Patent Application
20230184116
This specification is based upon and claims the benefit of priority from United Kingdom patent application GB 2118072.4 filed on Dec. 14, 2021, the entire contents of which is incorporated herein by reference.
The present disclosure relates to an aerofoil component of a rotor, the aerofoil component being formed of continuous fibre-reinforced polymer composite and having a pattern of Z-pins embedded therein.
Rotating aerofoil components formed of continuous fibre-reinforced polymer composites, such as gas turbine engine fan or compressor blades, wind turbine blades, helicopter rotor blades, propellors etc., are exposed to a high risk of delamination during operation due to, for example, bird-strike type events. Large amounts of delamination can lead to a loss of structural stiffness and potential catastrophic failure. Z-pins are conventionally used for increasing a delamination resistance of composite laminates, thus delaying or arresting the propagation of interlaminar cracks in the laminates. Typically, they are formed of metallic materials, carbon, carbon-fibre-reinforced resins, glass, ceramic materials, or composites based on such materials. The Z-pins extend across layers of the composite to bridge interlaminar cracks and oppose opening and sliding displacements, significantly increasing the energy necessary for delamination to progress. Depending on requirements, the Z-pins may be embedded in the composite so that they are perpendicular to the layers of the composite, or at an angle from this perpendicular direction. Either way, because Z-pins extend across layers, a Z-pin enters the wake of a growing interlaminar crack to exert a bridging action until it either ruptures or is completely pulled out from the embedding composite laminate. Thus, the damage resulting from in-service events affecting composite laminates and the components they form can be reduced, decreasing the risk of catastrophic failure.
United Kingdom patent application GB 2539249 A, United States patent application US 2018/0141287 and United States patent U.S. Ser. No. 10/307,973 disclose methods of inserting Z-pins into composite components, and United Kingdom patent application GB 2542893 A and United States patent U.S. Ser. No. 10/578,115 disclose composite components reinforced with Z-pins.
Conventionally, Z-pins are embedded in patterns which may require tens of thousands of Z-pins to realise their delamination-resistance function. Embedding this number of pins requires significant manufacturing time and increases the cost of each component. Furthermore, conventional patterns of Z-pins, despite including large numbers of Z-pins, do not always ensure reliable delamination-resistance performance.
Therefore, it is desirable to provide a composite material rotor aerofoil component having an enhanced and reliable delamination resistance achieved by a pattern of Z-pins embedded into the component structure, which pattern is both cost- and time-efficient to manufacture.
The present disclosure has been devised in light of the above considerations.
In a first aspect, the present disclosure provides an aerofoil component of a rotor, the aerofoil component being formed of continuous fibre-reinforced polymer composite created by curing laid up pre-preg layers extending in radial and chordal directions of the aerofoil component, the aerofoil component further including a plurality of Z-pins extending across the layers of the aerofoil component to pin the layers together, wherein: the Z-pins are arranged in a pattern forming a chevron on the pressure and/or suction surface of the aerofoil component, the chevron having a vertex and two arms extending at an angle from each other away from the vertex, and the chevron is pointed either (i) towards the radially inner root of the aerofoil component such that one of the arms extends radially outwardly from the vertex towards a leading edge of the aerofoil component, and the other arm extends radially outwardly from the vertex towards a trailing edge of the aerofoil component, or (ii) towards the radially outer tip of the aerofoil component such that one of the arms extends radially inwardly from the vertex towards a trailing edge of the aerofoil component, and the other arm extends radially inwardly from the vertex towards a leading edge of the aerofoil component.
Advantageously, the chevron pattern according to option (i) can guide a delamination initially travelling through the aerofoil component radially from its root to its tip towards the component's leading and trailing edges, away from the core of the blade. Keeping the core of the blade protected, even at the cost of possible delamination the peripheral edges of the blade, is important, as it is the core of the blade that forms the main mass, and therefore structural integrity of the blade. Furthermore, the stiffness of the blade core determines the lower order resonant frequencies of the blade, which much be kept as constant as possible, even after (for example) impact events such as bird strikes. The chevron pattern achieves this by providing a highly delamination-resistant barrier to continued delamination travel in the radial direction, but a less delamination-resistant barrier towards the leading and trailing edges. Similarly, a chevron pattern according to option (ii) can guide a delamination initially travelling through the aerofoil component radially from its tip to its root towards the component's leading and trailing edges. Furthermore, the chevron pattern can achieve this effect using substantially fewer Z-pins than conventional patterns.
