GB2572753A

GB2572753A - A fuel system for an internal combustion engine, an internal combustion engine and a method of operating a fuel system for an internal combustion engine

Info

Publication number: GB2572753A
Application number: GB1805505.3A
Authority: GB
Inventors: Wei Huang Hua; Carlos Roman Casado Juan; Tentorio Luca; Di Chiaro Giacomo; Knapton Jonathan
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2018-04-04
Filing date: 2018-04-04
Publication date: 2019-10-16
Also published as: GB201805505D0

Abstract

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. The fuel system 50 has at least one thermo electric device 60, each thermo-electric device 60 has a first surface (68, figure 5) arranged in heat transfer arrangement with the fuel system 50, 52, 54, 56 and a second surface (80) arranged in heat transfer arrangement with a heat sink (82) and each thermo-electric device 60 is electrically connected to an electric power supply 72. Thermoelectric devices 60 may be arranged in a heat transfer arrangement with one or more of the fuel manifold 56, a fuel injector 52 and a fuel delivery pipe 54. Thermoelectric devices 60 may be arranged circumferentially around the fuel manifold in the same plane and may comprise two half annular thermoelectric devices 60. Two half annular thermoelectric devices may be arranged circumferentially around each fuel delivery pipe 54 in the same plane. Thermoelectric devices 60 may be arranged around an internal conduit of or on the flange of each fuel injector 52.

Description

A FUEL SYSTEM FOR AN INTERNAL COMBUSTION ENGINE, AN INTERNAL COMBUSTION ENGINE AND A METHOD OF OPERATING A FUEL SYSTEM FOR AN INTERNAL COMBUSTION ENGINE

The present disclosure relates to a fuel system for an internal combustion engine and in particular to a fuel system for a gas turbine engine.

A gas turbine engine has one or more combustion chambers in which combustion occurs. The 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 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.

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.

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

-2the 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 an internal combustion engine which reduces or overcomes the above mentioned problem.

According to a first aspect there is provided a fuel system for an internal combustion engine, 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, the fuel system has at least one thermo-electric device, each thermo-electric device has a first surface arranged in heat transfer arrangement with the fuel system and a second surface arranged in heat transfer arrangement with a heat sink and each thermo-electric device is electrically connected to an electric power supply.

The thermo-electric device uses the Peltier effect to cool the fuel system.

The thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold. The thermo-electric device may be arranged in heat transfer arrangement with a fuel injector. The thermo-electric device may be arranged in heat transfer arrangement with an associated fuel delivery pipe of a fuel injector.

There may be a plurality of thermo-electric devices. At least one thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold and at least one thermo-electric device is arranged in heat transfer arrangement with a fuel injector. At least one thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold and at least one thermo-electric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a fuel injector. At least one thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric device is arranged in heat transfer arrangement with a corresponding one of the fuel injectors. At least one thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric

-3 device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a corresponding one of the fuel injectors. A plurality of thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric device is arranged in heat transfer arrangement with a corresponding one of the fuel injectors. A plurality of thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a corresponding one of the fuel injectors.

At least one thermo-electric device may be arranged in heat transfer arrangement with the fuel manifold, at least one thermo-electric device is arranged in heat transfer arrangement with a fuel injector and at least one thermo-electric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a fuel injector.

There may be a plurality of thermo-electric devices arranged circumferentially around the fuel manifold in the same plane. There may be two half annular thermo-electric devices arranged circumferentially around the fuel manifold in the same plane.

There may be a plurality of thermo-electric devices arranged circumferentially around each associated fuel delivery pipe in the same plane. There may be two half annular thermo-electric devices arranged circumferentially around each associated fuel delivery pipe in the same plane.

There may be a thermo-electric device arranged around an internal conduit within a fuel injector. There may be a thermo-electric device arranged on the flange of a fuel injector. There may be a thermo-electric device arranged around an internal conduit within each fuel injector. There may be a thermo-electric device arranged on the flange of each fuel injector.

The fuel system may comprise a fuel pipe arranged to supply fuel from a fuel supply to the fuel manifold. At least one thermo-electric device may be arranged in heat transfer arrangement with the fuel pipe. A plurality of thermo-electric devices may be arranged in heat transfer arrangement with the fuel pipe.

