GB2615314A

GB2615314A - Combination of a gas turbine engine, a heat exchanger, and a power electronics

Info

Publication number: GB2615314A
Application number: GB2201312.2A
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: GB202201312D0

Abstract

A gas turbine engine includes a closed loop cooling circuit 50 for cooling power electronics 51, having a pump 56, such as an electric pump for circulating a flow of cooling fluid, such as engine oil, water-glycol, or 1, 1, 1, 2-tetrafluoroethane around the circuit. A heat exchanger 53 is located to transfer heat from the cooling flow to an aircraft fluid heat sink 54, such as an airflow generated by the engine.

Description

COMBINATION OF A GAS TURBINE ENGINE, A HEAT EXCHANGER, AND A POWER ELECTRONICS

Field of the Invention

The present invention relates to a combination of a gas turbine engine, a heat exchanger, and a power electronics for powering aircraft and/or engine systems

Background

Gas turbine engines include sophisticated thermal management systems to control the temperatures of components. In particular, heat is rejected into the oil of the engine oil system used for cooling and lubricating engine components. The oil in the oil system is cooled in turn by transferring heat to engine fuel and/or air flows.

A further source of cooling demand can derive from the thermal management of electrical components, such as power electronics, which form an increasingly important part of aircraft and/or engine systems. Failure to meet increased cooling demands can result in less reliable or worsened performance of such systems. In particular, the performance and reliability of power electronics for powering aircraft and/or engine systems (e.g. the cabin blower system) can be affected by temperature changes and therefore reliably controlling its temperature during all phases of aircraft operation is important.

Conventional heat management systems meet the power electronics cooling demands during above-idle engine operation conditions by rejecting heat into engine fluid heat sinks, such as fuel flow to the engine combustor. However, during sub-idle engine operation conditions, which may for example occur at engine start-up and windmill relight, and also during post-shutdown heat soak back conditions, such heat sinks may be unavailable or insufficient to meet the cooling demands of the power electronics, if still active. Sub-idle engine operation conditions typically apply from 0 rpm to idle, which is the steady state engine operating condition with no load applied.

These problems can be exacerbated when the gas turbine engine includes a power gear box (PGB) to drive a fan, the PGB putting a further cooling demand on the thermal management system of the engine.

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 power electronics for powering aircraft and/or engine systems can be met by making use of a closed loop cooling circuit, even during sub-idle engine operation conditions.

Thus, in a first aspect, the present invention provides a combination of a gas turbine engine, a heat exchanger, and a power electronics for powering aircraft and/or engine systems, wherein.

the engine includes a closed loop cooling circuit for the power electronics, the circuit having a pump for circulating a flow of cooling fluid around the circuit; the heat exchanger is located to transfer heat from the cooling flow to an aircraft fluid heat sink; and the power electronics is configured to transfer heat produced by the power electronics to the flow of cooling fluid circulating around the circuit to prevent the power electronics overheating.

Thus, by adopting this arrangement, despite the unavailability or insufficiency of conventional heat sinks during sub-idle engine operation (e.g. at engine start-up and windmill relight) or post-shutdown heat soak back, the cooling demand of the power electronics can be met by implementing a closed loop cooling circuit. In particular, the pump can continuously circulate a cooling flow around the power electronics during all engine operation conditions so as to cool it. Furthermore, the pump can regulate the flow rate of the cooling fluid around the circuit to adjust the cooling rate according to changes in the cooling demand of the power electronics. Thus, the cooling demand of the power electronics can be adequately met during all engine operation conditions.

Optional features of the invention will now be set out. These are applicable singly or in any combination with any aspect of the invention.

The power electronics may be for powering an aircraft cabin blower system.

The pump may be an electric pump. For example, the electric pump may be powered directly by the engine or by an aircraft system (e.g. a cabin blower system). Such a pump can help to more reliably control the flow rate of the cooling flow depending on the cooling demand of the power electronics, and thus more reliably and effectively control the power electronics' temperature and rate of temperature change.

The cooling fluid may be water-glycol. Advantageously, water-glycol is uninflammable which contributes to its reliability and safety when used for cooling the power electronics.

However, the cooling fluid may be aircraft fuel or engine oil. Such fluids are conveniently available from the engine, and the closed loop cooling circuit can be fed by and integrated with fuel or oil circuits.

