Patent Application
20210254489
This specification is based upon and claims the benefit of priority from UK Patent Application Number 1916957.2 filed on 21 Nov. 2019, the entire contents of which are incorporated herein by reference.
The present disclosure concerns methods of forming wall structures on substrates, articles comprising wall structures formed on substrates, and gas turbine engines comprising such articles.
Wall structures are formed on substrates in many different contexts and for many different types of application.
For example, wall structures may be formed on support substrates in order to form sealing elements, such as abradable sealing elements as used in gas turbine engines. It is known to provide such abradable sealing elements in a sealing structure which surrounds aerofoil blades (and, in particular turbine blades), the sealing elements being abradable to enable tips of the aerofoil blades to wear away the sealing elements to optimal sizes and shapes without causing damage to the aerofoil blades.
Such abradable sealing elements may comprise one of more walls arranged to define open cells which may be filled with an abradable material. The one or more walls may be formed by machining (for example, by electro-discharge machining (EDM)) a sealing segment of the engine. Alternatively, the one or more walls may be formed by depositing material (for example by laser-weld deposition methods) onto a surface of a sealing segment of the engine. Wall formation through deposition of material may be used to retrofit abradable sealing elements to engines, for example following excessive wear of existing machined sealing elements.
According to a first aspect, there is provided a method of forming a wall structure on a substrate, the method comprising: (e.g. sequentially) depositing, by additive-layer, powder-fed, laser-weld deposition apparatus, a plurality of material layers overlying one another on the substrate to form the wall structure; wherein each material layer of the plurality of material layers has (a) a (e.g. maximum) layer thickness, measured in a direction locally perpendicular to a profile of the substrate, of no greater than about 350 μm and (b) a layer width, measured in a direction locally parallel to the profile of the substrate, of no greater than about 1200 μm.
The skilled person will appreciate that additive-layer, powder-fed, laser-weld deposition is a method of depositing a material on a substrate comprising supplying powdered material to the substrate (for example, by way of a nozzle) and melting the powdered material, as well as a portion of the substrate, by way of a laser beam to form a weld pool which solidifies to form a weld layer of material on the substrate. Additive-layer, powder-fed, laser-weld deposition is also known as laser weld deposition, laser metal deposition, laser blown powder, directed metal deposition and/or directed energy deposition.
The skilled person will also appreciate that a width of the wall structure, as measured in a direction locally parallel to the profile of the substrate, is determined (i.e. defined by) the layer width of each material layer forming the wall structure. Accordingly, the wall structure deposited by the method of the present invention may have a width (i.e. a wall width) of no greater than about 1200 μm.
The inventors have found that, when a wall structure having a width no greater than about 1200 μm is formed by depositing material layers having (e.g. maximum) layer thicknesses greater than about 350 μm, the mechanical properties of the wall structure are compromised. For example, such wall structures may exhibit a tendency to undergo brittle fracture, i.e. crack, for example when exposed to high temperatures. Such behaviour is particularly observable in wall structures formed by depositing material layers of superalloys, such as nickel-based superalloys, as are commonly used in high-temperature applications such as in components of gas turbine engines. However, the inventors have also found that when a wall structure having a width no greater than about 1200 μm is formed by depositing material layers having (e.g. maximum) layer thicknesses no greater than about 350 μm, the mechanical properties of the wall structure are improved. In particular, the high-temperature mechanical, thermophysical and/or tribological properties of the wall structure are improved. For example, the tendency of the wall structure to crack when exposed to high temperatures is reduced or even eliminated. Without wishing to be bound by theory, the inventors posit that such an improvement may be the result of a reduction in residual stresses and microcracking in the wall structure when the wall structure is formed from thinner material layers. Thinner material layers also enable wall structure properties or dimensions to be varied gently along the direction perpendicular to the substrate (for example, in the case of tapered wall structures), thereby further reducing the tendency to crack at high temperatures due to thermo-mechanical driven stresses. In any case, regardless of the mechanism by which the improvement occurs, the inventors have found that wall structures manufactured by the method of the present invention exhibit improved mechanical properties and a reduced tendency to crack. Components incorporating these wall structures may therefore also have enhanced usable lifespans before component maintenance or replacement is required.
