Patent Publication Number: US-2023134139-A1

Title: Combustor arrangement

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority pursuant to 35 U.S.C. 119(a) to United Kingdom Application No. 2115850.6, filed Nov. 4, 2021, which application is incorporated herein by reference in its entirety. 
     BACKGROUND 
     Field of the Disclosure 
     The present disclosure relates to a combustor arrangement and a gas turbine engine including the combustor arrangement. 
     Description of the Related Art 
     A combustor arrangement associated with a gas turbine engine typically includes one or more fuel injectors for supplying fuel into a combustion chamber during an operation of the gas turbine engine. The fuel injector is generally disposed in an aperture formed in a meter panel of the combustor arrangement. Further, the combustor arrangement includes one or more seals. Each seal surrounds and abuts with a corresponding fuel injector. Conventional seals include a generally annular structure and a flared portion. The seal is a floating component that ensures a seal is maintained with the meter panel during the operation of the gas turbine engine. The seal is subjected to hot combustion gases requiring the seal to be cooled using a cooling fluid for achieving a desired temperature of the seal. Thus, the seal includes a number of cooling apertures defined therein. In some examples, the cooling fluid may be air received from a compressor. The cooling fluid may reduce the temperature of the seal and/or a temperature of the fuel injector. The cooling apertures generally extend angularly with respect to a central axis of the seal. 
     The current design of the cooling apertures introduces the cooling fluid with an axial velocity component and a radial velocity component. Accordingly, the cooling fluid may be directed angularly through the cooling apertures, such that the cooling fluid is directed towards the flared portion thereby providing an impingement cooling of the flared portion. Further, the cooling fluid may be turned radially outwards by the flared portion. Moreover, a leakage flow may be achieved between the seal and the fuel injector providing a secondary cooling mechanism. The secondary cooling mechanism may not be effective as the leakage flow may be negligible and may close-off during an operation of the gas turbine engine. 
     Further, conventional seals may not provide an integrated cooling/mixing solution. More particularly, the design of conventional seals may not provide sufficient cooling of the seal as well as the fuel injector. In some examples, insufficient cooling of the seal may cause deterioration of the seal. More particularly, the impingement cooling on the flared portion may not be effective in reducing the temperature of the seal sufficiently to prevent melting and burn-back of the flared portion. Melting of the flared portion may cause release of material that may be deposited onto one or more components of the combustor arrangement. The deposition of molten material may lead to a blockage of the cooling apertures, which may further increase the temperature of the seal and/or the fuel injector, thereby reducing component life. Furthermore, melting of the flared portion may also lead to exposure of the fuel injector to elevated temperatures resulting in premature erosion and cracking of the fuel injector in service. In some examples, the exposure of the fuel injector to hot combustion gases may result in erosion of a tip of the fuel injector. Further, an oxidation of the flared portion of the seal due to insufficient cooling may also increase emission levels of the combustion chamber. In some examples, oxidation of the seal may result in a change in a flame position, which may in turn expose a heatshield of the combustor arrangement to hot combustion gases. 
     In some examples, the cooling fluid exiting the seal may mix with the fuel and air mixture delivered by the fuel injector. In some examples, the cooling fluid may adversely affect the fuel and air mixture supplied by the fuel injector, which may in turn affect a combustion of the fuel and air mixture in the combustion chamber. For example, the cooling fluid may cause an undesirable increase in emission levels of the combustion chamber. 
     SUMMARY 
     In a first aspect, there is provided a combustor arrangement. The combustor arrangement includes a fuel injector. The combustor arrangement also includes a seal arranged around the fuel injector and having an upstream end and a downstream end. The seal includes an annular body at least partially abutting the fuel injector. The annular body extends circumferentially about a central axis and axially extends along the central axis from the upstream end to the downstream end. The annular body includes an inner surface facing the fuel injector. The inner surface axially extends from the upstream end to the downstream end. The annular body also includes an outer surface radially spaced apart from the inner surface relative to the central axis and facing away from the fuel injector. The outer surface axially extends from the upstream end to the downstream end. The annular body further includes a plurality of slots disposed on the inner surface and circumferentially spaced apart from each other relative to the central axis. Each slot axially extends at least partially from the downstream end to the upstream end. Each slot is disposed in fluid contact across its maximum length with the fuel injector. The seal also includes a flange radially extending from the outer surface of the annular body at the upstream end. 
     The seal associated with the combustor arrangement may include a flare-less design having the plurality of slots. The slots associated with the seal may be embodied as substantially full-length slots. Further, a cooling fluid flowing through the substantially full-length slots may allow cooling of the seal and the fuel injector. Moreover, the spent cooling fluid may be then delivered locally to a fuel spray cone of the fuel injector which may in turn provide a benefit to engine emissions in terms of smoke. The seal described herein may provide improved cooling of the seal and the fuel injector. Further, the seal may reduce smoke emissions, thereby addressing issues associated with seal durability and engine emission certifications. The slots are disposed on the inner surface of the seal which may enable high levels of convective heat transfer to take place along a length of the seal and may provide a means to effectively cool hot portions of the seal. Further, the cooling fluid flowing through the slots may also enhance heat transfer at an outer diameter of the fuel injector. Moreover, a placement of the slots on the inner surface may also generate a radial positive pressure at an interface between the seal and the outer diameter of the fuel injector, which may in turn drive the cooling fluid by pressure. This feature may also improve durability of the seal which may prevent rapid oxidisation of the seal. In some examples, a pressure differential across the seal may create an aerodynamic bearing, which may reduce contact load and operational wear of the seal and/or the fuel injector. 
