GB2595529A

GB2595529A - A gas turbine engine

Info

Publication number: GB2595529A
Application number: GB2010311.5A
Authority: GB
Inventors: Goulos Ioannis; G Macmanus David; T J Sheaf Christopher
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2020-05-28
Filing date: 2020-07-06
Publication date: 2021-12-01
Also published as: GB202010311D0

Abstract

A gas turbine engine 200 for an aircraft (300, figure 1) comprising a secondary air system terminating at a vent 212 on a rear-facing surface located radially inwards of a bypass flow. The vent (212, figure 9) is non-axisymmetric about a central longitudinal axis of the engine. The vent may extend a distance in a circumferential around the central longitudinal axis of the engine, wherein the vent has a height in a radial direction which is non uniform. The vent may comprise an inboard portion and an outboard portion, the inboard portion being disposed either side of the central longitudinal axis. A radial height of the inboard portion may be different to a radial height of the outboard portion. The vent may be disposed on a surface of a core cowl 210. The vent may be an exhaust vent. The secondary air system may comprise a ventilation system.

Description

A GAS TURBINE ENGINE

Field of the Disclosure

The present disclosure relates to a gas turbine engine for an aircraft. The disclosure relates to an aft portion for a gas turbine engine for an aircraft. The disclosure also relates to an aircraft comprising a gas turbine engine.

Background

It is known for a gas turbine engine for an aircraft to comprise a secondary air system that terminates at a vent in a rear-facing surface. For example, the secondary air system may be a ventilation system. The flow of air exiting the vent in the rear-facing surface can affect one or more other exhaust gas flows from the gas turbine engine, e.g. the exhaust gas flow from an engine core and the exhaust gas flow from a bypass duct.

Exhaust flows from a gas turbine engine may be affected, during operation, by aerodynamic effects arising from the gas turbine engine's relative proximity to a fuselage and/or a wing of an aircraft. Resulting effects can include, for example, non-symmetric exhaust flows and suppression effects on exhaust and vent flow. Overall engine efficiency and/or operability can be negatively affected.

zo Summary of the disclosure

A first aspect provides a gas turbine engine for an aircraft comprising: a secondary air system terminating at a vent on a rear-facing surface located radially inwards of a bypass flow; wherein the vent is non-axisymmetric about a central longitudinal axis of the engine.

Advantageously, having a non-axisymmetric vent provides a means for varying air flow characteristics through the vent, e.g. to adjust air flow to a desired uniformity or non-uniformity, as compared with a standard axisym metric vent.

The vent may extend a distance in a circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through at least 45° of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through at least 90° of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through more than 1800 of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through more than 270° of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through 3600 of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine, i.e. the vent may be annular.

The rear-facing surface may be orientated relative to the central longitudinal axis of the gas turbine engine at an angle of up to 90°, up to or at least 60°, up to or at least 45°, up to or at least 30°, up to or at least 20°, up to or at least 100 or at least 5°.

The vent may have a height in a radial direction from the central longitudinal axis of the gas turbine engine.

The height in the radial direction may be the radial distance between a vent inner annulus exit and a bypass nozzle after-body trailing edge. The vent inner annulus exit may be in the same axial plane as the bypass nozzle after-body trailing edge.

The vent inner annulus exit may define a substantially circular shape. The bypass nozzle after-body trailing edge may define a substantially circular shape. The circular shape of the vent inner annulus may have a smaller diameter than the circular shape of the bypass nozzle after-body trailing edge circular shape. The vent inner annulus exit may comprise a central axis located away from, e.g. parallel to, a central axis of the bypass nozzle after-body trailing edge.

The vent may have a maximum height in the radial direction and a minimum height in the radial direction. The maximum height may be up to 10 times greater than the minimum height, up to five times greater than the minimum height or up to three times greater than the minimum height. The maximum height may be up to 9, 8, 7, 6, 5, 4, 3, 2 or 1.5 times greater than the minimum height.

The maximum height and minimum height may be located at any suitable positions about the vent.

The vent may comprise an inboard portion and an outboard portion, the inboard portion and the outboard portion being disposed either side of the central longitudinal axis of the engine. When the gas turbine engine is installed on an aircraft, the inboard portion and the outboard portion may be disposed respectively inboard and outboard relative to a fuselage. The inboard portion may have a height, e.g. a maximum height and a minimum height, in the radial direction and the outboard portion may have a height, e.g. a maximum height and a minimum height, in the radial direction. The height in the radial direction of the inboard portion may be different from the height in the radial direction of the outboard portion.

