GB2572783A

GB2572783A - A Gas Turbine Engine Comprising A Fuel Purging System And A Method of Purging Fuel From A Gas Turbine Engine

Info

Publication number: GB2572783A
Application number: GB1805931.1A
Authority: GB
Inventors: Tentorio Luca; Wei Huang Hua; Carlos Roman Casado Juan; Di Chiaro Giacomo; Knapton Jonathan
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2018-04-10
Filing date: 2018-04-10
Publication date: 2019-10-16
Also published as: GB201805931D0

Abstract

A gas turbine engine 10 comprises a compressor 15, a combustion chamber 16, a fuel system 50 and a fuel purging system 60. The fuel system comprises a fuel injector 52, an associated fuel delivery pipe 54 and a fuel manifold 56 which supplies fuel to the fuel injector through the associated fuel delivery pipe. The fuel purging system comprises a compressed air storage vessel 62, a first passage 64 to supply compressed air from the compressor to the compressed air storage vessel and a second passage 66 to supply compressed air from the compressed air storage vessel to the fuel manifold and/or the fuel injector. The fuel purging system has a first valve 68 in the first passage and a second valve 70 in the second passage. The fuel purging system is used to prevent coking of fuel within the fuel system. Also disclosed is a method of purging fuel and a geared turbofan engine each utilising the engine as set out above.

Description

A GAS TURBINE ENGINE COMPRISING A FUEL PURGING SYSTEM AND A METHOD OF PURGING FUEL FROM A GAS TURBINE ENGINE

The present disclosure relates to a gas turbine engine comprising a fuel purging system and a method of purging fuel from a gas turbine engine.

A gas turbine engine has one or more combustion chambers in which combustion occurs. A gas turbine engine has a fuel system which has fuel injectors to supply fuel into the combustion chambers and a fuel manifold to supply fuel to the fuel nozzles of the fuel injectors. The fuel in the fuel injectors and the fuel manifold are exposed to high temperatures in the vicinity of the combustion chamber and the turbines of the gas turbine engine which results in an increase in the temperature of the fuel. Another gas turbine engine has a lean burn fuel system which has lean burn fuel injectors to supply fuel into the combustion chambers, a pilot fuel manifold to supply fuel to pilot fuel nozzles of the lean burn fuel injectors and a main fuel manifold to supply fuel to main fuel nozzles of the lean bum fuel injectors. The fuel in the lean burn fuel injectors, the pilot fuel manifold and the main fuel manifold are exposed to high temperatures in the vicinity of the combustion chamber and the turbines of the gas turbine engine which results in an increase in the temperature of the fuel.

Gas turbine engines, in particular aero gas turbine engines, may also use the fuel as a hydraulic fluid to operate an actuator for variable vanes of a compressor of the gas turbine engine. The use of the fuel as a hydraulic fuel increases the temperature of the fuel. Gas turbine engines, in particular aero gas turbine engines, may also have a heat exchanger in the fuel system such that the fuel is used to cool oil in the gas turbine engine lubrication system. The use of the fuel system as a heat sink for the lubrication system also results in an increase in the temperature of the fuel.

The fuel in the fuel system may stagnate when the gas turbine engine stops operating and as mentioned above any fuel in the fuel injectors and fuel manifold are exposed to the high temperatures in the vicinity of the combustion chamber and the turbines of the gas turbine engine which results in an increase in the temperature of the fuel. The fuel in the lean burn fuel system may stagnate when the gas turbine engine stops operating and as mentioned above any fuel in the lean burn fuel injectors, the main fuel manifold and the pilot fuel manifold are exposed to the high temperatures in the vicinity of the combustion chamber and the turbines of the gas turbine engine which results in an increase in the temperature of the fuel. The fuel in the main fuel manifold and main fuel nozzles of the lean bum fuel system may stagnate when the gas turbine engine is operating with fuel supplied to the pilot fuel nozzles but not supplied to the main fuel nozzles and as mentioned above any fuel in the main fuel nozzles of the lean burn fuel injectors and main fuel manifold are exposed to the high temperatures in the vicinity of the combustion chamber and the turbines of the gas turbine engine which results in an increase in the temperature of the fuel.

High fuel temperatures stimulate the fuel auto-oxidation reactions, which lead to the formation of gums and other insoluble materials (including carbon) that tend to deposit on the walls of the fuel passages and fuel metering orifices within the fuel system, e.g. the fuel manifold and the fuel injectors. The deposition of carbon on the on the walls of the fuel passages and fuel metering orifices within the fuel injectors is known as coking.