It may be that the (e.g. maximum) layer thickness of each material layer of the plurality of material layers is no less than about 50 μm, for example, no less than about 100 μm. It may be that the (e.g. maximum) layer thickness of each material layer of the plurality of material layers is no greater than about 250 μm. Accordingly, it may be that the (e.g. maximum) layer thickness of each material layer of the plurality of material layers is from about 50 μm to about 350 μm, for example, from about 100 μm to about 350 μm, or from about 50 μm to about 250 μm, or from about 100 μm to about 250 μm.
It will be appreciated that the layer thickness of each of the material layers is the layer thickness as measured following deposition of the entire wall structure.
It may be that the layer width of each material layer of the plurality of material layers is no less than about 50 μm, for example, no less than about 100 μm, or no less than about 200 μm, or no less than about 300 μm, or no less than about 400 μm. It may be that the layer width of each material layer of the plurality of material layers is no greater than about 1000 μm, for example, no greater than about 900 μm, or no greater than about 800 μm, or no greater than about 700 μm, or no greater than about 600 μm. It may be that the layer width of each material layer of the plurality of material layers is from about 50 μm to about 1200 μm, for example, from about 50 μm to about 1000 μm, or from about 50 μm to about 900 μm, or from about 50 μm to about 800 μm, or from about 50 μm to about 700 μm, or from about 50 μm to about 600 μm, or from about 100 μm to about 1200 μm, or from about 100 μm to about 1000 μm, or from about 100 μm to about 900 μm, or from about 100 μm to about 800 μm, or from about 100 μm to about 700 μm, or from about 100 μm to about 600 μm, or from about 200 μm to about 1200 μm, or from about 200 μm to about 1000 μm, or from about 200 μm to about 900 μm, or from about 200 μm to about 800 μm, or from about 200 μm to about 700 μm, or from about 200 μm to about 600 μm, or from about 300 μm to about 1200 μm, or from about 300 μm to about 1000 μm, or from about 300 μm to about 900 μm, or from about 300 μm to about 800 μm, or from about 300 μm to about 700 μm, or from about 300 μm to about 600 μm, or from about 400 μm to about 1200 μm, or from about 400 μm to about 1000 μm, or from about 400 μm to about 900 μm, or from about 400 μm to about 800 μm, or from about 400 μm to about 700 μm, or from about 400 μm to about 600 μm.
It may be that the method comprises, during additive-layer, powder-fed, laser-weld deposition of the plurality of material layers, controlling one or more deposition parameters of the additive-layer, powder-fed, laser-weld deposition apparatus to deposit the plurality of material layers such that each material layer has (a) the layer thickness, measured in the direction locally perpendicular to a profile of the substrate, of no greater than about 350 μm and (b) the layer width, measured in the direction locally parallel to the profile of the substrate, of no greater than about 1200 μm.
It may be that the method comprises, during additive-layer, powder-fed, laser-weld deposition of the plurality of material layers, controlling a powder spot size to be from about 0.1 mm to about 3 mm, for example, from about 0.2 mm to about 0.5 mm. Additionally or alternatively, it may be that the method comprises, during additive-layer, powder-fed, laser-weld deposition of the plurality of material layers, controlling a laser spot size to be from about 50 μm to about 1000 μm, for example, from about 200 μm to about 600 μm. Additionally or alternatively, it may be that the method comprises, during additive-layer, powder-fed, laser-weld deposition of the plurality of material layers, controlling a laser scanning speed to be from about 400 mm/minute to about 2000 mm/minute, for example, from about 1000 mm/minute to about 1400 mm/minute. Additionally or alternatively, it may be that the method comprises, during additive-layer, powder-fed, laser-weld deposition of the plurality of material layers, controlling a powder feed rate to be from about 0.25 g/minute to about 10 g/minute, for example, from about 1 g/minute to about 3 g/minute. The inventors have found that such deposition parameters enable deposition of material layers having maximum thicknesses of no greater than about 350 μm, thereby reducing residual stresses and reducing the tendency of wall structures to crack.
