GB2587644A - Diffusion bonded vane - Google Patents

Abstract

A method of manufacturing a component 42 for an aircraft engine (which may be a vane), comprising; i) holding preforms 51-54 of different metal compositions (which may be different titanium alloys) in a target formation against each other; ii) applying a laser beam to weld the preforms together, and, iii) diffusion bonding the preforms together, wherein the laser beam is applied in vacuo. The preforms may be layered to form a sandwich structure. The method may further comprise the steps of; a) adding leachable material to the preforms such that the leachable material is trapped by the laser welded preforms, and, b) leaching out the leachable material after the preforms have been diffusion bonded. The preforms may be lap welded together. The method may comprise machining at least one of the preforms into a target shape before diffusion bonding, or, pressing at least one of the preforms into a target shape while diffusion bonding it to another preform. Further aspects are directed towards; a vane for an aircraft engine, a vane manufactured by a method according to the first aspect and an assembly comprising a gas turbine engine.

Description

DIFFUSION BONDED VANE

Field of the disclosure

The present disclosure relates to a method of manufacturing a component for an aircraft engine, a vane, an outlet guide vane assembly comprising the vane and an assembly comprising a gas turbine engine and the outlet guide vane assembly.

Background

A gas turbine engine such as a turbofan engine may comprise outlet guide vanes located behind the propulsive fan in a bypass duct of the gas turbine engine. The outlet guide vanes have two functions. An aerofoil profile of the outlet guide vane straightens airflow through the bypass duct to improve engine efficiency and therefore fuel consumption. The outlet guide vanes also act as structural components in order to transmit engine loads to the nacelle and casing of the gas turbine engine and so support that nacelle structure upon the core of the gas turbine engine.

Typically, outlet guide vanes are manufactured from sheet material, for example a titanium alloy such as Ti6A14V. The main structural factor is flutter margin which in turn is related to aerofoil curvature and its maximum chordal thickness.

Other alloy compositions can be chosen. Each alloy composition and grade has specific characteristics that provide advantages and disadvantages relative to other compositions. A particularly high quality alloy may be chosen so as to provide higher performance, with the disadvantage of increasing the material cost. An alloy may be chosen having high strength but may require a higher temperature for superplastic forming.

It is known for a vane to be manufactured by bonding two plates (or a folded plate) of material along edges and then superplastically deforming by inflation to create a hollow structure.

It is desirable to provide high performance of a component such as a vane without unduly increasing material costs.

Summary

According to a first aspect there is provided a method of manufacturing a component for an aircraft engine, the method comprising the steps of: holding preforms of different metal compositions in a target formation against each other; and applying heat via a laser beam so as to promote diffusion bonding between the preforms; wherein the laser beam is applied in vacuo.

Optionally, the method comprises removing laser welds formed by the application of the laser beam at the surface of the combined preforms.

Optionally, the method comprises processing at least one of the preforms to have a curved surface such that when it is laser welded to another of the preforms, a pocket is formed between the facing surfaces of the two preforms.

Optionally, more than two preforms are layered and diffusion bonded together to 25 form a sandwich structure. Optionally, the two outermost preforms are welded to each other but not to one or more preforms layered between the two outermost preforms.

Optionally, the method comprises adding a leachable material to the preforms such that the leachable material is trapped by the laser welded preforms; and leaching out the leachable material after the preforms have been diffusion bonded together.

Optionally, each metal composition comprises a titanium alloy. Optionally, the different metal compositions comprise different grades of titanium. Optionally, the different metal compositions have different strength and/or stiffness.

Optionally, the component is a vane that comprises a foot bulk part and an aerofoil part configured to be radially inward of the foot bulk part in an outlet guide vane assembly comprising the vane, wherein a preform corresponding to the foot bulk part has a metal composition that has a higher melting temperature than that of a preform corresponding to the aerofoil part.

Optionally, the vacuum has a pressure of at most 100kPa.

Optionally, the preforms are lap welded together by the application of the laser beam.

Optionally, the method comprises machining at least one of the preforms into a target shape before diffusion bonding it to another preform. Optionally, the method comprises pressing at least one of the preforms into a target shape while diffusion bonding it to another preform. Optionally, the preforms are near net shape.

