This specification is based upon and claims the benefit of UK Patent Application No. GB 2014423.4, filed on 14 Sep. 2020, which is hereby incorporated herein in its entirety.
The present disclosure relates to a combustion chamber for a gas turbine engine. In particular, it relates to the connection of combustion chamber segments within a gas turbine engine.
Typically, each gas turbine engine combustion chamber wall comprises an outer wall and an inner wall. The outer wall is either constructed from a fabricated sheet metal wall or a forged and machined wall. The inner wall comprises a plurality of cast metal tiles which are connected onto the outer wall using threaded studs, washers and nuts. The cast metal tiles either comprise a plurality of pedestals, projections or ribs on their outer-cooler-surface to provide convection cooling to the tiles. Alternatively, the cast metal tiles are provided with a plurality of apertures which extend form their outer-cooler-surfaces to their inner-hotter-surface to provide effusion cooling of the tiles. In both arrangements the cooling air is supplied through apertures in the outer wall into a space between the outer wall and inner wall. Another option to provide cooling is to use a cassette combustor architecture. Here the cassette combustor is an assembly consisting of segmented liner panels—named cassettes—that are manufactured by additive layer manufacturing and fastened together to form a continuous ring, each of which perform a function of the combustor assembly.
The conventional rear mounted combustor has a hoop-continuous combustion liner welded to a head at the front and to a hoop continuous support arm at the rear. The combustion liner can then be welded to a hoop continuous turbine interface ring at the rear—for a front mounted combustor. The compressor must be constructed so that it is robust. This is because the combustor is subject to ultimate load cases that affect the engine, such as a compressor surge or a combustor flame out. These events can result in a buckling of the hoop. The hoop rings can be clad with cast tiles to protect the liner from the thermal load of the combustion flame. However, the mating face that links the cassette to the combustor head needs to be tightly controlled so that the two faces are the same radius. Consequently, more bolts are required to seal the interface. During an engine surge, the pressure delta across the combustion walls increases momentarily. This pressure loads onto the combustion liners and the combustor head, which result in the combustor head experiencing a forward load. Consequently, the bolts are likely to fail, which makes this configuration difficult to resist surge loads. A similar effect occurs in the event of bird strikes as well.
Consequently, it is desired to have an improved construction to overcome the limitations of the prior art.
According to a first aspect of the disclosure there is provided a combustion chamber comprising a plurality of circumferentially arranged cassette segments coupled to a combustor head at one end and a wall section at the other end, each cassette segment extending the full length of the combustion chamber, and wherein the combustor head has an annular tongue structure on a mating surface, the tongue structure engages with a groove portion present in each of the cassettes so that when assembled the groove portions in the each of the plurality of cassette segments forms a substantially continuous groove.
This has the benefit that the combustor has a greater resistance to the loads applied to it during a flameout or a compressor surge.
The mating surface may be an axial mating surface. Thus, allowing for easier manufacture. The clearance between the tongue and groove portion along the circumferential length may be variable.
The combustor head may be provided with a hole, which aligns with a hole provided on an upper surface of the cassette segments for the insertion of a fastener. This allows the combustor head to be securely connected to the cassettes. The combustor head holes may be slotted allowing for thermal expansion of the combustion chamber. The fasteners may be aligned parallel with a centreline of the engine. The fastener may be aligned generally axially, but at an angle to a centreline of the engine. The combustor heads and the cassette segments may each have respective lugs through which a faster is inserted to connect the cassette segments and the combustor head. The lugs on the combustor head may be circumferentially spaced and align with a lug located at the centre of the cassette segments. The lugs on the combustor head may be circumferentially spaced and align with a lug located at the corners of the cassette segments. The lugs on a front edge of the cassette segment may be circumferentially aligned with the lugs on a rear edge of the cassette segment. The lugs may be discretely mounted on the combustor head and the cassette segments, so that they extend radially outwards the centre of the combustor. A single or double fastener may be used per lug.
