# US 2021/0340914 A1 · Thermal management system

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- Number: US 2021/0340914 A1
- Document ID: US20210340914A1
- Status: Published application
- Kind: A1
- Application number: 17/375,489
- Filed: 2021-07-14
- Published: 2021-11-04
- Earliest claimed: 2017-11-21
- Assignees: General Electric Company
- Applicants: General Electric Company
- Inventors: Daniel Alan Niergarth; Adon Delgado, JR.; Brandon Wayne Miller; Hendrik Pieter Jacobus de Bock
- Web page: https://patentgrep.com/patent/US20210340914A1

## CPC

- F02C 7/224: Heating fuel before feeding to the burner
- F02C 7/14: of fluids in the plant {, e.g. lubricant or fuel (F02C7/185 takes precedence)}
- F02C 7/141: of working fluid
- F28F 23/00: Features relating to the use of intermediate heat-exchange materials, e.g. selection of compositions (heat-transfer, heat-exchange or heat-storage materials C09K5/00)
- F05D 2220/323: for aircraft propulsion, e.g. jet engines
- F05D 2240/12: Fluid guiding means, e.g. vanes
- F05D 2260/213: by the provision of a heat exchanger within the cooling circuit
- F28F 27/02: for controlling the distribution of heat-exchange media between different channels ({static flow control means in header boxes F28F9/026}; arrangements of guide plates or guide vanes F28F9/22, F28F25/12)
- F28F 2250/06: Derivation channels, e.g. bypass
- Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft
- F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00)
- F28F: DETAILS OF HEAT-EXCHANGE AND HEAT-TRANSFER APPARATUS, OF GENERAL APPLICATION (water and air traps, air venting F16)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION

## Abstract

A gas turbine engine includes a turbomachine and a thermal management system. The thermal management system includes a heat source heat exchanger configured to collect heat from the turbomachine during operation; a heat sink heat exchanger; and a thermal transport bus having a heat exchange fluid configured to flow therethrough at a pressure within an operational pressure range. The thermal management system defines an operational temperature range for the heat exchange fluid, the operational temperature range having a lower temperature limit less than about zero degrees Fahrenheit at a pressure within the operational pressure range and an upper temperature limit of at least about 1000 degrees Fahrenheit at a pressure within the operational pressure range.

## Claims

### Claim 1

A gas turbine engine comprising:
a turbomachine; and
a thermal management system comprising
a first heat exchanger configured to collect heat from the turbomachine during certain operations;
a second heat exchanger; and
a thermal transport bus configured to have a heat exchange fluid flow therethrough at a pressure within an operational pressure range, the first heat exchanger and second heat exchanger each fluidly coupled to the thermal transport bus such that the first heat exchanger is operable to transfer heat to the heat exchange fluid during at least the certain operations and the second heat exchanger is operable to transfer heat from the heat exchange fluid during at least the certain operations, the thermal management system defining an operational temperature range for the heat exchange fluid, the operational temperature range having a lower temperature limit less than about zero degrees Fahrenheit at a pressure within the operational pressure range and an upper temperature limit of at least about 1000 degrees Fahrenheit at a pressure within the operational pressure range.

### Claim 2 (depends on 1)

The gas turbine engine of claim 1, wherein the first heat exchanger is a heat recovery heat exchanger.

### Claim 3 (depends on 2)

The gas turbine engine of claim 2, wherein the turbomachine comprises a turbine section and an exhaust section, the turbine section and exhaust section together defining at least in part a core air flowpath, and wherein the heat recovery heat exchanger is a waste heat recovery heat exchanger positioned to be in thermal communication with the core air flowpath within or downstream of the turbine section, the exhaust section, or both.

### Claim 4 (depends on 2)

The gas turbine engine of claim 2, wherein the turbomachine comprises a fuel delivery system, and wherein the second heat exchanger is a fuel heat exchanger thermally coupled to the fuel delivery system.

### Claim 5 (depends on 1)

The gas turbine engine of claim 1, wherein the turbomachine comprises a cooling air system, and wherein the first heat exchanger is thermally coupled to the cooling air system.

