GB2615315A

GB2615315A - Gas turbine engine

Info

Publication number: GB2615315A
Application number: GB2201315.5A
Authority: GB
Inventors: Minelli Andrea
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2022-02-02
Filing date: 2022-02-02
Publication date: 2023-08-09
Also published as: GB202201315D0

Abstract

A gas turbine engine 10 for an aircraft has hydraulic actuators, a fuel circuit 50 with a high and low 53 pressure pump and a fuel metering system. Wherein a hydraulic fluid line 60 is configured to convey hydraulic fluid to a hydraulic fluid circuit of the aircraft. The engine further includes a fuel-hydraulic fluid heat exchanger 63 which has a fuel side on the fuel circuit between the low- and high-pressure pumps. The heat exchanger transferring heat from the hydraulic fluid line to the fuel circuit. The engine may include a fuel-oil heat exchanger 51 to transfer heat from the oil to the fuel. The hydraulic fluid line may include a bypass loop 62 for bypassing the heat exchanger. The engine may include a gearbox 30(fig.1) receiving an input from a core shaft and outputs drive to a fan 23(fig.1) at a lower speed. Also disclosed is a gas turbine engine including a hydraulic fluid line and an air-hydraulic fluid heat exchanger 64(fig.4) configured to transfer heat from the hydraulic fluid line to an airflow 65(fig.4) produced by the engine e.g., a bypass duct airflow.

Description

GAS TURBINE ENGINE

Field of the Invention

The present invention relates to a gas turbine engine for an aircraft having hydraulic actuators.

Background

Aircraft include hydraulic fluid circuits providing hydraulic fluid to hydraulic actuators of the aircraft, e.g. for actuating aircraft flight control surfaces, brakes, nose wheel steering, and/or landing gear. The hydraulic fluid is driven around the circuit by one or more hydraulic pumps. During operation, the hydraulic pumps generate heat which can increase the temperature of the hydraulic fluid. This, in turn, can impair the performance of the hydraulic actuators of the aircraft.

To prevent this, conventional aircraft thermal management systems actively cool the hydraulic fluid by transferring heat from the hydraulic fluid to fuel stored in an aircraft fuel tank. However, transferring heat from the hydraulic fluid to fuel stored in the aircraft fuel tank requires the installation of a heat exchanger in the fuel tank to realise the heat transfer. Accessing the heat exchanger, e.g. for installation or maintenance purposes, can be difficult as this location is generally fuel-contaminated at all times.

Moreover, accessing the heat exchanger usually requires fully draining the fuel tank first, which can be undesirable. In addition, because the fuel in the fuel tank is substantially stagnant, heat transfer effectiveness to the fuel is low, necessitating a relatively large, and therefore heavy, heat exchanger.

The present invention has been devised in light of the above considerations.

Summary of the Invention

The present invention is at least partly based on a recognition that the cooling demand of a hydraulic fluid circuit for providing hydraulic fluid to aircraft hydraulic actuators can be met by making use of thermal management architecture of an aircraft engine (e.g. an engine fuel circuit or an airflow produced by the engine). This provides a wider choice of heat sinks available to receive heat from the hydraulic fluid circuit and can thus increase the design space for optimising the aircraft architecture. Advantageously, such an approach can also address disadvantages associated with conventional hydraulic circuit cooling solutions discussed above, e.g. limited accessibility to a heat exchanger in the fuel tank and low heat transfer effectiveness.

Accordingly, in a first aspect, the present invention provides a gas turbine engine for an aircraft having hydraulic actuators, the engine including: an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor; a fuel circuit configured to supply fuel to the combustor, the fuel circuit including: a low pressure pump configured to pressurise the fuel to a low pressure and a high pressure pump configured to receive the low pressure fuel and increase the pressure to a high pressure for supply to a fuel metering system and thence to the combustor: a hydraulic fluid line configured to receive hydraulic fluid from and convey hydraulic fluid to a hydraulic fluid circuit of the aircraft which provides hydraulic fluid for the hydraulic actuators; and a fuel-hydraulic fluid heat exchanger which has a fuel side on the fuel circuit between an outlet of the low pressure pump and an inlet of the high pressure pump, and a hydraulic fluid side on the hydraulic fluid line, whereby the fuel-hydraulic fluid heat exchanger transfers heat from the hydraulic fluid line to the fuel circuit.