The fuel passages and the fuel metering orifices within the fuel injector have relatively small cross-sectional dimensions and thus the deposition of the carbon in the fuel passages and the fuel metering orifices within the fuel injector may result in the fuel passages and the fuel metering orifices becoming restricted or blocked. The carbon deposits within the fuel passages and the fuel metering orifices may distort the fuel spray from the fuel injector and create appreciable non-uniformities in the fuel spray pattern.

The present invention seeks to provide a fuel system for a gas turbine engine which reduces or overcomes the above mentioned problem.

According to a first aspect there is provided a gas turbine engine comprising a compressor, at least one combustion chamber, a fuel system and a fuel purging system, the fuel system comprises at least one fuel injector, an associated fuel delivery pipe and a fuel manifold arranged to supply fuel to the at least one fuel injector through the associated fuel delivery pipe, each combustion chamber has at least one fuel injector, the fuel purging system comprises a compressed air storage vessel, a first passage to supply compressed air from the compressor to the compressed air storage vessel and a second passage to supply compressed air from the compressed air storage vessel to the fuel manifold and/or the at least one fuel injector, a first valve in the first passage and a second valve in the second passage.

The gas turbine engine may comprise an inner combustion chamber casing, an outer combustion chamber casing surrounding the inner combustion chamber casing and the at least one combustion chamber, the at least one combustion chamber being arranged between the inner combustion chamber casing and the outer combustion chamber casing, a diffuser to supply compressed air from the compressor to the at least one combustion chamber, the inner combustion chamber casing and the outer combustion chamber casing being sealingly secured to the diffuser to define a chamber around the at least one combustion chamber.

The compressed air storage vessel may be arranged outside the outer combustion chamber casing and the first passage is fluidly connected to the chamber.

The compressed air storage vessel may be arranged inside the outer combustion chamber casing. The compressed air storage vessel may be defined by the outer combustion chamber casing, the diffuser and a further wall sealingly secured to the diffuser and the outer combustion chamber casing and the first passage is an opening in the further wall.

The first valve may be a non-return valve. The first valve may comprise a valve member movable against a spring between a closed position and an open position. The first valve may be an active valve. The first valve may be an electrically operated valve or a hydraulically operated valve.

The second valve may be an active valve. The second valve may be an electrically operated valve or a hydraulically operated valve.

The at least one fuel injector may be a rich burn fuel injector comprising a main fuel nozzle, the fuel manifold is arranged to supply fuel to the main fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe.

The at least one fuel injector may be a lean burn injector comprising a pilot fuel nozzle and a main fuel nozzle. The fuel manifold is arranged to supply fuel to the main fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe. The fuel manifold is arranged to supply fuel to the pilot fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe.

The fuel system may comprise a fuel shut off valve arranged to supply fuel to the fuel manifold. The fuel system may comprise a fuel drainage valve to drain fuel from the fuel manifold.

According to a second aspect there is provided a method of purging fuel from a gas turbine engine, the gas turbine engine comprising a compressor, at least one combustion chamber, a fuel system and a fuel purging system, the fuel system comprises at least one fuel injector, an associated fuel delivery pipe and a fuel manifold arranged to supply fuel to the at least one fuel injector through the associated fuel delivery pipe, each combustion chamber has at least one fuel injector, the method comprising supplying compressed air from the compressor to a compressed air storage vessel, storing the compressed air in the storage vessel, and selectively supplying compressed air from the compressed air storage vessel to the fuel manifold and/or the at least one fuel injector.

The method may comprise closing the fuel shut off valve and closing the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the fuel manifold and the at least one fuel injector.

The method may comprise closing the fuel shut off valve and opening the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the fuel manifold.

The method may comprise supplying compressed air from the compressed air storage vessel through the at least one fuel injector.

The at least one fuel injector may be a lean burn fuel injector comprising a pilot fuel nozzle and a main fuel nozzle, a main fuel manifold arranged to supply fuel to the main fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe and a pilot fuel nozzle arranged to supply fuel to the pilot fuel nozzle of the at least one fuel injector.

The method may comprise closing the fuel shut off valve and closing the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the main fuel manifold and the main fuel nozzle of the at least one fuel injector.

The method may comprise closing the fuel shut off valve and closing the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the pilot fuel manifold and the pilot fuel nozzle of the at least one fuel injector.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

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

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

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

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

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

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390 cm (around 155 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 320 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip², where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and U_ti_P is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg'¹K'¹/(ms'¹)²). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

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

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

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg’¹s, 105 Nkg¹s, 100 Nkg’¹s, 95 Nkg¹s, 90 Nkg¹s, 85 Nkg¹s or 80 Nkg'¹s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 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). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 deg C (ambient pressure 101.3kPa, temperature 30 deg 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). The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

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

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent.