-4There may be a plurality of thermo-electric devices arranged circumferentially around the fuel pipe in the same plane. There may be two half annular thermo-electric devices arranged circumferentially around the fuel pipe in the same plane.

A heat conductor may be arranged between the thermo-electric device and the fuel system to facilitate the transfer of heat between the fuel system and the thermoelectric device.

The electrical power supply may be an electrical generator. The electrical generator may be mounted on and driven by the internal combustion engine. The electrical generator may be mounted on and driven by an auxiliary internal combustion engine. The internal combustion engine and the auxiliary internal combustion engine may be mounted on an associated vehicle. The electrical power supply may be a battery. The battery may be mounted on an associated vehicle.

The internal combustion engine may be a gas turbine engine. The auxiliary internal combustion engine may be an auxiliary power unit. The associated vehicle may be an aircraft.

The gas turbine engine may have a bypass duct and the heat sink may be arranged in the bypass duct or the heat sink may be arranged to release heat to air in the bypass duct. The gas turbine engine may have a casing, e.g. a turbine casing, and the heat sink may be the casing.

The fuel system may have a temperature sensor to detect the temperature of the fuel within the fuel system and a controller arranged to receive a signal from the temperature sensor and to control the supply of an electrical voltage to the thermoelectric device.

The controller may be arranged to supply an electrical voltage to the thermo-electric device such that heat is removed from the fuel system and supplied to the heat sink when the controller determines that the temperature provided by the temperature sensor is above a predetermined temperature.

- 5 The controller may be arranged to supply an electrical voltage to the thermo-electric device such that heat is removed from the heat sink and supplied to the fuel system when the controller determines that the temperature provided by the temperature sensor is below a predetermined temperature.

According to a second aspect there is provided an internal combustion engine comprising at least one combustion chamber and a fuel 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 arranged in operation to supply fuel into the corresponding combustion chamber, the fuel system has at least one thermo-electric device, each thermo-electric device has a first surface arranged in heat transfer arrangement with the fuel system and a second surface arranged in heat transfer arrangement with a heat sink and each thermo-electric device is electrically connected to an electric power supply.

According to a third aspect there is provided a method of operating a fuel system for an internal combustion engine, the internal combustion engine has at least one combustion chamber, 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 arranged in operation to supply fuel into the corresponding combustion chamber, the fuel system has at least one thermo-electric device, each thermo-electric device has a first surface arranged in heat transfer arrangement with the fuel system and a second surface arranged in heat transfer arrangement with a heat sink and each thermo-electric device is electrically connected to an electric power supply, the method comprising providing an electrical voltage to the at least one thermo-electric device to transfer heat from the fuel system to the heat sink or to transfer from heat from the heat sink to the fuel system.

-6The method may comprise measuring the temperature of the fuel within the system, and if the temperature of the fuel within the fuel system is above a first predetermined temperature providing an electrical voltage to the at least one thermoelectric device to transfer heat from the fuel system to the heat sink.

The method may comprise measuring the temperature of the fuel within the system, and if the temperature of the fuel within the fuel system is below a second predetermined temperature providing an electrical voltage to the at least one thermoelectric device to transfer from heat from the heat sink to the fuel system.

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

-7 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.

- 8 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

-9the 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/Uti_P ², where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg'¹K'¹/(ms'¹)²). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

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

- 10overall 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

- 11 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

- 12the 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

- 13 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 a fuel system for a gas turbine engine.

Figure 5 is a schematic diagram of a thermo-electric device for the fuel system for the gas turbine engine shown in Figure 4.

Figure 6 is a schematic diagram showing an arrangement of thermo-electric devices for the fuel system for the gas turbine engine shown in Figure 4.

- 14Figure 7 is a schematic diagram showing an alternative arrangement of thermoelectric devices for the fuel system for the gas turbine engine shown in Figure 4.

Figure 8 is a schematic diagram of an alternative fuel system for a gas turbine engine.

Figure 9 is a schematic diagram of an alternative fuel system for a gas turbine engine.

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.

- 15 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)

- 16gear 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

- 1720. 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), a turbojet engine 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 for the gas turbine engine 10 is shown in figure 4. 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, figure 4, the combustion equipment 16 of the gas turbine engine 10 comprises an annular combustion chamber 42. 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. The annular combustion chamber 42 has a plurality of circumferentially spaced fuel injectors 52 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.