The fuel-hydraulic fluid heat exchanger can have a significantly reduced size and weight compared to a conventional fuel tank-based heat exchanger. This is because fuel in the engine fuel circuit flows continuously therethrough and has a higher Reynolds number and therefore a higher heat transfer coefficient compared to the stagnant fuel in an aircraft fuel tank. The fuel-hydraulic fluid heat exchanger can also be much more accessible for repair and maintenance than a heat exchanger located in a fuel tank. By having the fuel side of the fuel-hydraulic fluid heat exchanger downstream of the outlet of the low pressure pump, any adverse effects of a raised fuel temperature on the pumping effectiveness of the low pressure pump can be avoided.

The gas turbine engine may further include: an oil circuit configured to cool and lubricate bearings of the engine core; and a fuel-oil heat exchanger having a fuel side on the fuel circuit between the fuel-hydraulic fluid heat exchanger and the inlet of the high pressure pump, and an oil side on an oil circuit to transfer heat from the oil circuit to the fuel circuit. Thus, in this way the heat transfer from the hydraulic fluid line into the fuel is performed before the generally greater heat transfer into the fuel from the fuel-oil heat exchanger. Were these two heat transfer processes performed in the opposite order, the cooling of the hydraulic fluid circuit might be compromised.

The hydraulic fluid line may include a hydraulic pump configured to drive the hydraulic fluid around the hydraulic fluid circuit, the hydraulic pump being adjacent to, and preferably downstream of, the fuel-hydraulic fluid heat exchanger. Conveniently, the hydraulic pump may be driven by an accessory gearbox of the engine. By having the fuel-hydraulic fluid heat exchanger in close physical proximity to the pump, the heat exchanger can more efficiently manage temperature increases in the hydraulic fluid produced by the pump. It can also reduce the amount of pipe routing required to connect the pump to the heat exchanger.

The hydraulic fluid line may include a bypass loop at the fuel-hydraulic fluid heat exchanger, the bypass loop being configured to allow hydraulic fluid to bypass the fuel-hydraulic fluid heat exchanger, thereby controlling the amount of hydraulic fluid entering the fuel-hydraulic fluid heat exchanger. Thus, under conditions when less cooling of the hydraulic fluid is required, e.g. on cold days, some or all of the hydraulic fluid can bypass the fuel-hydraulic fluid heat exchanger, thereby reducing a risk of overcooling the hydraulic fluid, which can also impair the performance of the hydraulic actuators of the aircraft.

In a second aspect, the present invention provides a gas turbine engine for an aircraft having hydraulic actuators, the engine including: a hydraulic fluid line configured to receive hydraulic fluid from and convey hydraulic fluid to a hydraulic fluid circuit of the aircraft which provides hydraulic fluid for the hydraulic actuators; and an air-hydraulic fluid heat exchanger configured to transfer heat from the hydraulic fluid line to an airflow produced by the engine.

This arrangement uses a non-fuel heat sink to receive heat from the hydraulic fluid circuit, and thereby reduces the cooling burden on an engine fuel circuit. It can enable a decrease of the total weight of the aircraft thermal management architecture and thus of the aircraft itself The air-hydraulic fluid heat exchanger can also be more accessible for repair and maintenance than a conventional fuel-hydraulic fluid heat exchanger located in a fuel tank.

The hydraulic fluid line may include a bypass loop at the air-hydraulic fluid heat exchanger, the bypass loop being configured to allow hydraulic fluid to bypass the air-hydraulic fluid heat exchanger, thereby controlling the amount of hydraulic fluid entering the air-hydraulic fluid heat exchanger. Thus, under conditions when less cooling of the hydraulic fluid is required, e.g. on cold days, some or all of the hydraulic fluid can bypass the air-hydraulic fluid heat exchanger, thereby reducing a risk of overcooling the hydraulic fluid, which can also impair the performance of the hydraulic actuators of the aircraft.

