GB2590659A - Nozzle for gas turbine engine and method of manufacture thereof - Google Patents

Abstract

A nozzle 100 for a gas turbine (10 Fig. 1) has an outer member 101 and an inner member 201. The outer member includes a main body 102 which has a first end 105, a second end 106 and an outer passage (107 Fig 5A). An internal surface 109 of the main body has at least one internal thread 108. The inner member has an elongate body 202 and a first end 203, a second end 204 and an inner passage 205. An external surface 209 of the elongate body has at least one external thread 208, which cooperates with the internal thread of the main body. The cooperation of the at least one internal thread and the at least one external thread together forms a plurality of labyrinth grooves, which are arranged adjacent a gap 304 formed between the inner and outer members. The nozzle is manufactured my inserting the inner member into the outer member (902 Fig 10) and rotating to engage the cooperating threads (904 Fig. 10), and then fastening the inner member to the outer member (906 Fig. 10) by, for example, by laser welding. The arrangement provides tolerance to thermal expansion and carbon formation.

Description

NOZZLE FOR GAS TURBINE ENGINE AND METHOD OF MANUFACTURE

THEREOF

FIELD OF THE DISCLOSURE

The present disclosure relates to a nozzle, and in particular to a nozzle for a gas turbine engine and a method of manufacturing the nozzle.

BACKGROUND

Fuel spray nozzles are used for delivering a fuel into a combustor of a gas turbine. As the fuel spray nozzles are generally positioned downstream of a compressor of the gas turbine, the nozzles are surrounded by high temperature gases exiting the compressor. Such a high temperature can cause the fuel within the nozzle to break down, and form lacquers and carbon deposits on surfaces of the nozzle. These deposits can reduce flow through the nozzle or block it completely, affecting engine performance.

In order to keep the temperature of fuel-wetted surfaces in the nozzles at an acceptable level, cavities and air gaps are formed in a structure associated with the nozzles to reduce the conduction of heat into the fuel. The structures that form these cavities are typically anchored at one end and free at the other to allow for the differences in thermal expansion of external walls of the nozzle in contact with the hot gases and internal walls of the nozzle in contact with the fuel. It may be possible for the fuel to enter the insulating cavities through the free end of the structure, where it may start to form lacquers and carbon when it contacts the hot external walls of the nozzle. Carbon is not as good an insulator as air, so its formation can increase heat transfer into the fuel. The cyclic expansion and contraction of these cavities under thermal load can also lead to a phenomenon known as carbon jacking, which can cause deformation, crushing and cracking of components.

SUMMARY

According to a first aspect there is provided a nozzle having an outer member and an inner member. The outer member includes a main body having a first end, a second end opposite to the first end, and an outer passage extending between the first end and the second end. The outer member further includes at least one internal thread disposed on an internal surface of the main body. The inner member is at least partially received within the outer passage of the outer member. The inner member includes an elongate body having a first end, a second end opposite to the first end, an inner passage extending between the first end and the second end, an inlet opening disposed in fluid communication with the inner passage, and an outlet opening disposed in fluid communication with the inner passage. The inner member further includes at least one external thread disposed on an external surface of the elongate body and cooperating with the at least one internal thread of the outer member. Further, at least a portion of the internal surface of the main body and at least a portion of the external surface of the elongate body define a gap therebetween. The at least one internal thread and the at least one external thread together form a plurality of labyrinth grooves adjacent to the gap.

The labyrinth grooves formed by the at least one internal thread and the at least one external thread may provide a tortuous or convoluted path to any fluid leaking into the nozzle from the second end of the main body. The labyrinth grooves may capture the fluid and thereby prevent the fluid from entering the gap between the inner member and the outer member. The labyrinth grooves may therefore form a double-sided labyrinth seal adjacent to the gap. The fluid that can leak into the nozzle may be a fuel or a fuel air mixture.

The labyrinth grooves may prevent any leakage of fuel into the gap between the inner member and the outer member, thereby avoiding any formation of deposits on the internal surface of the outer member and/or the external surface of the inner member. An insulation provided by the gap may therefore not be adversely impacted by deposits during operation of the nozzle.

In some embodiments, the at least one internal thread is disposed proximal to the second end of the main body.

In some embodiments, the at least one external thread is disposed proximal to the second end of the elongate body.