Optionally, the method comprises the component is a vane and the method comprises: holding the vane and a vane foot in a target formation against each other, wherein the vane foot is configured to connect the vane to a ring of an outlet guide vane assembly; and applying heat via a laser beam to where the vane and the vane foot join so as to promote diffusion bonding between the vane and the vane foot, wherein the laser beam is applied in vacuo.

According to a second aspect there is provided a vane for an aircraft engine, the vane comprising a plurality of preforms of different metal compositions diffusion bonded together. Optionally, the vane is manufactured by the method described above.

According to a third aspect there is provided an outlet guide vane assembly comprising the vane described above.

According to a fourth aspect there is provided an assembly comprising a gas turbine engine, a nacelle and the outlet guide vane assembly described above.

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 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 U. 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/11,02, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip 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 (all units in this paragraph being Jkg-1K-1/(.ns-) ) 1.2%. 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, or 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, 01 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. 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: 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 95 Nkg-ls, 90 Nkg-ls, 85 Nkg-ls or 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 85 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 aluminium-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 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. 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 35000 ft (10668 m).

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.

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 a schematic view of a sandwich structure formed of preforms diffusion bonded together; Figure 5 shows different types of weld that may be formed; Figure 6 schematically shows a laser beam being used to join preforms together; Figure 7 schematically shows welds outside the target shape of the vane; Figure 8 shows a vane attached to a vane foot; Figure 9 shows a vane formed of a sandwich structure attached to a vane foot; Figure 10 shows an alternative vane formed by a sandwich structure attached to a vane foot; and Figure 11 shows an outlet guide vane assembly.

Detailed description

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.

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.

At least one outlet guide vane assembly 41 is located behind the fan 23 in the bypass duct 22. The outlet guide vane assembly 41 has two functions. An aerofoil profile of the outlet guide vane assembly 41 straightens air flow through the bypass duct 22 to improve engine efficiency and therefore fuel consumption. The outlet guide vane assembly 41 also acts as a structural component in order to transmit engine loads to the nacelle 21 and casing of the gas turbine engine 10 and so support that nacelle structure upon the engine core 11.

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.

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 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 exhaust nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area.

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.

In the description below, the invention will be described with reference to a vane 42 being the component that is produced. However, the invention can be applied to other components. In general, the invention uses laser welding in vacuum to permit diffusion bonding. This creates multi-metallic forms from primitive geometric precursors. An embodiment of the invention is expected to eliminate or at least reduce the need for expensive hand welding (which requires high expertise).

As mentioned above, at least one outlet guide vane assembly 41 may be located behind the fan 23 in the bypass duct 22. As shown in Figure 11, the outlet guide vane assembly 41 comprises a plurality of vanes 42 extending radially between an inner ring 71 and an outer ring 72. A method of manufacturing a vane 42 will be described below with reference to Figures 4 to 10.

Figure 4 schematically shows a sandwich structure 50 made of different preforms 51 to 54. Optionally, at least two of the preforms 51 to 54 are of different metal compositions. For example, the second preform 52 may have a different metal composition compared to that of the first preform 51 or the third preform 53. In one example, the preforms 52 to 54 have different alloy compositions. The alloy composition may be very similar to each other but with different titanium grades for the different preforms 52 to 54. Hence, the processing characteristics of the preforms are different.

Some of the preforms may have the same processing characteristics and the same metal composition. For example, the first preform 51 and the third preform 53 may have the same composition. Optionally, each metal composition comprises a titanium alloy. Optionally, the different metal compositions have a different strength rating from each other. Optionally, the different metal compositions have a different stiffness rating compared to each other. Optionally, the different metal compositions have different melting points relative to each other.

Optionally, the method of manufacturing the vane 42 comprises holding preforms 51, 52 of different metal compositions in a target formation against each other. In the example shown in Figure 4, the first preform 51 is positioned on top of the second preform 52. It is not necessary for the preforms to completely overlap in the manner shown in Figure 4. Other formations are possible.

The preforms 51, 52 etc. may be held together using assembly aids. For example, simple, rugged assembly aids can position, support and clamp the preforms.