The maximum height may be disposed at or near the or an outboard portion. The 15 minimum height may be disposed at or near the or an inboard portion. The maximum height may be disposed at or near the or an upper portion or a lower portion of the vent.

The vent may comprise at least two distinct regions comprising a maximum height. 20 For example, the vent may comprise an oval arranged such that a maximum height is disposed at or near opposing portions of the vent.

The vent may be disposed in a rear-facing surface of a core cowl. The vent may be disposed at any suitable position along the core cowl.

The rear-facing surface of the core cowl may have an upstream end and a downstream end and a core cowl length (Lem) defined as a distance from the upstream end to the downstream end. The vent may be located in the rear-facing surface at a distance (Lvent) from the upstream end, Lent being up to 95% of Lum.

Lvent may be up to 80% of L"wl, up to 70% of L"""1, up to 60% of Lo",,,, up to 50% of Lowl, up to 45% of LGOWi or up to 40% of I-cowl-[vent may be up to or at least 30% of Low', up to or at least 25% of Lcowi, up to or at least 20% of Low or up to or at least 15% of Low,. Lent may be at least 5% of Lowl. Lvent may be at least 10% of Low,. In an example, L"nt may be at least 10% of L",",, and up to 45% of Loom. In another example, Lent may be at least 10% of [cowl and up to 40% of Low,. In another example, Lent may be at least 10% of [cowl and up to 30% of Lcowl. In an example implementation, Lent may be approximately 20% of [cowl. A ratio of L",,t/L"""I may be up to 0.95. Lvenacowl may be up to 0.8, up to 0.7, up to 0.6, up to 0.5, up to 0.45 or up to 0.4. Lent/I-cowl may be up to or at least 0.3, up to or at least 0.25, up to or at least 0.2 or up to or at least 0.15. Lvent/Lem may be at least 0.05. Lvent/Lowi may be at least 0.1. In an example, Lvent/L",, may be at least 0.1 and up to 0.45. In another example, Lvenr/Lco,,, may be at least 0.1 and up to 0.4. In another example, Lvent/Lcowl may be at least 0.1 and up to 0.3.

The vent may be disposed downstream or upstream of a bypass duct throat. The vent may be disposed substantially in line with the bypass duct throat.

The secondary air system may comprise a ventilation system. The secondary air system may convey a flow of fluid, in use, from any one or more of a compression, combustion or exhaust region of the gas turbine engine.

The vent may be an exhaust vent.

The secondary air system may terminate at any number of vents in the rear-facing surface, e.g. the rear-facing surface of the core cowl. The secondary air system may terminate at more than one vent in the rear-facing surface, e.g. the rear-facing surface of the core cowl.

The gas turbine engine may have a relatively low specific thrust, e.g. a specific thrust of less than 150 Nkg-ls or less than 120 Nkg-ls, or less than 110 Nkg-ls, or less than 100 Nkg-ls, or less than 90 Nkg-ls, or less than 80 Nkg-ls, or less than 70 Nkg-ls, or less than 60 Nkg-ls, or less than 50 Nkg-ls, or less than 40 Nkg-ls.

The gas turbine engine may have any desired fan nozzle pressure ratio. The fan nozzle pressure ratio may be defined as the average total pressure downstream of the fan outlet guide vane divided by the ambient static pressure. The gas turbine engine may have a relatively low fan nozzle pressure ratio, e.g. 2.5 or less, at mid-cruise conditions. For instance, the fan nozzle pressure ratio may be approximately 2.3 or less at mid-cruise conditions. The fan nozzle pressure ratio may be approximately 2.2 at mid-cruise conditions. In some embodiments, the gas turbine engine may have a fan nozzle pressure ratio of between 1.9 and 2.3 at mid-cruise conditions.

The gas turbine engine may have any desired bypass ratio. The gas turbine engine may have a very high bypass ratio (VHBR) or an ultra-high bypass ratio (UHBR).

