GAS TURBINE ENGINE

Abstract

A gas turbine engine including an engine core with a duct, nacelle, and bypass duct receiving fan-accelerated bypass air flow. The core duct is located radially inside the bypass duct and receives the core air flow. A housing between the core and bypass ducts has an outer wall, which is the bypass duct inner wall, and inner wall which is core duct outer wall. The housing extends axially from the gas turbine engine, and splits fan accelerated air flow axially forward into the bypass and core ducts. At least two heat exchangers for cooling engine based oil are mounted in the housing. A flow passage inside the housing delivers air flow to the heat exchangers, and returns air flow from the heat exchangers. The at least two heat exchangers extend circumferentially, and a flow divider is between the heat exchanger ends and diverts air flow to the heat exchangers.

Claims

1. A gas turbine engine including: an engine core having an engine core duct surrounded by an annular outer wall and having a radially annular inner wall; a nacelle surrounding the engine core and a propulsive fan; a bypass duct for receiving a bypass air flow accelerated by the propulsive fan, the bypass duct being defined between the nacelle and the core duct which is designed for receiving a core air flow accelerated by the propulsive fan, and which is located radially inside the bypass duct; a housing having an outer wall which is an inner wall of the bypass duct, and having an inner wall which is the outer wall of the core duct; whereby the housing is extending axially from an area of the gas turbine engine downstream of the fan and is splitting the air flow accelerated by the fan in an axially forward area of the housing so that the air flows partly into the bypass duct and partly into the core duct; at least two heat exchangers for cooling engine based oil using an air flow which is passed from the fan to the heat exchangers and from the heat exchangers into the bypass duct, the heat exchangers being mounted radially between the radially inner annular wall and the radially outer annular wall of the housing; and a flow passage inside the housing delivering the air flow from an annular inlet of the flow passage to the heat exchangers, and returning the air flow from the heat exchangers through an annular outlet in the radially outer annular wall of the housing into the bypass duct; wherein the inlet of the flow passage is located upstream the heat exchangers, and is located in the forward area of the housing between the radially inner annular wall and the radially outer annular wall of the housing; and/or in the radially inner annular wall of the housing downstream of the forward area of the housing and of outer guide vanes extending radially through the bypass duct, and downstream of engine section stators extending radially through the core duct, so air enters the inlet from a stage of an IP-compressor unit; and/or in the radially outer annular wall of the housing directly downstream of the outer guide vanes and of the forward area of the housing; wherein the inlet surrounds the engine core circumferentially; and wherein at least two heat exchangers extend circumferentially, whereby a flow divider is provided between facing ends of the heat exchangers that diverts the air flow in direction to the heat exchangers respectively.

2. The gas turbine engine of claim 1, wherein the flow passage comprises guide vanes upstream the heat exchangers to straighten out the swirl air from the fan.

3. The gas turbine engine of claim 1, wherein the inlet of the flow passage located in the radial outer annular wall of the housing comprises a ram scoop that extends radially and circumferentially into the bypass duct, and guides the air flow into the flow passage.

4. The gas turbine engine of claim 1, wherein the flow passage comprises means which are arranged to direct the air flow in the flow passage in such a way that the air flow streams towards the heat exchangers at a defined angle.

5. The gas turbine engine of claim 1, wherein at least two heat exchangers are provided that extend circumferentially, whereby a flow divider is provided between facing ends of the heat exchangers that diverts the air flow in direction to the heat exchangers respectively.

6. The gas turbine engine of claim 1, wherein a firewall is provided in the housing downstream the heat exchangers that separates the heat exchangers from a rear core fire zone.

7. The gas turbine engine of claim 1, wherein each heat exchanger is mounted to the housing or to the firewall.

8. The gas turbine engine of claim 1, wherein excess air from the circumferential inlet is fed to other systems of the gas turbine engine, whereby inlets for these systems are placed between the heat exchangers.

9. The gas turbine engine of claim 1, wherein an ejector is provided in the flow passage downstream of the heat exchangers that accelerates the air flow downstream the heat exchangers.

10. The gas turbine engine of claim 1, wherein the ejector is fluidly coupled with a cabin blower unit, a high pressure bleed or a compressor offtake.

11. The gas turbine engine of claim 1, wherein facing ends of the heat exchangers are spaced apart from each other circumferentially, supporting thermal growth and oil porting.