In some embodiments, the outer member includes a first part defining the first end of the main body, and a second part fixedly attached to the first part and defining the second end of the main body. The second part of the main body includes the at least one internal thread. The two-part construction of the outer member may facilitate assembly of the nozzle.

In some embodiments, the first part is welded to the second part.

In some embodiments, the at least one internal thread is helically disposed on the internal surface of the main body.

In some embodiments, the at least one external thread is helically disposed on the outer surface of the elongate body.

In some embodiments, the main body is further configured to receive a cooling fluid in the gap between the internal surface of the main body and the external surface of the elongate body. The cooling fluid may be air. The cooling fluid may cool the nozzle and insulate a fuel within the nozzle from high temperature gases.

In some embodiments, the inlet opening is disposed at the first end of the elongate body.

In some embodiments, the outlet opening is disposed at the second end of the elongate body.

In some embodiments, the inlet opening of the elongate body is configured to receive a fuel.

In some embodiments, a gas turbine for an aircraft is provided. The gas turbine includes a combustor and the nozzle of the first aspect. The nozzle is configured to supply fuel to the combustor.

According to a second aspect, there is provided a method of manufacturing a nozzle. The method incudes inserting an inner member within a first part of an outer member. The method further includes rotating the inner member and a second part of the outer member relative to one another in order to cooperate at least one external thread of the inner member with at least one internal thread of the second part of the outer member. The method further includes fixedly attaching the first part of the outer member to the second part of the outer member.

In some embodiments, fixedly attaching the first part to the second part further includes welding at an interface between the first part and the second pad.

In some embodiments, welding at the interface between the first part and the second part includes laser welding.

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 10 downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (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 nonlim itative 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 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity U. 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/Ut1p2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg-1K-1/(ms-1)2). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5,13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the engine core. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 95 Nkg- 15, 90 Nkg-ls, 85 Nkg-ls or 80 Nkg-ls. 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-ls to 100 Nkg-1s, or 85 Nkg-ls to 95 Nkg-ls. 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: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. 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 330kN to 420 kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K. The maximum TEl may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example, at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc.

By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be 5 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 10 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, m id-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 m id-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). 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 50kN to 65kN) 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 m id-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: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close up sectional side view of an upstream portion of a gas turbine engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine 5 engine; Figure 4 is a sectional side view of a combustor of the gas turbine engine including a nozzle; Figure 5A is a sectional side view of an outer member of the nozzle; Figure 5B is another sectional side view of the outer member of Figure 5a; Figure 6A is a perspective view of an inner member of the nozzle; Figure 6B is a side view of the inner member of Figure 6a; Figure 6C is a sectional side view of the inner member of Figure 6a; Figure 7 is a partial sectional view of the nozzle; Figure 8 is a cutaway view of the nozzle; Figures 9A, 9B, 9C and 9D depict various types of external and internal threads that can be used with the nozzle; and Figure 10 is a flowchart of a method of manufacturing a nozzle.

DETAILED DESCRIPTION

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.

Figure 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, combustor 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 5 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 combustor 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 10 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 Figure 2. The low pressure turbine 19 (see Figure 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 engine 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 (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 Figure 3. Each of the sun gear 28, planet gears 32 and 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 Figure 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 Figures 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 Figures 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 Figure 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 (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 Figure 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 Figure 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 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 Figure 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 Figure 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular.

In addition, the present invention is equally applicable to aero gas turbine 5 engines, marine gas turbine engines and land-based gas turbine engines.

Figure 4 illustrates an enlarged side view of the combustor 16, the high pressure compressor 15 and the high pressure turbine 17. For injecting fuel or an air-fuel mixture into the combustor 16, a plurality of nozzles 100 (only one shown in Figure 4) are provided. At an upstream end of the combustor 16, a combustor ring R is provided, on which the plurality of nozzles 100 are arranged along a circular line to inject the fuel or fuel air mixture into the combustor 16. Each nozzle 100 is connected to the combustor ring R by screws or any other mechanical arrangement. Each nozzle 100 is configured to supply fuel or air-fuel mixture to the combustor 16.

Referring again to Figure 4, compressed air from the high pressure compressor 15 enters the combustor 16. In the combustor 16, the compressed air is mixed with fuel and the mixture is ignited to create energy. The air-fuel mixture then 20 passes through the high pressure turbine 17 to generate mechanical power.