     Moreover, as mentioned above, the design of the seal described herein may address engine emission issues by realising an interaction of the cooling fluid with a fuel spray cone. Specifically, the cooling fluid may exit the slots in close proximity to the fuel injector which may allow control over a fuel spray cone angle. For example, by introducing the cooling fluid locally to the fuel spray cone, the fuel spray cone angle may be narrowed, which may in turn reduce a residence time of combustion processes in a primary zone (i.e., a front section of a combustion chamber). The primary zone may be primarily responsible for production of smoke emissions. Further, the seal described herein may be compact and lightweight. Furthermore, the seal may be manufactured via a range of processes, such as, casting, electrical discharge machining (EDM), grinding, additive layer manufacturing (ALM), broaching etc., providing supply chain flexibility. Moreover, the seal may be retrofitted in existing combustor arrangements without any changes to the combustor arrangement. 
     In some embodiments, each slot extends along a slot axis. The slot axis of each slot is parallel to or circumferentially angled relative to the central axis by an oblique angle. When the slot axis of each slot is circumferentially angled relative to the central axis, the slots may include an increased length which may in turn increase convective heat transfer between the cooling fluid and one or more portions of the seal and/or the fuel injector. Moreover, this feature may also be used to control an interaction of the spent cooling fluid with the fuel spray cone, i.e., co-swirling for reduced interaction and counter swirling for maximum interaction. 
     In some embodiments, an angular width of each slot with respect to the central axis is uniform or variable along the slot axis. 
     In some embodiments, the angular width of each slot progressively increases or decreases from the upstream end to the downstream end. More particularly, in some cases, the slot may have a converging cross-section that may accelerate a flow of the cooling fluid at an exit of the seal. Such a feature may ensure increased heat transfer in high temperature regions of the seal and/or the fuel injector. 
     In some embodiments, each slot has a maximum radial height along a radial direction with respect to the central axis. The maximum radial height of each slot is at most 90% of a maximum radial thickness of the annular body from the inner surface to the outer surface. 
     In some embodiments, each slot has a maximum angular width along a circumferential direction with respect to the central axis. The maximum angular width is equal to or different from the maximum radial height. 
     In some embodiments, the maximum angular width is greater than the maximum radial height by a factor of less than or equal to 90. 
     In some embodiments, the maximum angular width is less than the maximum radial height. 
     In some embodiments, each slot has a maximum slot length along the central axis. The maximum slot length of each slot is at least 90% of a maximum axial length of the annular body from the upstream end to the downstream end. Accordingly, the slots extending axially along a major portion of the maximum axial length may allow increase in a rate of heat transfer between the cooling fluid and one or more portions of the seal or the fuel injector. 
     In some embodiments, the maximum slot length of each slot is less than the maximum axial length of the annular body. 
     In some embodiments, the maximum slot length of each slot is substantially equal to the maximum axial length of the annular body. This feature may allow increase in a rate of heat transfer between the cooling fluid and one or more portions of the seal or the fuel injector. 
     In some embodiments, each slot extends from an upstream slot end proximal to the upstream end of the seal to a downstream slot end disposed at the downstream end of the seal. The upstream slot end is axially spaced apart from the upstream end of the seal with respect to the central axis. 
     In some embodiments, the inner surface at least partially contacts the fuel injector. Each slot radially extends from the inner surface partially towards the outer surface. 
     In some embodiments, the annular body further includes a plurality of slot walls. Each slot wall defines a corresponding slot from the plurality of slots. The slot wall includes at least one of a plurality of recesses and a plurality of projections. The recesses and/or the projections may increase a rate of heat transfer between the cooling fluid and one or more portions of the seal. 
     In some embodiments, the annular body further includes a plurality of wall portions extending from the inner surface and circumferentially spaced apart from each other relative to the central axis. Each wall portion at least partially contacts the fuel injector. Each slot is defined by the inner surface and a pair of corresponding adjacent wall portions from the plurality of wall portions. 
     In some embodiments, each slot has a maximum angular extent relative to the central axis. The maximum angular extent is less than 120 degrees. 
     In some embodiments, the plurality of slots includes at least three slots. 
     In some embodiments, each slot has a cross-sectional shape that is at least one of semi-circular, rectangular, concave, square, and trapezoidal. In some examples, the cross-sectional shape may include a curved portion to aid in manufacturing of the seal and reduce contact wear from sharp edges. Moreover, the cross-sectional shape may include a high aspect ratio which may increase heat transfer areas within the slots. 
     In some embodiments, the outer surface of the annular body has a substantially uniform outer diameter along the central axis. Thus, the seal described herein may eliminate a flared portion associated with conventional seals thereby reducing a complexity of manufacturing the seal. Moreover, an absence of the flared portion may reduce a possibility of material accumulation at an exit of the slots due to oxidation and/or melting of the flared portion as in conventional seals. 
     In some embodiments, the combustor arrangement further includes a meter panel having a cold side and a hot side. The meter panel has an aperture extending through the meter panel between the hot and cold sides. The seal is sized to fit through the aperture of the meter panel, such that the upstream end of the seal is proximal to the cold side of the meter panel and the downstream end of the seal is proximal to the hot side of the meter panel. Each slot is configured to receive the cooling fluid at the upstream end of the seal and discharge the cooling fluid at the downstream end of the seal, such that the cooling fluid contacts the fuel injector while flowing through each slot 
     In a second aspect, there is provided a gas turbine engine comprising the combustor arrangement of the first aspect. 