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

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

Purely by way of example, the cruise conditions may correspond to: a forward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of -55 deg C.

As used anywhere herein, “cruise” or “cruise conditions” may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Embodiments will now be described by way of example only, with reference to the Figures, in which:

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 engine;

Figure 4 is a schematic diagram of the combustion equipment of the gas turbine engine showing a fuel system and a fuel purging system.

Figure 5 is an enlarged cross-sectional view of the combustion equipment of the gas turbine engine showing a fuel system and a fuel purging system.

Figure 6 is an enlarged cross-sectional view of the combustion equipment of the gas turbine engine showing a fuel system and an alternative fuel purging system.

Figure 7 is a schematic diagram of the combustion equipment of the gas turbine engine showing an alternative fuel system and an alternative fuel purging system.

Figure 8 is a schematic diagram of the combustion equipment of the gas turbine engine showing another fuel system and another fuel purging system.

Figure 9 is an enlarged cross-sectional view of the combustion equipment of the gas turbine engine showing a fuel system and an alternative fuel purging system.

Figure 10 is an enlarged cross-sectional view of the combustion equipment of the gas turbine engine showing a fuel system and an alternative fuel purging system.

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, 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 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 precess 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 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine 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.

A fuel system 50 and a fuel purging system 60 for the gas turbine engine 10 is shown in figures 4 and 5. The combustion equipment 16 of the gas turbine engine 10 comprises at least one combustion chamber 42. The fuel system 50 comprises at least one fuel injector 52, an associated fuel delivery pipe 54 and a fuel manifold 56 arranged to supply fuel to the at least one fuel injector 52 through the associated fuel delivery pipe 54. Each combustion chamber 42 has at least one fuel injector 52 arranged in operation to supply fuel into the corresponding combustion chamber 42.

In this example the combustion equipment 16 of the gas turbine engine 10 comprises an annular combustion chamber 42. The annular combustion chamber 42 comprises a radially inner annular wall 42A, a radially outer annular wall 42B and an upstream end wall 42C. The fuel system 50 comprises a plurality of fuel injectors 52, a plurality of fuel delivery pipes 54 and a fuel manifold 56 arranged to supply fuel to the at least one fuel injector 52 through the associated fuel delivery pipe 54. Each fuel injector 52 has an associated fuel delivery pipe 54 fluidly inter-connecting the fuel injector 52 and the fuel manifold 56 to supply fuel from the fuel manifold 56 to the fuel injector 52. The annular combustion chamber 42 has a plurality of circumferentially spaced fuel injectors 52 arranged in apertures in the upstream end wall 42C and arranged in operation to supply fuel into the annular combustion chamber 42. The fuel system 50 also comprises a fuel pipe 58 arranged to supply fuel from a fuel supply (not shown) to the fuel manifold 56 and a fuel shut off valve 59 arranged in the fuel pipe 58 upstream of the fuel manifold 56, e.g. between the fuel supply and the fuel manifold 56. The fuel system 50 also has a fuel drainage valve 57 connected to the fuel manifold 56 to drain fuel from the fuel system 50. The fuel injectors 52 may each have a fuel scheduling valve.

The combustion equipment 16 of the gas turbine engine 10, as shown in figure 5, also comprises an inner combustion chamber casing 44 and an outer combustion chamber casing 46 surrounding the inner combustion chamber casing 46 and the at least one combustion chamber 42. The at least one combustion chamber 42 is arranged between the inner combustion chamber casing 44 and the outer combustion chamber casing 46. A diffuser 48 is arranged to supply compressed air from the high pressure compressor 15 to the at least one combustion chamber 42. The diffuser 48 is arranged in flow series between the high pressure compressor 15 and the at least one combustion chamber 42. The inner combustion chamber casing 44 and the outer combustion chamber casing 46 are sealingly secured to the diffuser 48 to define a chamber 49 around the at least one combustion chamber 42. An upstream end of the inner combustion chamber casing 44 is directly secured to the diffuser 48 and an upstream end of the outer combustion chamber casing 46 is secured to the diffuser 48 by a radially extending wall 47. Thus, the chamber 49 is defined by the inner combustion chamber casing 44, the outer combustion chamber casing 46, the radially extending wall 47 and the diffuser 48.

The fuel purging system 60 comprises a compressed air storage vessel 62, a first passage 64 to supply compressed air from the high pressure compressor 15 to the compressed air storage vessel 62 and a second passage 66 to supply compressed air from the compressed air storage vessel 62 to the fuel manifold 56, a first valve 68 in the first passage 64 and a second valve 70 in the second passage 66. The compressed air storage vessel 62 is arranged outside the outer combustion chamber casing 46 and the first passage 64 is fluidly connected to the chamber 49 defined by the inner combustion chamber casing 44, the outer combustion chamber casing 46, the radially extending wall 47 and the diffuser 48.