Referring to Figures 5A, 5B, 6A, 6B, 6C and 7, the nozzle 100 includes an outer member 101 and an inner member 201. The outer member 101 and the inner member 201 are positioned concentrically with respect to each other. The outer member 101 includes a main body 102. The inner member 201 includes an elongate body 202.

Referring now to Figures 5A-5B, the main body 102 of the outer member 101 defines a first end 105 and a second end 106. The main body 102 further 30 defines an outer passage 107 extending between the first end 105 and the second end 106. Specifically, the outer passage 107 extends along a longitudinal axis 110 of the main body 102. The main body 102 further includes an internal surface 109. The outer member 101 further includes at least one internal thread 108 disposed on the internal surface 109 of the main body 102.

The at least one internal thread 108 is disposed proximal to the second end 106 of the main body 102.

Furthermore, as shown in Figures 5A-5B, the main body 102 of the outer 5 member 101 includes a first part 103 and a second part 104. The first part 103 defines the first end 105 of the main body 102 and the second pad 104 defines the second end 106 of the main body 102. The at least one internal thread 108 is provided in the internal surface 109 of the second part 104 of the main body 102. The at least one internal thread 108 is helically disposed on the internal 10 surface 109 of the main body 102. In the illustrated embodiment, the outer member 101 includes one internal thread 108. An end 108A of the internal thread 108 is shown in Figure 5B. The end 108A is spaced apart from the second end 106 of the main body 102. However, the outer member 101 may include multiple internal threads (not shown) with same or different pitches.

In some embodiments, the first part 103 is fixedly attached to the second part 104. In some embodiments, the first part 103 is welded to the second part 104. In some embodiments, welding the first part 103 to the second part 104 includes laser welding. The two-part construction of the outer member 101 may facilitate assembly of the nozzle 100. In some alternative embodiments, main body 102 may be a single piece component.

Figures 6A, 6B and 6C illustrate perspective views of the inner member 201 of the nozzle 100. The inner member 201 is at least partially received within the outer passage 107 (shown in Figure 5A) of the main body 102. The elongate body 202 of the inner member 201 defines a first end 203 and a second end 204 opposite to the first end 203. The elongate body 202 further defines an inner passage 205 (shown in Figure 6C) extending between the first end 203 and the second end 204. The elongate body 202 further defines an inlet opening 206 (shown in Figure 6C) disposed in fluid communication with the inner passage 205. The elongate body 202 further defines an outlet opening 207 disposed in fluid communication with the inner passage 205. In the illustrated embodiment, the inlet opening 206 is disposed at the first end 203 of the elongate body 202. Further, the outlet opening 207 is disposed at the second end 204 of the elongate body 202. In some embodiments, the inlet opening 206 of the elongate body 202 is configured to receive a fuel. In some other embodiments, the inlet opening 206 of the elongate body 202 may be configured to receive air.

The elongate body 202 further includes an external surface 209. The inner member 201 further includes at least one external thread 208 disposed on the external surface 209 of the elongate body 202. The at least one external thread 208 is disposed proximal to the second end 204 of the elongate body 202. The at least one external thread 208 is helically disposed on the external surface 209 of the elongate body 202. The at least one external thread 208 disposed on the external surface 209 of the elongate body 202 cooperates with the at least one internal thread 108 (shown in Figure 5A) disposed on the internal surface 109 of the main body 102. The cooperation may occur upon assembly of the inner member 201 with the outer member 101.

In the illustrated embodiment, the inner member 201 includes one external thread 208. However, the outer member 201 may include multiple external threads (not shown) with same or different pitches. The multiple external threads may cooperate with multiple internal threads of the elongate body 202.

Figure 7 shows the nozzle 100. The inner member 201 is at least partially received within the outer passage 107 (shown in Figure 5A) of the main body 102. A portion 109A of the internal surface 109 of the main body 102 and a portion 209A of the external surface 209 of the elongate body 202 define a gap 304 therebetween. The gap 304 may be an annular gap. The at least one internal thread 108 cooperates with the at least one external thread 208 to form a plurality of labyrinth grooves 402 adjacent to the gap 304. The labyrinth grooves 402 may be axially spaced apart from each other relative to the longitudinal axis 110 Further, each labyrinth groove 402 may extend circumferentially about the longitudinal axis 110. Specifically, the cooperation between the at least one internal thread 108 and the at least one external thread 208 forms a double-sided labyrinth seal between the main body 102 and the elongate body 202. Further, the labyrinth grooves 402 are axially disposed between the gap 304 and the second end 106 of the main body 102.