A cowl may be provided to connect to the combustor head on an opposing side to the mating surface that connects with the cassette segments. The cowl may be provided with a plurality of scallops, or cut-backs, on both a radially outer axially extending flange and a radially inner axially extending flange. The cowl may be provided with holes, which are arranged to align with the holes provided in the combustor head and cassette segments. The cowl may be provided with an edge, so it radially sits outboard of the interface between the combustor head and Cassette segments. The cowl may be connected only to the combustor head via a separate lug to that connects the combustor head with the cassette segments.
According to a second aspect of the disclosure there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and 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, wherein: The turbine may be a first turbine, the compressor is a first compressor, and the core shaft may be a first core shaft; the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.
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.
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). 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: 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 or 390 cm (around 155 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).
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 250 cm to 300 cm (for example 250 cm to 280 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 320 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 1600 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/Utip, 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: 0.3, 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).
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, or 17. 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). 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).
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). 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). 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: 1400 K, 1450 K, 1500 K, 1550 K, 1600 K or 1650 K. 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: 1700 K, 1750 K, 1800 K, 1850 K, 1900 K, 1950 K or 2000 K. 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). 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 blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. 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 16, 18, 20, or 22 fan blades.
As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent.
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 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 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: a forward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of −55 degrees C.
As used anywhere herein, “cruise” or “cruise conditions” may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.
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.
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:
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 combustion chamber 16, as shown more clearly in
The annular combustion chamber 16 is positioned radially between a radially outer combustion chamber casing 110 and a radially inner combustion chamber casing 112. The radially inner combustion chamber casing 112 comprises a first, upstream, portion 112A, a second, intermediate, portion 112B and a third, downstream, portion 112C. The upstream end of the first portion 112A of the radially inner combustion chamber casing 112 is removably secured to the upstream end of the radially outer combustion chamber casing 110. In this example a flange at the upstream end of the first portion 112A of the radially inner combustion chamber casing 112 is removably secured to a flange at the upstream end of the radially outer combustion chamber casing 110 by suitable fasteners, e.g. nuts and bolts, passing through the flanges. The downstream end of the first portion 112A of the radially inner combustion chamber casing 112 is removably secured to the upstream end of the second portion 112B of the radially inner combustion chamber casing 112. In this example a flange at the upstream end of the second portion 112B of the radially inner combustion chamber casing 112 is removably secured to a flange at the downstream end of the first portion 112A of the radially inner combustion chamber casing 112 by suitable fasteners, e.g. nuts and bolts, passing through the flanges. The downstream end of the second portion 112B of the radially inner combustion chamber casing 112 is removably secured to the upstream end of the third portion 112C of the radially inner combustion chamber casing 112 and the downstream end of the third portion 112C of the radially inner combustion chamber casing 112 is removably secured to the radially inner ends of the turbine nozzle guide vanes 52. In this example a flange at the upstream end of the third portion 112C of the radially inner combustion chamber casing 112 is removably secured to a flange at the downstream end of the second portion 112B of the radially inner combustion chamber casing 112 by nuts and bolts passing through the flanges and flanges on the turbine nozzle guide vanes 52 are removably secured to a flange at the downstream end of the third portion 112C of the radially inner combustion chamber casing 112 by nuts and bolts passing through the flanges.
The first portion 112A of the radially inner combustion chamber casing 112 is generally frustoconical and extends radially inwardly and axially downstream from its upstream end to the radially outer ends of the compressor outlet guide vanes 39 and extends radially inwardly and axially downstream from the radially inner ends of the compressor outlet guide vanes 39 to its downstream end. The second portion 112B of the radially inner combustion chamber casing 112 is generally cylindrical. The third portion 112C of the radially inner combustion casing 112 is generally frustoconical and extends radially outwardly and axially downstream from its upstream end to the radially inner ends of the turbine nozzle guide vanes 52.