     The present disclosure may generally relate to seals associated with the combustor arrangement of gas turbine engines. 
     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 the 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, the second compressor, and the 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 the 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 the 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, or 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, or for example in the range of from 1400 rpm to 1800 rpm. 
     In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity U tip . 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/U tip   2 , where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and U tip  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 -1 K -   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, or 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e., the values may form upper or lower bounds), for example, in the range of from 50 to 70. 
     Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg -1 s, 105 Nkg -1 s, 100 Nkg -1 s, 95 Nkg -1 s, 90 Nkg -1 s, 85 Nkg -1 s, or80 Nkg -1 s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e., the values may form upper or lower bounds), for example, in the range of from 80 Nkg -1 s to 100 Nkg -1 s, or 85 Nkg -1 s to 95 Nkg -1 s. Such engines may be particularly efficient in comparison with conventional gas turbine engines. 
     A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160 kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, 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° C. (ambient pressure 101.3 kPa, temperature 30° 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), for example, in the range of from 1800 K to 1950 K. The maximum TET may occur, for example, at a high thrust condition, or 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, or 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. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Embodiments will now be described by way of example only, with reference to the Figures, in which: 
         FIG.  1    is a sectional side view of a gas turbine engine; 
         FIG.  2    is a close-up sectional side view of an upstream portion of the gas turbine engine of  FIG.  1   ; 
         FIG.  3    is a partially cut-away view of a gearbox for the gas turbine engine of  FIG.  1   ; 
         FIG.  4    is a sectional side view of a combustor arrangement associated with the gas turbine engine of  FIG.  1    according to an embodiment of the present disclosure; 
         FIG.  5    illustrates a partial perspective view of a seal and a fuel injector associated with the combustor arrangement of  FIG.  4    according to an embodiment of the present disclosure; 
         FIG.  6    illustrates a perspective view of the seal of  FIG.  5    according to an embodiment of the present disclosure; 
         FIG.  7    is a schematic side view of a portion of the seal of  FIG.  6    according to an embodiment of the present disclosure; 
         FIG.  8    illustrates a perspective view of a portion of the seal of  FIG.  6    according to an embodiment of the present disclosure; 
         FIG.  9 A  is a schematic view illustrating a slot associated with the seal of  FIG.  6    wherein an angular width of the slot progressively increases from an upstream end of the seal to a downstream end of the seal according to an embodiment of the present disclosure; 
         FIG.  9 B  is a schematic view illustrating a slot associated with the seal of  FIG.  6    wherein an angular width of the slot progressively decreases from the upstream end of the seal to the downstream end of the seal according to an embodiment of the present disclosure 
         FIGS.  10 A,  10 B, and  10 C  illustrate respective schematic views of different longitudinal configurations of the slots of the seal of  FIG.  6    according to various embodiments of the present disclosure; 
         FIG.  11 A  is a schematic view illustrating a slot having a number of recesses according to an embodiment of the present disclosure; 
         FIGS.  11 B to  11 D  are respective schematic views illustrating slots having different types of projections according to various embodiments of the present disclosure; 
         FIGS.  12 A to  12 E  are respective schematic views illustrating slots having different cross-sectional shapes according to various embodiments of the present disclosure; and 
         FIG.  13    illustrates a perspective view of a seal associated with the combustor arrangement of  FIG.  4    according to another embodiment of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION OF THE DISCLOSURE 
     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. 
       FIG.  1    illustrates a gas turbine engine  10  having a principal rotational axis  9 . The engine  10  comprises an air intake  12  and a propulsive fan  23  that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine  10  comprises a core  11  that receives the core airflow A. The engine core  11  comprises, in axial flow series, a low-pressure compressor  14 , a high-pressure compressor  15 , a combustion equipment  16 , a high-pressure turbine  17 , a low pressure turbine  19 , and a core exhaust nozzle  20 . A nacelle  21  surrounds the gas turbine engine  10  and defines a bypass duct  22  and a bypass exhaust nozzle  18 . The bypass airflow B flows through the bypass duct  22 . The fan  23  is attached to and driven by the low pressure turbine  19  via a shaft  26  and an epicyclic gearbox  30 . 
     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  FIG.  2   . The low pressure turbine  19  (see  FIG.  1   ) drives the shaft  26 , which is coupled to a sun wheel, or sun gear,  28  of the epicyclic gear arrangement  30 . Radially outwardly of the sun gear  28  and intermeshing therewith is a plurality of planet gears  32  that are coupled together by a planet carrier  34 . The planet carrier  34  constrains the planet gears  32  to process around the sun gear  28  in synchronicity whilst enabling each planet gear  32  to rotate about its own axis. The planet carrier  34  is coupled via linkages  36  to the fan  23  in order to drive its rotation about the rotational axis  9 . Radially outwardly of the planet gears  32  and intermeshing therewith is an annulus or ring gear  38  that is coupled, via linkages  40 , to a stationary supporting structure  24 . 
     Note that the terms “low pressure turbine” and “low-pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e., not including the fan  23 ) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft  26  with the lowest rotational speed in the engine  10  (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  FIG.  3   . Each of the sun gear  28 , the planet gears  32 , and the ring gear  38  comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in  FIG.  3   . There are four planet gears  32  illustrated, although it will be apparent to the skilled reader that more or fewer planet gears  32  may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox  30  generally comprise at least three planet gears  32 . 
