US-12618366-B2 - Thermal management system for an aircraft

US12618366B2US 12618366 B2US12618366 B2US 12618366B2US-12618366-B2

Abstract

A thermal management system for an aircraft includes a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, and a first heat exchanger. Waste heat energy generated by at least one first gas turbine engine, and first electric machine, transfers to the first heat transfer fluid. The first heat exchanger directs a first proportion of the first heat transfer fluid through a first heat dissipation portion wherein a first proportion of the waste heat energy transfers to a first dissipation medium dependent on the first dissipation medium temperature and mass flow rate. The first heat exchanger directs a second proportion of the first heat transfer fluid through a second heat dissipation portion wherein the second proportion of waste heat energy transfers to a second dissipation medium dependent on the second dissipation medium temperature and mass flow rate.

Inventors

Paul R DAVIES
Richard G MOCHRIE
David A Jones

Assignees

ROLLS-ROYCE PLC

Dates

Publication Date: 20260505
Application Date: 20230814
Priority Date: 20220906

Claims (11)

1 . A thermal management system for an aircraft, the thermal management system comprising: a first gas turbine engine; one or more first electric machines rotatably coupled to the first gas turbine engine; a first thermal bus; a first heat exchanger; and one or more ancillary systems, wherein the first thermal bus comprises a first heat transfer fluid, the first heat transfer fluid being in fluid communication, in a closed loop flow, between the first gas turbine engine, each of the one or more first electric machines, and the first heat exchanger, such that waste heat energy generated by at least one of the first gas turbine engine, and each of the one or more first electric machines, is transferred to the first heat transfer fluid, the first heat exchanger is configured to direct a first proportion of the flow of the first heat transfer fluid through a first heat dissipation portion in which a first proportion Q A of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium being an air flow, and a second proportion of the flow of the first heat transfer fluid through a second heat dissipation portion in which a second proportion Q B of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium being a fuel flow, and the first heat exchanger selectively directs the first proportion through the first heat dissipation portion depending on a temperature of the first dissipation medium and a mass flow rate of the first dissipation medium, and/or directs the second proportion through the second heat dissipation portion depending on a temperature of the second dissipation medium and a mass flow rate of the second dissipation medium, and in response to a temperature of the first heat transfer fluid leaving the first heat dissipation portion being less than the temperature of the first heat dissipation fluid, the first heat transfer fluid upstream of the first heat exchanger is controlled to pass through only the second heat dissipation portion; and in response to the temperature of the first heat transfer fluid leaving the first heat dissipation portion being greater than the temperature of the first heat dissipation fluid, the first heat transfer fluid upstream of the first heat exchanger is controlled to pass through the second heat dissipation portion and then through portion the first heat dissipation, via a cross-over path extending from an outlet of the second dissipation portion to an inlet of the first dissipation portion.
2 . The thermal management system as claimed in claim 1 , wherein the first thermal bus is arranged in a recirculatory ring configuration with the first heat transfer fluid passing through each of the first gas turbine engine, each of the one or more first electric machines, the first heat exchanger, and each of the one or more ancillary systems.
3 . The thermal management system as claimed in claim 1 , wherein the first heat exchanger is configured to direct the first proportion of the flow of the first heat transfer fluid through the first heat dissipation portion in which the first proportion Q A of the waste heat energy from the first heat transfer fluid is transferred to the first dissipation medium depending on a first temperature differential between the temperature of the first heat transfer fluid and the temperature of the first dissipation medium.
4 . The thermal management system as claimed in claim 1 , wherein the first heat exchanger comprises a bypass flow path, and wherein the first heat exchanger is configured to direct a third proportion of the flow of the first heat transfer fluid through the bypass flow path depending the temperature of the first dissipation medium and the mass flow rate of the first dissipation medium, and/or the temperature of the second dissipation medium and the mass flow rate of the second dissipation medium.
5 . The thermal management system as claimed in claim 1 , wherein the first gas turbine engine comprises, in axial flow sequence, the first heat exchanger, a compressor module, a combustor module, and a turbine module, and the first dissipation medium is the air flow passing through the first heat exchanger and entering the compressor module.
6 . The thermal management system as claimed in claim 1 , wherein the first gas turbine engine comprises, in axial flow sequence, a compressor module, a combustor module, and a turbine module, and the second dissipation medium is the fuel flow passing through the first heat exchanger and subsequently being directed to the combustor module.
7 . The thermal management system as claimed in claim 1 , wherein the first gas turbine engine is a first turbofan gas turbine engine, the turbofan gas turbine engine comprising, in axial flow sequential arrangement, a fan module, a compressor module, a combustor module, and a turbine module, the fan module comprising at least one fan stage having a plurality of fan blades extending radially from a hub, the plurality of fan blades defining a fan diameter (D FAN ), and wherein the fan diameter D FAN is within the range of 0.3 m to 2.0 m.
8 . The thermal management system as claimed in claim 7 , wherein the first turbofan gas turbine engine further comprises an outer casing, the outer casing enclosing the sequential arrangement of the fan assembly, the compressor module, the combustor module, and the turbine module, an annular bypass duct being defined between the outer casing and the sequential arrangement of the compressor module, the combustor module, and the turbine module, a bypass ratio being defined as a ratio of a mass air flow rate through the bypass duct to a mass air flow rate through the sequential arrangement of the compressor module, the combustor module, and the turbine module, and wherein the bypass ratio is less than 4.0.
9 . The thermal management system as claimed in claim 7 , wherein the at least one fan stage is two or more fan stages.
10 . A method of operating a thermal management system for an aircraft, the thermal management system comprising a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, and a first heat exchanger, the method comprising: (i) providing a first thermal bus comprising a first heat transfer fluid with the first heat transfer fluid providing fluid communication, in a closed loop flow, between the first gas turbine engine, each of the one or more first electric machines and the first heat exchanger, such that waste heat energy generated by at least one of the first gas turbine engine, and each of the one or more first electric machines; (ii) directing a first proportion of the flow of the first heat transfer fluid through a first heat dissipation portion in which a first proportion Q A of the waste heat energy from the first heat transfer fluid is transferred to a first dissipation medium being an air flow; and (iii) directing a second proportion of the flow of the first heat transfer fluid through a second heat dissipation portion in which a second proportion Q B of the waste heat energy from the first heat transfer fluid is transferred to a second dissipation medium being a fuel flow, wherein the first heat exchanger selectively directs the first proportion through the first heat dissipation portion depending on a temperature of the first dissipation medium and a mass flow rate of the first dissipation medium, and/or directs the second proportion through the second heat dissipation portion depending on a temperature of the second dissipation medium and a mass flow rate of the second dissipation medium, and in response to a temperature of the first heat transfer fluid leaving the first heat dissipation portion being less than the temperature of the first heat dissipation fluid, the first heat transfer fluid upstream of the first heat exchanger is controlled to pass through only the second heat dissipation portion; and in response to the temperature of the first heat transfer fluid leaving the first heat dissipation portion being greater than the temperature of the first heat dissipation fluid, the first heat transfer fluid upstream of the first heat exchanger is controlled to pass through the second heat dissipation portion and then through portion the first heat dissipation, via a cross-over path extending from an outlet of the second dissipation portion to an inlet of the first dissipation portion.
11 . The method as claimed in claim 10 , further comprising: (iv) directing a third proportion of the flow of the first heat transfer fluid through the bypass flow path depending the temperature of the first dissipation medium and the mass flow rate of the first dissipation medium, and/or the temperature of the second dissipation medium and the mass flow rate of the second dissipation medium.

