THERMAL MANAGEMENT SYSTEM AND A GAS TURBINE ENGINE

Abstract

There is disclosed a gas turbine engine having a thermal management system, the gas turbine engine comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and a fan located upstream of the engine core. The thermal management system comprises an oil tank; a heat exchanger; an oil coolant circuit that connects the oil tank and the heat exchanger; and an oil pump that pumps oil around the oil coolant circuit. The oil tank is located within the engine core, and the oil pump is electrically driven such that it is operable independently of the core shaft.

Claims

1. A gas turbine engine having a thermal management system, the gas turbine engine comprising an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core; the thermal management system comprising: an oil tank; a heat exchanger; an oil coolant circuit that connects the oil tank and the heat exchanger; and an oil pump that pumps oil around the oil coolant circuit; wherein the oil tank is located within the engine core, and the oil pump is electrically driven such that it is operable independently of the core shaft.

2. The gas turbine engine of claim 1, wherein the oil tank and the heat exchanger are integral with each other.

3. The gas turbine engine of claim 1, wherein the oil tank is located adjacent a power gearbox in the gas turbine engine.

4. The gas turbine engine of claim 1, wherein the heat exchanger is located directly downstream of the outlet guide vanes of the fan in the gas turbine engine.

5. The gas turbine engine of claim 1, wherein a portion of the heat exchanger is located within the engine core.

6. The gas turbine engine of claim 1, further comprising a plurality of heat exchangers connected to the oil coolant circuit.

7. The gas turbine engine of claim 1, further comprising 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.

8. The gas turbine engine of claim 1, further comprising a controller which is configured to vary the speed of oil through the thermal management system based on the prevailing conditions of the engine.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0043] Embodiments will now be described by way of example only, with reference to the Figures, in which:

[0044] FIG. 1 schematically shows a sectional side view of a gas turbine engine;

[0045] FIG. 2 schematically shows a close up sectional side view of an upstream portion of a gas turbine engine;

[0046] FIG. 3 schematically shows a partially cut-away view of a gearbox for a gas turbine engine; and

[0047] FIG. 4 schematically shows a simplified sectional side view of a gas turbine engine with a thermal management system.

DETAILED DESCRIPTION OF THE DISCLOSURE

[0048] Aspects and embodiments of the present disclosure will now be discussed with reference to the accompanying figures. Further aspects and embodiments will be apparent to those skilled in the art.

[0049] FIG. 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.

[0050] 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 core exhaust nozzle 20 to provide some propulsive thrust.

[0051] 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.

[0052] An exemplary arrangement for a geared fan gas turbine engine 10 is shown in

[0053] FIG. 2. The low pressure turbine 19 (see FIG. 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.

[0054] 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.

[0055] The epicyclic gearbox 30 is shown by way of example in greater detail in FIG. 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 FIG. 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. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

[0056] The epicyclic gearbox 30 illustrated by way of example in FIGS. 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.

[0057] It will be appreciated that the arrangement shown in FIGS. 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 FIG. 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 FIG. 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 FIG. 2.

[0058] 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.

[0059] Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

[0060] 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 FIG. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core exhaust nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

[0061] 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 FIG. 1), and a circumferential direction (perpendicular to the page in the FIG. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

[0062] FIG. 4 shows a sectional side view of the gas turbine engine 10 comprising a thermal management system 50. The gas turbine engine may be as described with respect to FIGS. 1-3.

[0063] The gas turbine engine 10 comprises a fan casing 52 enclosing the fan 23, and a plurality of fan outlet guide vanes 54 directly downstream (or aft) of the fan 23, and enclosed within the fan casing 52.

[0064] The fan 23 comprises a plurality of fan blades, which extend radially from a hub 56. In this example, the epicyclic gearbox 30 (power gearbox) is located directly downstream of the hub 56 and is connected to the hub 56 so that the fan 23 is driven via the gearbox 30. In other examples, there may be no gearbox through which the fan is driven, such that the fan 23 is driven directly by the shaft 26 that rotates with a corresponding turbine 15, 17, 19.

[0065] The thermal management system 50 comprises an oil tank 60, a heat exchanger 62, and an oil pump 64. It also comprises an oil coolant circuit including a plurality of pipes for connecting the oil tank 60, heat exchanger 62, and oil pump 64 with each other and for connecting other components to be cooled in the engine 10 to the oil tank 60, oil pump 64 and heat exchanger 62. The oil coolant circuit may also connect further heat exchangers to the oil tank and oil pump.