The air-hydraulic fluid heat exchanger may be located in an air intake of the engine, or may be located externally on a nacelle of the engine. Alternatively, when the gas turbine engine is a turbofan engine having a bypass duct for a bypass airflow outside the engine core, the air-hydraulic fluid heat exchanger may be located in the bypass duct to transfer heat from the hydraulic fluid line to the bypass duct airflow produced by the engine fan.

The air-hydraulic fluid heat exchanger may be a matrix cooler or a surface cooler. A matrix cooler typically has multiple stacked layers of plates spaced apart by fins projecting from the layers. The spaces between the layers contain alternating flows of hydraulic fluid and air in intimate contact with the fins. The flows are typically arranged in a crossflow configuration. In contrast, a surface cooler is typically a more plate-like structure, containing a flow channel or channels for flow of hydraulic fluid, and an array of fins which project therefrom into the airflow. A surface cooler may be integrated in an aircraft structure to perform a structural function as well as a cooling function, while a matrix cooler can generally provide a more compact cooling solution.

As noted elsewhere herein, the present disclosure relates 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.

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

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

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

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

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

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-!imitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity 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/U11p2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all values being dimensionless). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the 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), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 95 Nkg-ls, 90 Nkg-ls, 85 Nkg-'s or 80 Nkg-ls. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg-'s to 100 Nkg-1s, 01 85 Nkg-1s to 95 Nkg-1s. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.

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

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

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

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

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint -in terms of time and/or distance -between top of climb and start of descent. Cruise conditions thus define an operating point of, the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide -in combination with any other engines on the aircraft -steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

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

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 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 an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m).

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

According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.

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

Summary of the Figures

Embodiments illustrating the principles of the invention will now be discussed with reference to the accompanying 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 view of part of a fuel circuit and a hydraulic fluid line of a gas turbine engine; Figure 5 is a schematic view of a variant of the fuel circuit and the hydraulic fluid line of Figure 4; and Figure 6 is a comparative bar chart showing number of coolers for different engine configurations and for different operating conditions.

Detailed Description of the Invention

Aspects and embodiments of the present invention will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art. All documents mentioned in this text are incorporated herein by reference.

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 pads of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in Figure 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core 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.

The gas turbine engine 10 of Figures 1 and 2 further has an oil circuit configured to cool and lubricate bearings of the engine core 11, and gears and bearings of the power gear box (PGB) 30, a fuel circuit for supplying fuel to the combustor 16, and a hydraulic fluid line for receiving hydraulic fluid from and conveying hydraulic fluid to an aircraft hydraulic fluid circuit which provides hydraulic fluid for hydraulic actuators of the aircraft. For example, the hydraulic actuators can be for actuating flight control surfaces and/or landing gear of the aircraft. To prevent the oil in the oil circuit overheating during operation, the oil circuit transfers heat accumulated while lubricating the bearings of the engine core and the gears and bearings of the PGB to fluid heat sinks, i.e. fuel from the fuel circuit and airflows at respective heat exchangers. The transfer of heat from the oil circuit to the fuel circuit also protects the fuel from forming fuel-borne ice particles. Exchange of heat between the oil and fuel circuits helps to reduce thermodynamic losses in the engine 10.

Figure 4 shows schematically a part of the fuel circuit 50 and the hydraulic fluid line 60 of the gas turbine engine 10. The hydraulic fluid line 60 includes a hydraulic pump 61 configured to drive the hydraulic fluid around the hydraulic fluid circuit and the hydraulic fluid side of a fuel-hydraulic fluid heat exchanger (i.e. a fuel-cooled hydraulics cooler, or FCHC) 63 preferably located adjacent to and downstream of the pump. The hydraulic pump is typically a main source of heat into the hydraulic fluid, and the FCHC's location on the hydraulic fluid line in close physical proximity to the hydraulic pump allows it to efficiently manage the temperature of hydraulic fluid without significant additional pipe routing. Conveniently, the pump can be powered by an accessory gearbox of the engine. The hydraulic circuit also includes an optional bypass loop 62 at the FCHC 63 to allow hydraulic fluid to bypass the FCHC when it requires no or less cooling. For example, on cold days, the hydraulic fluid in the hydraulic fluid line may require less cooling and thus some or all of the hydraulic fluid can follow the bypass loop to prevent the hydraulic fluid overcooling in the FCHC 63. Management of flow through the bypass can be controlled, for example, by a thermostatic valve, a pressure relief valve, or a more active fully modulating valve with actuation based on measurement of hydraulic fluid temperature.