### Claim 6 (depends on 1)

The gas turbine engine of claim 1, wherein the lower temperature limit of the operational temperature range at a pressure within the operational pressure range is between about −100 degrees Fahrenheit and about −5 degrees Fahrenheit.

### Claim 7 (depends on 6)

The gas turbine engine of claim 6, wherein the upper temperature limit of the operational temperature range at a pressure within the operational pressure range is between about 1000 degrees Fahrenheit and about 1800 degrees Fahrenheit.

### Claim 8 (depends on 1)

The gas turbine engine of claim 1, wherein the heat exchange fluid is at least one of a liquid metal alloy, a molten salt, a silicone oil, an ionic fluid, a pressurized gas, or a supercritical gas.

### Claim 9 (depends on 1)

The gas turbine engine of claim 1, wherein the heat exchange fluid is a liquid metal alloy, and wherein the operational pressure range is greater than zero pounds per square inch and less than about 500 pounds per square inch.

### Claim 10 (depends on 1)

The gas turbine engine of claim 1, wherein the heat exchange fluid is a eutectic metal alloy comprising gallium, indium, and tin.

### Claim 11 (depends on 1)

The gas turbine engine of claim 1, wherein the heat exchange fluid is a pressurized gas.

### Claim 12 (depends on 1)

The gas turbine engine of claim 1, wherein the heat exchange fluid is a supercritical gas, wherein the supercritical gas defines critical point pressure, wherein the operational pressure range is greater than the critical point pressure of the supercritical gas and up to about 8000 pounds per square inch.

### Claim 13 (depends on 12)

The gas turbine engine of claim 12, wherein the heat exchange fluid is a supercritical gas, and wherein the operational pressure range is greater than about 1000 pounds per square inch and less than about 8000 pounds per square inch.

### Claim 14 (depends on 1)

The gas turbine engine of claim 1, further comprising:
an outer nacelle at least partially surrounding the turbomachine; and
an outlet guide vane extending between the outer nacelle and the turbomachine, wherein the second heat exchanger is integrated into, or coupled to, the outlet guide vane.

### Claim 15 (depends on 1)

The gas turbine engine of claim 1, wherein the thermal management system further comprises a heater thermally coupled to the thermal transfer bus for heating the thermal transfer fluid within the thermal transfer bus.

### Claim 16 (depends on 1)

The gas turbine engine of claim 1, wherein the first heat exchanger is a first heat source heat exchanger, wherein the thermal management system further comprises a second heat source heat exchanger, wherein the first heat source heat exchanger is a waste heat recovery heat exchanger, and wherein the second heat source heat exchanger is a cooling air system heat exchanger.

### Claim 17 (depends on 1)

The gas turbine engine of claim 1, wherein the second heat exchanger is a fuel heat exchanger, and wherein the thermal management system further comprises a bypass airflow passage heat exchanger.

### Claim 18

A gas turbine engine comprising:
a compressor section, a combustion section, a turbine section, and an exhaust section arranged in serial flow order and together defining a core air flowpath;
a fuel delivery system for providing a flow of fuel to the combustion section; and
a thermal management system comprising
a first heat exchanger positioned to be in thermal communication with a flow through the core air flowpath within or downstream of the turbine section, the exhaust section, or both;
a second heat exchanger in thermal communication with the fuel delivery system for transferring heat to the flow of fuel provided to the combustion section; and
a thermal transport bus having a heat exchange fluid flowing therethrough, the first heat exchanger and second heat exchanger each fluidly coupled to the thermal transport bus.

### Claim 19 (depends on 18)

The gas turbine engine of claim 18, wherein the thermal management system defines an operational temperature range for the heat exchange fluid, the operational temperature range having a lower temperature limit less than about zero degrees Fahrenheit at a pressure within the operational pressure range and an upper temperature limit of at least about 1000 degrees Fahrenheit at a pressure within the operational pressure range.

### Claim 20 (depends on 19)

The gas turbine engine of claim 19, wherein the heat exchange fluid is at least one of a molten salt, a silicone oil, an ionic fluid, a liquid metal alloy, or a supercritical gas.