The method may comprise changing the direction of travel of the additive-layer, powder-fed, laser-weld deposition apparatus (i.e. of the nozzle of the additive-layer, powder-fed, laser-weld deposition apparatus) and/or changing the direction of travel of the substrate between deposition of each nickel-based superalloy layer. For example, the method may comprise depositing sequential layers in alternating directions of travel. The inventors have found that alternating the direction of travel between layers assists in reducing the tendency of wall structures to crack.
It may be that the method comprises: varying one or more deposition parameters of the additive-layer, powder-fed, laser-weld deposition apparatus during deposition of the plurality of material layers such that two or more material layers of the plurality of material layers have different layer widths. For example, it may be that the method comprises: varying one or more deposition parameters of the additive-layer, powder-fed, laser-weld deposition apparatus during deposition of the plurality of material layers such that each material layer deposited has a layer width which is less than or equal to the layer width of any previously deposited material layer of the plurality of material layers (i.e. and wherein at least one, for example some, of the material layers have layer widths less than the layer width of a previously deposited material layer), thereby forming a wall structure having a tapered width profile along a direction locally perpendicular to the profile of the substrate. The method may further comprise varying the one or more deposition parameters such that a (e.g. maximum) difference in layer width between sequentially (i.e. consecutively) deposited layers is no greater than about 50 μm, for example, no greater than about 25 μm, or no greater than about 10 μm (thereby forming a wall structure having a gently tapered width profile).
The additive-layer, powder-fed, laser-weld deposition apparatus may comprise a nozzle for supplying powdered material and a laser for generating a laser beam to fuse the powdered material to form each material layer.
It may be that, during deposition by the additive-layer, powder-fed, laser-weld deposition apparatus, the profile of the substrate is inclined at an oblique angle with respect to the laser beam. The oblique angle may be no greater than about 45°, for example, no greater than about 30° or no greater than about 15°. The oblique angle may be no less than about 5°. The oblique angle may be from about 5° to about 45°, for example, from about 5° to about 30°, or from about 5° to about 15°. In embodiments in which the profile of the substrate is inclined at an oblique angle with respect to the laser beam, it may be that the method further comprises: adjusting the relative position of the substrate and the additive-layer, powder-fed, laser-weld deposition apparatus between deposition of each material layer of the plurality of material layers such that the wall structure formed extends substantially perpendicular to the profile of the substrate.
It may be that one or more material layer (e.g. each material layer) of the plurality of material layers comprises (e.g. is formed (e.g. entirely) from) superalloy. The superalloy may be a nickel-based superalloy. The nickel-based superalloy may comprise from about 50 wt. % to about 85 wt. % Ni, from about 2 wt. % to about 8 wt. Al, and the usual impurities. The nickel-based superalloy may further comprise: from about 2 wt. % to about 15 wt. % Co, from about 3 wt. % to about 10 wt. % Cr, from about 1 wt. % to about 7 wt. % W, up to about 5 wt. % Re, about 4 wt. % to about 8 wt. % Ta, up to about 1 wt. % Si, up to about 3 wt. % Hf, up to about 3 wt. % Mo, up to about 1 wt. % Fe, up to about 1 wt. % Ti, up to about 1 wt. % Cu, up to about 0.04 wt. % C, and/or up to about 0.03 wt. B.
It may be that the substrate comprises (e.g. is formed (e.g. entirely) from) superalloy. The superalloy may be a nickel-based superalloy. The nickel-based superalloy may be monocrystalline, i.e. single-crystal.
The method may comprise forming the wall structure at room temperature.
The substrate may be a gas turbine engine component. For example, the substrate may be a sealing structure for or of a gas turbine engine.
The wall structure may be a wall structure of a sealing element, for example an abradable sealing element. The wall structure may be one of a plurality of wall structures of the sealing element, for example, the abradable sealing element.