Figure 6 schematically shows a laser beam 60 being applied to the preforms 51, 52. In the image of Figure 6, the first preform 51 is not shown so that the laser beam 60 can be clearly viewed. As shown in Figure 6, optionally the method of manufacturing the vane 42 comprises applying a laser beam 60. Optionally the laser beam 60 is applied to the join between two preforms 51, 52. However, this is not necessarily the case and the laser beam 60 could be applied to positions away from the join. As shown in Figure 6, the laser beam 60 may follow a path 61 along one or more preforms 51, 52.

Optionally, the application of the laser beam 60 produces a weld 55 (shown in Figure 5) at the surface of the combined preforms 51, 52. Figure 5 schematically shows different types of weld that may be formed. The left-most drawing in Figure 5 shows an end weld 55 (also known as an edge weld). The second drawing in Figure 5 shows a fillet weld 55. The third drawing shows a lap weld 55. The fourth drawing shows end welds 55.

Optionally, the preforms 51, 52 are lap welded together by the application of the laser beam 60. A lap weld 55 can be formed by simple laser manipulation, without challenging assembly tolerances. A lap weld 55 results in a laser weld that is under compression during the diffusion bonding process. The package is thus more securely sealed during the pressing process compared to other types of welds.

Although Figure 4 shows a sandwich structure 50 in which the preforms 51 to 54 are laid on top of each other in the same orientation, this is not necessarily the case. In an alternative arrangement, the preforms 51, 52 may be angled relative to each other before being bonded together. For example, assembly aids could be used to position, support and clamp angled ply assemblies.

By applying the laser beam 60, the preforms 51, 52 are welded together into a secure package. The laser beam creates a seal between the preforms 51, 52. The package of preforms 51, 52 can then undergo a diffusion bonding process. For example, optionally hot isostatic pressing (e.g. in a furnace filled with argon gas) is performed whereby diffusion bonding between the preforms 51, 52 is promoted. The preforms 51, 52 may bond to each other not just via welds 55 at the exterior surface of the combined preforms, but also by diffusion bonding at the surfaces of the preforms 51, 52 that are pressed against each other. Optionally, pressure is applied to press the preforms 51, 52 against each other during the diffusion bonding process.

Optionally, the laser beam 60 is applied in vacuo. By applying the laser beam 60 in a vacuum, the pressure inside the package that undergoes diffusion bonding is reduced. A packet comprising multiple different materials is diffusion bonded simultaneously. A plurality of diffusion bonds can be formed at the same time.

Additionally, the penetration depth of the laser beam 60 can be increased. Additionally, by applying the laser beam 60 in a vacuum, the weld geometry can be improved. Specifically, a weld that is deeper and more parallel can be achieved. In other words a similar desirable geometry to that formed when using electron beam welding can be achieved by applying a laser beam in a vacuum.

Of course, the vacuum may not be a perfect vacuum. Optionally, the vacuum has a pressure of at most 100 kPa. Optionally the vacuum has a pressure of at most 50 kPa. Optionally, at most 20 kPa, optionally at most 10 kPa and optionally at most 5 kPa.

Figure 7 schematically shows a stack of preforms 51, 52 bonded together and welds 55 at the surface of the combined preforms. Figure 7 also shows the desired shape of the vane 42 that is to be produced. As shown in Figure 7, some parts of the combined preform structure lies outside of the target shape for the vane 42 to be produced. These sections can be removed.

Optionally, the method comprises removing at least one laser weld 55 that was formed by the application of the laser beam 60 at the surface of the combined preforms. This helps to produce the vane 42 to have the desired shape. By using laser welds in the method of manufacturing the vane 42, the size of the heat-affected zone can be reduced (e.g. compared to a tungsten inert gas and electron beam weld). This helps to reduce the required amount of material because a smaller amount of material (corresponding to a smaller heat-affected zone) is required to be removed from the finished geometry.

Optionally, the method comprises processing at least one of the preforms 51 to have a curved surface such that when it is laser welded to another of the preforms 52, a pocket is formed between the facing surfaces of the two preforms 51, 52. It is desirable to have pockets in the package of preforms 51, 52 that is to undergo the diffusion bonding. Such pockets collapse during the diffusion bonding process. This provides a visual check that the process is performed correctly. Additionally, the presence of pockets reduces the possibility of unwanted media particles being trapped between the preforms 51, 52 during the vacuum.