- 18 The fuel system 50 comprises at least one thermo-electric device 60, as shown in figure 5. A thermoelectric device 60 produces a voltage when exposed to a temperature gradient or on the contrary, a thermo-electric 60 generates a temperature difference when an electrical voltage is applied to the thermo-electric device 60. One form of thermo-electric device 60 comprises two different semiconductors, a first semiconductor 62 is and n-type and the second semiconductor 64 is p-type and each has different electron densities. The semiconductors 62 and 64 are arranged thermally in parallel and are electrically connected in series by an electrical conductor 66. The semiconductors 62 and 64 are also joined by thermally conducting members 68 and 70 on opposite sides. The semiconductors 62 and 64 are electrically connected to an electrical power supply 72 by conductors 74 and 76. The electrical power supply 72 applies a voltage across two semiconductors 62 and 64 and causes an electrical current to flow across the semiconductors’ junction, which produces a temperature difference. A first surface 78 of the thermally conducting member 68 absorbs heat which is then transferred to a second surface 80 of the thermally conducting member 70.

Each thermo-electric device 60 has the first surface 78 arranged in heat transfer arrangement with the fuel system 50 and a second surface 80 arranged in heat transfer arrangement with a heat sink 82 and each thermo-electric device 60 is electrically connected to an electric power supply 72. The thermo-electric device 60 uses the Peltier effect to cool the fuel system.

The thermo-electric device 60 may be arranged in heat transfer arrangement with the fuel manifold 56, the thermo-electric device 60 may be arranged in heat transfer arrangement with a fuel injector 52, the thermo-electric device 60 may be arranged in heat transfer arrangement with an associated fuel delivery pipe 54 of a fuel injector 52 or the thermo-electric device 60 may be arranged in heat transfer arrangement with the fuel pipe 58.

The fuel system 50 may comprise a plurality of thermo-electric devices 60. In one arrangement at least one thermo-electric device 60 is arranged in heat transfer arrangement with the fuel manifold 56 and at least one thermo-electric device 60 is arranged in heat transfer arrangement with a fuel injector 52. In another

- 19arrangement at least one thermo-electric device 60 is arranged in heat transfer arrangement with the fuel manifold 56 and at least one thermo-electric device 60 is arranged in heat transfer arrangement with an associated fuel delivery pipe 54 of a fuel injector 52. In an additional arrangement at least one thermo-electric device 60 is arranged in heat transfer arrangement with the fuel manifold 56, at least one thermo-electric device 60 is arranged in heat transfer arrangement with a fuel injector 52 and at least one thermo-electric device 60 is arranged in heat transfer arrangement with an associated fuel delivery pipe 54 of a fuel injector 52. In a further arrangement at least one thermo-electric device 60 is arranged in heat transfer arrangement with the fuel manifold 56 and each one of a plurality of thermoelectric devices 60 is arranged in heat transfer arrangement with a corresponding one of the fuel injectors 52. In another arrangement at least one thermo-electric 60 is arranged in heat transfer arrangement with the fuel manifold 56 and each one of a plurality of thermo-electric devices 60 is arranged in heat transfer arrangement with an associated fuel delivery pipe 54 of a corresponding one of the fuel injectors 52. In an additional arrangement a plurality of thermo-electric devices 60 are arranged in heat transfer arrangement with the fuel manifold 56 and each one of a plurality of thermo-electric devices 60 is arranged in heat transfer arrangement with a corresponding one of the fuel injectors 53. A plurality of thermo-electric devices 60 are arranged in heat transfer arrangement with the fuel manifold 56 and each one of a plurality of thermo-electric devices 60 is arranged in heat transfer arrangement with an associated fuel delivery pipe 54 of a corresponding one of the fuel injectors 52.

A plurality of thermo-electric 60 devices may be arranged in heat transfer arrangement with the fuel pipe 58.

In the example shown in figure 4 the fuel system 50 comprises at least one thermoelectric device 60 arranged in heat transfer arrangement with the fuel pipe 58, at least one thermo-electric device 60 arranged in heat transfer arrangement with the fuel manifold 56 and at least one thermo-electric device 60 is arranged in heat transfer arrangement with the associated fuel delivery pipe 54 of each fuel injector 52.