In a second aspect, there is provided an article manufactured by the method according to the first aspect.
The article may be a sealing element, for example a sealing element for a gas turbine engine. The sealing element may be an abradable sealing element. The wall structure may define at least one cell of the sealing element. The at least one cell may be fillable with an abradable material. The at least cell may be filled with an abradable material. The abradable material may comprise one or more ceramics, metals or intermetallic compounds. For example, the abradable material be selected from: yttria-stabilised zirconia (YSZ), alumina, a nickel-aluminium intermetallic compound such as nickel aluminide (Ni3Al), a nickel-aluminium (Ni—Al) alloy, or combinations thereof. The abradable material may be a sintered material.
In a third aspect, there is provided an article comprising a support and a wall structure extending from the support, wherein the wall structure has a wall width of no greater than about 1200 μm and a multi-layered microstructure, observed in cross-section in a plane locally perpendicular to a profile of the support, comprising a plurality of stacked weld layers, each of the weld layers having a layer thickness, measured in a stacking direction, of no greater than about 350 μm.
The multi-layered microstructure comprising the plurality of stacked weld layers may be observable by imaging a cross-section of the wall structure, for example in a Scanning Electron Microscope (SEM) or optical microscope (for example, following standard metallurgical sample processing (e.g. sectioning, grinding, polishing and etching)). Each weld layer may be visible as a region of relatively homogenous contrast in the wall structure. Adjacent weld layers may be separated in the SEM or optical image by weld interfaces visible as (generally linear) regions of differing (e.g. stronger) contrast between weld layers. The location of weld interfaces may also be identified by corresponding indentations on the external surfaces of the wall structures.
It will be appreciated that the layer thickness of a given weld layer is the distance, in the stacking direction, between weld interfaces bounding the said weld layer as measured in the thickest portion of the weld layer. That is to say, the layer thickness of a given weld layer is the maximum thickness (i.e. measured in the stacking direction) of the said weld layer. Accordingly, it may be that each of the weld layers has a maximum thickness, measured in the stacking direction, of no greater than about 350 μm.
The stacking direction may be substantially perpendicular to each of the weld interfaces in the stack.
It may be that the (i.e. maximum) layer thickness of each weld layer of the plurality of stacked weld layers is no less than about 50 μm, for example, no less than about 100 μm. It may be that the (i.e. maximum) layer thickness of each weld layer of the plurality of stacked weld layers is no greater than about 250 μm. Accordingly, the (i.e. maximum) layer thickness of each weld layer of the plurality of stacked weld layers may be from about 50 μm to about 350 μm, for example, from about 100 μm to about 350 μm, or from about 50 μm to about 250 μm, or from about 100 μm to about 250 μm.
It may be that the wall structure has a tapered width profile in the direction extending away from the support. A tapered wall structure width profile may be achieved when some or all of the weld layers in the stack have different layer widths. In particular, it may be that the width of the weld layers reduces with increasing distance from the support. A tapered width profile has been found to reduce occurrence of horizontal crack formation (i.e. formation of cracks parallel to the support profile) during use.
It may be that the tapered width profile of the tapered wall structure is a gently tapered width profile. A gently tapered width profile may be achieved when the difference in layer width between adjacent weld layers in a given stack is low. For example, it may be that the (e.g. maximum) difference in layer width between adjacent weld layers in the stack is no greater than about 50 μm, for example, no greater than about 25 μm, or no greater than about 10 μm. A gently tapered width profile has been found to be particularly effective in reducing the occurrence of horizontal crack formation during use.
It may be that (i.e. maximum) thickness of each weld layer in the stack varies with distance from the support. For example, it may be that the (i.e. maximum) thickness of one or more weld layers further away from the support is less than the (i.e. maximum) thickness of one or more weld layers closer to the support. It may be that the (i.e. maximum) thickness of the weld layers decreases with distance from the support.