     The epicyclic gearbox  30  illustrated by way of example in  FIGS.  2  and  3    is of the planetary type, in that the planet carrier  34  is coupled to an output shaft via linkages  36 , with the ring gear  38  fixed. However, any other suitable type of epicyclic gearbox  30  may be used. By way of further example, the epicyclic gearbox  30  may be a star arrangement, in which the planet carrier  34  is held fixed, with the ring (or annulus) gear  38  allowed to rotate. In such an arrangement, the fan  23  is driven by the ring gear  38 . By way of further alternative example, the gearbox  30  may be a differential gearbox in which the ring gear  38  and the planet carrier  34  are both allowed to rotate. 
     It will be appreciated that the arrangement shown in  FIGS.  2  and  3    is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox  30  in the engine  10  and/or for connecting the gearbox  30  to the engine  10 . By way of further example, the connections (such as the linkages  36 ,  40  in the  FIG.  2    example) between the gearbox  30  and other parts of the engine  10  (such as the input shaft  26 , the output shaft and the fixed structure  24 ) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine  10  (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of  FIG.  2   . For example, where the gearbox  30  has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in  FIG.  2   . 
     Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangements, 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  10  shown in  FIG.  1    has a split flow nozzle  18 ,  20  meaning that the flow through the bypass duct  22  has its own nozzle  18  that is separate to and radially outside the core exhaust nozzle  20 . However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct  22  and the flow through the core  11  are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as, an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine  10  may not comprise a gearbox  30 . 
     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  FIG.  1   ), and a circumferential direction (perpendicular to the page in the  FIG.  1    view). The axial, radial, and circumferential directions are mutually perpendicular. 
     In addition, the present invention is equally applicable to aero gas turbine engines, marine gas turbine engines, and land-based gas turbine engines. 
       FIG.  4    illustrates an exemplary embodiment of a combustor arrangement  100  associated with the gas turbine engine  10  of  FIG.  1   . The combustor arrangement  100  defines a combustion chamber  102 . Specifically, the combustion chamber  102  is defined by a combustor wall  104  and an annular meter panel  106 . 
     The combustor arrangement  100  includes the meter panel  106  having a cold side  108  and a hot side  110 . The meter panel  106  is disposed proximate to the high-pressure compressor  15  (see  FIG.  1   ). The meter panel  106  has an aperture  112  extending through the meter panel  106  between the hot and cold sides  108 ,  110 . The upstream cold side  108  faces the high-pressure compressor  15  and the downstream hot side  110  faces the combustion chamber  102 . An annular location ring (not shown) may be coupled at an inside of the aperture  112  in the meter panel  106  for locating a seal  200  and a fuel injector  114 . It should be noted that the meter panel  106  may include a number of apertures  112 , such that each aperture  112  may receive a corresponding location ring, a corresponding seal  200 , and a corresponding fuel injector  114  therethrough. The meter panel  106  may include at least one retaining feature (not shown) at its cold side  108  and at least one slot (not shown) that extends through the meter panel  106  from the cold side  108  to the hot side  110  and is open to the aperture  112 . The combustor arrangement  100  further includes a heatshield  116  provided on the hot side  110  of the meter panel  106 . 
     Further, the combustor arrangement  100  includes the fuel injector  114 . The fuel injector  114  is arranged to supply fuel into the combustion chamber  102  during operation of the gas turbine engine  10 . The fuel injector  114  defines an outer diameter D 1 . As shown in  FIG.  5   , the combustor arrangement  100  also includes the seal  200  arranged around the fuel injector  114  and having an upstream end  202  (shown in  FIG.  6   ) and a downstream end  204 . The seal  200  is sized to fit through the aperture  112  (see  FIG.  4   ) of the meter panel  106  (see  FIG.  4   ), such that the upstream end  202  of the seal  200  is proximal to the cold side  108  (see  FIG.  4   ) of the meter panel  106  and the downstream end  204  of the seal  200  is proximal to the hot side  110  (see  FIG.  4   ) of the meter panel  106 . The seal  200  may be mounted from the cold side  108  of the meter panel  106 . Further, the fuel injector  114  may be mounted through the seal  200  from the cold side  108  of the meter panel  106 . 
     As shown in  FIG.  6   , the seal  200  defines a central axis A 1  extending along an axial direction A 2 . Further, the seal  200  includes a generally annular shape. The seal  200  includes an annular body  206  at least partially abutting the fuel injector  114  (see  FIG.  5   ). The seal  200  further includes a flange  208  radially extending from an outer surface  210  of the annular body  206  at the upstream end  202 . The flange  208  defines a diameter D 2 . The seal  200  may include a pair of tangs (not shown) extending axially from the flange  208  proximal to the upstream end  202  of the seal  200 . The pair of tangs may be arranged diametrically opposite to each other. Each tang may be sized to pass through a corresponding slot in the meter panel  106  (see  FIG.  4   ). It should be noted that the meter panel  106  has the same number and distribution of slots as the seal  200  has the tangs. 
     The annular body  206  extends circumferentially about the central axis A 1  and extends axially along the central axis A 1  from the upstream end  202  to the downstream end  204 . The annular body  206  defines a maximum axial length L 1  from the upstream end  202  to the downstream end  204 . Further, the annular body  206  includes an inner surface  212  facing the fuel injector  114 . The inner surface  212  axially extends from the upstream end  202  to the downstream end  204 . The inner surface  212  at least partially contacts the fuel injector  114 . In some examples, the inner surface  212  may abut with the fuel injector  114 . 