The first valve 68 may be a non-return valve and for example the first valve 68 may comprise a valve member movable against a spring between a closed position and an open position. Alternatively, the first valve may be an active valve and for example the first valve may be an electrically operated valve or a hydraulically operated valve. The second valve 70 may be an active valve and for example the second valve 70 may be an electrically operated valve or a hydraulically operated valve.

In operation of the gas turbine engine, as discussed above, compressed air is supplied by the high pressure compressor 15 via the diffuser 48 into the chamber 49 and then to the combustion equipment 16 and fuel is supplied via the fuel pipe 58, the fuel manifold 56 and the delivery pipes 54 to the fuel injectors 52 for the combustion of fuel in the combustion equipment 16. During the operation of the gas turbine engine 10 the valve 68 is open/or opens so that some of the compressed air is supplied from the chamber 49 via the passage 64 and the valve 68 into the compressed air storage vessel 62 and the valve 70 is closed to prevent compressed air being supplied to the fuel system 50. When the gas turbine engine 10 stops operating the fuel supply is terminated but fuel remains in the fuel manifold 56, the fuel delivery pipes 54 and the fuel injectors 52. In order to prevent the fuel overheating and causing coking and potentially blockage within the fuel manifold 56 and the fuel injectors 52 the fuel shut off valve 59 is closed, the fuel drainage valve 57 is maintained closed, and then the second valve 70 is opened to allow the compressed air in the compressed air storage vessel 62 to flow through the passage 66 and flush, or purge, the fuel from the fuel manifold 56 and the fuel injectors 52. Alternatively, to prevent the fuel overheating and causing coking and potentially blockage within the fuel manifold 56 the fuel shut off valve 59 is closed, the fuel drainage valve 57 is opened, and then the second valve 70 is opened to allow the compressed air in the compressed air storage vessel 62 to flow through the passage 66 and flush, or purge, the fuel from the fuel manifold 56. As mentioned above the valve 68 is a non-return valve or the valve 68 is closed before purging so that the compressed air in the compressed air storage vessel 62 is supplied to the fuel system 50.

An alternative fuel purging system 160 for the gas turbine engine 10 is shown in figure 6. The fuel system 50 is substantially the same as the fuel system 50, and like parts are denoted by like numerals. The fuel purging system 160 is substantially the same as the fuel purging system 60, and like parts are denoted by like numerals. The fuel purging system 160 differs in that the compressed air storage vessel 162 is arranged inside the outer combustion chamber casing 46. In this example the compressed air storage vessel 160 is defined by the outer combustion chamber casing 46, the radially extending wall 47, the diffuser 48 and a further wall 51 sealingly secured to the diffuser 48 and the outer combustion chamber casing 46 and the first passage 64 is an opening in the further wall 51. The further wall 51 is also a radially extending wall and is arranged axially spaced from the radially extending wall 47. The radially extending wall 47 may be arranged in an axial midregion of the diffuser 48 and the further wall 51 is arranged at the downstream end of the diffuser 48. The first valve 68 may be arranged in the opening 64. The fuel purging system 160 works in substantially the same way the fuel purging system 60.

As shown in figures 4, 5 and 6 the at least one fuel injector 52 is a rich bum fuel injector comprising a main fuel nozzle, the fuel manifold 56 is arranged to supply fuel to the main fuel nozzle of the at least one fuel injector 52 through the associated fuel delivery pipe 54. In particular the fuel manifold 56 is arranged to supply fuel to the main fuel nozzle of a plurality of fuel injectors 52 through the associated fuel delivery pipes 54.