It may be that the wall width of the wall structure is no less than about 50 μm, for example, no less than about 100 μm, or no less than about 200 μm, or no less than about 300 μm, or no less than about 400 μm. It may be that the wall width of the wall structure is no greater than about 1000 μm, for example, no greater than about 900 μm, or no greater than about 800 μm, or no greater than about 700 μm, or no greater than about 600 μm. It may be that the wall width of the wall structure is from about 50 μm to about 1200 μm, for example, from about 50 μm to about 1000 μm, or from about 50 μm to about 900 μm, or from about 50 μm to about 800 μm, or from about 50 μm to about 700 μm, or from about 50 μm to about 600 μm, or from about 100 μm to about 1200 μm, or from about 100 μm to about 1000 μm, or from about 100 μm to about 900 μm, or from about 100 μm to about 800 μm, or from about 100 μm to about 700 μm, or from about 100 μm to about 600 μm, or from about 200 μm to about 1200 μm, or from about 200 μm to about 1000 μm, or from about 200 μm to about 900 μm, or from about 200 μm to about 800 μm, or from about 200 μm to about 700 μm, or from about 200 μm to about 600 μm, or from about 300 μm to about 1200 μm, or from about 300 μm to about 1000 μm, or from about 300 μm to about 900 μm, or from about 300 μm to about 800 μm, or from about 300 μm to about 700 μm, or from about 300 μm to about 600 μm, or from about 400 μm to about 1200 μm, or from about 400 μm to about 1000 μm, or from about 400 μm to about 900 μm, or from about 400 μm to about 800 μm, or from about 400 μm to about 700 μm, or from about 400 μm to about 600 μm.
The multi-layered microstructure of each of the one or more wall structures may comprise at least three, for example, at least five, or at least ten, or at least twenty, or at least thirty, or at least forty, stacked weld layers.
As an alternative to a tapered width profile, it may be that the wall structure has a substantially uniform width profile in a direction extending away from the support. A substantially uniform wall structure width profile may be achieved when each of the weld layers in the stack has substantially the same layer width.
It may be that the wall structure is locally inclined at an oblique angle with respect to a profile of the support. The oblique angle may be no greater than about 45°, for example, no greater than about 30° or no greater than about 15°. The oblique angle may be no less than about 5°. The oblique angle may be from about 5° to about 45°, for example, from about 5° to about 30°, or from about 5° to about 15°. Alternatively, it may be that the wall structure extends in a direction which is substantially perpendicular to the profile of the support.
It may be that one or more (e.g. each) weld layer comprises (e.g. is formed (e.g. entirely) from) superalloy. The superalloy may be a nickel-based superalloy. The nickel-based superalloy may comprise from about 50 wt. % to about 85 wt. % Ni, from about 2 wt. % to about 8 wt. % Al, and the usual impurities. The nickel-based superalloy may further comprise: from about 2 wt. % to about 15 wt. % Co, from about 3 wt. % to about 10 wt. % Cr, from about 1 wt. % to about 7 wt. % W, up to about 5 wt. % Re, about 4 wt. % to about 8 wt. % Ta, up to about 1 wt. % Si, up to about 3 wt. Hf.
The article may be a sealing element. The sealing element may be an abradable sealing element.
It may be that the support comprises (e.g. is formed (e.g. entirely) from) superalloy. The superalloy may be a nickel-based superalloy. The nickel-based superalloy may be monocrystalline, i.e. single-crystal. The support may be a component for or of a gas turbine engine.
The wall structure may define at least one cell (e.g. at least one cell of the sealing element (e.g. the abradable sealing element)). The at least one cell may be fillable with an abradable material. The at least one cell may be filled with abradable material. The abradable material may comprise one or more ceramics, metals or intermetallic compounds. For example, the abradable material be selected from: yttria-stabilised zirconia (YSZ), alumina, a nickel-aluminium intermetallic compound such as nickel aluminide (Ni3Al), a nickel-aluminium (Ni—Al) alloy, or combinations thereof. The abradable material may be a sintered material.
In a fourth aspect, there is provided a gas turbine engine comprising an article according to the third or fourth aspects.
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 platform.