Another aspect provides an aft portion for a gas turbine engine, the aft portion comprising a vent on a rear-facing surface located so as to be radially inwards of a bypass flow when the aft portion is installed in the gas turbine engine, wherein, when installed in the gas turbine engine, the vent is non-axisymmetric about a central longitudinal axis of the engine.

The central longitudinal axis of the engine may be a principal rotation axis of the engine.

When installed in a gas turbine engine, the vent may extend a distance in a circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through at least 45° of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through at least 90° of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through more than 180° of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through more than 2700 of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine. The vent may extend through 360° of arc in the circumferential direction around the central longitudinal axis of the gas turbine engine, i.e. the vent may be annular.

When installed in a gas turbine engine, the rear-facing surface may be orientated relative to the central longitudinal axis of the gas turbine engine at an angle of up to 90°, up to or at least 600, up to or at least 45°, up to or at least 300, up to or at least 200, up to or at least 10° or at least 5°.

The vent may have a maximum height in the radial direction and a minimum height in the radial direction. The maximum height may be up to 10 times greater than the minimum height, up to five times greater than the minimum height or up to three times greater than the minimum height.

The maximum height may be up to 9, 8, 7" 4, 3, 2 or 1.5 times greater than the minimum height.

The vent may comprise an inboard portion and an outboard portion, the inboard portion and the outboard portion being disposed either side of the central longitudinal axis of the engine. When the aft portion is installed on an aircraft, the inboard portion and the outboard portion may be disposed respectively inboard and outboard relative to a fuselage. The inboard portion may have a height, e.g. a maximum height and a minimum height, in the radial direction and the outboard portion may have a height, e.g. a maximum height and a minimum height, in the radial direction.

The height in the radial direction of the inboard portion may be different from the height in the radial direction of the outboard portion.

The maximum height may be disposed at or near the or an outboard portion. The 5 minimum height may be disposed at or near the or an inboard portion. The maximum height may be disposed at or near the or an upper portion or a lower portion of the vent.

The vent may comprise at least two distinct regions comprising a maximum height. 10 For example, the vent may comprise an oval arranged such that a maximum height is disposed at or near opposing portions of the vent.

The rear-facing surface of the core cowl may have an upstream end and a downstream end and a core cowl length (Law) defined as a distance from the upstream end to the downstream end. The vent may be located in the rear-facing surface at a distance (Lvent) from the upstream end, Lent being up to 95% of Lami.

Lent may be up to 80% of Loom, up to 70% of Law, up to 60% of Low], up to 50% of up to 45% of LGOWI or up to 40% of Lova. Lvent may be up to or at least 30% of 1_,"1, up to or at least 25% of Lamb up to or at least 20% of Leowl or up to or at least 15% of Lcowi Lvent may be at least 5% of LG0,,,n Lvent may be at least 10% of 1-Cowl In an example, [vent may be at least 10% of Low, and up to 45% of In another example, Lvent may be at least 10% of [cowl and up to 40% of [Gem. In another example, Lvent may be at least 10% of [cowl and up to 30% of [cowl. In an example implementation, [vent may be approximately 20% of Loom. A ratio of Lventinnwi may be up to 0.95. Lvent/[cowl may be up to 0.8, up to 0.7, up to 0.6, up to 0.5, up to 0.45 or up to 0.4. Lvent/Lcowl may be up to or at least 0.3, up to or at least 0.25, up to or at least 0.2 or up to or at least 0.15. Lvent/L",, may be at least 0.05. Lvent/L"wi may be at least 0.1. In an example, Lventicow may be at least 0.1 and up to 0.45. In another example, [vent/cowl may be at least 0.1 and up to 0.4. In another example, Lvent/Lcowi may be at least 0.1 and up to 0.3.

When installed in a gas turbine engine, the vent may be disposed downstream or upstream of a bypass duct throat. The vent may be disposed substantially in line with the bypass duct throat.

When installed in a gas turbine engine, a secondary air system may terminate at the vent. The secondary air system may comprise a ventilation system. The secondary air system may convey a flow of fluid, in use, from any one or more of a compression, combustion or exhaust region of the gas turbine engine.

The vent may be an exhaust vent.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise 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. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used.