[0066] The oil tank 60 is located within the engine core 11, close to or adjacent the component in the gas turbine engine which generates the most heat in use, and therefore requires the most cooling. In this example, the gearbox 30 may be the component which requires the most cooling in use. Therefore, the oil tank 60 is located downstream of the gearbox 30 and in thermal contact with the gearbox 30, such that no pipework (i.e. no pipes extending between spaced-apart components) is required to fluidically connect the oil tank 60 with the gearbox 30. In other words, an aperture in the oil tank 60 is contiguous with an aperture in the gearbox 30, so that they are fluidically connected through the apertures. In some examples, the oil tank is merely located close to the gearbox, such that a pipe of the coolant circuit fluidically connects the gearbox with the oil tank, which may be of relatively short length nonetheless.

[0067] Providing the oil tank 60 in the engine core ensures that the oil tank does not need to be mounted in the body of the nacelle of the gas turbine engine 10. This enables the use of a slim nacelle, which is particularly advantageous for use with a large fan.

[0068] The heat exchanger 62 is located close to the oil tank 60. In this example, the heat exchanger 62 is integrated with the oil tank 60, such that no pipework (i.e. no pipes extending between spaced-apart components) is required to connect the heat exchanger 62 and the oil tank 60. In other examples, the heat exchanger is co-located with the oil tank i.e. located close to, or adjacent the oil tank with at least a portion of the heat exchanger being located within the engine core, to reduce the amount of pipework required to connect the heat exchanger and oil tank compared with previously known systems. In such examples, the pipe length between the oil tank and heat exchanger may be less than 3 m, such as less than 2.5 m, 2 m, 1.5 m, 1 m, 0.75 m, 0.5 m or 0.25 m. A further reduction in pipework may be achieved in this case if the oil tank 60 is used as the header tank for the heat exchanger 62.

[0069] In previously known systems, the heat exchanger was not located close to the oil tank, because classic offtake duct theory has always taught that a heat exchanger should have a long inlet duct and exhaust duct to maximise efficiency of the heat exchanger e.g. 5-10 times the height of the heat exchanger.

[0070] However, the applicant has concluded that such a duct is not required, and that a heat exchanger may work with sufficiently high efficiency even with a small inlet duct and exhaust duct length. Therefore, by placing the heat exchanger as close to the oil tank as possible, the pipework in the coolant circuit of the thermal management system 50 can be reduced.

[0071] This reduction of pipework reduces the total weight of the thermal management system 50 due to the reduced amount of piping, and therefore also due to the reduced amount of oil required in the pipes of the thermal management system 50. Therefore the total weight of the gas turbine engine 10 is reduced so that the specific fuel consumption (SFC) of the gas turbine engine 10 is also reduced, without compromising the efficiency of the heat exchanger 62.

[0072] The amount of pipework of the thermal management system 50 can be reduced as described above so that a ratio of volume of oil in the pipes to volume of oil in the tank is less than 1:10, such as a ratio of 1:20. This reduction in pipework also reduces the gulp in the oil tank (i.e. on engine start-up, the sudden reduction in the fluid level in the oil tank to fill all the pipes and volumes within the system).

[0073] The oil pump 64 in this example is an electrical oil pump. Previously known systems have used mechanical oil pumps which are driven by a shaft of the engine. Electric pumps are heavier than mechanical pumps, but the applicant has found that using an electric oil pump can result in efficiency gains despite the additional weight.

[0074] By using an electric oil pump, the oil can be pumped around at any speed controlled by a controller, in contrast to mechanical oil pumps, in which speed and power is a function of the rotational speed of the engine shaft which drives it. This means that the oil can be pumped around the coolant circuit at an optimal speed for the prevailing conditions in the engine, rather than at the speed dictated by the mechanical coupling of the pump to the engine shaft.

[0075] For example, the oil flow rate can be increased during peak load with the electric oil pump 64, and at high oil flow requirements the electric oil pump 64 can actually be operated at higher speeds than are currently possible with a mechanical pump.

[0076] Further, the oil flow rate can be reduced during cruise. Operating an electric pump so that it runs at reduced speeds during cruise reduces the hydrodynamic load of the oil system, reducing the load on turbomachinery and hence reducing the SFC of the gas turbine engine 10. This reduction in SFC offsets the losses in efficiency due to the additional weight of the electrical pump.

[0077] As explained above, reducing the pipework in the thermal management system 50 also means that there is less oil in the system 50. Furthermore, the reduced pipework also results in less excrescence drag (drag from the roughness of the inside surface of pipes). As such, a less powerful pump 64 is needed to pump the oil around the coolant circuit. Therefore, the electric pump 64 can be smaller (and thus lighter), so that the gains in efficiency of using an electric pump further outweigh the losses due to the additional weight of the pump.

[0078] It will be understood that the gas turbine engine of the present disclosure is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.