A fuel supply conduit 301 is in fluid communication with the inlet opening 206 disposed at the first end 203 of the elongate body 202. The fuel supply conduit 301 supplies fuel or air-fuel mixture to the inner passage 205 of the elongate body 202 of the inner member 201. The fuel or air-fuel mixture travels through the inner passage 205 and exits through the outlet opening 207 disposed at the second end 204 of the elongate body 202.

An air supply conduit 302 is in fluid communication with the gap 304 between the portion 109A of the internal surface 109 of the main body 102 and the portion 209A of the external surface 209 of the elongate body 202. The main body 102 is further configured to receive a cooling fluid in the gap 304 between the internal surface 109 of the main body 102 and the external surface 209 of the elongate body 202. The air supply conduit 302 may supply compressed cooled air from a cooler part (not shown) of the gas turbine engine 10 (shown in Figure 1) to the gap 304. Air in the gap 304 may provide insulation, and reduce transmission of heat to the main body 102 and hence the inner passage 205 from surrounding parts. Insulation of the main body 102 may at least partially reduce deposition of fuel on walls of the elongate body 202.

The first end 105 of the main body 102 is disposed at an upstream end of the nozzle 100 and the second end 106 of the main body 102 is disposed at a downstream end of the nozzle 100. In the same manner, the first end 203 of the elongate body 202 is disposed at the upstream end and the second end 204 of the elongate body 202 is disposed at the downstream end of the nozzle 100.

Referring to Figure 8, the formation of the labyrinth grooves 402 by the cooperation of the at least internal thread 108 and the at least one external thread 208 creates a tortuous path 305 for a fluid, for example, fuel or air-fuel mixture. The fluid may leak into the nozzle 100 from the second end 106 of the main body 102. The tortuous path 305 may act restrict a flow of the fluid from the second end 106 to the first end 105 of the main body 102. Specifically, the tortuous path 305 may act as a restriction for the flow of the fluid along the longitudinal axis 110. Thus, the fluid (fuel or air-fuel mixture) may get trapped or captured in the tortuous path 305 as it travels from the second end 106.

The labyrinth grooves 402 may prevent any leakage of fuel into the gap 304 between the inner member 201 and the outer member 101, thereby avoiding any formation of deposits on the internal surface 109 of the outer member 101 and/or the external surface 209 of the inner member 201. An insulation provided by the gap 304 may therefore not be adversely impacted by deposits during operation of the nozzle 100.

In some embodiments, a swirler (not shown) may be disposed in the combustor 16 (shown in Figure 4). During operation, a gas pressure drops as gas passes through vanes of the swirler. Consequently, a pressure at an exit of the nozzle 100 may be lower than a pressure inside the nozzle 100. This may drive a flow of the cooling fluid (i.e., air) across the double-sided labyrinth seal formed by the labyrinth grooves 402. In some embodiments, a hole may be provided in the outer member 101 upstream of the labyrinth grooves 402 to provide a pressure drop or gradient across the labyrinth grooves 402.

As shown in Figures 5A-5B, 6A-6B and 7, each of the internal thread 108 and the external thread 208 has a trapezoidal shape. However, alternative shapes of the internal thread 108 and the external thread 208 are possible within the scope of the present disclosure. Figure 9A shows an internal thread 708A and an external thread 808A having sawtooth profiles. Figure 9B shows an internal thread 708B and an external thread 808B having a triangular profile. Figure 9C shows an internal thread 708C and an external thread 808C having a rectangular profile. Figure 9D also shows an internal thread 708D and an external thread 808D having a rectangular profile. However, a pitch of the internal thread 708D and the external thread 808D of Figure 9D is greater than a pitch of the internal thread 708C and the external thread 808C of Figure 9C. A shape and a pitch of internal and external threads may be varied as per application requirements. An increase in pitch may increase a volume of resultant labyrinth grooves and vice versa.

As shown in Figure 7 and already described above with reference to Figures 5A-5B, the outer member 101 of the nozzle 100 has a two-piece design. The two-piece design includes the first part 103 and the second part 104.

Figure 10 illustrates a method 900 of manufacturing the nozzle 100. At step 902 the method 900 includes inserting the inner member 201 within the first part 103 of the outer member 101.