For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any plafform. The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the 11.

radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/U2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Ufip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, 01 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 150 Nkg-is, 130 Nkg-ls, 110 Nkes, 105 Nkg-ls, 100 Nkes, 95 Nkg-ls, Nkg-ls, 85 Nkg-ls or 80 Nkg-ls. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg-ls to 100 Nkg-ls, or Nkg-ls to 95 Nkg-ls. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an alum inium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint -in terms of time and/or distance-between top of climb and start of descent. Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide -in combination with any other engines on the aircraft -steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m).

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached.

Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

zo Brief description of the drawings

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 4 is an example of a portion of an aircraft comprising a gas turbine engine; Figure 5 is a side view of an exhaust flow distribution from a gas turbine engine; Figure 6 is a rear view of a mass-flux distribution from an axisymmetric vent in a rear-facing surface of a core cowl of the gas turbine engine of Figure 5; Figure 7 is a schematic cross-sectional side view of an upper portion of a gas turbine engine; Figure 8 is a close up of region A as shown in Figure 7; Figure 9 is a schematic end view of a gas turbine engine; and Figure 10 is a schematic view of an aircraft.

Detailed description

Figure 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.

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.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2. The low pressure turbine 19 (see Figure 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.

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.

The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 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 Figure 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.

The epicyclic gearbox 30 illustrated by way of example in Figures 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.

It will be appreciated that the arrangement shown in Figures 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 Figure 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 Figure 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 Figure 2.

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.

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

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 Figure 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 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.

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 Figure 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular.

Figure 4 shows a portion of an aircraft 50 comprising a fuselage 52, a wing 54 and a gas turbine engine 56 attached to an underside of the wing 54. The gas turbine engine 56 comprises an inboard side 58 and an outboard side 60. The inboard side 58 is located closer to the fuselage 52 than the outboard side 60.

When a gas turbine engine is installed on an aircraft, the presence of the wing and fuselage results in a non-uniform local static pressure field around the gas turbine engine (indicated by the shading in Figure 4). The non-uniform local static pressure field may have an adverse effect on the bypass and core airflows and, hence, overall engine performance. These adverse effects may be worse for gas turbine engines with close-coupled installations and lower specific thrust. For instance, for a given airframe, a gas turbine engine with a larger fan diameter may be coupled relatively closely to the wing and/or fuselage, as compared with a gas turbine engine with a smaller fan diameter.

The non-uniform local static pressure field also results in a non-uniform distribution of flow from a rear-facing vent (or vents) exhausting air from a secondary air system, as will be described below. This may adversely affect performance of the secondary air system, which may adversely affect the overall performance and/or operability of the gas turbine engine.

Figure 5 is a cross-sectional top view of an aft portion of another conventional, axisymmetric gas turbine engine 100 comprising an engine core 102, 102a and a nacelle 104, 104a. Air speed is indicated by the shading in Figure 5.

The nacelle 104, 104a is disposed radially outwards from the engine core 102, 102a. Located between the engine core 102, 102a and nacelle 104, 104a is a bypass duct 106, the bypass duct being arranged to convey a bypass flow 107, 107a. The engine core 102, 102a comprises a core engine nozzle 109, 109a disposed at an aft region arranged to convey a core engine nozzle flow 108, 108a. The bypass duct 106, 106a is configured to convey the bypass flow 107, 107a towards the aft portion of the gas turbine engine 100.

The gas turbine engine 100 comprises a core cowl 110, 110a. The core cowl 110, 110a is arranged radially outwards from the core engine nozzle flow 108, 108a and radially inwards of the bypass flow 107, 107a. Disposed upon a rear-facing surface of the core cowl 110, 110a is a vent 113, 113a. The vent 113, 113a is arranged to convey a flow of gas, a vent flow 112, 112a, from a secondary air system.

The opening of the vent 113, 113a is defined by a bypass nozzle after-body trailing edge 114, 114a and a vent inner annulus exit 116, 116a. The nacelle 104, 104a comprises a nacelle trailing edge 118, 118a disposed at a downstream end.

The gas turbine engine 100 has a bypass ratio greater than 10 and may be considered to be an ultra-high bypass ratio engine. In some embodiments, the gas turbine engine 100 may have a bypass ratio within the range of from 5 to 20.

The gas turbine engine 100 has a fan nozzle pressure ratio of 2.2 at mid-cruise conditions. In some embodiments, the gas turbine engine 100 may have a fan nozzle pressure ratio within the range from 1.9 to 2.6 at mid-cruise conditions.