     Further, the annular body  206  includes the outer surface  210  radially spaced apart from the inner surface  212  relative to the central axis A 1  and facing away from the fuel injector  114 . The outer surface  210  axially extends from the upstream end  202  to the downstream end  204 . Further, in the illustrated example of  FIG.  6   , the outer surface  210  of the annular body  206  has a substantially uniform outer diameter D 3  along the central axis A 1 . Specifically, the seal  200  described herein may not include a flared portion at the downstream end  204  thereof. Further, the diameter D 2  of the flange  208  may be greater than the outer diameter D 3  defined by the outer surface  210  of the annular body  206 . Moreover, the annular body  206  defines a maximum radial thickness T 1  from the inner surface  212  to the outer surface  210 . 
     As illustrated in  FIG.  6   , the annular body  206  includes a plurality of slots  214  disposed on the inner surface  212  and circumferentially spaced apart from each other relative to the central axis A 1 . Each slot  214  axially extends at least partially from the downstream end  204  to the upstream end  202 . Further, each slot  214  is disposed in fluid contact with the fuel injector  114 . Moreover, each slot  214  is configured to receive a cooling fluid at the upstream end  202  of the seal  200  and discharge the cooling fluid at the downstream end  204  of the seal  200 , such that the cooling fluid contacts the fuel injector  114  while flowing through each slot  214 . In some examples, the plurality of slots  214  may include at least three slots  214 . The slots  214  may be equally spaced apart from each other along a circumferential direction C 1 . 
     As illustrated in  FIG.  7   , each slot  214  radially extends from the inner surface  212  partially towards the outer surface  210 . Specifically, each slot  214  has a maximum radial height H 1  along a radial direction R 1  with respect to the central axis A 1 . In the illustrated example of  FIG.  7   , the maximum radial height H 1  is defined proximate to a downstream slot end  218  of the slot  214 . Further, the maximum radial height H 1  of each slot  214  may be at most 90% of the maximum radial thickness T 1  of the annular body  206  from the inner surface  212  to the outer surface  210 . In various examples, the maximum radial height H 1  of each slot  214  may be at most 75%, 50%, 25%, and the like, of the maximum radial thickness T 1 . 
     Referring now to  FIG.  8   , each slot  214  extends from an upstream slot end  216  proximal to the upstream end  202  of the seal  200  to the downstream slot end  218  disposed at the downstream end  204  of the seal  200 . The upstream slot end  216  is axially spaced apart from the upstream end  202  of the seal  200  with respect to the central axis A 1 . The upstream slot end  216  includes a rounded design herein. In other examples, the upstream slot end  216  may be disposed at the upstream end  202  of the seal  200 . The downstream end  204  is open to allow discharge of the cooling fluid from the corresponding slot  214 . 
     Moreover, each slot  214  has a maximum slot length L 2 - 1  along the central axis A 1 . The maximum slot length L 2 - 1  of each slot  214  may be at least 90% of the maximum axial length L 1  of the annular body  206  from the upstream end  202  to the downstream end  204 . In various examples, the maximum slot length L 2 - 1  of each slot  214  may be at least 92%, 95%, 97%, 99%, and the like, of the maximum axial length L 1 . In the illustrated example of  FIG.  8   , the maximum slot length L 2 - 1  of each slot  214  may be less than the maximum axial length L 1  of the annular body  206 . Specifically, as the upstream slot end  216  may be axially spaced apart from the upstream end  202  of the seal  200  with respect to the central axis A 1 , the maximum slot length L 2 - 1  may be less than the maximum axial length L 1 . Alternatively, the maximum slot length L 2 - 1  of each slot  214  may be substantially equal to the maximum axial length L 1  of the annular body  206 . Specifically, in examples wherein the upstream slot end  216  may be disposed at the upstream end  202  of the seal  200 , the maximum slot length L 2 - 1  may be substantially equal to the maximum axial length L 1 . 
     Further, each slot  214  defines an angular width W 1 - 1 . It should be noted that the angular width W 1 - 1  may vary at different portions of the slot  214  along the central axis A 1 . Thus, each slot  214  has a maximum angular width W 1 - 2  along the circumferential direction C 1  with respect to the central axis A 1 . The maximum angular width W 1 - 2  may be defined at a portion of the slot  214  having a highest value of the maximum angular width W 1 - 1 . In the illustrated example of  FIG.  8   , the angular width W 1 - 1  may be substantially similar to the maximum angular width W 1 - 2 . Further, in some examples, the maximum angular width W 1 - 2  of the slots  214  may be based on a total number of the slots  214  and the outer diameter D 3  (see  FIG.  6   ). In some exemplary embodiments, the maximum angular width W 1 - 2  may be from about 0.5 mm to about 4 mm. The maximum angular width W 1 - 2  may be equal to or different from the maximum radial height H 1 . In an example, the maximum angular width W 1 - 2  may be greater than the maximum radial height H 1  by a factor of less than or equal to  90 . In another example, the maximum angular width W 1 - 2  may be less than the maximum radial height H 1 . In some examples, a circumferential spacing between adjacent slots  214  may not be less than the maximum angular width W 1 - 2  of the slots  214 . In other words, a circumferential spacing between adjacent slots  214  may be greater than or equal to the maximum angular width W 1 - 2  of the slots  214 . 