Although Figure 4 shows only one hydraulic pump 61, in general two pumps can be provided per engine for redundancy. These pumps can be connected on the hydraulic fluid line 60 in parallel.

The fuel circuit 50 supplies fuel from an aircraft fuel tank 52 to the combustor 16 for combustion. During operation, fuel drawn from the aircraft fuel tank is pressurised to a low pressure by a low pressure (LP) pump 53. The low pressure fuel is then directed to a high pressure (HP) pump (not shown) which increases the pressure of the fuel to a high pressure and supplies it to a fuel metering system (not shown) which controls the engine-consumed flow in response to a fuel demand indicated by an electronic engine controller (EEC) (not shown). The fuel circuit also includes a fuel-oil heat exchanger 51, i.e. a fuel-cooled oil cooler (FCOC), having a fuel side on the fuel circuit between an outlet of the LP pump and an inlet of the HP pump. During operation, the FCOC transfers heat from the oil circuit to the fuel circuit so as to prevent the oil in the oil circuit overheating. The fuel circuit further includes a fuel side of the FCHC 63 between the outlet of the LP pump 53 and the fuel side of the FCOC 51. Thus, during operation, the FCHC transfers heat from the hydraulic fluid line to the fuel circuit. Advantageously, the heat transfer from the hydraulic fluid circuit into the fuel at the FCHC is performed before the generally greater heat transfer into the fuel from the FCOC The arrangement of Figure 4 can meet the cooling demand of the hydraulic fluid circuit using a FCHC 63 having a significantly reduced size and weight compared to a conventional fuel tank-based heat exchanger. This is because fuel in the fuel circuit flows continuously therethrough and has a higher Reynolds number and therefore a higher heat transfer coefficient compared to the substantially stagnant (sloshing) fuel in the aircraft fuel tank. The FCHC can also be fitted in such a way as to be more accessible for repair and maintenance than a heat exchanger located in a fuel tank. By having the fuel side of the FCHC downstream of the outlet of the LP pump 53, any adverse effects of a raised fuel temperature on the pumping effectiveness of the LP pump can be avoided.

Figure 5 schematically shows a variant of the fuel circuit 50 and the hydraulic fluid line 60 of Figure 4. The fuel circuit of Figure 5 is identical to its counterpart of Figure 4 except that it does not include the fuel side of an FCHC 63. Thus in this variant, the hydraulic fluid line does not transfer heat to the fuel circuit. Instead, the hydraulic fluid line 60 includes a hydraulic fluid side of an air-hydraulic fluid heat exchanger 64, such as a matrix air-cooled hydraulics cooler (MACHC), preferably located downstream of and in physical proximity to the hydraulic pump 61. The MACHC 64 utilises an airflow 65 produced by the engine 10 such that it transfers heat from the hydraulic fluid line to the airflow. For example, the MACHC can be located in the bypass duct (BPD) 22 to transfer heat from the hydraulic fluid line 60 to the bypass duct airflow B produced by the fan 23. Alternatively, the MACHC can be located in the air intake 12 of the engine or externally on the nacelle 21 to transfer heat to an airflow respectively in the air intake or around the nacelle. Advantageously, implementing the air-hydraulic fluid heat exchanger as a MACHC can achieve a compact cooling solution.

However, a different option for the air-hydraulic fluid heat exchanger 64 is to realise it as a surface air-cooled hydraulics cooler (SACHC) rather than an MACHC. Advantageously, an SACHC can be integrated in the engine to perform a structural function as well as a cooling function.

Similarly to the arrangement shown in Figure 4, the hydraulic circuit also includes an optional bypass loop 66 at the MACHC 64 to allow hydraulic fluid to bypass the MACHC when it requires no or less cooling.