In an alternative arrangement, as shown in figure 7, the at least one fuel injector 252 is a lean bum fuel injector comprising a pilot fuel nozzle and a main fuel nozzle. The fuel system 250 for the at least one lean bum fuel injector 252 comprises a pilot fuel pipe 258, a pilot shut off valve 259, a pilot fuel manifold 256, a pilot fuel drainage valve 257 and an associated fuel delivery pipe 254 to supply fuel to the pilot fuel nozzle for the at least one lean bum fuel injector 252 and a main fuel pipe 58, a main fuel shut off valve 59, a main fuel manifold 56, a main fuel drainage valve 57 and an associated fuel delivery pipe 54 to supply fuel to the main fuel nozzle for the at least one lean bum fuel injector 252. In particular the pilot fuel manifold 256 is arranged to supply fuel to the pilot fuel nozzle of a plurality of lean burn fuel injectors 252 through the associated fuel delivery pipes 254 and the main fuel manifold 56 is arranged to supply fuel to the main fuel nozzle of a plurality of lean bum fuel injectors 252 through the associated fuel delivery pipes 54. A fuel purging system 60, 260 may be used with the main fuel manifold 56 and the main fuel nozzle(s) of the lean bum fuel injector(s) 252 or the pilot fuel manifold 256 and the pilot fuel nozzle(s) of the lean bum fuel injector(s) 252. In particular, as shown in figure 7 a main fuel purging system 60 may be used with the main fuel manifold 56 and the main fuel nozzle(s) of the lean bum fuel injector(s) 252 and a pilot fuel purging system 260 may be used with the pilot fuel manifold 256 and the pilot fuel nozzle(s) of the lean bum fuel injector(s) 252.

The main fuel purging system 60 comprises a compressed air storage vessel 62, a first passage 64, a second passage 66, a first valve 68 and a second valve 70 as discussed above and the compressed air storage vessel 62 may be arranged outside the outer combustion chamber casing 46 as in figure 5 or inside the outer combustion chamber casing 46 as in figure 6. The pilot fuel purging system 260 comprises a compressed air storage vessel 262, a first passage 264, a second passage 266, a first valve 268 and a second valve 270 as discussed above and the compressed air storage vessel 262 may be arranged outside the outer combustion chamber casing 46 as in figure 5 or inside the outer combustion chamber casing 46 as in figure 6.

In operation of a combustion chamber with lean burn fuel injectors 252 there are operating conditions, at engine idle and low power conditions, in which the main fuel nozzles are not supplied with fuel while the pilot fuel nozzles are supplied with fuel and the use of a main fuel purging system 60 to purge fuel from the main fuel manifold 56 and main fuel nozzles of the lean bum fuel injectors 252 is desirable. In order to prevent the fuel overheating and causing coking and potentially blockage within the main fuel manifold 56 and the main fuel nozzles of the lean burn fuel injectors 252 the main fuel shut off valve 59 is closed, the main fuel drainage valve 57 is maintained closed, and then the second valve 70 is opened to allow the compressed air in the compressed air storage vessel 62 to flow through the passage 66 and flush, or purge, the fuel from the main fuel manifold 56 and the main fuel nozzles of the lean burn fuel injectors 252. Alternatively, to prevent the fuel overheating and causing coking and potentially blockage within the main fuel manifold 56 the main fuel shut off valve 59 is closed, the main fuel drainage valve 57 is opened, and then the second valve 70 is opened to allow the compressed air in the compressed air storage vessel 62 to flow through the passage 66 and flush, or purge, the fuel from the main fuel manifold 56.

Both fuel purging systems 60 and 260 may be used when the gas turbine engine 10 ceases operation, e.g. the main fuel shut off valve 59 and main fuel drainage valve 57 are closed and then the valve 70 is opened to allow compressed air to be supplied from the compressed air storage vessel 62 through the main fuel manifold 56 and the main fuel nozzles of the lean bum injectors 252 and the pilot fuel shut off valve 259 and pilot fuel drainage valve 257 are closed and then the valve 270 is opened to allow compressed air to be supplied from the compressed air storage vessel 262 through the pilot fuel manifold 256 and the pilot fuel nozzles of the lean burn injectors 252. Alternatively, the main fuel shut off valve 59 is closed and the main fuel drainage valve 57 is opened and then the valve 70 is opened to allow compressed air to be supplied from the compressed air storage vessel 62 through the main fuel manifold 56 and the pilot fuel shut off valve 259 is closed and the pilot fuel drainage valve 257 is opened and then the valve 270 is opened to allow compressed air to be supplied from the compressed air storage vessel 262 through the pilot fuel manifold 256. It may also be possible to purge the fuel from the main fuel manifold 56 and main fuel nozzles of the lean burn fuel injectors 252 and only purge fuel from the pilot fuel manifold 256 or to only purge the fuel from the main fuel manifold 56 and purge fuel from the pilot fuel manifold 256 and the pilot fuel nozzles of the lean burn fuel injectors 252.

As mentioned above the first valves 68 and 268 are non-return valves or the valves 68 and 268 are closed before purging so that the compressed air in the compressed air storage vessels 62 and 262 is supplied from the associated fuel purging system 60 and 260 to the main fuel manifold 56 and pilot fuel manifold 256 respectively.