During operation of the gas turbine engine 100, the vent flow 112, 112a interacts with the bypass flow 107, 107a. The impact on the local bypass flow 107, 107a may be significant enough to have an appreciable negative effect on the overall efficiency of the gas turbine engine 100.

Due to the gas turbine engine 100, in use, being disposed in close proximity to both a wing and a fuselage of an aircraft, the air speed of the vent flow 112, 112a is non-axisymmetric about a central longitudinal axis of the engine 100. This is apparent from the differences in air speed between the outboard vent flow 112 and the inboard vent flow 112a.

The interaction of the outboard vent flow 112 and the inboard vent flow 112a on the corresponding bypass flow 107, 107a affects the exhaust flow characteristics, one characteristic being efficiency of the bypass flow 107, 107a. It is apparent that the inboard vent flow 112a comprises a faster flow speed than the outboard vent flow 112. The non-uniformity of the vent flows 112, 112a result in non-uniform effects on the exhaust flow characteristics.

Figure 6 illustrates the mass flux distribution at the exit of the vent, including the mass flux of the inboard vent flow 112a and the outboard vent flow 112. It is apparent that the inboard vent flow 112a comprises a greater mass flux than the outboard vent flow 112. The non-uniformity of the mass flux of the vent flows 112, 112a result in non-uniform effects on the exhaust flow characteristics. Figure 7 illustrates a cross-sectional side view of an upper portion of another gas turbine engine 200 comprising an engine core 202 and a nacelle 204, and Figure 8 is a close up of region C as shown in Figure 7.

The nacelle 204 is disposed radially outwards from the engine core 202. Located between the engine core 202 and nacelle 204 is a bypass duct 206. The engine core 202 comprises a core engine nozzle 208 disposed at an aft region. The bypass duct 206 is configured to convey a flow of air towards an aft portion of the gas turbine engine 200.

The gas turbine engine 200 comprises a core cowl 210. The core cowl 210 is arranged radially outwards from the core engine nozzle 208 and radially inwards of the bypass duct 206. Disposed upon a rear-facing surface 220 of the core cowl 210 is a vent 212. The vent 212 is arranged to convey a flow of gas from a secondary air system. During operation of the gas turbine engine 200, the flow of gas dispelled from the vent 212 interacts with the flow of gas conveyed through the bypass duct 206.

The core cowl 210 comprises a rear-facing surface 220. The rear-facing surface 220 extends from an upstream end 222 to a downstream end 224. The vent 212 is defined by a bypass nozzle after-body trailing edge 214 and a vent inner annulus exit 216. The vent 212 extends in an arc extending, for example, at least 270° and up to 350° or 360° in a circumferential direction around a principal rotational axis of the gas turbine engine 200. For example, the vent 212 may extend in an arc extending around 300° in a circumferential direction around a principal rotational axis of the gas turbine engine 200. The nacelle 204 comprises a nacelle trailing edge 218 disposed radially outward of the vent 212.

The height of the vent 212 is defined by the bypass nozzle after-body trailing edge 214 and the vent inner annulus exit 216. The nacelle 204 comprises a nacelle trailing edge 218 and is radially outward of the vent 212. The core cowl 210 comprises a core aft trailing edge 220.

Figure 9 illustrates a schematic rear view of the gas turbine engine 200 comprising: the core aft trailing edge 220; the bypass nozzle after-body trailing edge 214; the nacelle trailing edge 218; the vent inner annulus exit 216; and the vent 212.

The vent 212 is operable to convey a flow of gas from a secondary air system arranged within the gas turbine engine 200. The height of the vent 212 in a radial direction is defined by the bypass nozzle after-body trailing edge 214 and the vent inner annulus exit 216, the bypass nozzle after-body trailing edge 214 being radially outwards from the vent inner annulus exit 216.

The vent 212 is non-axisymmetric. In the embodiment shown, the height of the vent 212 comprises a maximum height at an outboard side 222 of the gas turbine engine 200 and comprises a minimum height at an opposing inboard side 224 of the gas turbine engine 200. The maximum height of the vent 212 is approximately four times greater than the minimum height. In other embodiments, the maximum height of the vent 212 may be located at an inboard side 224. Providing a non-axisymmetric vent 212 allows the speed and mass flux of the gas flow from the vent 212 to be adjusted to a desired uniformity or non-uniformity, as compared with a conventional axisymmetric vent. Hence, the effects of the non-uniform local static pressure field due to installation of the engine may be modified, as compared with a conventional axisymmetric set-up. For instance, the effects of the non-uniform local static pressure field due to installation of the engine may be mitigated or countered to at least some extent.