This can prevent the hydraulic fluid overcooling in the MACHC. Management of flow through the bypass can be controlled, for example, by a thermostatic valve, a pressure relief valve, or a more active fully modulating valve with actuation based on measurement of hydraulic fluid temperature.

The cooling arrangements of both Figures 4 and 5 address disadvantages associated with conventional hydraulic circuit cooling solutions discussed previously, i.e. limited accessibility to a fuel tank-based heat exchanger for maintenance and repair purpose, and low heat transfer effectiveness. Additionally, both arrangements enable an up to 30% reduction in the total number of air coolers, e.g. matrix air-cooled oil coolers (MACOCs), required to meet the respective cooling demands of the oil and fuel circuits. In the variant of Figure 4, this is because including an FCHC leads to a reduction in the temperature of the fuel in the fuel tank, thereby increasing the heat capacity of the fuel and allowing a reduction in the number and/or size of the required MACOCs. In the variant of Figure 5, the hydraulic fluid line does not transfer heat to the fuel circuit and therefore the temperature of the fuel in the fuel circuit is reduced. This increases the heat capacity of the fuel and in turn also allows a reduction in the number and/or size of the required MACOCs.

This is evidenced by Figure 6 which is a comparative bar chart showing the results of modelling to determine the number of air coolers needed for three different engine configurations and for six different operating conditions. The engine configurations are: a (Baseline) configuration having a conventional fuel tank-based hydraulic fluid cooler, a configuration (NB) based on the arrangement of Figure 4; and a configuration (C) based on the arrangement of Figure 5. The engine operating conditions include: design cruise, top of climb condition at international standard atmosphere +15°C (TOC ISA+15), top of climb condition at international standard atmosphere +10°C (TOC ISA+10), maximum take-off weight at sea level and at end of runway (MTO SL EoR), maximum continuous thrust at engine inoperative altitude capability (MCI EIAC), and maximum continuous thrust at extended-range twin-engine operational performance standard (MCT ETOPS). The number of air coolers is calculated based on a baseline heat exchanger design having a known size and capacity. Thus, for each operating condition, the number of baseline-design air coolers, required to meet the corresponding cooling demand, is calculated. The chart shows that implementing the arrangement of Figure 4 consistently results in a reduction in the number of air coolers required for all the studied operation conditions. Implementing the arrangement of Figure 5 allows an even greater reduction, which can be attributed to the reduced cooling burden on the fuel circuit, as discussed above.

Overall, both arrangements are advantageous as they can enable a decrease in the total weight of the aircraft thermal management architecture and thus of the aircraft itself, and facilitate servicing and maintenance. * ;The features disclosed in the foregoing description, or in the following claims, or in the accompanying drawings, expressed in their specific forms or in terms of a means for performing the disclosed function, or a method or process for obtaining the disclosed results, as appropriate, may, separately, or in any combination of such features, be utilised for realising the invention in diverse forms thereof While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure. Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention. ;For the avoidance of any doubt, any theoretical explanations provided herein are provided for the purposes of improving the understanding of a reader. The inventors do not wish to be bound by any of these theoretical explanations. ;Any section headings used herein are for organizational purposes only and are not to be construed as limiting the subject matter described. ;Throughout this specification, including the claims which follow, unless the context requires otherwise, the word "comprise" and "include", and variations such as "comprises", "comprising", and "including" will be understood to imply the inclusion of a stated integer or step or group of integers or steps but not the exclusion of any other integer or step or group of integers or steps. ;It must be noted that, as used in the specification and the appended claims, the singular forms a, an, and "the" include plural referents unless the context clearly dictates otherwise. Ranges may be expressed herein as from "about" one particular value, and/or to "about" another particular value. When such a range is expressed, another embodiment includes from the one particular value and/or to the other particular value. Similarly, when values are expressed as approximations, by the use of the antecedent "about," it will be understood that the particular value forms another embodiment. The term "about" in relation to a numerical value is optional and means for example +/-10%. *