In another arrangement, as shown in figure 8, the at least one fuel injector 352 is a rich burn fuel injector comprising a main fuel nozzle. The fuel system 350 for the at least one rich burn fuel injector 352 comprises a first fuel pipe 358A, a first fuel shut off valve 359A, a first fuel manifold 356A, a first fuel drainage valve 357A and an associated fuel delivery pipe 354A to supply fuel to the main fuel nozzle of at least one rich bum fuel injector 352A and a second fuel pipe 358B, a second fuel shut of valve 359B, a second fuel manifold 356B, a second fuel drainage valve 357B and an associated fuel delivery pipe 354B to supply fuel to the main fuel nozzle of at least one rich burn fuel injector 352B. In particular the first fuel manifold 356A is arranged to supply fuel to the main fuel nozzle of a plurality of rich bum fuel injectors 352A through the associated fuel delivery pipes 354A and the second fuel manifold 356B is arranged to supply fuel to the main fuel nozzle of a plurality of rich bum fuel injectors 352B through the associated fuel delivery pipes 354B. A fuel purging system 360A, 360B may be used with the main fuel manifold 356A and the main fuel nozzle(s) of the rich bum fuel injector(s) 352A or the main fuel manifold 356B and the main fuel nozzle(s) of the rich bum fuel injector(s) 352B. In particular, as shown in figure 8 a first fuel purging system 360A may be used with the first fuel manifold

356A and the main fuel nozzle(s) of the rich burn fuel injector(s) 352A and a second fuel purging system 360B may be used with the second fuel manifold 356B and the main fuel nozzle(s) of the rich burn fuel injector(s) 352B.

The first and second fuel purging systems 360A and 360B each comprises a compressed air storage vessel 362, a first passage 364, a second passage 366, a first valve 368 and a second valve 370 as discussed above and the compressed air storage vessel 362 may be arranged outside the outer combustion chamber casing 46 as in figure 5 or inside the outer combustion chamber casing 46 as in figure 6.

In operation of a combustion chamber with rich burn fuel injectors 352A and 352B there are operating conditions, at engine idle and low power conditions known as staged combustion, in which the main fuel nozzles of the rich bum fuel injectors 352B are not supplied with fuel while the main fuel nozzles of the rich burn fuel injectors 352A are supplied with fuel and the use of a purging system 360B to purge fuel from the second fuel manifold 356B and main fuel nozzles of the rich burn fuel injectors 352B is desirable. The second fuel shut off valve 359B and the second fuel drainage valve 357B are closed and then the valve 370 is opened to allow compressed air to be supplied from the compressed air storage vessel 362 of fuel purging system 360B through the second fuel manifold 356B and the main fuel nozzles of the rich burn fuel injectors 352B. Alternatively, the second fuel shut off valve 359B is closed and the second fuel drainage valve 357B is opened and then the valve 370 is opened to allow compressed air to be supplied from the compressed air storage vessel 362 of fuel purging system 360B only through the second fuel manifold 356B.

Both fuel purging systems 360A, 360B may be used when the gas turbine engine 10 ceases operation, e.g. the second fuel shut off valve 359B and the second fuel drainage valve 357B are closed and then the valve 370 is opened to allow compressed air to be supplied from the compressed air storage vessel 362 of fuel purging system 360B through the second fuel manifold 356B and the main fuel nozzles of the rich burn fuel injectors 352B and the first fuel shut off valve 359A and the first fuel drainage valve 357A are closed and then the valve 370 is opened to allow compressed air to be supplied from the compressed air storage vessel 362 of fuel purging system 360A through the first fuel manifold 356A and the main fuel nozzles of the rich burn fuel injectors 352A. Alternatively, the second fuel shut off valve 359B is closed and the second fuel drainage valve 357B is opened and then the valve 370 is opened to allow compressed air to be supplied from the compressed air storage vessel 362 of fuel purging system 360B only through the second fuel manifold 356B and the first fuel shut off valve 359A is closed and the first fuel drainage valve 357A is opened and then the valve 370 is opened to allow compressed air to be supplied from the compressed air storage vessel 362 of fuel purging system 360A only through the first fuel manifold 356A.

As mentioned above the first valves 368 are non-return valves or are closed before purging so that the compressed air in the compressed air storage vessels 362 is supplied from the associated fuel purging system 360A and 360B to the first fuel manifold 356A and the second fuel manifold 356B respectively.

In an example a compressed air storage vessel has a volume of 0.01 m³ where air is stored at a pressure of 22bar and at a temperature of 500K, this has taken into account the reduction in pressure due to the temperature decreasing after the compressed air storage vessel has been filled at MTO condition, e.g. 900K and 40 bar. It is possible to supply the whole of the air stored in the compressed air storage vessel through the existing fuel system, e.g. fuel manifold, fuel delivery pipes and fuel injectors for nearly 3 seconds achieving fuel purging of any fuel trapped in the fuel system.