Optionally, one or more preforms 51, 52 are hot pressed, stamped or machined to create the pockets. This processing is performed before the laser welds create the seal.

Optionally, the two outermost preforms 51, 54 are welded to each other but not to one or more preforms 52, 53 layered between the two outermost preforms 51, 54. The inner preforms 52, 53 and held in position by being trapped by the outer preforms 51, 54. The outer preforms 51, 54 are sealed together to form a packet that is filled by the inner preforms 52, 53. This can reduce the amount of laser beam welding required. It is not necessary to laser beam weld each adjacent pair of preforms. Instead only the outermost preforms 51, 54 may be welded and sealed together. For example, for a stack of four preforms, only one laser beam weld is made, instead of three laser beam welds. This reduces the burden of preparing and inspecting the laser beam welds.

Optionally, the inner preforms 52, 53 are smaller in plan view than the outer preforms 51, 54. This reduces the amount of the preforms that is machined away after the diffusion bonds have been formed.

Optionally, a leachable material is added to the preforms such that the leachable material is trapped by the laser welded preforms. Optionally, the leachable material has an ionic composition. Optionally, the leachable material comprises an ion of zinc, magnesium, aluminium or calcium. Optionally, the leachable material is a phosphate. For example, the leachable material may be sodium aluminium phosphate. Optionally, the leachable material is a ceramic casting core material.

The leachable material is trapped inside the package of preforms that is to be diffusion bonded together. Optionally, the leachable material can be leached out after the preforms have been diffusion bonded together. For example, the leachable material may be leached out after a hot isostatic pressing operation. This makes it easier for the diffusion bonded shape to be inflated. This makes it easier for more complex shapes to be produced.

Optionally, the leachable material is added to the package as pellets. Optionally, the pellets have a rounded shape. This reduces the possibility of the pellets damages the preforms and/or concentrating stresses in the preforms. Optionally, a hole in one of the preforms is provided for leaching out the leachable material. Optionally, the hole is perpendicular to the preforms. Optionally, the hole is rounded. This reduces the possibility of a stress concentration feature developing.

Optionally, a different material having a higher density than the preforms 51-54 is added to the package. The further material may be embedded in the package. This can help to control the position of the centre of mass of the component. Optionally, the further material may have a lower density than that of the preforms.

Optionally, the component may be inflated at the same time as undergoing the bot isostatic pressing. This can reduce the joints in the component, thereby reducing its 15 mass.

Figure 8 schematically shows a vane 42 made according to the invention. As shown in Figure 8, the vane 42 may be attached to a vane foot 43. The vane foot 43 is configured to connect the vane 42 to a ring of the outlet guide vane assembly 41.

For example, the vane foot 43 may be configured to connect the vane 42 to the inner ring 71 or to the outer ring 72 (shown in Figure 11).

As shown in Figure 8, the vane 42 comprises different sections 44 to 46 that may have different optimal physical requirements. Optionally, the vane 42 comprises a foot bulk part 44, 45. The foot bulk part 44, 45 is positioned close the vane foot 43.

The foot bulk part 44, 45 may be built up of multiple sheets (i.e. preforms). This is shown in more detail in Figure 9. In Figure 8, the section with the reference numeral 45 represents a preform of the foot bulk part that extends longitudinally further than that of the part with the reference numeral 44 They both form part of the foot bulk part 44, 45.

As shown in Figure 8, the vane 42 may further comprise an aerofoil part 46. The aerofoil part 46 is further away from the vane foot 43. The aerofoil part 46 may have an aerofoil profile to straighten air flow through the bypass duct 22. This is to improve engine efficiency and therefore reduce fuel consumption. In the example shown in Figure 8, the vane foot 43 is configured to connect the vane 42 to the outer ring 72. Hence, the foot bulk part 44, 45 is configured to be radially outward of the aerofoil part 46 in the outlet guide vane assembly 41. However, a corresponding foot bulk part may be provided at the radially inner side of the aerofoil part 46, near to another vane foot that connects the vane 42 to the inner ring 71 of the guide vane assembly 41.