The upstream end wall 43 has an inner annular flange 43A extending in an axially downstream direction therefrom and an outer annular flange 43B extending in an axially downstream direction therefrom. The upstream end wall 43 forms a radially inner upstream ring structure and a radially outer upstream ring structure. A radially inner downstream ring structure 54 is mounted off the radially inner combustion chamber casing 112 and a radially outer downstream ring structure 56 is mounted off the radially outer combustion chamber casing 110. The radially inner annular wall structure 41 of the annular combustion chamber 16 and the radially outer annular wall structure 42 of the annular combustion chamber 16 comprise a plurality of circumferentially arranged combustion chamber segments 58 and 60 respectively. It is to be noted that the combustion chamber segments 58, 60 extend the full axial, longitudinal, length of the annular combustion chamber 16.
The circumferential arrangement of combustion chamber segments 58 and 60 of the radially inner and radially outer annular wall structures 41 and 42 of the annular combustion chamber 16 are clearly shown in
The outer wall 66 of each combustion chamber segment 58, 60 has at least one dilution aperture 100, the inner wall 66 of each combustion chamber segment 58, 60 has at least one dilution aperture 102 aligned with the corresponding dilution aperture 100 in the outer wall 64. At least one dilution wall 104 extends from the periphery of the corresponding dilution aperture 100 in the outer wall 64 to the periphery of the corresponding dilution aperture 102 in the inner wall 66. The inner wall 66 of each combustion chamber segment 58, 60 has at least one dilution chute 106, the at least one dilution chute 106 extends from the inner wall 66 in a radial direction away from the inner wall 66 and the outer wall 66 and each dilution chute 106 can be aligned with a corresponding one of the dilution apertures 104 in the inner wall 66, as shown in
If the combustion chamber is a lean burn combustion chamber the combustion chamber segments 58, 60 are not provided with dilution apertures, dilution walls and dilution chutes.
Each combustion chamber segment 58 and 60, as shown in
The upstream end of each combustion chamber segment 58, 60 is secured to the upstream ring structure and the downstream end of each combustion chamber segment is mounted on the downstream ring structure. Thus, the upstream end of each combustion chamber segment 58 is secured to the upstream ring structure, e.g. the upstream end wall structure, 44 and the downstream end of each combustion chamber segment 58 is mounted on the radially inner downstream ring structure, e.g. the radially inner discharge nozzle, 54. Similarly, the upstream end of each combustion chamber segment 60 is secured to the upstream ring structure, e.g. the upstream end wall structure, 44 and the downstream end of each combustion chamber segment 60 is mounted on the radially outer downstream ring structure, e.g. the radially outer discharge nozzle, 56.
The first hook 70 extends the length of the box like structure 62 between a securing arrangement and a mounting arrangement and the second hook 74 also extends the length of the box like structure 62 between the securing arrangement and the mounting arrangement. The securing arrangement and the mounting arrangement are discussed further below.
However, it may be possible for the first hook to extend the full length of the box like structure and for the second hook to extend the full length of the box like structure. Alternatively, it may be possible for the first hook to extend only a part of the full length of the box like structure and for the second hook to extend only a part of the full length of the box like structure. Additionally, it may be possible for there to be a plurality of first hooks arranged along the length of the box like structure and for there to be a number of second hooks arranged along the length of the box like structure.
The box like structure 62 of each combustion chamber segment 58, 60 has a first end wall 76 extending from a first, upstream, end of the outer wall 64 to a first, upstream, end of the inner wall 66, a second end wall 78 extending from a second, downstream and opposite, end of the outer wall 64 to a second, downstream and opposite, end of the inner wall 66. A first edge wall 80 extending from a first circumferential edge of the outer wall 64 to a first circumferential edge of the inner wall 66, a second edge wall 82 extending from a second, opposite circumferential, edge of the outer wall 64 to a second, opposite circumferential, edge of the inner wall 66 to form the box like structure 62.
