Patent Application
20210270136
This specification is based upon and claims the benefit of priority from IN Patent Application Number 202011006225 filed on 13 Feb. 2020, and UK Patent Application Number 2004925.0 filed on 3 Apr. 2020, the entire contents of which are incorporated herein by reference.
The present disclosure is related to an aerofoil assembly and a method of reducing losses in the aerofoil assembly.
Gas turbine engines typically employ rows of blades on wheels/disks of a rotor assembly, which alternate with rows of stationary vanes on a stator or nozzle assembly. Axial and/or radial openings at an interface between rotating blades and stationary vanes can allow hot combustion gases to exit a hot gas path and enter an intervening wheel space between the rows.
To limit such incursion of hot gases, cooling air or purge air is often introduced into a wheel space between the rows. This purge air serves to cool components and spaces within the wheel spaces as well as providing a counter flow of cooling air to further restrict incursion of hot gases into the wheel space.
Nevertheless, conventional gas turbine engines exhibit a significant amount of purge air escape into the hot gas path. The consequent mixing of cooler purge air with hot gas results in large mixing losses, due not only to the differences in temperature but also due to the differences in flow direction or swirl of purge air and hot gases. Losses may also result from the formation of strong pressure side horseshoe vortices. Such losses may decrease the efficiency of the gas turbine engine.
In one aspect, an aerofoil assembly includes a platform and a plurality of aerofoils extending radially outward from the platform. The platform has a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. Each aerofoil has a leading edge proximal to the first edge and a trailing edge distal to the first edge. A pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface. A mid-pitch location is defined midway along the pitch spacing. The platform defines one or more recesses disposed between the leading edges of the plurality of aerofoils and the first edge. Each of the one or more recesses is disposed proximal to the mid-pitch location between adjacent aerofoils.
The recesses may provide an easier escape path for purge air to leak into a hot gas region with minimal interaction with the leading edges. Maintaining a gap or distance between a flow of purge air and the leading edges may mitigate the formation of pressure side horseshoe vortices. Therefore, the recesses may reduce secondary losses in a passage between adjacent aerofoils. Consequently, the recesses may improve an efficiency of the aerofoil assembly. Further, the recesses may result in weight reduction of the aerofoil assembly.
In some embodiments, each of the one or more recesses may have a maximum depth of between about 0.1% to about 6% of a maximum height of each aerofoil relative to the platform surface.
The maximum depth of each recess relative to the maximum height of the aerofoil may be chosen so as to minimise interaction between purge air and the leading edge of the aerofoil.
In some embodiments, the aerofoil assembly may include a plurality of blade segments disposed adjacent to each other. Each blade segment may include a pair of aerofoils from the plurality of aerofoils and a platform portion that forms part of the platform.
A number of the blade segments may be based on assembly requirements.
In some embodiments, the platform portion of each blade segment may define a recess from the one or more recesses between the pair of adjacent aerofoils.
A number of the recesses may depend upon a number of the aerofoils such that one recess is disposed between two adjacent aerofoils.
In some embodiments, a mid-pitch region, at least partly defined between adjacent aerofoils, may extend from the first edge to the second edge through the mid-pitch location. Each of the one or more recesses may be disposed on the mid-pitch region.
The location of each recess on the mid-pitch region may provide an optimal gap between the flow of purge air and the leading edges of adjacent aerofoils, thereby mitigating the formation of pressure side horseshoe vortices.
In some embodiments, each of the one or more recesses may include a first lobe and a second lobe adjoining the first lobe.
The first lobe and the second lobe of each recess may facilitate the flow of purge air.
In some embodiments, a minimum distance between the mid-pitch location and each of the one or more recesses may be between about 0% to about 70% of a distance between the first edge and the leading edge of each aerofoil.
The location of each recess relative to the mid-pitch location may be chosen so as to minimise secondary losses.
In some embodiments, a minimum distance between the first edge and each of the one or more recesses may be between about 0% to about 70% of the distance between the first edge and the leading edge of each aerofoil.
The location of each recess relative to the first edge may be chosen so as to minimise secondary losses.
In some embodiments, a minimum distance between each of the one or more recesses and the leading edge of each of the adjacent aerofoils may be between about 10% to about 60% of the pitch spacing between the leading edges of the adjacent aerofoils.
The location of each recess relative to the pitch spacing may be chosen so as to minimise secondary losses.