In operation each thermo-electric device 60 in the fuel system 50 of the gas turbine engine 10 extracts heat from the fuel and discharges the heat to the heat sink 82 to

-20reduce the temperature of the fuel within the fuel system 50 to levels low enough to prevent 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 56, the fuel delivery pipes 54 and the fuel injectors 52 and therefore extend the working life of the component parts of the fuel system 52 of the gas turbine engine 10. The fuel also circulates though the fuel system 52 and is exposed to high temperatures because it is used as hydraulic medium to operate compressor variable stator vanes and to act as a heat sink for the lubrication system. The thermo-electric devices 60 thus reduce the temperature of the fuel, slowing or stopping the rate of coking and deposit formation in sensitive elements, such as in the passages within the fuel injectors 52. The thermo-electric devices 60 are arranged to discharge the heat to the heat sink 82 and the heat sink 82 may be bypass air in the bypass duct 22 or a casing, e.g. a compressor casing or a turbine casing, of the gas turbine engine 10.

Figure 6 shows an arrangement of thermo-electric devices 60 arranged circumferentially around a portion of the fuel system 52 in the same plane. The thermo-electric devices 60 may be arranged in the same plane circumferentially around the fuel delivery pipe 54, the fuel manifold 56 or the fuel pipe 58. There may be a plurality of thermo-electric devices 60 arranged in the same plane circumferentially around each associated fuel delivery pipe 54.

Figure 7 shows another arrangement of thermo-electric devices 60 arranged circumferentially around a portion of the fuel system 52 in the same plane. In this arrangement there are two half annular thermo-electric devices 60 arranged circumferentially around the fuel system 52 in the same plane. The thermo-electric devices 60 may be arranged in the same plane circumferentially around the fuel delivery pipe 54, the fuel manifold 56 or the fuel pipe 58. There may be two half annular thermo-electric devices 60 arranged in the same plane circumferentially around each associated fuel delivery pipe 54.

There may be a thermo-electric device arranged around an internal conduit within a fuel injector. There may be a thermo-electric device arranged on the flange of a fuel injector. There may be a thermo-electric device arranged around an internal conduit

-21 within each fuel injector. There may be a thermo-electric device arranged on the flange of each fuel injector.

An alternative fuel system 150 for the gas turbine engine 10 is shown in figure 8. The fuel system 150 is substantially the same as the fuel system 50, and like parts are denoted by like numerals. The fuel system 150 differs in that the fuel system 150 has at least one temperature sensor 84 to detect the temperature of the fuel within the fuel system 150 and a controller 86 arranged to receive a signal from the at least one temperature sensor 84 and to control the electric power supply 72 which supplies an electrical voltage to the thermo-electric device 60 or each thermo-electric device 60. The controller 86 may be arranged to control the electrical power supply 72 which supplies an electrical voltage to the thermo-electric device 60, or each thermo-electric device 60, such that heat is removed from the fuel system 150 and is supplied to the heat sink 82 when the controller 86 determines that the temperature provided by the at least one temperature sensor 84 is above a predetermined temperature. The controller 86 may be arranged to control the electrical power supply 72 which supplies an electrical voltage to the thermo-electric device 6, or each thermo-electric device 60, such that heat is removed from the heat sink 82 and is supplied to the fuel system 150 when the controller 86 determines that the temperature provided by the at least one temperature sensor 84 is below a predetermined temperature.

Figure 8 also shows thermo-electric devices 60 on the flanges 53 of the fuel injectors 52 rather than on the fuel delivery pipes 54 of the fuel injectors 52.

An alternative fuel system 250 for the gas turbine engine 10 is shown in figure 9. The fuel system 250 is substantially the same as the fuel system 50, and like parts are denoted by like numerals. The fuel system 250 differs in that the fuel system 250 comprises lean bum fuel injectors 252 whereas the fuel system 50 comprises rich bum fuel injectors. The fuel system 250 comprises a plurality of lean fuel

-22injectors 252, a plurality of pilot fuel delivery pipes 54A and a pilot fuel manifold 56A arranged to supply fuel to the at least one lean bum fuel injector 252 through the associated fuel delivery pipe 54A. Each lean burn fuel injector 252 has an associated pilot fuel delivery pipe 54A. The fuel system 250 also comprises a plurality of main fuel delivery pipes 54B and a main fuel manifold 56B arranged to supply fuel to the at least one lean burn fuel injector 252 through the associated fuel delivery pipe 54B. Each lean bum fuel injector 252 has an associated main fuel delivery pipe 54B. The annular combustion chamber 42 has a plurality of circumferentially spaced lean bum fuel injectors 252 arranged in operation to supply fuel into the annular combustion chamber 42. The fuel system 250 also comprises a pilot fuel pipe 58A arranged to supply fuel from a fuel supply (not shown) to the pilot fuel manifold 56A and a main fuel pipe 58B arranged to supply fuel from the fuel supply (not shown) to the main fuel manifold 58B.