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), 350 cm, 360 cm (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 270 cm) 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/Utip2, 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/(ms−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, or 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−1 s, 105 Nkg−1 s, 100 Nkg−1 s, 95 Nkg−1 s, 90 Nkg−1 s, 85 Nkg−1 s or 80 Nkg−1 s. 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−1 s to 100 Nkg−1 s, or 85 Nkg−1 s to 95 Nkg−1 s. 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: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. 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 330 kN to 420 kN, for example 350 kN to 400 kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C. (ambient pressure 101.3 kPa, 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 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000 ft (11582 m). 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 50 kN to 65 kN) 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 described in relation to any one of the above aspects may be applied mutatis mutandis 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.
Embodiments will now be described by way of example only, with reference to the Figures, in which:
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.
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
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
The epicyclic gearbox 30 illustrated by way of example in
It will be appreciated that the arrangement shown in
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
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
An open cell structure 48 is formed integrally with the turbine seal segment 41 in the region of its radially inner surface 43. The open cell structure 48 includes upstanding walls 50 which project radially inwards. The walls 50 define therebetween a plurality of open cells 54, the open cells 54 having generally circumferential bases 52. The open cells 54 are able to receive and support an abradable material.
The skilled person will appreciate that many different wall arrangements, and therefore many different open cell shapes, are possible. For example,
Wall arrangements having few or no wall intersections or joins (for example, as seen in
Wall arrangements lacking enclosing lateral sidewalls (for example, where abradable material forms one or more lateral sides of the abradable sealing element, for example as seen in
The skilled person will further appreciate that different wall and cell dimensions are possible. For example, each upstanding wall may have a wall height (i.e. measured in a direction locally perpendicular to the substrate) of up to about 10 mm and a wall width (i.e. measured in a direction locally parallel to the substrate) of from about 50 μm to about 1.2 mm. Adjacent walls may, on average, be spaced apart from one another by about 1 mm to about 2.5 mm. Each wall may extend away from the substrate in a direction which is substantially locally perpendicular to the substrate. Alternatively, one or more of the walls may be angled with respect to the substrate. For example, one or more of the wall may extend away from the substrate in a direction which is locally inclined at an oblique angle with respect to the substrate. The skilled person will appreciate that the relative geometry of the walls and the substrate is defined locally because the substrate may not be flat; for example, the substrate may have a curved surface.
As discussed hereinabove, the upstanding walls are formed from superalloy. The skilled person will appreciate that any suitable superalloy known in the art may be used. However, the superalloy selected is commonly a nickel-based superalloy. Suitable nickel-based superalloys include CM186, Rene 142, Haynes 214, Amdry 955, IN792 and Haynes 282, among others. The nickel-based superalloy will commonly include at least Ni and Al. One particularly suitable nickel-based superalloy has the composition defined in Table 1.
The substrate is also made of a superalloy. The skilled person will appreciate that any suitable superalloy known in the art may be used. However, the superalloy selected is commonly a nickel-based superalloy. Suitable nickel-based superalloys include CMSX-3, MarM247, CMSX-4, MM002, C1023, IN713LC and CM186, among others. The substrate may be a monocrystalline (i.e. single-crystal) nickel-based superalloy.
The skilled person will also appreciate that any suitable abradable material known in the art may be used to fill the open cells. The abradable material may be a ceramic, a metal or an intermetallic compound. For example, the abradable material may be selected from: yttria-stabilised zirconia (YSZ), alumina, nickel aluminide (Ni3Al), or any combination thereof. The abradable material may be particulate material. The abradable material may be a sintered material such as a sintered powder. The abradable material may be a ceramic, a metal or an intermetallic compound following sintering. For example, the abradable material may be or comprise nickel aluminide, formed by sintering nickel-aluminium powder. The abradable material may be porous or may be formed from porous or hollow powder (such as porous YZS, porous alumina or hollow NiAl powder (e.g. Metco 2101ZB, Metco 2110 or Metco 2501 available from OC Oerlikon)).