     Each slot  214  defines a slot axis A 3 - 1  along the maximum slot length L 2 - 1 . Therefore, each slot  214  extends along the corresponding slot axis A 3 - 1 . In some examples, the angular width W 1 - 1  of each slot  214  with respect to the central axis A 1  may be uniform or variable along the slot axis A 3 - 1 . As shown in  FIG.  8   , the angular width W 1 - 1  of each slot  214  with respect to the central axis A 1  may be uniform along the slot axis A 3 - 1 . Alternatively, as shown in  FIGS.  9 A and  9 B , an angular width W 2 - 1 , W 3 - 1  of each slot  230 ,  232  (shown in  FIGS.  9 A and  9 B , respectively) with respect to the central axis A 1  (see  FIG.  6   ) may be variable along a slot axis A 3 - 2 , A 3 - 3  (shown in  FIGS.  9 A and  9   b   , respectively). Further, the angular width W 2 - 1 , W 3 - 1  (shown in  FIGS.  9 A and  9   b   , respectively) of each slot  230 ,  232  may progressively increase or decrease from the upstream end  202  to the downstream end  204 . Specifically, the angular width W 2 - 1  of the slot  230  varies along the corresponding slot axis A 3 - 2 . Further, the angular width W 3 - 1  of the slot  232  varies along the corresponding slot axis A 3 - 3 . The slots  230 ,  232  including the progressively increasing or decreasing annular widths W 2 - 1 , W 3 - 1 , respectively, may allow a variation in the flow of the cooling fluid through the slots  230 ,  232 . 
     As shown in  FIG.  9 A , the angular width W 2 - 1  of each slot  230  may progressively increase from the upstream end  202  to the downstream end  204  such that the slot  230  may include a diverging cross-section. Further, the slot  230  has a maximum angular width W 2 - 2  at the downstream end  204 . The maximum angular width W 2 - 2  may be equal to or different from the maximum radial height H 1  (see  FIG.  6   ). In another example, as shown in  FIG.  9 B , the angular width W 3 - 1  of each slot  232  may progressively decrease from the upstream end  202  to the downstream end  204  such that the slot  232  may include a converging cross-section. For example, the slots  232  including the progressively increasing annular width W 3 - 1  may accelerate the flow through the slots  232  to ensure maximum heat transfer in high temperature regions. Further, the slot  232  has a maximum angular width W 3 - 2  at or proximal to the upstream end  202 . The maximum angular width W 3 - 2  may be equal to or different from the maximum radial height H 1  (see  FIG.  6   ). 
     Referring to  FIG.  10 A , each slot  214  may extend along the slot axis A 3 - 1 . The slot axis A 3 - 1  of each slot  214  may be parallel to or circumferentially angled relative to the central axis A 1  by an oblique angle. As illustrated in  FIG.  10 A , the slot axis A 3 - 1  of each slot  214  may be parallel to the central axis A 1 . Referring to  FIGS.  10 B and  10 C , each slot  234 ,  236  (shown in  FIGS.  10 B and  10 C , respectively) may extend along slot axes A 3 - 4 , A 3 - 5  (shown in  FIGS.  10 B and  10 C , respectively). Further, as illustrated in  FIGS.  10 B and  10 C , the slot axis A 3 - 4 , A 3 - 5  of each slot  234 ,  236  may be circumferentially angled relative to the central axis A 1  by an oblique angle O 1 , O 2  (shown in  FIGS.  10 B and  10 C , respectively). More specifically, in some examples, each slot 234may be circumferentially angled relative to the central axis A 1  by the oblique angle O 1 . Further, each slot  236  may be circumferentially angled relative to the central axis A 1  by the oblique angle O 2 . Furthermore, each of the slots  234 ,  236  may be obliquely inclined relative to the circumferential direction C 1 . Such slots  234 ,  236  may increase corresponding maximum slot lengths L 2 - 2 , L 2 - 3  of the corresponding slots  234 ,  236 , which may in turn improve a rate of convective heat transfer. Specifically, the maximum slot length L 2 - 2  of each slot  234  may be greater than the maximum slot length L 2 - 1  of each slot  214  (shown in  FIG.  10 A ). Similarly, the maximum slot length L 2 - 3  of each slot  236  may be greater than the maximum slot length L 2 - 1  of each slot  214 . Further, the circumferentially angled slots  234 ,  236  may allow control of an interaction between the cooling fluid flowing through the corresponding slots  234 ,  236  and a fuel spray cone, i.e., co-swirling for reduced interaction and/or counter swirling for maximum interaction. In some examples, the oblique angle O 1 , O 2  may be from about 30 degrees to about 75 degrees. In an example, the oblique angle O 1 , O 2  may be substantially equal to 60 degrees. In an example, as illustrated in  FIG.  10 B , the slots  234  may be circumferentially angled such that the oblique angle O 1  defined between the central axis A 1  and the slot axis A 3 - 4  may be measured in a clockwise direction CD1. In another example, as illustrated in  FIG.  10 C , the slots  236  may be circumferentially angled such that the oblique angle O 2   defined between the central axis A 1  and the slot axis A 3 - 5  may be measured in a counterclockwise direction CD2. 
     As shown in  FIG.  8   , the annular body  206  further includes a plurality of slot walls  220 . Each slot wall  220  defines a corresponding slot  214  from the plurality of slots  214 . As shown in  FIGS.  11 A,  11 B,  11 C, and  11 D , the slot wall  220  includes at least one of a plurality of recesses  222  (shown in  FIG.  11 A ) and a plurality of projections  224 ,  226 ,  228  (shown in  FIGS.  11 B,  11 C, and  11 D , respectively). The plurality of recesses  222  and the plurality of projections  224 ,  226 ,  228  may increase a rate of heat transfer based on an increase in a surface area defined by the slots  214 . In some examples, the slot wall  220  may include the plurality of recesses  222  or the plurality of projections  224 ,  226 ,  228 . Alternatively, the slot wall  220  may include a combination of the plurality of recesses  222  and the plurality of projections  224 ,  226 ,  228 , without any limitations. 
