the second row 501 of turbine blades.

[0071] An inlet 503 of the passage 502 is positioned to entrain airflow at a maximum pressure exiting from the first row 505 of turbine blades. As shown in FIG. 6, the air pressure 601 varies across the span 512 of the vane 511, reaching a maximum at around halfway across the span. The inlet 503 may therefore be positioned such that it extends between around 20% and 80% of the span of the vane 511, or alternatively between around 30% and 70% of the span of the vane 511, or further alternatively between around 40% and 60% of the span of the vane 511.

[0072] The tip diameter D.sub.2 of the second row 501 of turbine blades is substantially smaller than the tip diameter D.sub.1 of the first row 505 of turbine blades. The tip diameter D.sub.2 of the second row 501 of turbine blades may for example be less than around 70% of the tip diameter of the first row 505 of turbine blades, and may be between around 30% and 70% of the tip diameter of the first row 505.

[0073] The span 513 of each blade in the second row 501 is substantially smaller than the span 512 of each blade in the first row 505, and may for example be between around 20% and 50% of the span 512 of each blade in the first row 505.

[0074] In the example shown in FIG. 5, the second row 501 of turbine blades drives a rotating machine 504, which may for example be an electrical generator or a hydraulic pump. A hydraulic pump may be used for pumping lubricant and/or cooling fluid through the engine, whereas an electrical generator may provide electrical power for powering aircraft electrical systems. In some examples a gearbox 506 may be connected between the rotating machine and the second row 501 of turbine blades.

[0075] The second row 501 of turbine blades may be mounted to a disc 503, as with the first row 505 and other rows of turbine blades in the turbine 402.

[0076] In alternative examples to that shown in FIG. 5, the first and second rows 505, 501 of turbine blades may be mounted for rotation on a common shaft, for example the turbine shaft 408. A connection between the second row 501 and the turbine shaft 408 may be made via a clutch and/or a gearbox to enable rotation of the second row 501 to be engaged and disengaged and its rotational speed matched to that of the turbine shaft 408. In other examples the connection may be fixed, so that a rotational speed of the first and second rows 505, 501 is the same. The gearbox 504 may allow the second row 501 of turbine blades to run faster than those of the turbine 402.

[0077] An adjustable valve 507 may be provided in the passage 502, for example positioned at or adjacent the inlet 503, to allow airflow through the passage 502 and through to the second row 501 of turbine blades to be controlled.

[0078] In each of the examples described and illustrated herein, a row of turbine blades comprises a plurality of turbine blades that are mounted around the circumference of a common disc at a common axial position along a length of the engine core.

[0079] The passage 502 through each of the guide vanes 511 may lead to a duct 508 that extends around an inner circumference of the exhaust 403, forming an annular gallery that distributes air flowing from each passage through to the second row 501 of turbine blades, the duct 508 and passages 502 together forming a manifold that collects and distributes airflow downstream of the first row 505 of turbine blades to the second row 501 of turbine blades. Inlet guide vanes 509 may be provided in the duct 508 immediately upstream of the second row 501 of turbine blades to direct airflow. The inlet guide vanes 509 may be variable to control the airflow entering the second row 501 of turbine blades. The variable vane 509 can be used to control the speed of a generator independently from other stages of the turbine 402, and may allow for a reduction in the exit airflow Mach number.

[0080] The rotating machine 504 may be positioned fore or aft in relation to the position of the second row 501 of turbine blades. The rotating machine 504 may for example be positioned within the plug 412 downstream of the exhaust 403. The plug 412 may be mounted to rotate along with the second row 501 of turbine blades or may be fixed in relation to the engine support structure 409.

[0081] The power drawn by the second row 501 of turbine blades as a proportion of the total power generated by the turbine 402 may be around 10%. For a 2,000 HP (1.5 MW) turbine, the power drawn by the second row 501 may be around 200 HP (150 kW). This power may be provided as electrical power via an electrical generator or hydraulic power via a hydraulic pump or as additional power fed back along the turbine shaft.

[0082] A clutch may be provided that enables an electrical or hydraulic generator to be engaged when needed and for the second row 501 of turbine blades to be otherwise engaged with the other rows of turbine blades in the turbine. Electrical power may be extracted for example during descent of the aircraft.

[0083] A further advantage is being able to controllably share power between shafts so that power generated on the low pressure turbine can be redistributed to other shafts on the engine via electrical power.

[0084] 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.