Figure 9 shows an alternative fuel purging system 460 for a fuel system, and is substantially the same as that shown in figure 5. The fuel purging system 460 differs in that the second passage 66 is connected to the fuel delivery pipe 54 of one or more of the fuel injectors 52 to purge fuel from the fuel nozzle within the fuel injector 52 and not to purge fuel from the associated fuel manifold. In particular the second passage 66 is connected to the fuel delivery pipe 54 downstream of the fuel scheduling valve within the fuel injector 52. This arrangement may also be used for the main fuel nozzles of the lean bum fuel injectors of figure 7 and/or for the pilot fuel nozzles of the lean burn fuel injectors of figure 7. This arrangement may also be used for the either one or both of the first and second fuel injectors of figure 8.

Figure 10 shows an alternative fuel purging system 560 for a fuel system, and is substantially the same as that shown in figure 6. The fuel purging system 560 differs in that the second passage 66 is connected to the fuel delivery pipe 54 of one or more of the fuel injectors 52 to purge fuel from the fuel nozzle within the fuel injector 52 and not to purge fuel from the associated fuel manifold. In particular the second passage 66 is connected to the fuel delivery pipe 54 downstream of the fuel scheduling valve within the fuel injector 52.This arrangement may also be used for the main fuel nozzles of the lean bum fuel injectors of figure 7 and/or for the pilot fuel nozzles of the lean burn fuel injectors of figure 7. This arrangement may also be used for the either one or both of the first and second fuel injectors of figure 8.

The fuel purging system reduces or avoids coking of the fuel injectors. The fuel purging system uses air at high pressure stored during the high power operation of the gas turbine engine. The compressed air storage vessel of the fuel purging system is a high delta pressure system with high purging efficiency and provides almost instantaneous purging of fuel from the fuel system and a long operating time is not required. The fuel purging system may be used for lean burn fuel injectors and rich burn fuel injectors. The fuel purging system may be used on aero gas turbine engines, marine gas turbine engines, automotive gas turbine engines, locomotive gas turbine engines or industrial gas turbine engines. The compressed air storage vessel of the fuel purging system may be integrated into the existing architecture of the gas turbine engine, e.g. integrated into the combustion chamber casing, diffuser structure which increases the structural robustness, strength. The fuel purging system does not require other substances, e.g. materials or liquids to clean coking deposits from within the fuel system. The fuel purging system does not require a battery and a blower to purge the fuel system.

The fuel purging system may be used in three different methods, e.g. firstly the fuel purging system may be used to purge fuel from the fuel manifold and each fuel injector supplied with fuel by the fuel manifold, secondly the fuel purging system may be used to purge fuel from the fuel manifold but not purge fuel from the fuel injectors and thirdly the fuel purging system may be used to purge fuel from one or more of the fuel injectors but not purge fuel from the fuel manifold.

It will be understood that the invention is not limited to the embodiments abovedescribed 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 subcombinations of one or more features described herein.

Claims

1. A gas turbine engine comprising a compressor, at least one combustion chamber, a fuel system and a fuel purging system, the fuel system comprises at least one fuel injector, an associated fuel delivery pipe and a fuel manifold arranged to supply fuel to the at least one fuel injector through the associated fuel delivery pipe, each combustion chamber has at least one fuel injector, the fuel purging system comprises a compressed air storage vessel, a first passage to supply compressed air from the compressor to the compressed air storage vessel and a second passage to supply compressed air from the compressed air storage vessel to the fuel manifold and/or the at least one fuel injector, a first valve in the first passage and a second valve in the second passage.

2. A gas turbine engine as claimed in claim 1 wherein the gas turbine engine comprises an inner combustion chamber casing, an outer combustion chamber casing surrounding the inner combustion chamber casing and the at least one combustion chamber, the at least one combustion chamber being arranged between the inner combustion chamber casing and the outer combustion chamber casing, a diffuser to supply compressed air from the compressor to the at least one combustion chamber, the inner combustion chamber casing and the outer combustion chamber casing being sealingly secured to the diffuser to define a chamber around the at least one combustion chamber.

3. A gas turbine engine as claimed in claim 2 wherein the compressed air storage vessel is arranged outside the outer combustion chamber casing and the first passage is fluidly connected to the chamber.

4. A gas turbine engine as claimed in claim 2 wherein the compressed air storage vessel is arranged inside the outer combustion chamber casing.