In some embodiments, each of the one or more recesses may extend from the first edge of the platform to the mid-pitch location.
The extent of each recess from the first edge of the platform to the mid-pitch location may minimise secondary losses.
In some embodiments, the aerofoil assembly may be a turbine blade assembly.
In one aspect, a gas turbine engine may include the aerofoil assembly.
In another aspect, a method of reducing losses in an aerofoil assembly is provided. The method includes providing a platform having a first edge, a second edge, and a platform surface disposed between the first edge and the second edge. The method further includes providing a plurality of aerofoils extending radially outward from the platform and disposed between the first edge and the second edge. Each aerofoil has a leading edge proximal to the first edge and a trailing edge distal to the first edge. A pitch spacing is defined between the leading edges of adjacent aerofoils along the platform surface. A mid-pitch location is defined midway along the pitch spacing. The method further includes forming one or more recesses disposed between the leading edges of the plurality of aerofoils and the first edge. Each of the one or more recesses is disposed proximal to the mid-pitch location between adjacent aerofoils.
The method may improve the stage efficiency of a turbine by mitigating the formation of pressure side horseshoe vortices. The method may also result in weight reduction of the turbine.
In some embodiments, each of the one or more recesses may be formed by removing material from the platform surface. In some other embodiments, each of the one or more recesses may be formed by casting. In other words, each of the one or more recesses may be a cast-in feature.
Each recess may be formed by any suitable process that is chosen based on ease of manufacture while maintaining strength of the aerofoil assembly.
In some embodiments, each of the one or more recesses may have a maximum depth of between about 0.1% to about 6% of a maximum height of each aerofoil relative to the platform surface.
In some embodiments, a mid-pitch region at least partly defined between adjacent aerofoils may extend from the first edge to the second edge through the mid-pitch location. Each of the one or more recesses may be disposed on the mid-pitch region.
In some embodiments, each of the one or more recesses may include a first lobe and a second lobe adjoining the first lobe.
In some embodiments, a minimum distance between the mid-pitch location and each of the one or more recesses may be between about 0% to about 70% of a distance between the first edge and the leading edge of each aerofoil.
In some embodiments, a minimum distance between the first edge and each of the one or more recesses may be between about 0% to about 70% of the distance between the first edge and the leading edge of each aerofoil.
In some embodiments, a minimum distance between each of the one or more recesses and the leading edge of each of the adjacent aerofoils may be between about 10% to about 60% of the pitch spacing between the leading edges of the adjacent aerofoils.
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 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−1s, 105 Nkg−1s, 100 Nkg−1s, 95 Nkg−1s, 90 Nkg−1s, 85 Nkg−1s or 80 Nkg−1s. 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−1s to 100 Nkg−1s, or 85 Nkg−1s to 95 Nkg−1s. 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 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 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
Referring to
The stator vanes 202 extend from a static wall 206. The blades 204 extend from a rotating wall 208. A wheel space 210 is defined between the static wall 206 and the rotating wall 208. In operation, cooling air or purge air 212 is introduced into the wheel space 210. Purge air 212 may cool components and spaces within the wheel space 210. Purge air 212 may be tapped from a compressor, for example, the low pressure compressor 14 and/or the high pressure compressor 15.
Further, a hot gas region 213 is defined between the stator vanes 202 and the blades 204. The hot gas region 213 receives hot gas 214. Purge air 212 may restrict incursion of hot gas 214 into the wheel space 210. Specifically, a flow of purge air 212 may be used to purge the wheel space 210 into the hot gas region 213 such that purge air 212 restricts hot gas 214 from flowing into the wheel space 210. Purge air 212 may therefore provide a rim seal flow in the aerofoil assembly 200.
An ejection of purge air 212 out of the wheel space 210 and interaction with hot gas 214 may result in a reduction of efficiency of the aerofoil assembly 200. The reduction of efficiency may be due to various types of losses, for example, mixing losses, penetration losses, secondary vortices, etc.