     Referring to  FIG.  11 A , the slot wall  220  may include the plurality of recesses  222 . The recesses  222  in the slot wall  220  may have a rectangular shape, a circular shape, an arrow shape, a circular shape, a semi-circular shape, an oval shape, a triangular shape, and the like, without any limitations. In the illustrated embodiment of  FIG.  11 A , each recess  222  has a substantially rectangular shape. Referring to  FIGS.  11 B,  11 C, and  11 D , the slot wall  220  may include the plurality of projections  224 ,  226 ,  228  (see  FIGS.  11 B,  11 C, and  11 D , respectively). In some examples, the projections  224 ,  226 ,  228  may be embodied as ribs, tabs, or textures. As illustrated in  FIG.  11 B , the projections  224  extending from the slot wall  220  may include a generally rectangular shape. As illustrated in  FIG.  11 C , the projections  226  extending from the slot wall  220  may include a generally square shape. As illustrated in  FIG.  11 D , the projections  228  extending from the slot wall  220  may be generally arrow shaped. Alternatively, the projections  224 ,  226 ,  228  extending from the slot wall  220  may include a circular shape, a semi-circular shape, an oval shape, a triangular shape, and the like, without any limitations. 
     Further, referring to  FIGS.  12 A to  12 E , each slot  214  may have a cross-sectional shape S 1 , S 2 , S 3 , S 4 , S 5  (shown in  FIGS.  12 A,  12 B,  12 C,  12 D, and  12 E , respectively) that is at least one of semi-circular, rectangular, concave, square, and trapezoidal. In some examples, the cross-sectional shape S 1 , S 2 , S 3 , S 4 , S 5  may have a high aspect ratio for increasing a heat transfer area within the slot  214 . In some examples, the cross-sectional shapes S 1 , S 2 , S 3 , S 4 , S 5  may include rounded corners that may simplify a manufacturing process of the slots  214  and may also reduce contact wear issues that may otherwise exist due to sharp edges. As shown in  FIG.  12 A , the cross-sectional shape S 1  of each slot  214  may be semi-circular. As shown in  FIG.  12 B , the cross-sectional shape S 2  of each slot  214  may be concave. Further, in the illustrated embodiment of  FIG.  12 B , the inner surface  212  of the annular body  206  may have a curved profile. As shown in  FIG.  12 C , the cross-sectional shape S 3  of each slot  214  may be rectangular with rounded corners. As shown in  FIG.  12 D , the cross-sectional shape S 4  of each slot  214  may be square with rounded corners. As shown in  FIG.  12 E , the cross-sectional shape S 5  of each slot  214  may be trapezoidal with rounded corners. It should be noted that the present disclosure is not limited by a cross-sectional shape of the slots  214 . Accordingly, the slots  214  may include any other cross-sectional shape, without any limitations. 
       FIG.  13    illustrates a seal  1300  associated with the combustor arrangement  100  (see  FIG.  4   ), according to another embodiment of the present disclosure. The seal  1300  includes an upstream end  1302  and a downstream end  1304  similar to the upstream and downstream ends  202 ,  204  of the seal  200  shown in  FIG.  6   . Further, the seal  1300  includes an annular body  1306  at least partially abutting the fuel injector  114  (see  FIG.  4   ) similar to the annular body  206  of the seal  200  shown in  FIG.  6   . The seal  1300  also includes a flange  1308  radially extending from an outer surface  1310  of the annular body  1306  at the upstream end  1302 . The flange  1308  is similar to the flange  208  of the seal  200  shown in  FIG.  6   . 
     Further, the annular body  1306  defines an inner surface  1312  facing the fuel injector  114 . The inner surface  1312  axially extends from the upstream end  1302  to the downstream end  1304 . The annular body  1306  also includes the outer surface  1310  radially spaced apart from the inner surface  1312  relative to a central axis A 4  and facing away from the fuel injector  114 . The outer surface  1310  axially extends from the upstream end  1302  to the downstream end  1304 . Further, the outer surface  1310  of the annular body  1306  has a substantially uniform outer diameter D 4  along the central axis A 4 . 
     The annular body  1306  further includes a plurality of wall portions  1338  extending from the inner surface  1312  and circumferentially spaced apart from each other relative to the central axis A 4 . In other words, the wall portions  1338  are spaced apart from each other along a circumferential direction C 2  defined with respect to the central axis A 4 . Each wall portion  1338  at least partially contacts the fuel injector  114 . Moreover, the annular body  1306  includes a plurality of slots  1314  disposed on the inner surface  1312  and circumferentially spaced apart from each other relative to the central axis A 4 . Specifically, the slots  1314  are spaced apart from each other along the circumferential direction C 2 . Each slot  1314  is defined by the inner surface  1312  and a pair of corresponding adjacent wall portions  1338  from the plurality of wall portions  1338 . Each slot  1314  axially extends at least partially from the downstream end  1304  to the upstream end  1302 . Further, each slot  1314  is disposed in fluid contact with the fuel injector  114 . In the illustrated embodiment of  FIG.  13   , the annular body  1306  includes three slots  1314 . Alternatively, the annular body  1306  may include six slots  1314 , nine slots  1314 , twelve slots  1314 , and the like, without any limitations. 