Optionally, a preform corresponding to the foot bulk part 44, 45 has a metal composition that undergoes a superplastic phase change at a higher temperature or has a higher melting temperature than that of a preform corresponding to the aerofoil part 46. A low temperature forming alloy (i.e. that undergoes a superplastic phase change at a lower temperature) may be used for the aerofoil part 46. This is to improve the superplastic forming used to form the shape of the vane 42. Meanwhile, the foot bulk 44, 45 may not be required to have a low temperature forming alloy because it may not undergo superplastic forming to the same extent as the aerofoil part 46. Hence, the metal composition used for the foot bulk part 44, 45 may be chosen depending on the optimal characteristics for the foot bulk part 44, 45. For example, a stronger alloy may be used. The metal compositions may differ in terms of grain sizes.

Hence, different metal compositions can be used for different sections of the vane 42 so as to provide optimal processing at each part. Meanwhile, the cost of the materials for the vane 42 may not be unduly increased because a cheaper material can be used where it is not necessary to use the more expensive material. An embodiment of the invention is expected to be less wasteful in terms of discarding material which is trimmed and machined away. An embodiment of the invention is also expected to be less wasteful for material which is over-specified for the localised region on the vane 42. An embodiment of the invention is expected to enable an outlet guide vane assembly 41 to be optimised for weight and aerodynamic performance. This helps to reduce the specific fuel consumption of the gas turbine engine 10.

The present invention is particularly suited to the manufacturing of components of a gas turbine engine 10, such as outlet guide vane assemblies 41, which have geometries that make conventional welding difficult or especially time consuming.

The outlet guide vane assembly 41 has two functions. An aerofoil profile straightens airflow through the bypass duct to improve engine efficiency and therefore fuel consumption. The vanes 42 also act as structural components in order to transmit engine loads to the nacelle and casing of the gas turbine engine and so support that nacelle structure upon the core of the gas turbine engine. The outlet guide vane assembly 41 may comprise of the order of 50 vanes 42. Optionally, the vanes 42 have different curvatures depending on their position around the perimeter of the outlet guide vane assembly 41. For example, some vanes 42 may be made to be stiffer if its function is more structural. Other vanes 42 may be curved to be more aerodynamic if their function is more for straightening the airflow. The invention is expected to make it easier and cheaper to make such an outlet guide vane assembly 41.

Optionally, the method comprises holding the vane 42 and the vane foot 43 in a target formation against each other, e.g. the formation shown in Figure 8 or Figure 9.

Heat may be applied via the laser beam 60 to where the vane 42 and the vane 43 join so as to promote diffusion bonding between the vane 42 and the vane foot 43. The laser beam 60 may be applied in a vacuum environment.

Figure 9 is a closer up view of the foot bulk part of the vane 42 connected to the vane foot 43. In the example shown in Figure 9, the foot bulk part is made of the sandwich structure 50 shown in Figure 4, for example. A plurality (e.g. four) which preforms 51 to 54 are selected for their shape and metal compositions and are bonded together according to the invention. The sandwich structure 15 formed by the preforms 52 to 54 is welded to the vane foot 43.

As shown in Figure 9, optionally the vane foot 43 comprises one or more bosses 48. The bosses 48 can help to secure the vane 42 to the inner ring 71 and/or the outer ring 72 of the outlet guide vane assembly 41. As shown in Figure 9, optionally the vane foot 43 comprises a protrusion 47 on an opposite side of the vane foot 43 from the vane 42. As shown in Figure 9, optionally the protrusion 47 is formed of a plurality of preforms sandwiched and bonded together according to the invention. Assemblies can have varying degrees of form and sophistication of welding.

Figure 10 is a schematic view of the foot bulk part of another vane 42 attached to a vane foot 43. Only the differences in Figure 10 compared to Figure 9 are described below.

In the example shown in Figure 10, the third preform 53 is replaced by two other preforms 56, 57. The preforms 56, 57 may be bonded together. The preforms 56, 57 may have different metallic compositions. For example, the preform 57 that is further from the vane foot 43 may have a lower melting temperature, may be less stiff, and may be less strong.