The aerofoil assembly 200 further includes a platform 302 having a first edge 304, a second edge 306, and a platform surface 308 disposed between the first edge 304 and the second edge 306. The first edge 304 faces the stator vanes 202. The platform surface 308 is the radially outward surface of the platform 302. Each blade 204 includes an aerofoil 310 extending radially outward from the platform 302 and disposed between the first edge 304 and the second edge 306. The aerofoil assembly 200 includes a plurality of such aerofoils 310 arranged in an array. Each aerofoil 310 includes a leading edge 312 proximal to the first edge 304 of the platform 302 and a trailing edge 314 distal to the first edge 304. Each aerofoil 310 defines a maximum height “HM” relative to the platform surface 308. The maximum height “HM” is the maximum radial height between the platform surface 308 and a tip 315 of the aerofoil 310. The maximum height “HM” may be defined between the platform surface 308 and the tip 315 of the aerofoil 310 adjacent to the trailing edge 314.
The platform 302 defines one or more recesses 316 (only one shown in
The aerofoil 310 further defines a chord length “CL” between the leading edge 312 and the trailing edge 314. The chord length “CL” is a length of a straight line connecting the leading and trailing edges 312, 314. In some embodiments, a minimum distance “D2” between the recess 316 and the leading edge 312 is 0 percent (%) to 5% of the chord length “CL”. The minimum distance “D2” may be a minimum distance between the boundary 317 of the recess 316 and the leading edge 312.
The recesses 316 may reduce losses due to purge air 212 and improve the efficiency of the aerofoil assembly 200, and hence the gas turbine engine 10.
The platform portion 404 of each blade segment 402 includes a pair of longitudinal edges 406. The longitudinal edge 406 of the platform portion 404 of each blade segment 402 is aligned with the longitudinal edge 406 of the platform portion 404 of the adjacent blade segment 402. In the illustrated embodiment of
The platform portions 404 together form the platform 302. The first edge 304 of the platform 302 may be formed together by first edge segments (not shown) of the blade segments 402. Similarly, the second edge 306 may be formed together by second edge segments (not shown) of the blade segments 402. The longitudinal edge 406 of one blade segment 402 may be joined to the longitudinal edge 406 of the adjacent blade segment 402 by various methods, for example, but not limited to, welding, brazing, mechanical fasteners, mechanical joints, or combinations thereof.
Referring to
In an embodiment, each of the one or more recesses 316 extends from the first edge 304 of the platform 302 to the mid-pitch location “PL”. In such a case, the first edge 304 includes a portion of the recess 316. Further, each of the minimum distances “D1”, “D2” and “D3” is zero.
In some embodiments, a minimum distance “D4”, “D5” between each of the one or more recesses 316 and the leading edge 312 of each of the adjacent aerofoils 304 is between about 10% to about 60% of the pitch spacing “PD” between the leading edges 312 of the adjacent aerofoils 304. The minimum distance “D4” may be defined between the recess 316 and one of the adjacent aerofoils 304 on one side of the recess 316. The minimum distance “D5” may be defined between the recess 316 and the other of the adjacent aerofoils 304 on another side of the recess 316. In some embodiments, the minimum distance “D4” is equal to the minimum distance “D5”. In some other embodiments, the minimum distance “D4” is different from the minimum distance “D5”. The minimum distance “D4” may be a minimum circumferentially projected distance from the boundary 317 of the recess 316 to the leading 312 of one of the adjacent aerofoils 304. Similarly, the minimum distance “D5” may be a minimum circumferentially projected distance from the boundary 317 of the recess 316 to the leading edge 312 of the other adjacent aerofoil 304.
Further, a mid-pitch region “PR” is at least partly defined between the adjacent aerofoils 310. The mid-pitch region “PR” extends from the first edge 304 to the second edge 306 of the platform 302. Further, the mid-pitch region “PR” extends through the mid-pitch location “PL”. The mid-pitch region “PR” may be a line that is a locus of mid-points between the adjacent aerofoils 310 on the platform surface 308. Specifically, the mid-pitch region “PR” may be a line that joins all mid-points between a pressure surface 318 of one aerofoil 310 and a suction surface 320 of the adjacent aerofoil 310 along the platform surface 308. The line may be straight, curved or a combination of both. Further, the mid-pitch region “PR” intersects the mid-pitch location “PL”.
The recess 316 includes a first lobe 322 and a second lobe 324 disposed adjoining the first lobe 322. The boundary 317 of the recess 316 may therefore define two curved regions that are joined by a pair of rounded regions. In some embodiments, an area of the first lobe 322 may be substantially equal to an area of the second lobe 324. In alternative embodiments, the area of the first lobe 322 may be different from the area of the second lobe 324. Each of the first lobe 322 and the second lobe 324 may have any suitable shape, for example, but not limited to, circular, elliptical, oval or any curved shape.