The first and second edges 68 and 72 of the combustion chamber segments 58, 60 are axially profiled so that the at least some of the apertures 67 in the inner wall 66 direct coolant over at least a portion of one of the edges 68 and 72 of the combustion chamber segment 58, 60, as shown in
Alternatively, the first and second edges 68, 72 of the combustion chamber segments 58, 60 may extend with axial and circumferential components, as shown in
The box like structure 62 of each combustion chamber segment 58, 60 comprises a frame. The frame comprises the first and second end walls 76 and 78 and the first and second edge walls 80. The first and second end walls 76 and 78 and the first and second edge walls 80 are integral, e.g. one piece. The frame of each combustion chamber segment 58, 60 is radially thicker, and stiffer, than the outer wall 64 and the inner wall 66 and the first and second end walls 76 and 78 and the first and second edge walls 80 are thicker axially and thicker circumferentially respectively than the radial thickness of the outer and inner walls 64 and 66 in order to carry loads and interface with adjacent combustion chamber segments 58, 60 and the upstream ring structure and the downstream ring structure. The frame of each combustion chamber segment 58, 60 is arranged to carry the structural loads, the thermal loads, surge loads and flameout loads. The first hook 70 is provided on the first edge wall 80 and the second hook 74 is provided on the second edge wall. In other words, the box like structure 62 of each combustion chamber segment 58, 60 comprises the frame and portions of the outer and inner walls 64 and 66 extending axially, longitudinally, between the first and second end walls 76 and 78 and extending circumferentially, laterally, between the first and second edge walls 80.
The first and second edge walls 80 and 82 of the combustion chamber segments 58, 60 are arranged at a non-perpendicular angle to the outer wall 64 and/or the inner wall 66, as shown in
The first, upstream, end of the outer wall 64 of each combustion chamber segment 58, 60 has a flange 84 and the flange 84 has at least one locally thicker region 88, each locally thicker region 88 of the outer wall 64 has an aperture 92 extending there-through. The first, upstream, end of the inner wall 66 has a flange 86 and the flange 86 has at least one locally thicker region 90, each locally thicker region 90 of the inner wall 66 has an aperture 94 extending there-through. The at least one locally thicker region 88 at the first end of the outer wall 64 is arranged such that the aperture 92 is aligned with the aperture 94 through the corresponding locally thicker region 90 of the inner wall 66 and an annular slot 95 is formed between the flange 84 of the first end of the inner wall 66 and the flange 86 of the first end of the outer wall 66. The flange 84 at the first end of the outer wall 64 and the flange 86 at the first end of the inner wall 66 of each combustion chamber segment 58, 60 have a plurality of locally thickened regions 88, 90 respectively and the locally thicker regions 88, 90 are spaced apart circumferentially, laterally, between the first and second edges 68, 70 of the outer and inner walls 64 and 66 of the combustion chamber segments 58, 60. The aperture 94 in the at least one, or each, locally thickened region 90 of the inner wall 66 of each combustion chamber segment 58, 60 is threaded.
Each combustion chamber segment 58, 60 is secured to the upstream end wall structure 44 by one or more bolts 96. Each combustion chamber segment 58 is positioned such that the inner annular flange 44A of the upstream end wall structure 44 is located radially between the flanges 84 and 86 at the upstream end of the combustion segment 58 and such that the apertures 92 and 94 in the flanges 84 and 86 are aligned with a corresponding one of a plurality of circumferentially spaced apertures 45A in the flange 44A of the upstream end wall structure 44. Bolts 96 are inserted through the aligned apertures 92 and 45A and threaded into the apertures 94 to secure the combustion chamber segment 58 to the upstream end wall structure 44. Similarly, each combustion chamber segment 60 is positioned such that the inner annular flange 44B of the upstream end wall structure 44 is located radially between the flanges 84 and 86 at the upstream end of the combustion segment 60 and such that the apertures 92 and 94 in the flanges 84 and 86 are aligned with a corresponding one of a plurality of circumferentially spaced apertures 45B in the flange 44B of the upstream end wall structure 44. Bolts 96 are inserted through the aligned apertures 92 and 45A and threaded into the apertures 94 to secure the combustion chamber segment 60 to the upstream end wall structure 44. Alternatively, rivets may be inserted through the aligned apertures 92 and 45A and the apertures 94 to secure the combustion chamber segment 60 to the upstream end wall structure 44.