In a second aspect, the present disclosure provides a method of forming the aerofoil component of the first aspect, the method including: providing laid up pre-preg layers of fibre-reinforced polymer precursor, the layers extending in radial and chordal directions of the aerofoil component; forming guide holes in the pre-preg layers extending across the planes of the layers, the guide holes forming a chevron pattern on the pressure and/or suction surface of the aerofoil component, the chevron having a vertex and two arms extending at an angle from each other away from the vertex, wherein the chevron is pointed either (i) towards the root of the aerofoil component such that one of the arms extends radially outwardly from the vertex towards a leading edge of the aerofoil component, and the other arm extends radially outwardly from the vertex towards a trailing edge of the aerofoil component, or (ii) towards the tip of the aerofoil component such that one of the arms extends radially inwardly from the vertex towards a trailing edge of the aerofoil component, the other arm extends radially inwardly from the vertex towards a leading edge of the aerofoil component; and inserting the Z-pins into the guide holes, for example using a hollow needle located in each guide hole as an insertion guide for each Z-pin.
The method may further include a subsequent step of curing (typically by applying heat and pressure) the pre-preg layers containing the inserted Z-pins.
Optional features of the disclosure will now be set out. These are applicable singly or in any combination with any aspect of the disclosure.
At least one of the arms of the chevron may extend fully to the leading edge or fully to the trailing edge. Advantageously, the delamination can be thus routed directly to the leading or trailing edges of the component, thereby reducing the risk of delamination within the central core of the aerofoil component to maintain acceptable structural integrity and vibration frequencies.
Both the arms of the chevron may extend fully to the leading edge and fully to the trailing edge.
The arms of the chevron may form an angle therebetween at the vertex of 20° or more, and preferably of 40° or more, or 60° or more.
The arms of the chevron may form an angle therebetween at the vertex of 160° or less, and preferably of 140° or less, or 120° or less.
The areal density of the Z-pins on the pressure and/or suction surface in the chevron may be 0.5% or more, and preferably of 1% or more, or 2% or more.
The areal density of the Z-pins on the pressure and/or suction surface in the chevron may be 10% or less, and preferably of 8% or less, or 6% or less.
The cross-sectional area of each Z-pin may be 0.008 mm2 or more, and preferably of 0.04 mm2 or more, or 0.1 mm2 or more.
The cross-sectional area of each Z-pin may be 7 mm2 or less, and preferably of 4 mm2 or less, or 2 mm2 or less.
The width of each of the arms on the pressure and/or suction surface of the chevron may be 1% or more of the component radial span, and preferably of 5% or more.
The width of each of the arms on the pressure and/or suction surface of the chevron may be 20% or less of the component radial span, and preferably of 15% or less.
The Z-pins may be arranged in a pattern forming plural nested chevrons on the pressure and/or suction surface of the aerofoil component. This pattern can provide additional protection against delamination growth through the component.
The aerofoil component may be a gas turbine fan blade, a gas turbine compressor blade, a wind turbine blade, a propellor blade, or a helicopter rotor blade.
In a third aspect, the present disclosure provides a gas turbine engine having a propulsive fan or a compressor including a circumferential row of blades, each blade being the aerofoil component of the first aspect.
As noted elsewhere herein, 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 values being dimensionless). 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.
A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured at least in part from a composite, for example an organic matrix composite, such as carbon fibre. However, 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 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 may be 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 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.
Embodiments illustrating the principles of the disclosure will now be discussed with reference to the accompanying 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. All documents mentioned in this text are incorporated herein by reference.
In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.
An exemplary arrangement for a geared fan gas turbine engine 10 is shown in
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
The modelled, simplified component 25 has a length of 290 mm and a width of 140 mm, with the clamped region being 30 mm wide. The thickness of the component is 12 mm at the root, tapering gradually to 9 mm at the tip by means of single-sided laminate tapering.
In
In
In
Although the chevrons of
More generally, the areal density of Z-pins in the chevrons 39 may be 0.5% or more and/or 10% or less. The cross-sectional area of the individual Z-pins in the patterns may be 008 mm2 or more and/or 7 mm2 or less. The arms of the chevron 39 may form an angle therebetween at the vertex of 20° or more and/or 160° or less. The width of each of the arms on the pressure and/or suction surface of the chevron may be 5 mm or more and/or 100 mm or less.