The azimuthal variation in geometry due to the non-axisymmetric vent may also enable control of the local effective pressure ratio which affects the azimuthal massflow distribution through the vent from the secondary air system and the interaction between the vent flow and the bypass duct flow and the core flow.

Accordingly, for a given engine installation, overall engine performance may be improved, as compared with a conventional axisymmetric set-up.

Figure 10 shows an aircraft 300. The aircraft 300 comprises a fuselage 301, a first wing 302 and a second wing 303, the first wing 302 and the second wing 303 extending from opposite sides of the fuselage 301. A first gas turbine engine 304 is connected to an underside of the first wing 302. A second gas turbine engine 305 is connected to an underside of the second wing 303. One or both of the first gas turbine engine 304 and the second gas turbine engine 305 are a gas turbine engine as described herein.

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.

Claims

CLAIMS1. A gas turbine engine (10, 100, 200) for an aircraft (300) comprising: a secondary air system terminating at a vent (113, 113a, 212) on a rear-facing surface located radially inwards of a bypass flow; wherein the vent (113, 113a, 212) is non-axisymmetric about a central longitudinal axis of the engine.
2. The gas turbine engine (10, 100, 200) of claim 1, wherein the vent (113, 113a, 212) extends a distance in a circumferential direction around the central longitudinal axis of the engine, wherein the vent (113, 113a, 212) has a height in a radial direction which is non-uniform.
3. The gas turbine engine (10, 100, 200) as claimed in any previous claim, wherein the vent (113, 113a, 212) has a maximum height in a radial direction and a minimum height in a radial direction, the maximum height being up to ten times greater than the minimum height.
4. The gas turbine engine (10, 100, 200) as claimed in claim 4, wherein the maximum height is up to five times greater than the minimum height.
5. The gas turbine engine (10, 100, 200) as claimed in any previous claim, wherein the vent (113, 113a, 212) comprises an inboard portion and an outboard portion, the inboard portion and the outboard portion being disposed either side of the central longitudinal axis of the engine; wherein the inboard portion has a height in a radial direction and the outboard portion has a height in a radial direction, the height in the radial direction of the inboard portion being different from the height in the radial direction of the outboard portion.
6. The gas turbine engine (10, 100, 200) as claimed in any previous claim, wherein the vent (113, 113a, 212) is disposed upon a surface of a core cowl (110, 110a, 210).
7. The gas turbine engine (10, 100, 200) as claimed in any previous claim, wherein the vent (113, 113a, 212) is disposed downstream or upstream of a bypass duct throat.
8. The gas turbine engine (10, 100, 200) as claimed in any previous claim, wherein the vent (113, 113a, 212) is an exhaust vent.
9. The gas turbine engine (10, 100, 200) as claimed in any previous claim, wherein the secondary air system comprises a ventilation system. 10
10. A gas turbine engine (10, 100, 200) as claimed in any previous claim comprising: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.zo
11. The gas turbine engine (10, 100, 200) as claimed in claim 10, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); the engine core further comprises a second turbine (17), a second compressor (15), and a second core shaft (27) connecting the second turbine to the 25 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.
12. An aft portion for a gas turbine engine, the aft portion comprising a vent on a rear-facing surface located so as to be radially inwards of a bypass flow when the aft portion is installed in the gas turbine engine, wherein, when installed in the gas turbine engine, the vent is non-axisymmetric about a central longitudinal axis of the engine. An aft portion for a gas turbine engine (10, 100, 200), the aft portion comprising: a vent (113, 113a, 212) on a rear-facing surface located so as to be radially inwards of a bypass flow (107, 107a) when the aft portion is installed in the gas turbine engine; wherein, when installed in the gas turbine engine, the vent (113, 113a, 212) is non-axisymmetric about a central longitudinal axis of the engine.
13. An aircraft (300) comprising at least one gas turbine engine (10, 100, 200) as claimed in any one of claims 1 to 11.