The fuel system 250 comprises at least one thermo-electric device 60A, 60B. The thermo-electric device 60A may be arranged in heat transfer arrangement with the pilot fuel manifold 56A, the thermo-electric device 60A may be arranged in heat transfer arrangement with a lean bum fuel injector 252, the thermo-electric device 60A may be arranged in heat transfer arrangement with an associated pilot fuel delivery pipe 54A of a lean burn fuel injector 252 or the thermo-electric device 60A may be arranged in heat transfer arrangement with the pilot fuel pipe 58A. The thermo-electric device 60B may be arranged in heat transfer arrangement with the main fuel manifold 56B, the thermo-electric device 60B may be arranged in heat transfer arrangement with a lean bum fuel injector 252, the thermo-electric device 60B may be arranged in heat transfer arrangement with an associated main fuel delivery pipe 54B of a lean bum fuel injector 252 or the thermo-electric device 60B may be arranged in heat transfer arrangement with the main fuel pipe 58B.

It may be possible to provide a thermo-electric device 60A in heat transfer arrangement with the pilot fuel manifold 56A and a thermo-electric device 60A arranged in heat transfer arrangement with a lean bum fuel injector 252 or to provide a thermo-electric device 60A in heat transfer arrangement with the pilot fuel manifold 56A and a thermo-electric device 60A arranged in heat transfer arrangement with an associated pilot fuel delivery pipe 54A of a lean bum fuel injector 252. A thermo

-23 electric device 60A may be arranged in heat transfer arrangement with the pilot fuel pipe 58A.

It may be possible to provide a thermo-electric device 60B in heat transfer arrangement with the main fuel manifold 56B and a thermo-electric device 60B arranged in heat transfer arrangement with a lean burn fuel injector 252 or to provide a thermo-electric device 60B in heat transfer arrangement with the main fuel manifold 56B and a thermo-electric device 60B arranged in heat transfer arrangement with an associated main fuel delivery pipe 54B of a lean burn fuel injector 252. A thermo-electric device 60B may be arranged in heat transfer arrangement with the main fuel pipe 58B.

It may be possible to provide one or more thermo-electric device 60B in heat transfer arrangement with the main fuel system and none in heat transfer arrangement with the pilot fuel system.

The electrical power supply 72 may be an electrical generator. The electrical generator may be mounted on and driven by the gas turbine engine 10. The electrical generator may be mounted on and driven by an auxiliary gas turbine engine. The gas turbine engine and the auxiliary gas turbine engine may be mounted on an associated vehicle. The electrical power supply may be a battery. The battery may be mounted on an associated vehicle. The auxiliary internal gas turbine engine may be an auxiliary power unit. The associated vehicle may be an aircraft.

At 160°C, jet fuel has specific heat capacity of 2.5 KJ/Kg K, according to the Handbook of Aviation Fuels. Standard efficiency figures for Peltier effect cooling systems are in the order of 10%. As a result, up to 225 MJ of electrical energy are provided to the thermo-electric devices 60 of the fuel system 50 of a single gas turbine engine 10 to achieve a fuel temperature reduction of 20°C during a specified period.

Lowering the fuel temperature within the fuel system significantly increases the working life of the fuel injectors and reduces the risk of blockage due to carbon

-24deposition. It is believed that lowering the fuel temperature by 10°C doubles the working life of the fuel injectors. It is believed that a linear reduction in temperature between 250°C and 200°C results in an exponential increase in fuel injector life. Therefore, the engine maintenance and operating costs are reduced, because less replacement parts and labour are required at engine maintenance. The thermoelectric devices may be packaged into a small space and they are high reliability due to the lack of moving parts.