At step 904, the method 900 includes rotating the inner member 201 and the 10 second part 104 of the outer member 101 relative to one another to cooperate the at least one external thread 208 of the inner member 201 with the at least one internal thread 108 of the second part 104 of the outer member 101.

At step 906, the method 900 further includes fixedly attaching the first part 103 to the second part 104. In some embodiments, fixedly attaching the first part 103 to the second part 104 further includes welding at an interface 303 between the first part 103 and the second part 104. The interface 303 may be a circumferential interface. In some embodiments, welding at the interface 303 between the first part 103 and the second part 104 includes laser welding. In some other embodiments, brazing may be used to join the first part 103 to the second part 104.

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.

Claims (15)

1. CLAIMS: 1 A nozzle (100) comprising: an outer member (101) comprising: a main body (102) defining a first end (105), a second end (106) opposite to the first end (105), and an outer passage (107) extending between the first end (105) and the second end (106); and at least one internal thread (108) disposed on an internal surface (109) of the main body (102); and an inner member (201) at least partially received within the outer passage (107) of the outer member (101), the inner member (201) corn prising: an elongate body (202) defining a first end (203), a second end (204) opposite to the first end (203), an inner passage (205) extending between the first end (203) and the second end (204), an inlet opening (206) disposed in fluid communication with the inner passage (205) and an outlet opening (207) disposed in fluid communication with the inner passage (205); and at least one external thread (208) disposed on an external surface (209) of the elongate body (202) and cooperating with the at least one internal thread (108) of the outer member (101); wherein at least a portion (109A) of the internal surface (109) of the main body (102) and at least a portion (209A) of the external surface (209) of the elongate body (202) define a gap (304) therebetween, and wherein the at least one internal thread (108) and the at least one external thread (208) together form a plurality of labyrinth grooves (402) adjacent to the gap.
2. 2. The nozzle (100) of claim 1, wherein the at least one internal thread (108) is disposed proximal to the second end (106) of the main body (102).
3. 3. The nozzle (100) of claim 1 01 2, wherein the at least one external thread (208) is disposed proximal to the second end (204) of the elongate body (202).
4. 4. The nozzle (100) of any preceding claim, wherein the outer member (101) comprises a first part (103) defining the first end (105) of the main body (102) and a second part (104) fixedly attached to the first part (103) and defining the second end (106) of the main body (102), wherein the second part (104) comprises the at least one internal thread (108).
5. 5. The nozzle (100) of claim 4, wherein the first pad (103) is welded to the second part (104).
6. 6. The nozzle (100) of any preceding claim, wherein at least one internal thread (108) is helically disposed on the internal surface (109) of the main body (102).
7. 7. The nozzle (100) of any preceding claim, wherein at least one external thread (208) is helically disposed on the outer surface (209) of the elongate body (202).
8. 8. The nozzle (100) of any preceding claim, wherein the main body (102) is further configured to receive a cooling fluid in the gap (304) between the internal surface (109) of the main body (102) and the external surface (209) of the elongate body (202).
9. 9. The nozzle (100) of any preceding claim, wherein the inlet opening (206) is disposed at the first end (203) of the elongate body (202).
10. 10. The nozzle (100) of any preceding claim, wherein the outlet opening (207) is disposed at the second end (204) of the elongate body (202).
11. 11. The nozzle (100) of any preceding claim, wherein the inlet opening (206) of the elongate body (202) is configured to receive a fuel.
12. 12.A gas turbine engine (10) for an aircraft, the gas turbine engine (10) comprising: a combustor (16); and the nozzle (100) of any preceding claim, wherein the nozzle (100) is configured to supply fuel to the combustor (16).
13. 13.A method (900) of manufacturing a nozzle (100), the method (900) comprising: inserting (902) an inner member (201) within a first part (103) of an outer member (101); rotating (904) the inner member (201) and a second part (104) of the outer member (101) relative to one another in order to cooperate at least one external thread (208) of the inner member (201) with at least one internal thread (108) of the second part (204) of the outer member (101); and fixedly attaching (906) the first part (103) to the second part (104).
14. 14. The method (900) of claim 13, wherein fixedly attaching the first part (103) to the second part (104) further comprises welding at an interface (303) between the first part (103) and the second part (104).
15. 15. The method (900) of claim 14, wherein welding at the interface (303) between the first part (103) and the second part (104) comprises laser welding