The combination of a gas turbine engine, a heat exchanger, and a power electronics may be further configured such that: the cooling circuit forms a refrigeration cycle in which the cooling fluid is a refrigerant fluid; the power electronics is configured to transfer heat produced by the power electronics to the refrigerant fluid in an evaporator that evaporates liquid in the refrigerant fluid to form a saturated vapour; the pump is a compressor which compresses the saturated vapour to a superheated vapour and sends it to the heat exchanger; the heat exchanger is a condenser in which the transfer of heat from the compressed vapour to the aircraft fluid heat sink converts the superheated vapour to a saturated liquid; and the cooling circuit further includes an expansion valve between the condenser and the power electronics to convert the saturated liquid to a reduced temperature liquid and vapour mixture which returns to the evaporator. Advantageously, forming the cooling circuit as a refrigeration cycle can decrease the size of the cooling circuit required to meet the cooling demand of the power electronics. Thus, the engine footprint of the cooling circuit can be reduced.

The refrigerant fluid may be 1,1,1,2-tetrafluoroethane, known commercially as refrigerant R1 34a.

The power electronics may further include internal passages for flow therethrough of the cooling fluid. Alternatively or additionally, the power electronics may further include one or more cold plates cooled by the flow of cooling fluid for extracting heat produced by the power electronics. Such internal passages and/or cold plates can enhance the transfer of heat from the power electronics to the cooling flow, thereby more reliably and effectively controlling the power electronics' temperature.

The cooling circuit may further include a cooling fluid reservoir. Such a cooling fluid reservoir can compensate for any volumetric changes of the cooling fluid due to temperature and density changes such that the cooling flow can perform its cooling function reliably and consistently during all engine operation conditions.

The aircraft fluid heat sink may be an airflow produced by the engine, and the heat exchanger may be an air-cooled heat exchanger. Such a heat exchanger can transfer heat from the cooling flow to the airflow produced by the engine during all times at which such an airflow is available. For example, the air-cooled heat exchanger may be an air-cooled surface cooler. An air-cooled surface cooler typically has a platelike structure, containing a flow channel or channels for flow of cooling fluid, and an array of fins which project therefrom into the airflow. Advantageously, even when the airflow produced by the engine is at or close to zero (e.g. during engine post-shutdown soak back conditions), the air-cooled surface cooler is generally still able to provide air cooling by natural convection. Additionally, the air-cooled surface cooler can be integrated into an aircraft or engine structure to perform a structural function as well as a cooling function.

The air-cooled heat exchanger may be located in an air intake of the engine, or may be located externally on a nacelle of the engine. When the gas turbine engine is a turbofan engine having a bypass duct for a bypass airflow outside the engine core, another option is for the air-cooled heat exchanger to be located in the bypass duct to transfer heat from the cooling flow to the bypass duct airflow. Advantageously, the air-cooled heat exchanger located in these positions is in contact with airflows produced by the engine and can transfer heat from the power electronics into the airflows.

A different option for the aircraft fluid heat sink is to use fuel contained in an aircraft fuel tank. The heat exchanger may then be a fuel-cooled heat exchanger immersed in the fuel tank. Fuel in the aircraft fuel tank has a large heat transfer capacity due to its large volume and therefore can adequately meet the cooling demand of the power electronics during all engine operation conditions.

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 (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

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

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

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 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/Ut1p2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all 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-limitafive 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-1s, 105 Nkg-1s, 100 Nkg-1s, 95 Nkg-1s, 90 Nkg-1s, 85 Nkg-'s 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), for example in the range of from 80 Nkg-1s 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; and Figure 4 to 6 are schematic diagrams of respective variants of a closed loop cooling circuit of the gas turbine engine.

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

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

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

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in Figure 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core 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 fin 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 (not shown) configured to cool and lubricate bearings of the engine core 11, and gears and bearings of the power gear box (PGB) 30, and a fuel circuit (not shown) for supplying fuel to the combustor 16.

The aircraft has power electronics (PE), such as for powering a cabin blower (CB) system, or powering aircraft control actuators. The PE generates heat and requires thermal management, in the form of active cooling, to ensure it does not exceed a temperature which could impair its performance. The PE can be mounted to the engine 10 or more preferably, it may be mounted at a convenient location on the aircraft at a distance from the engine. Thus, it can be ensured that during engine post-shutdown heat soak back conditions, less heat reaches the PE.