Optionally, the method of manufacturing the vane 42 comprises machining at least one of the preforms 51 to 54 into a target shape before diffusion bonding it to another preform. Hence, optionally the preforms 51 to 54 are near net shape.

Additionally or alternatively, the method of manufacturing the vane 42 may comprise pressing at least one of the preforms 51 to 54 into a target shape while diffusion bonding it to another preform. For example, some sections might be pressed into shape by differential pressure against fixed tooling during the bonding cycle.

Optionally, a combination of these two ways of shaping the preforms may be implemented. For example, some preforms may be machined prior to the bonding process, while others may be pressed into shape during the bonding process.

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 (21)

1. CLAIMS1 A method of manufacturing a component (42) for an aircraft engine (10), the method comprising the steps of: holding preforms (51-54) of different metal compositions in a target formation against each other; applying a laser beam (60) so as to weld the preforms together; and diffusion bonding the preforms together; wherein the laser beam is applied in vacuo.
2. The method of claim 1, further comprising removing laser welds (55) formed by the application of the laser beam.
3. The method of claim 1 or 2, further comprising: processing at least one of the preforms to have a curved surface such that when it is laser welded to another of the preforms, a pocket is formed between the facing surfaces of the two preforms.
4. The method of any preceding claim, wherein more than two preforms are layered and diffusion bonded together to form a sandwich structure (50).
5. The method of claim 4, wherein the two outermost preforms are welded to each other but not to one or more preforms layered between the two outermost preforms.
6. The method of any preceding claim, further comprising: adding a leachable material to the preforms such that the leachable material is trapped by the laser welded preforms; and leaching out the leachable material after the preforms have been diffusion bonded together. 2. 4. 5.
7. 7. The method of any preceding claim, wherein each metal composition comprises a titanium alloy.
8. 8. The method of any preceding claim, wherein the different metal compositions comprise different grades of titanium.
9. 9 The method of any preceding claim, wherein the different metal compositions have different strength and/or stiffness.
10. 10.The method of any preceding claim, wherein the component is a vane (42) that comprises a foot bulk part (44, 45) and an aerofoil part (46) configured to be radially inward of the foot bulk part in an outlet guide vane assembly (41) comprising the vane, wherein a preform corresponding to the foot bulk part has a metal composition that has a higher melting temperature than that of a preform corresponding to the aerofoil part.
11. 11.The method of any preceding claim, wherein the vacuum has a pressure of at most 100kPa.
12. 12. The method of any preceding claim, wherein the preforms are lap welded together by the application of the laser beam.
13. 13. The method of any preceding claim, further comprising machining at least one of the preforms into a target shape before diffusion bonding it to another preform.
14. 14. The method of any preceding claim, further comprising pressing at least one of the preforms into a target shape while diffusion bonding it to another preform.
15. 15. The method of any preceding claim, wherein the preforms are near net shape.
16. 16.The method of any preceding claim, wherein the component is a vane (42) and the method comprises: holding the vane and a vane foot (43) in a target formation against each other, wherein the vane foot is configured to connect the vane to a ring (71, 72) of an outlet guide vane assembly (41); and applying a laser beam to where the vane and the vane foot join so as to weld the vane and the vane foot together, wherein the laser beam is applied in vacuo.
17. 17.A vane (42) for an aircraft engine (10), the vane comprising a plurality of preforms (51-54) of different metal compositions diffusion bonded together.
18. 18.A vane (42) for an aircraft engine (10) manufactured by the method of any one of claims 1 to 16.
19. 19. An outlet guide vane assembly (41) comprising the vane of claim 17 or 18. 15
20. 20. An assembly comprising: a gas turbine engine (10) for an aircraft that comprises: an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine (19) to the compressor (14); a fan (23) located upstream of the engine core (11), the fan (23) 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 (23) so as to drive the fan (23) at a lower rotational speed than the core shaft (26); a nacelle (21) surrounding the gas turbine engine (10) and defining a bypass duct (22); and the outlet guide vane assembly of claim 19 located in the bypass duct (22) and configured to straighten air flow through the bypass duct (22).
21. 21. The assembly of claim 20, 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 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.