5. A gas turbine engine as claimed in claim 4 wherein the compressed air storage vessel is defined by the outer combustion chamber casing, the diffuser and a further wall sealingly secured to the diffuser and the outer combustion chamber casing and the first passage is an opening in the further wall.

6. A gas turbine engine as claimed in any of claims 1 to 5 wherein the first valve is a non-return valve.

7. A gas turbine engine as claimed in any of claim 1 to 6 wherein the first valve comprises a valve member movable against a spring between a closed position and an open position.

8. A gas turbine engine as claimed in any of claims 1 to 5 wherein the first valve is an active valve.

9. A gas turbine engine as claimed in claim 8 wherein the first valve is an electrically operated valve or a hydraulically operated valve.

10. A gas turbine engine as claimed in any of claims 1 to 9 wherein the second valve is an active valve.

11. A gas turbine engine as claimed in claim 10 wherein the second valve is an electrically operated valve or a hydraulically operated valve.

12. A gas turbine engine as claimed in any of claims 1 to 11 wherein the at least one fuel injector is a rich burn fuel injector comprising a main fuel nozzle, the fuel manifold is arranged to supply fuel to the main fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe.

13. A gas turbine engine as claimed in any of claims 1 to 11 wherein the at least one fuel injector is a lean burn injector comprising a pilot fuel nozzle and a main fuel nozzle.

14. A gas turbine engine as claimed in claim 13 wherein the fuel manifold is arranged to supply fuel to the main fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe.

15. A gas turbine engine as claimed in claim 13 or claim 14 wherein the fuel manifold is arranged to supply fuel to the pilot fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe.

16. A gas turbine engine as claimed in any of claims 1 to 15 wherein the fuel system comprises a fuel shut off valve arranged to supply fuel to the fuel manifold.

17. A gas turbine engine as claimed in any of claims 1 to 16 wherein the fuel system comprises a fuel drainage valve arranged to drain fuel from the fuel manifold.

18. A method of purging fuel from a gas turbine engine, the gas turbine engine comprising a compressor, at least one combustion chamber, a fuel system and a fuel purging system, the fuel system comprises at least one fuel injector, an associated fuel delivery pipe and a fuel manifold arranged to supply fuel to the at least one fuel injector through the associated fuel delivery pipe, each combustion chamber has at least one fuel injector, the method comprising supplying compressed air from the compressor to a compressed air storage vessel, storing the compressed air in the storage vessel, and selectively supplying compressed air from the compressed air storage vessel to the fuel manifold and/or the at least one fuel injector.

19. A method as claimed in claim 18 wherein the fuel system comprises a fuel shut off valve arranged to supply fuel to the fuel manifold and a fuel drainage valve arranged to drain fuel from the fuel manifold.

20. A method as claimed in claim 19 wherein the method comprises closing the fuel shut off valve and closing the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the fuel manifold and the at least one fuel injector.

21. A method as claimed in claim 19 wherein the method comprises closing the fuel shut off valve and opening the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the fuel manifold.

22. A method as claimed in claim 18 or claim 19 wherein the method comprises supplying compressed air from the compressed air storage vessel through the at least one fuel injector.

23. A method as claimed in claim 18 or claim 19 wherein the at least one fuel injector may be a lean bum fuel injector comprising a pilot fuel nozzle and a main fuel nozzle, a main fuel manifold arranged to supply fuel to the main fuel nozzle of the at least one fuel injector through the associated fuel delivery pipe and a pilot fuel nozzle arranged to supply fuel to the pilot fuel nozzle of the at least one fuel injector.

24. A method as claimed in claim 23 comprising closing the fuel shut off valve and closing the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the main fuel manifold and the main fuel nozzle of the at least one fuel injector.

25. A method as claimed in claim 23 or claim 24 comprising closing the fuel shut off valve and closing the fuel drainage valve and supplying compressed air from the compressed air storage vessel through the pilot fuel manifold and the pilot fuel nozzle of the at least one fuel injector.

26. A gas turbine engine for an aircraft comprising:

an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein: the engine core further comprising at least one combustion chamber, a fuel system and a fuel purging system, the fuel system comprises at least one fuel injector, an associated fuel delivery pipe and a fuel manifold arranged to supply fuel to the at least one fuel injector through the associated fuel delivery pipe, each combustion chamber has at least one fuel injector, the fuel purging system comprises a compressed air storage vessel, a first passage to supply compressed air from the compressor to the compressed air storage vessel and a second passage to supply compressed air from the compressed air storage vessel to the fuel manifold and/or the at least one fuel injector, a first valve in the first passage and a second valve in the second passage.

27. The gas turbine engine according to claim 23 wherein:

the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;

the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.