Although the disclosure has referred to a fuel system for a gas turbine engine, a type of continuous combustion internal combustion engine, it is equally applicable to other internal combustion engines, for example reciprocating internal combustion engines, rotary internal combustion engines and continuous combustion internal combustion engines e.g. gasoline engines, diesel engines and Wankel engines. Although the disclosure has referred to an aircraft as an associated vehicle it is equally applicable for the associated vehicle to be a marine vessel, a locomotive or an automotive vehicle. Thus, it is applicable to provide a continuous combustion internal combustion, a reciprocating internal combustion engine or a rotary internal combustion engine onto any one of an aircraft, a marine vessel, a locomotive or an automotive vehicle.

The gas turbine engine fuel injectors may be rich bum fuel injectors or lean bum fuel injectors.

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 fuel system for an internal combustion engine, 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, the fuel system has at least one thermo-electric device, each thermo-electric device has a first surface arranged in heat transfer arrangement with the fuel system and a second surface arranged in heat transfer arrangement with a heat sink and each thermo-electric device is electrically connected to an electric power supply.

2. A fuel system as claimed in claim 1 wherein the thermo-electric device is arranged in heat transfer arrangement with the fuel manifold, the thermo-electric device is arranged in heat transfer arrangement with a fuel injector or the thermoelectric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a fuel injector.

3. A fuel system as claimed in claim 1 or claim 2 wherein there are a plurality of thermo-electric devices.

4. A fuel system as claimed in claim 3 wherein at least one thermo-electric device is arranged in heat transfer arrangement with the fuel manifold and at least one thermo-electric device is arranged in heat transfer arrangement with a fuel injector.

5. A fuel system as claimed in claim 3 wherein at least one thermo-electric device is arranged in heat transfer arrangement with the fuel manifold and at least one thermo-electric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a fuel injector.

6. A fuel system as claimed I claim 3 wherein at least one thermo-electric device is arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric device is arranged in heat transfer arrangement with a corresponding one of the fuel injectors.

7. A fuel system as claimed in claim 3 wherein at least one thermo-electric device is arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a corresponding one of the fuel injectors.

8. A fuel system as claimed in claim 3 wherein a plurality of thermo-electric device are arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric device is arranged in heat transfer arrangement with a corresponding one of the fuel injectors.

9. A fuel system as claimed in claim 3 wherein a plurality of thermo-electric device are arranged in heat transfer arrangement with the fuel manifold and each one of a plurality of thermo-electric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a corresponding one of the fuel injectors.

10. A fuel system as claimed in claim 3 wherein at least one thermo-electric device is arranged in heat transfer arrangement with the fuel manifold, at least one thermo-electric device is arranged in heat transfer arrangement with a fuel injector and at least one thermo-electric device is arranged in heat transfer arrangement with an associated fuel delivery pipe of a fuel injector.

11. A fuel system as claimed in any of claims 1 to 10 wherein there are a plurality of thermo-electric devices arranged circumferentially around the fuel manifold in the same plane.

12. A fuel system as claimed in claim 11 wherein there are two half annular thermo-electric devices arranged circumferentially around the fuel manifold in the same plane.

13. A fuel system as claimed in any of claims 1 to 12 wherein there are a plurality of thermo-electric devices arranged circumferentially around each associated fuel delivery pipe in the same plane.

14. A fuel system as claimed in claim 13 wherein there are two half annular thermo-electric devices arranged circumferentially around each associated fuel delivery pipe in the same plane.

15. A fuel system as claimed in any of claims 1 to 14 wherein there is a thermoelectric device arranged around an internal conduit within a fuel injector.

16. A fuel system as claimed in claim 15 wherein there is a thermo-electric device arranged around an internal conduit within each fuel injector.

17. A fuel system as claimed in any of claims 1 to 15 wherein there is a thermoelectric device arranged on the flange of a fuel injector.

18. A fuel system as claimed in claim 17 wherein there is a thermo-electric device arranged on the flange of each fuel injector.

19. A fuel system as claimed in any of claims 1 to 18 wherein the fuel system comprises a fuel pipe arranged to supply fuel from a fuel supply to the fuel manifold.

20. A fuel system as claimed in claim 19 wherein at least one thermo-electric device is arranged in heat transfer arrangement with the fuel pipe.