The cassettes 58 and 60 are provided with a slot/groove 150 for engagement shaped with the mating surface 152. The slot in the cassette is designed to fit a corresponding tongue 154 which is shaped in the combustor head. Abutting the rear of the combustor head may be a cowl. The cowl and the combustor head are provided with a hole. This hole corresponds with a tapped hole that is provided in an upper surface of the cassette, such that a bolt 156 can be inserted, as shown in
The tongue in the hoop may be machined into a continuous tongue. For example, the tongue may be turned, and if it is a broken tongue these further sections can be machined out. The combustor head may be made from a machined forging, or a semi-machined casting. The tongue may be machined on a casting. A continuous hoop may be used as it is able to form an unbroken connection with the cassette. The hoop is connected the grooves in the cassette. The cassette sections each have a slot machined into them. The slot in the cassettes when they are all assembled therefore forms a continuous slot into which the continuous tongue on the combustor head fits. Alternatively, the hoop section may be noncontinuous and can be arranged either to fit in a continuous groove or a non-continuous groove in which the cut outs are aligned with the portions of the non-continuous tongue. The downstream end of the groove may be overhanging during the manufacturing of the cassettes. It may have an arch geometry to allow it to be manufactured. The tongue or groove maybe configured to be parallel or tapered relative to a central axis. Alternatively, they tongue and/or groove may be dovetailed, involute, or have a compressible/crushable sealing feature integrated. The cassette may be manufactured using additive layer manufacture (ALM) as shown in
The presence of an axial mating face simplifies the manufacture of the cassette. The axial mating face may be produced by a wire electro discharge machine process to remove the cassette component from the build plate as shown in
Where the fasteners are positioned, they may be formed as part of a lug 160, which corresponds to a raised section of the combustor head. This is shown in
It is to be noted that the radially outer downstream ring structure 56 is a separate structure to the upstream end wall 44 and the radially inner downstream ring structure 54 is a separate structure to the upstream end wall, upstream ring structure.
A further benefit is that the combustion chamber loads are transmitted into the frame structure of the combustion chamber segments and not into the inner wall and/or outer wall of the combustion chamber segments. Load transmission from the frame of the cassette maybe augmented by stiffening the cassette panel or adding features such as ribs to the cold side.
An additional benefit is that the combustion chamber segments are removably secured to the corresponding downstream ring structure which allows the combustion chamber segments to be repaired or replaced. Thus, the combustion chamber segments may have a shorter working life than the corresponding downstream ring structure.
An advantage of the present disclosure is that the fasteners at the upstream ends of the combustion chamber segments radially and axially restrain the combustion chamber segments relative to the upstream end wall of the combustion chamber during normal operation and also during ultimate load situations, e.g. during compressor surge or combustion chamber flame out, when relatively high radial loads are exerted onto the combustion chamber segments tending to force the combustion chamber segments of the radially outer annular wall of the annular combustion chamber radially outwardly and to force the combustion chamber segments of the radially inner annular wall of the annular combustion chamber radially inwardly.
A further benefit is that the fasteners at the upstream ends of the combustion chamber segments allow the combustion chamber segments to be removed from the upstream end wall of the combustion chamber and replaced if the combustion chamber segments are damaged or to be repaired and reinserted into the combustion chamber.
Another benefit of the fastener arrangement is that there are low stresses in the portions of the combustion chamber segments which have cooling arrangements. Furthermore, the combination of radial and axial bolts allows accommodation of the different cassette lengths as determined by their build tolerances.
It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.
Number | Date | Country | Kind |
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2014423.4 | Sep 2020 | GB | national |