Accordingly, the gas turbine engine 10 of Figures 1 and 2 further has a closed loop cooling circuit for cooling the PE. Respective variants of the closed loop cooling circuit are shown in the schematic diagrams of Figures 4 to 6. The variants of Figures 4 and 5 differ from each other in the location of a bypass loop 61 on the cooling circuit 50, while the variant of Figure 6 shows an implementation of the cooling circuit 50' as a refrigerant cycle (discussed in more detail below).

Referring to Figures 4 and 5, the cooling circuit 50 has a pump 56 for circulating a flow of cooling fluid around the circuit, and a heat exchanger 53 for transferring heat from the cooling flow to an aircraft fluid heat sink 54. The pump is preferably an electric pump and can be powered directly by the engine or by the CB system. The pump controls the flow rate of the cooling fluid through the cooling circuit according to changes in the cooling demand of the PE. The cooling fluid can be water-glycol which, advantageously, is uninflammable. Alternatively, the cooling fluid may be fuel or engine oil. Advantageously, those fluids are already used in the aircraft and therefore are readily available for use in the closed loop cooling circuit.

According to a first option, the aircraft fluid heat sink is an airflow produced by the engine and correspondingly the heat exchanger is an air-cooled heat exchanger. For example, the air-cooled heat exchanger can be located in the bypass duct 22 to transfer heat from the cooling flow to the bypass duct airflow B produced by the fan 23. As another option, the air-cooled heat exchanger 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. Preferably, the air-cooled heat exchanger is implemented as an air-cooled surface cooler. An air-cooled surface cooler is typically a plate-like structure, containing a flow channel or channels for flow of cooling fluid, and an array of fins which project therefrom into the airflow. Advantageously, this type of air-cooled heat exchanger is generally able to provide air cooling by natural convection even when the produced airflow by the engine is at or close to zero, which allows the PE to be actively cooled during all engine operation conditions, including sub-idle operating conditions and engine shutdown. Additionally, an air-cooled surface cooler can be integrated in an aircraft or engine structure to perform a structural function as well as a cooling function, However, if cooling by natural convention is not required, the air-cooled heat exchanger can be a different type, such as a matrix cooler.

According to a second option, the aircraft fluid heat sink 54 is fuel contained in an aircraft fuel tank, and correspondingly the heat exchanger is a fuel-cooled heat exchanger immersed in the fuel tank. Fuel in the aircraft fuel tank has a large heat transfer capacity due to its large volume and therefore is a suitable heat sink for adequately meeting the cooling demand of the power electronics during all engine operation conditions.

During operation, the pump 56 circulates the flow of cooling fluid around the circuit 50 such that the PE 51 transfers heat that it produces to the cooling flow via a further heat exchanger 52 so that the PE does not overheat. The direction of the cooling flow around the closed loop is indicated by arrows. The further heat exchanger is preferably configured as internal passages in the PE for flow therethrough of the cooling fluid and/or one or more cold plates in/on the PE, the cold plates being cooled by the cooling flow. The heat rejected by the PE is then transferred from the cooling fluid to the aircraft fluid heat sink 54 via the heat exchanger 53.

The cooling circuit 50 of Figures 4 and 5 includes an optional cooling fluid reservoir 55 and an optional bypass loop 61. During operation, the cooling fluid may experience volumetric and density changes as a result of thermal changes related to the cooling of the PE 51. The cooling fluid reservoir compensates for such volumetric changes of the cooling fluid such that the cooling circuit can perform its cooling function reliably and consistently during all engine operation conditions. The location of the pump 56 in the direction of flow around the loop is preferably downstream of the reservoir 55, but it can also be located upstream of it (as indicated by the dash-line pump in Figures 4 and 5). The bypass loop 61 is preferably located at the air-cooled surface cooler 53, as shown in Figure 4. During low-temperature external conditions, such as on cold-days, the extraction of heat from the PE 51 can be reduced by having the cooling flow bypass the heat exchanger, to avoid excessively decreasing the PE's temperature which could also impair its performance. Alternatively, the bypass loop can be located at the further heat exchanger 52, as shown in Figure 5, to perform the same function.

Modifications of the cooling circuit 50 are possible. For example, the function of the bypass loop 61 could be realised by one or more active and/or passive control valves (e.g. thermostatic and/or pressure-relief valves) located on the cooling circuit.