     Further, each slot  1314  extends from an upstream slot end  1316  disposed at the upstream end  1302  of the seal  1300  to a downstream slot end  1318  disposed at the downstream end  1304  of the seal  1300 . Furthermore, a maximum slot length L 2 - 4  of each slot  1314  may be substantially equal to a maximum axial length L 1 - 1  of the annular body  1306 . Specifically, as the upstream slot end  1316  of the slots  1314  may be disposed at the upstream end  1302  of the seal  1300 , the maximum slot length L 2 - 4  may be substantially equal to the maximum axial length L 1 - 1 . Moreover, each slot  1314  has a maximum radial height H 2  along a radial direction R 2  with respect to the central axis A 4 . 
     It should be noted that the details related to design, dimensions, and applications provided for the slots  214 ,  230 ,  232 ,  234 ,  236  (see  FIGS.  8 ,  9 A,  9 B,  10 B, and  10 C , respectively) of the seal  200  may be equally applicable to the slots  1314  of the seal  1300 , without any limitations. 
     In the illustrated embodiment of  FIG.  13   , each slot  1314  has a maximum angular extent E 1  relative to the central axis A 4 . The maximum angular extent E 1  may be less than 120 degrees. In various examples, the maximum angular extent E 1  may be about 90 degrees, 60 degrees, 30 degrees, and the like. It should be noted that the maximum angular extent E 1  may vary based on a total number of the slots  1314  associated with the annular body  1306 . In the illustrated embodiment of  FIG.  13   , the maximum angular extent E 1  may be slightly less than 120 degrees as the annular body  1306  includes three slots  1314 . Further, each slot  1314  may define a maximum angular width W 4 - 2  along the circumferential direction C 2 . In some examples, the maximum angular width W 4 - 2  may be greater than the maximum radial height H 2  by a factor of less than or equal to  90 . 
     Referring to  FIGS.  5  to  13   , the seals  200 , 1300  (see  FIGS.  6  and  13   ) may be manufactured, for example, by casting. Subsequently, the seals  200 ,  1300  may be drilled using techniques, such as, electrochemical machining (ECM), electrical discharge machining (EDM), or laser drilling for forming the corresponding slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  (see  FIGS.  8 ,  9 A,  9 B,  10 B,  10 C, and  13    respectively). In some examples, grinding or broaching may be used to manufacture the seals  200 ,  1300 . In another example, the seals  200 ,  1300  may be manufactured by casting using cores to define the corresponding slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  and then removing the cores, such as, by dissolving. Alternatively, the seals  200 ,  1300  may be manufactured by additive layer manufacturing (ALM), e.g., three-dimensional printing, powder bed laser deposition, and the like. Thus, the seals  200 ,  1300  may be manufactured via a range of processes, thereby providing supply chain flexibility. 
     The seals  200 ,  1300  include the number of corresponding slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314 . The slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  may include an improved design such that the flow of the cooling fluid through the slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  may provide efficient cooling of the corresponding seals  200 ,  1300 , and the fuel injector  114 . Further, the seals  200 ,  1300  include a flare-less design. Elimination of the flared portion may reduce a possibility of accumulation of molten seal material proximal to the downstream ends  204 ,  1304  of the corresponding seals  200 ,  1300  and may also reduce smoke emissions. The slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  associated with the corresponding seals  200 ,  1300  may be embodied as substantially full-length slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314 . Further, the cooling fluid flowing through the substantially full-length slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  may allow cooling of the corresponding seals  200 ,  1300  as well as the fuel injector  114 . Moreover, the spent cooling fluid may be then delivered locally to the fuel spray cone of the fuel injector  114  which may in turn provide a benefit to engine emissions in terms of smoke. 
     Further, in addition to reducing a temperature of the seals  200 ,  1300  and the fuel injector  114 , the slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  in the corresponding seals  200 ,  1300   may also reduce smoke emissions in order to address issues associated with seal durability and engine emission certifications. The slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  may be disposed proximate to the inner surface  212 ,  1312  of the corresponding seals  200 ,  1300  which may enable high levels of convective heat transfer to occur between the cooling fluid and the seal  200 ,  1300 . This feature may in turn provide a means to effectively cool hot portions of the seal  200 ,  1300  and the fuel injector  114 . Moreover, a placement of the slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  proximal to the corresponding inner surfaces  212 ,  1312  may also generate a radial positive pressure at corresponding interfaces between the corresponding seals  200 ,  1300  and the outer diameter D 1  of the fuel injector  114 , which may in turn drive the cooling fluid by pressure. In some examples, a pressure differential across the corresponding seals  200 ,  1300  may create an aerodynamic bearing, which may reduce contact load and operational wear of the corresponding seals  200 ,  1300  and/or the fuel injector  114 . This feature may also address durability issues associated with conventional seals that typically causes rapid oxidisation of conventional seals due to inadequate cooling. Further, the cooling fluid flowing through the slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  may enhance heat transfer at the outer diameter D 1  of the fuel injector  114 . 
     Moreover, as mentioned above, the design of the seals  200 ,  1300  described herein may address engine emission issues by realising an improved interaction of the cooling fluid with the fuel spray cone. Further, the cooling fluid may exit the slots  214 ,  230 ,  232 ,  234 ,  236 ,  1314  in close proximity to the fuel injector  114  which may allow control over a fuel spray cone angle. In some examples, by introducing the cooling fluid locally to the fuel spray cone, the fuel spray cone angle may be narrowed which may in turn reduce a residence time of the combustion process in a primary zone (i.e., a front section of the combustion chamber  102 ) that may be primarily responsible for production of smoke emissions. Further, the seals  200 ,  1300  described herein may be compact and light in weight. Moreover, the seals  200 ,  1300  may be retrofitted in existing combustor arrangements without making any changes to the combustor arrangements 
     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.