Abradable sealing elements such as those illustrated in
The upstanding walls may be formed on the substrate by additive-layer, powder-fed, laser-weld deposition (also known as laser weld deposition, laser metal deposition, laser blown powder, directed metal deposition or directed energy deposition).
Prior to deposition of the upstanding walls, the laser beam 602 may be scanned across the surface of the substrate 604, along paths where the walls are to be deposited, in order to remove contaminants, such as oxide films, from the substrate.
Following deposition of the upstanding walls using the laser-weld apparatus, the walls may be shaped using any suitable machining method, such as electrical discharge machining (EDM). In particular, machining may be used to reduce the height of the walls, render the height of the walls more uniform across the substrate, or otherwise adjust the profile of the walls. The surfaces of the walls and the open cells defined therebetween may be coated, for example by nickel plating.
The open cells may then be filled with the abradable material in powder form. The abradable material may then be sintered inside the open cells. A layer of less dense sinter material (i.e. a friable layer of sinter material) may then be removed by, e.g., manual dress, EDM or machining.
In particular, the inventors have found that it is preferable for each weld layer to have a thickness t (measured in a direction locally perpendicular to the surface of the substrate on which the wall is deposited) which is no greater than about 350 μm. When the thickness t of the weld layers exceeds about 350 μm, the inventors have found that the walls have a tendency to crack (i.e. undergo brittle failure) in use. In contrast, when the thickness t of the weld layers is less than about 350 μm, the mechanical properties (and particularly the high-temperature mechanical properties) of the walls are greatly improved.
Without wishing to be bound by theory, the inventors posit that reducing the weld layer thickness reduces residual stresses that are built up during deposition of the wall structure, resulting in reduced microcracking during manufacture and an increased resistance to temperature-induced cracking when the walls are exposed to the high temperatures found in the hot sections of gas turbine engines. The inventors have also found that a minimum weld layer thickness t of about 50 μm further improves the mechanical properties of the walls formed therefrom.
It should be noted that the width w of each weld layer forming the wall structure need not be the same. For example,
The inventors have found that a weld-layer thickness from about 50 μm to about 350 μm, thereby achieving the desired microstructure and reduction in residual stresses, can be obtained using the following laser-weld deposition parameters (where preferred values are indicated in parentheses):
Laser power: 20 to 500 Watts;
Laser scanning speed: 400 to 2000 mm/minute (e.g. about 1200 mm/minute);
Laser spot size: 50 to 1000 μm (e.g. about 200 to about 600 μm);
Powder feed rate: 0.25 to 10 g/minute (e.g. about 1.4 to about 3 g/minute);
Powder spot size: 0.1 to 3 mm (e.g. about 0.2 to about 0.5 mm).
Travel direction: alternating between layers.
Such laser-weld deposition parameters also result in a lower heat input and consequently a further reduction in residual stresses and microcracking.
The above parameters were determined using a CO2 laser, for example of the type TR1750/380 (Wegmann-Baasel Laser GmbH), but any other suitable type of laser (such as a solid state or fibre laser) may be used. The laser may be operated in a pulse mode, for example having a pulse frequency of about 1 kHz. The powder may be supplied to the substrate using a carrier gas, such as argon, for example supplied at a gas flow rate of about 1 to 5 l/minute, with a nozzle gas flow rate of about 5 to 12 l/minute.
Although the above discussion relates to the deposition of nickel-based superalloy walls (i.e. using nickel-based superalloy powder in the laser-weld deposition process), the skilled person will appreciate that the methods described herein may be used to deposit walls of any type of material depositable by laser-weld deposition. In particular, the method may be used to deposit walls having a maximum wall thickness of about 1200 μm. The mechanical properties of such walls may be improved by controlling the weld layer thickness to be no greater than about 350 μm, and preferably from about 50 μm to about 350 μm.