12. The gas turbine engine of claim 1, wherein the radially outer annular wall and radial inner annular wall of the housing comprise several removable panels to allow access underneath.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0073] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0074] FIG. 1 is a sectional side view of a gas turbine engine;

[0075] FIG. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

[0076] FIG. 3 is a partially cut-away view of a gearbox for a gas turbine engine;

[0077] FIG. 4 is an enlarged partial longitudinal sectional view of an embodiment of a gas turbine engine with a heat exchanger which can be supplied with an air flow via a so called ESS-style-annular inlet located at a core duct (the gas path) bypass duct splitter with an inlet flow path between a core duct (gas path) outer wall and a core fairing (bypass duct) inner wall and directly downstream of a propulsive fan of the engine;

[0078] FIG. 5 is an enlarged partial longitudinal sectional view of a further embodiment of a gas turbine engine with a heat exchanger which can be supplied with an annular air flow via a so called IPC-style-inlet located in the outer wall of the core duct (the gas path);

[0079] FIG. 6 is an enlarged partial longitudinal sectional view of a further embodiment of a gas turbine engine with a heat exchanger which can be supplied with an annular air flow via a so called OGV-style-inlet located in the core fairing (bypass duct inner wall) directly downstream of fan OGV and with a firewall;

[0080] FIG. 7 is a schematic cross-sectional view of a gas turbine engine with several heat exchangers extending in the circumferential direction; and

[0081] FIG. 8 is a three-dimensional partial view of an embodiment of a gas turbine engine which is comprising several heat exchangers.

[0082] The following table lists the reference numerals used in the drawings:

TABLE-US-00001 Ref no. Feature A Core airflow B Bypass airflow  9 Principal and rotational axis (of engine) 10 Gas turbine engine 11 Core 12 Air intake 14 Low pressure compressor 15 High pressure compressor 16 Combustion equipment 17 High pressure turbine 18 Bypass exhaust nozzle 19 Low pressure turbine 20 Core exhaust nozzle 21 Fan nacelle or fan case 22 Bypass duct 23 Fan 24 Stationary supporting structure 26 Shaft 27 Interconnecting shaft 28 Sun wheel or sun gear 30 Epicyclic gear arrangement 32 Planet gears 34 Planet carrier 36 Linkage 38 Ring gear 39 Forward area of the housing 40 Linkage 41 Core duct outer wall 42 Heat exchanger 42A to 42G Heat exchanger 43 Inlet of the flow passage, ESS-style-inlet 44 Core duct inner wall 45 Core fairing; bypass duct inner wall; housing outer wall 46 Housing 47 Core duct 48 Flow passage 49 Outlet of the flow passage 50 Air flow 51 Guide vane 53 Air flow 54 Inlet of the flow passage, IPC-style-inlet 55 Air flow 56 Inlet of the flow passage, OGV-style-inlet 58 Firewall 59 Zone 60 Further zone 64 Divider 65 Nacelle outer wall 66 Nacelle inner wall 67 OGV casing 72 Outer guide vane 75 Radially outer annular wall, annular cooling duct outer wall

DETAILED DESCRIPTION OF THE DISCLOSURE

[0083] 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.

[0084] 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.

[0085] 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.

[0086] 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.

[0087] 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.

[0088] 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.

[0089] The epicyclic gearbox 30 illustrated by way of example in FIG. 2 and FIG. 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.

[0090] It will be appreciated that the arrangement shown in FIG. 2 and FIG. 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.

[0091] 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.

[0092] Optionally, the gearbox may drive additional and/or alternative components (e.g., the intermediate pressure compressor and/or a booster compressor).

[0093] 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 engine core 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.

[0094] 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.

[0095] The engine core 11 is surrounded by a core duct (the gas path) 47. The core duct 47 comprises an inner wall 44 and an outer wall 41 which is radially spaced apart from an inner wall 45 of the bypass duct 22, also called a core fairing. Radially between the core duct 47 and the bypass duct 22 is a so called housing 46 provided, wherein the outer wall of the housing 46 is formed by the radially inner wall (core fairing) 45 of the bypass duct 22, and the radially inner wall of the housing 46 is formed by the radially outer wall 41 of the core duct 47.

[0096] The bypass duct 22 is receiving the bypass air flow B accelerated by the propulsive fan 23 and is defined between the nacelle 21 and the bypass duct inner wall (core fairing) 45. The core duct 47 for receiving the core airflow A accelerated by the propulsive fan 23 is defined radially inside the core duct outer wall 41 and radially outside the core duct inner wall 44. In a forward area 39 of the housing 46 the air flow accelerated by the fan 23 is split into the bypass duct airflow B and into the core duct airflow A so that the forward area 39 forms a so called bypass to core splitter. The nacelle 21 comprises a radially outer wall 65 and a radially inner wall 66 which defines also the outer wall of the bypass duct 22.

[0097] FIG. 4 to FIG. 8 show views of several embodiments of the gas turbine engine 10, which have essentially the same design structure and the same functionality as the gas turbine engine 10 according to FIG. 1 to FIG. 3, and differ from each other only in the manner described in more detail below. In order to avoid repetition, reference is made to the above description relating to FIG. 1 to FIG. 3 with regard to the basic construction and functionality of the gas turbine engines 10 shown in FIG. 4 to FIG. 8.

[0098] FIG. 4 shows an enlarged partial longitudinal sectional view of an embodiment of the gas turbine engine 10 with a heat exchanger 42 which can be supplied with an annular air flow via an inlet 43. The inlet 43 is a so called ESS-style-inlet and surrounds the engine core 11 circumferentially. The heat exchanger 42 is designed as a so called matrix air-cooled oil cooler (MACOC) and can also be designed in another suitable way.