Description

This disclosure claims the benefit of UK Patent Application No. GB 2212953.0, filed on 6 Sep. 2022, which is hereby incorporated herein in its entirety. FIELD OF THE DISCLOSURE The present disclosure relates to a thermal management system for an aircraft and particularly, but not exclusively, to a thermal management system for a gas turbine engine for an aircraft. BACKGROUND TO THE DISCLOSURE A conventional gas turbine engine for an aircraft includes 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 in turn cooled 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 of electrical components can result in less reliable or decreased performance of such electrical systems. In particular, the performance and reliability of power electronics for powering aircraft and/or engine systems (e.g., the aircraft environmental control 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 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. Similar problems may also occur during post-shutdown heat soak back conditions. Furthermore, during low-power conditions such as idle and descent, the rate of flow of fuel through the engine fuel system is reduced. Thus, continuing to transfer the heat produced by the oil system to the fuel system during those engine operation conditions can lead to high fuel temperature and the formation of solid fuel deposits (fuel lacquering/coking) e.g., inside engine fuel spray nozzles. Such deposits can impair the performance of the engine and might lead to component malfunctions and/or failures. The problem of high fuel temperature is exacerbated in lean burn fuel systems in which, to ensure margin against weak extinction during low-power operating conditions, the fuel system operates in pilot-only mode and stringent fuel temperature limits have to be imposed at the inlet of the fuel spray nozzles to avoid formation of deposits in the mains nozzle internal passages where stagnant fuel can be present. A further problem of high fuel temperature can arise during deceleration manoeuvres (e.g., transition from end of cruise to top of descent) when the fuel flow is suddenly reduced, but due to the thermal inertia of the oil system, the heat load lags behind. STATEMENTS OF DISCLOSURE According to a first aspect of the present disclosure there is provided a thermal management system for an aircraft, the thermal management system comprising a first gas turbine engine, one or more first electric machines rotatably coupled to the first gas turbine engine, a first thermal bus, a first heat exchanger, and one or more first ancillary systems; wherein the first thermal bus comprises a first heat transfer fluid, the first heat transfer fluid being in fluid communication, in a closed loop flow sequence, between the or each first electric machine, the first gas turbine engine, the first heat exchanger, and the or each first ancillary system, such that waste heat energy generated by at least one of the first gas turbine engine, the or each first electric machine, and the or each first ancillary system, is transferred to the first heat transfer fluid, andthe first heat exchanger is configured to transfer the waste heat energy from the first heat transfer fluid to a dissipation medium. The first thermal bus provides a thermal conduit between heat generating portions, such as a gas turbine engine, an electric machine, and an ancillary system, and heat dissipating portions, such as the heat exchanger. Consequently, waste heat energy generated by the gas turbine engine, electric machine and ancillary system can be transferred to the heat exchanger where it can be dissipated to a dissipation medium. Examples of first ancillary system may include, for example, power converters for use with the electric machine, and the electrical control and thermal regulation of an energy storage system, for example battery or capacit