Figure 6 shows a further variant of the closed loop cooling circuit 50'. In this variant, the cooling circuit forms a refrigeration cycle in which the cooling fluid is a refrigerant fluid. For example, the refrigerant fluid may be 1,1,1,2-tetrafluoroethane, although other refrigerant fluids can be used. During operation, the heat produced by the PE 51 is transferred to the refrigerant fluid in an evaporator 59 that evaporates liquid in the refrigerant fluid to form a saturated vapour. The evaporator may be integrated into the PE for a more compact cooling circuit 50'. The pump is configured as a compressor 56' which compresses the saturated vapour to a superheated vapour and sends it to the heat exchanger, which is configured as a condenser 53' in which the transfer of heat from the compressed vapour to the aircraft fluid heat sink 54 converts the superheated vapour to a saturated liquid. The cooling circuit further includes an expansion valve 58 between the condenser and the PE to convert the saturated liquid to a reduced-temperature liquid and vapour mixture which then returns to the evaporator.

Advantageously, the condenser 53' of the Figure 6 can be made smaller than the heat exchanger 53 of Figures 4 and 5 while providing the same heat transfer effectiveness, which reduces the footprint of the cooling circuit 50' in the engine 10.

Advantageously, the cooling circuit 50, 50' can meet a variable cooling demand from the PE 51 due to the ability to modulate the cooling flow rate by the pump 56, 56'. * ;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 combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) for powering aircraft and/or engine systems, wherein: the engine includes a closed loop cooling circuit (50, 50') for the power electronics, the circuit having a pump (56, 56') for circulating a flow of cooling fluid around the circuit; the heat exchanger (53, 53') is located to transfer heat from the cooling flow to an aircraft fluid heat sink (54); and the power electronics is configured to transfer heat produced by the power electronics to the flow of cooling fluid circulating around the circuit to prevent the power electronics overheating.
2. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 1, wherein the pump (56, 56') is an electric pump.
3. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 1 or 2, wherein the cooling fluid is water-glycol.
4. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 1 or 2, wherein the cooling fluid is aircraft fuel or engine oil.
5. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 1 or 2, wherein: the cooling circuit (50') forms a refrigeration cycle in which the cooling fluid is a refrigerant fluid; the power electronics is configured to transfer heat produced by the power electronics to the refrigerant fluid in an evaporator (59) that evaporates liquid in the refrigerant fluid to form a saturated 25 vapour; the pump (56') is a compressor which compresses the saturated vapour to a superheated vapour and sends it to the heat exchanger (53'); the heat exchanger is a condenser in which the transfer of heat from the compressed vapour to the aircraft fluid heat sink (54) converts the superheated vapour to a saturated liquid; and the cooling circuit further includes an expansion valve (58) between the condenser and the power electronics to convert the saturated liquid to a reduced temperature liquid and vapour mixture which returns to the evaporator.
6. The combination of a gas turbine engine (10), a heat exchanger (53'), and a power electronics (51) according to claims, wherein the refrigerant fluid is 1,1,12-tetrafluoroethane.
7. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to any one of the previous claims, wherein the power electronics further includes internal passages for flow therethrough of the cooling fluid.
8. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to any one of the previous claims, wherein the power electronics further includes one or more cold plates cooled by the flow of cooling fluid for extracting heat produced by the power electronics.
9. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to any one of the previous claims, wherein the cooling circuit (50, 50') further includes a cooling fluid reservoir (55).
10. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to any one of the previous claims, wherein the aircraft fluid heat sink (54) is an airflow produced by the engine and the heat exchanger is an air-cooled heat exchanger.
11. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 10, wherein the air-cooled heat exchanger is an air-cooled surface cooler.
12. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 10 or 11, wherein the air-cooled heat exchanger (53) is located in an air intake of the engine, or is located externally on a nacelle (21) of the engine.
13. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 10 or 11, 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 produced by the fan outside the engine core (11), the air-cooled heat exchanger (53) being located in the bypass duct to transfer heat from the cooling flow to the bypass airflow.
14. The combination of a geared gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to claim 13, 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.
15. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to any one of claims 1 to 9, wherein the aircraft fluid heat sink (54) is fuel contained in an aircraft fuel tank and the heat exchanger is a fuel-cooled heat exchanger immersed in the fuel tank.
16. The combination of a gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to any one of claims 1 to 12 and 15, 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.
17. The combination of a geared gas turbine engine (10), a heat exchanger (53, 53'), and a power electronics (51) according to any one of claims 13, 14 and 16, 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.