Claims

Claims: 1. A gas turbine engine (10) for an aircraft having hydraulic actuators, the engine including: an engine core (11) comprising a turbine (19), a combustor (16), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fuel circuit (50) configured to supply fuel to the combustor (16), the fuel circuit including: a low pressure pump (53) configured to pressurise the fuel to a low pressure; and a high pressure pump configured to receive the low pressure fuel and increase the pressure to a high pressure for supply to a fuel metering system and thence to the combustor; a hydraulic fluid line (60) configured to receive hydraulic fluid from and convey hydraulic fluid to a hydraulic fluid circuit of the aircraft which provides hydraulic fluid for the hydraulic actuators; and a fuel-hydraulic fluid heat exchanger (63) which has a fuel side on the fuel circuit between an outlet of the low pressure pump and an inlet of the high pressure pump, and a hydraulic fluid side on the hydraulic fluid line, whereby the fuel-hydraulic fluid heat exchanger transfers heat from the hydraulic fluid line to the fuel circuit.
2. The gas turbine engine (10) according to claim 1 further including: an oil circuit configured to cool and lubricate bearings of the engine core (11); and a fuel-oil heat exchanger (51) having a fuel side on the fuel circuit (50) between the fuel-hydraulic fluid heat exchanger and the inlet of the high pressure pump, and an oil side on an oil circuit to transfer heat from the oil circuit to the fuel circuit.
3. The gas turbine engine (10) according to claim 1 or 2, wherein the hydraulic fluid line (60) includes a hydraulic pump (61) configured to drive the hydraulic fluid around the hydraulic fluid circuit, the hydraulic pump being adjacent to the fuel-hydraulic fluid heat exchanger (63).
4. The gas turbine engine (10) according to claim 3, wherein the hydraulic pump (61) is driven by an accessory gearbox of the engine.
5. The gas turbine engine (10) according to any one of the previous claims, wherein the hydraulic fluid line includes a bypass loop (62) at the fuel-hydraulic fluid heat exchanger (63), the bypass loop being configured to allow hydraulic fluid to bypass the fuel-hydraulic fluid heat exchanger, thereby controlling the amount of hydraulic fluid entering the fuel-hydraulic fluid heat exchanger.
6. The gas turbine engine (10) according to any one of the previous claims further including a fan (23) located upstream of the engine core, and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.
7. A gas turbine engine (10) for an aircraft having hydraulic actuators, the engine including: a hydraulic fluid line (60) configured to receive hydraulic fluid from and convey hydraulic fluid to a hydraulic fluid circuit of the aircraft which provides hydraulic fluid for the hydraulic actuators; and an air-hydraulic fluid heat exchanger (64) configured to transfer heat from the hydraulic fluid line to an airflow (65) produced by the engine.
8. The gas turbine engine (10) according to claim 7, wherein the hydraulic fluid line includes a bypass loop (66) at the air-hydraulic fluid heat exchanger (64), the bypass loop being configured to allow hydraulic fluid to bypass the air-hydraulic fluid heat exchanger, thereby controlling the amount of hydraulic fluid entering the air-hydraulic fluid heat exchanger.
9. The gas turbine engine (10) according to claim 7 or 8, wherein the air-hydraulic fluid heat exchanger (64) is located in an air intake of the engine, or is located externally on a nacelle (21) of the engine.
10. The gas turbine engine (10) according to claim 7 or 8, wherein the engine further includes: an engine core (11) comprising a turbine (19), a combustor (16), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades; and a bypass duct (22) for a bypass airflow (B) produced by the fan outside the engine core (11), the air-hydraulic fluid heat exchanger (64) being located in the bypass duct to transfer heat from the hydraulic fluid line (60) to the bypass airflow.
11. The gas turbine engine (10) according to claim 10, wherein the engine further includes a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft
12. The gas turbine engine (10) according to any one of claims 7 to 9, wherein the engine further includes: an engine core (11) comprising a turbine (19), a combustor (16), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan blades; and a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.
13. The gas turbine engine (10) according to any one of claims 1 to 6 and 9 to 12, wherein: the turbine is a first turbine (19), the compressor is a first compressor (14), and the core shaft is a first core shaft (26); the engine core further includes a second turbine (17), a second compressor (15), and a second core shaft (27) 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.