In these examples, the impacting object is modelled as a smooth particle hydrodynamic (SPH) projectile having an average mass of 71.6 g and an average impact velocity of 165 m/s. The region 41 of impact is in the top right corner of the pressure or suction surface 31, near the tip of the component. The model integrates a mass-based damping coefficient having a value of 0.4, and attributes interlamellar fracture toughnesses of 1.0 N/mm and 7.0 N/mm respectively to the unpinned and pinned regions of the surface 31. The growth of the delamination with time is indicated by the increasing extent of the coloured area across the surface 31, the delamination typically initiating at bottom left near the root 33. The level of delamination is indicated by colour-coding with no delamination (level 0) being uncoloured, fully delaminated (level 1) being magenta and 1, and levels in between being coloured according to the legends provided above the plots.
The chevron pattern and the double chevron pattern both display significantly improved delamination-resistance performance compared to the double strip pattern. Experiments and the computational modelling described above suggest that, relative to simple straight strips of Z-pins and other known patterns, arranging the Z-pins in chevrons can provide similar levels of delamination protection but using about 40% fewer Z-pins. Thus the chevrons enable substantial reductions in manufacturing times and costs of Z-pin protected, polymer composite aerofoil components.
The chevron pattern 39 in
The example versions of the component 25 of
In the example version of the component 25 of
In the example versions of the component 25 shown in
Versions with the arrangement of Z-pins shown in
The example arrangement shown in
As previously noted, it will be appreciated that further chevrons can be added to further increase the reliability by which the Z-pins can delay or arrest radial delamination propagation through the component. Such a plurality of chevrons can be formed from any combination of those described in the present disclosure, including chevrons where both arms extend towards the edges of the aerofoil component without reaching it, chevrons where one arm extends towards the edge of the aerofoil component without reaching it and the other arm reaches the edge of the aerofoil component, and chevrons where both arms reach the edge of the aerofoil component, according to the expected operating conditions, predicted failure modes, and/or desired design parameters of the user.
It is envisaged that the Z-pins forming the chevron patterns can be made of any suitable reinforcement material, for example carbon, glass, metallic materials, ceramic materials, plastic materials, or composites based on such materials.
Preferably, the aerofoil component 25 is formed by providing laid up pre-preg layers of fibre-reinforced polymer precursor, the layers extending in radial and chordal directions of the aerofoil component. Then, guide holes are formed in the pre-preg layers extending across the planes of the layers. The guide holes form a chevron 39 pattern on the pressure and/or suction surface 31 of the aerofoil component 25, such as the chevron patterns of
However, it is also envisaged that the Z-pins may be inserted into the laid up pre-preg layers via any suitable method, including for example ultrasonically assisted Z-pin insertion methods, employing an ultrasonic hammer or ultrasonic gun. Another option is to insert the Z-pins into the guide holes using a hollow needle located in each guide hole as an insertion guide for each Z-pin.
Subsequent to Z-pin insertion, the pre-preg layers containing the inserted Z-pins are cured to form the component.
The features disclosed in the foregoing description, or in the following claims, or in the accompanying drawings, expressed in their specific forms or in terms of a means for performing the disclosed function, or a method or process for obtaining the disclosed results, as appropriate, may, separately, or in any combination of such features, be utilised for realising the disclosure in diverse forms thereof.
While the disclosure has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the disclosure set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the disclosure.
For the avoidance of any doubt, any theoretical explanations provided herein are provided for the purposes of improving the understanding of a reader. The inventors do not wish to be bound by any of these theoretical explanations.
Any section headings used herein are for organizational purposes only and are not to be construed as limiting the subject matter described.
Throughout this specification, including the claims which follow, unless the context requires otherwise, the word “comprise” and “include”, and variations such as “comprises”, “comprising”, and “including” will be understood to imply the inclusion of a stated integer or step or group of integers or steps but not the exclusion of any other integer or step or group of integers or steps.
It must be noted that, as used in the specification and the appended claims, the singular forms “a,” “an,” and “the” include plural referents unless the context clearly dictates otherwise. Ranges may be expressed herein as from “about” one particular value, and/or to “about” another particular value. When such a range is expressed, another embodiment includes from the one particular value and/or to the other particular value. Similarly, when values are expressed as approximations, by the use of the antecedent “about,” it will be understood that the particular value forms another embodiment. The term “about” in relation to a numerical value is optional and means for example +/−10%.
It will be understood that the disclosure is not limited to the embodiments above —described and various modifications and improvements can be made without departing from the concepts described herein.
| Number | Date | Country | Kind |
|---|---|---|---|
| 2118072.4 | Dec 2021 | GB | national |