[0099] The heat exchanger 42 is arranged for cooling engine based oil using an air flow 50 which is passed from the fan 23 to the heat exchanger 42 and from the heat exchanger 42 into the bypass duct 22.

[0100] A flow passage 48 is delivering the air flow 50 from the annular ESS-style-inlet 43 of the flow passage 48 to the heat exchanger 42, and is returning the air flow 50 from the heat exchanger 42 through an annular outlet 49 in the core fairing 45 into the bypass duct 22. The ESS-style-inlet 43 of the flow passage 48 is located upstream the heat exchanger 42, and is located in the axially forward area 39 of the housing 46 between the radially inner annular wall 44 and the radially outer annular wall (core fairing) 45 of the housing 46.

[0101] The flow passage 48 is comprising guide vanes 51 upstream the heat exchanger 42. The guides vanes 51 are designed to straighten out the swirl air from the fan 23 and they also transmit load between an OGV casing 67 and the core duct (the gas path) outer wall 46, i.e. they provide a structural load path from the fan OGV 72 to core casings.

[0102] FIG. 5 shows an enlarged partial longitudinal sectional view of a further embodiment of the gas turbine engine 10 with the heat exchanger 42 which can be supplied with an air flow 53 via an annular inlet 54 located in the core duct (the gas path) outer wall 41. The inlet 54 is a so called IPC (Intermediate Pressure Compressor)-style-inlet and is extending circumferentially over the entire circumference of the engine core 11. The annular IPC-style-inlet 54, which may be positioned anywhere down the IPC, is also located downstream of the stationary supporting structure 24 which includes so called engine sector stators (ESS) of the gas turbine engine 10.

[0103] The flow passage 48 is delivering the air flow 53 from the annular IPC-style-inlet 54 of the flow passage 48 to the heat exchanger 42, and is returning the air flow 53 from the heat exchanger 42 through the annular outlet 49 in the core fairing 45 into the bypass duct 22.

[0104] FIG. 6 shows an enlarged partial longitudinal sectional view of a further embodiment of the gas turbine engine 10 with the heat exchanger 42 which can be supplied with an air flow 55 via an annular inlet 56 located in the bypass duct inner wall 45. The inlet 56 is a so called OGV-style-inlet and is arranged directly downstream of fan outer guide vanes 72 (OGV), which radially run through the bypass duct 22. Moreover, the annular OGV-style-inlet 56 can comprise a circumferential ram scoop that extends radially into the bypass duct 22 and guides the air flow 55 into the flow passage 48. An annular inner wall of the flow passage 48 is formed by the inner wall (core fairing) 45 of the bypass duct 22. Further, a radially outer annular wall 75, or an annular cooling duct outer wall respectively, of the flow passage 48 extends axially between the annular inlet 56 and the annular outlet 49 which both lead into the bypass duct 22. The annular cooling duct outer wall 75 is fixed to a radially outer area of the heat exchanger 42. In addition, the heat exchanger 42 is mounted on the bypass duct inner wall (core fairing) 45 and a firewall 58 is provided. The firewall 58 separates a front core zone 59 or flammable fluid front core zone respectively, in which the heat exchanger 42 is located, from a further zone 60.

[0105] FIG. 7 shows a schematic cross-sectional view of a possible configuration of the gas turbine engine 10 with, by way of example, several heat exchangers 42A to 42G extending circumferentially. The heat exchangers 42A to 42G are designed as PGB (Power Gear Box) coolers 42A, 42B, 42C, 42D, TBM cooler 42E, and aircraft generator coolers 42F, 42G, with some of the PGB coolers 42A, 42B preferably being designed as MACOC (Matrix Air Cooled Oil Cooler). The heat exchangers 42A to 42G provide cooling for multiple engine based oil systems and can be combined with each of the inlets 43, 54 or 56 described in more detail above. The number of heat exchangers and the type of heat exchangers may vary in different embodiments.

[0106] As can be seen in FIG. 6 and FIG. 7, annular inlet flow dividers 64 or annular inlet bifurcations or annular inlet wedges are provided between facing ends of the heat exchangers 42A to 42G that divert the air flow 50, 53 or 55 in direction to the heat exchangers 42A to 42G respectively.

[0107] The excess air from the annular inlets 43, 54 or 56 can be fed to other systems of the gas turbine engine 10. Inlets for these systems can be placed between the facing ends of the heat exchangers 42A bis 42G.

[0108] FIG. 8 shows a three-dimensional partial view of the embodiment of the gas turbine engine 10 according to FIG. 7 and FIG. 6. All cooling of turbo machinery components, the power gear box 30 and other engine and/or aircraft systems like a cabin blower is provided through a circumferential array of heat exchangers 42A to 42G arranged in the core fairing 45. Ejectors can be provided if they are required.

[0109] 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.