The shape of the recess 316 is exemplary in nature. In alternative embodiments, the shape of the recess 316 can be, for example, but not limited to, polygonal, oval, circular, elliptical, or any other suitable shape.
Each aerofoil 310 may be made of any suitable material such as a metal, a metal alloy, a ceramic, a composite, or combinations thereof. Each aerofoil 310 may include one or more channels for allowing flow of a cooling fluid.
The platform 302 may be made of any suitable material such as a metal, a metal alloy, a ceramic, a composite, or combinations thereof. The platform 302 may include one or more channels for allowing flow of a cooling fluid.
The recesses 316 described above may provide an easier escape path for purge air 212 to leak into the hot gas region 213 with minimal interaction with the respective leading edges 312. Maintaining a gap or distance between the flow of purge air 212 and the respective leading edges 312 may mitigate the formation of pressure side horseshoe vortices. Therefore, the recesses 316, may reduce secondary losses in a passage between respective adjacent aerofoils 310. Consequently, the recesses 316 may improve the efficiency of the aerofoil assembly 200. In some cases, the recesses 316 may improve a stage efficiency of a turbine by at least 0.1%, at least 0.2%, at least 0.5%, at least 1%, at least 2%, or at least 3%. The recesses 316 may also result in weight reduction of the aerofoil assembly 200.
At step 602, the method 600 includes providing the platform 302 having the first edge 304, the second edge 306, and the platform surface 308 disposed between the first edge 304 and the second edge 306.
At step 604, the method 600 includes providing the plurality of aerofoils 310 extending radially outward from the platform 302, and disposed between the first edge 304 and the second edge 306. Each aerofoil 310 has the leading edge 312 proximal to the first edge 304 and the trailing edge 314 distal to the first edge 304. The pitch spacing “PD” is defined between the leading edges 312 of adjacent aerofoils 310 along the platform surface 308. The mid-pitch location “PL” is defined midway along the pitch spacing “PD”.
At step 606, the method 600 further includes forming the one or more recesses 316 disposed between the leading edges 312 of the plurality of aerofoils 310 and the first edge 304. Each of the one or more recesses 316 is disposed proximal to the mid-pitch location “PL” between adjacent aerofoils 310.
In some embodiments, each of the one or more recesses 316 is formed by removing material from the platform surface 308. The material can be removed from the platform surface 308 by various material removal processes, for example, but not limited to, milling, drilling, grinding, electrical discharge machining, ultrasonic machining, abrasive jet machining, electron beam machining, or combinations thereof.
In some other embodiments, each of the one or more recesses 316 is formed by casting. In other words, each of the one or more recesses 316 is a cast-in feature.
In some embodiments, each of the one or more recesses 316 has the maximum depth “DM” of between about 0.1% to about 6% of the maximum height “HM” of each aerofoil 310 relative to the platform surface 308. In some embodiments, the mid-pitch region “PR” at least partly defined between adjacent aerofoils 310 extends from the first edge 304 to the second edge 306 through the mid-pitch location “PL”. Each of the one or more recesses 316 is disposed on the mid-pitch region “PR”.
In some embodiments, each of the one or more recesses 316 includes the first lobe 322 and the second lobe 324 adjoining the first lobe 322.
In some embodiments, the minimum distance “D3” between the mid-pitch location “PL” and the recess 316 is between about 0% to about 70% of the distance “AL” between the first edge 304 and the leading edge 312 of each aerofoil 310.
In some embodiments, the minimum distance “D1” between the first edge 304 and the recess 316 is between about 0% to about 70% of the distance “AL” between the first edge 304 and the leading edge 312 of each aerofoil 310.
In some embodiments, the minimum distance “D4”, “D5” between the recess 316 and the leading edge 312 of each of the adjacent aerofoils 304 is between about 10% to about 60% of the pitch spacing “PD” between the leading edges 312 of the adjacent aerofoils 304.
In some embodiments, the recess 316 extends from the first edge 304 of the platform 302 to the mid-pitch location “PL”.
The method 600 may improve the stage efficiency of a turbine by mitigating the formation of pressure side horseshoe vortices. The method 600 may also result in weight reduction of the turbine.
It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.
| Number | Date | Country | Kind |
|---|---|---|---|
| 202011006225 | Feb 2020 | IN | national |
| 2004925.0 | Apr 2020 | GB | national |