21. A fuel system as claimed in claim 20 wherein a plurality of thermo-electric devices are arranged in heat transfer arrangement with the fuel pipe.

22. A fuel system as claimed in claim 21 wherein there are a plurality of thermoelectric devices arranged circumferentially around the fuel pipe in the same plane.

23. A fuel system as claimed in claim 22 wherein there are two half annular thermo-electric devices arranged circumferentially around the fuel pipe in the same plane.

24. A fuel system as claimed in any of claims 1 to 23 wherein a heat conductor is arranged between the thermo-electric device and the fuel system to facilitate the transfer of heat between the fuel system and the thermo-electric device.

25. A fuel system as claimed in any of claims 1 to 24 wherein the electrical power supply is an electrical generator.

26. A fuel system as claimed in claim 25 wherein the electrical generator is mounted on and driven by the internal combustion engine.

27. A fuel system as claimed in claim 25 wherein the electrical generator is mounted on and driven by an auxiliary internal combustion engine.

28. A fuel system as claimed in any of claims 1 to 28 wherein the electrical power supply is a battery.

29. A fuel system as claimed in claim 26, claim 27 or claim 27 wherein the internal combustion engine and/or the auxiliary internal combustion engine and/or the battery are mounted on an associated vehicle.

30. A fuel system as claimed in claim 29 wherein the associated vehicle is an aircraft.

31. A fuel system as claimed in claim 27 wherein the auxiliary internal combustion engine is an auxiliary power unit.

32. A fuel system as claimed in any of claims 1 to 31 wherein the internal combustion engine is a gas turbine engine.

33. A fuel system as claimed in any of claims 1 to 32 wherein the fuel system has a temperature sensor to detect the temperature of the fuel within the fuel system and a controller arranged to receive a signal from the temperature sensor and to control the supply of an electrical voltage to the thermo-electric device.

34. A fuel system as claimed in claim 33 wherein the controller is arranged to supply an electrical voltage to the thermo-electric device such that heat is removed from the fuel system and supplied to the heat sink when the controller determines that the temperature provided by the temperature sensor is above a predetermined temperature.

35. A fuel system as claimed in claim 33 wherein the controller is arranged to supply an electrical voltage to the thermo-electric device such that heat is removed from the heat sink and supplied to the fuel system when the controller determines that the temperature provided by the temperature sensor is below a predetermined temperature.

36. An internal combustion engine comprising at least one combustion chamber and a fuel 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 arranged in operation to supply fuel into the corresponding combustion chamber, the fuel system has at least one thermo-electric device, each thermo-electric device has a first surface arranged in heat transfer arrangement with the fuel system and a second surface arranged in heat transfer arrangement with a heat sink and each thermo-electric device is electrically connected to an electric power supply.

37. A method of operating a fuel system for an internal combustion engine, the internal combustion engine has at least one combustion chamber, 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 arranged in operation to supply fuel into the corresponding combustion chamber, the fuel system has at least one thermo-electric device, each thermo-electric device has a first surface arranged in heat transfer arrangement with the fuel system

-30and a second surface arranged in heat transfer arrangement with a heat sink and each thermo-electric device is electrically connected to an electric power supply, the method comprising providing an electrical voltage to the at least one thermo-electric device to transfer heat from the fuel system to the heat sink or to transfer from heat from the heat sink to the fuel system.

38. A method as claimed in claim 37 comprising measuring the temperature of the fuel within the system, and if the temperature of the fuel within the fuel system is above a first predetermined temperature providing an electrical voltage to the at least one thermo-electric device to transfer heat from the fuel system to the heat sink.

39. A method as claimed in claim 37 or claim 38 comprising measuring the temperature of the fuel within the system, and if the temperature of the fuel within the fuel system is below a second predetermined temperature providing an electrical voltage to the at least one thermo-electric device to transfer from heat from the heat sink to the fuel system.

40. 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 gas turbine engine has at least one combustion chamber, 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 arranged in operation to supply fuel into the corresponding combustion chamber, the fuel system has at least one thermo-electric device, each thermo-electric device has a first surface arranged in heat transfer arrangement with the fuel system and a second surface arranged in heat transfer arrangement with a heat sink and each thermo-electric device is electrically connected to an electric power supply.

41. The gas turbine engine according to claim 40, 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.