The inventors have also developed a method for depositing walls (such as the upstanding walls of abradable sealing elements) on substrates using laser-weld deposition in circumstances where pre-existing neighbouring features on the substrate might otherwise be considered to hinder or prevent laser-weld deposition. For example,
The deposition methods discussed hereinabove may be used to form abradable seal elements on clean (i.e. new) substrates. Alternatively, these deposition methods may be used to form abradable seal elements on repurposed substrates. For example, these deposition methods may be used to retrofit abradable seal elements to existing gas turbine engine components, for example by depositing new abradable seal elements to the existing gas turbine engine components following removal of previous abradable seal elements (which may have been damaged in use).
As discussed hereinabove, abradable seal elements, such as those illustrated in
Following testing of multiple abradable seal designs, the inventors have found that the high-temperature performance of the seal depends at least in part on the relative amounts of superalloy and abradable material present. In particular, the inventors have found improved behaviour when nickel-based superalloy constitutes from about 10% to about 50% of the total volume of the cellular region of the abradable seal element. The inventors have also found even further improved behaviour when nickel-based superalloy constitutes from about 25% to about 35% of the total volume of the cellular region of the abradable seal element. In particular, optimisation of the relative amounts of superalloy and abradable material enables the required tribological, thermophysical and mechanical properties for an abradable seal to be achieved while ensuring that the deposited structures do not undergo cracking during use at high temperatures.
The skilled person will appreciate that the cellular region of the abradable seal element is the region of the abradable seal element formed by the superalloy upstanding walls and the abradable material filling the open cells defined therebetween. The total volume of the cellular region is the total geometric volume of the cellular region, i.e. as defined by the outer surfaces of bounding upstanding walls and an outer surface of the abradable material filling each open cell. For example,
Since the total area of the cellular region is formed by areas of superalloy wall and areas of abradable material, when the volume of superalloy is from about 10% to about 50% of the total volume of the cellular region, the volume of abradable material is generally about 50% to about 90% of the total volume of the cellular region. Similarly, when the volume of superalloy is from about 25% to about 35% of the total volume of the cellular region, the volume of abradable material is generally about 65 to about 75% of the total volume of the cellular region. That is to say, the volume of superalloy wall and the volume of abradable material generally adds up to the total volume of the cellular region.
The inventors have found that these optimal volume ratios of nickel-based superalloy to abradable material are independent of the geometry (i.e. the cell wall arrangement and cell shape) of the abradable seal element. These results are based on tests carried out for abradable seal elements in which the walls are formed from a nickel-based superalloy having the composition shown in Table 1 and an abradable material formed by sintering a hollow NiAl powder. However, the inventors have found that the optimal volume ratio is substantially independent of the particular nickel-based superalloy or abradable material chosen.
An abradable seal element was manufactured by (a) using additive-layer, powder-fed, laser-weld deposition to form a plurality of nickel-based superalloy wall structures defining open cells on a surface of a substrate and (b) sintering an abradable material within the open cells. The additive-layer, powder-fed, laser-weld deposition parameters were set such that the wall structures included thick weld layers having layer thicknesses greater than 350 μm and the dimensions of each wall structure were selected such that nickel-based superalloy constituted from about 10% to about 50 of the total volume of the abradable seal element following filling of the open cells with abradable material.
A section through the abradable seal element, in a plane perpendicular to the surface of the substrate, was ground and polished according to standard metallurgical sampling procedures and imaged in an optical microscope.
An abradable seal element was manufactured by (a) using additive-layer, powder-fed, laser-weld deposition to form a plurality of nickel-based superalloy wall structures defining open cells on a surface of a substrate and (b) sintering an abradable material within the open cells. The additive-layer, powder-fed, laser-weld deposition parameters were set such that each wall structure was formed from weld layers having layer thicknesses no greater than 350 μm and the dimensions of each wall structure were selected such that nickel-based superalloy constituted from about 10% to about 50% of the total volume of the abradable seal element following filling of the open cells with abradable material.
A section through the abradable seal element, in a plane perpendicular to the surface of the substrate, was ground, polished and etched according to standard metallurgical sampling procedures and imaged in an optical microscope.
| Number | Date | Country | Kind |
|---|---|---|---|
| 1916957.2 | Nov 2019 | GB | national |
