Cooling air delivery system and methods thereof

Patent metadata

Field	Value
Publication number	US-12320299-B2
Application number	US-202217869060-A
Country	US
Kind code	B2
Filing date	Jul 20, 2022
Priority date	Jul 20, 2022
Publication date	Jun 3, 2025
Grant date	Jun 3, 2025

Abstract

A gas turbine engine having a compressor section, a combustion section, and a turbine section in serial flow arrangement to define a diffuser cavity between the compressor section and the combustor section, and an aft cavity. A compressor discharge pressure duct fluidly draws air from the diffuser cavity, passes it through a heat exchanger to cool the air, and then supplies the cooled air to the aft cavity.

First claim

We claim:

1. A gas turbine engine comprising:
a compressor, a combustion section, and a turbine arranged in serial flow,
wherein the combustion section comprises an outer combustor casing and an inner combustor casing defining at least in part a diffuser cavity therebetween,
wherein the compressor is in fluid communication with the diffuser cavity to provide a flow of compressed air from the compressor to the diffuser cavity;
a high-pressure shaft disposed radially inward from the inner combustor casing,
wherein the high-pressure shaft and the inner combustor casing at least partially define an aft cavity and a forward shaft outer cavity therebetween,
wherein a seal extends from the high pressure shaft to the inner combustor casing to separate the aft cavity from the forward shaft outer cavity, and
wherein the aft cavity is defined at least in part by the inner combustor casing, the seal, and the compressor; and
a cooling system, comprising:
a heat exchanger fluidly coupled to the diffuser cavity via a first fluid opening,
wherein the heat exchanger receives airflow from the diffuser cavity via the first fluid opening;
a first cooling duct fluidly coupling the heat exchanger to the aft cavity, wherein the first cooling duct includes a first pipe and a manifold extending in a circumferential direction from the first pipe about a longitudinal centerline of the gas turbine engine, the first pipe and the manifold extending through the diffuser cavity downstream from the heat exchanger;
a valve disposed within the first cooling duct downstream from the manifold and upstream from the aft cavity; and
a second cooling duct fluidly coupling the heat exchanger to the turbine via the forward shaft outer cavity, and
wherein the second cooling duct extends within the diffuser cavity.

2. The gas turbine engine of claim 1 , wherein the turbine of the gas turbine engine is part of a turbine section of the gas turbine engine, wherein the second cooling duct provides a portion of the airflow to the turbine of the turbine section.

3. The gas turbine engine of claim 2 , wherein the compressor comprises a material, wherein the material defines a material temperature limit in degrees Fahrenheit, and wherein the cooling system is configured to provide another portion of the airflow to the aft cavity at a temperature in degrees Fahrenheit less than or equal to 85% of the material temperature limit when the gas turbine engine is operated at a rated speed during standard day operating conditions.

4. The gas turbine engine of claim 2 , wherein the turbine of the turbine section is a high-pressure turbine.

5. The gas turbine engine of claim 4 , wherein the high-pressure turbine comprises an inlet guide vane and a first stage blade, wherein the inlet guide vane and the first stage blade define a forward wheelspace cavity, and wherein the gas turbine engine defines a purge air flowpath from the aft cavity to the forward wheelspace cavity.

6. The gas turbine engine of claim 1 , the cooling system further comprising: at least one sensor for sensing data indicative of an operative condition of the gas turbine engine, wherein the valve is configured to operate based on the operative condition.

7. The gas turbine engine of claim 6 , wherein the operative condition comprises at least one of: a high operating temperature condition; a high-pressure condition; a supersonic cruise condition; a takeoff condition; or a climb condition.

8. The gas turbine engine of claim 7 , wherein the operative condition is the high operating temperature condition, and wherein the high operating temperature comprises an operating condition wherein a compressor exit temperature is higher than 1000 degrees Fahrenheit.

9. The gas turbine engine of claim 6 , further comprising a controller, the controller comprising one or more computing devices in operable communication with the at least one sensor and the valve, the one or more computing devices of the controller being configured to receive the data indicative of the operative condition from the at least one sensor and control the valve to provide a portion of the airflow from the first cooling duct to the aft cavity in response to receiving the data indicative of the operative condition of the gas turbine engine.

10. The gas turbine engine of claim 6 , wherein the cooling system further comprises a controller, the controller comprising one or more computing devices operably coupled to the at least one sensor, for receiving the data indicative of the operative condition, and to the valve, for actuating the valve, wherein the one or more computing devices of the controller are configured to move the valve to an open position when the gas turbine engine is in the operative condition and are further configured to move the valve to a closed position when the gas turbine engine is not in the operative condition.

11. The gas turbine engine of claim 1 , wherein the manifold is divided into two segments, wherein each of the two segments extends 180 degrees about the longitudinal centerline of the gas turbine engine.

12. The gas turbine engine of claim 1 , wherein the heat exchanger is positioned outside of the outer combustor casing.

13. The gas turbine engine of claim 1 , wherein the heat exchanger is fluidly coupled to the diffuser cavity via a compressor discharge pressure duct.

14. A method of cooling one or more sections in a gas turbine engine, the method comprising:
receiving an airflow from a diffuser cavity defined by an outer combustor casing and an inner combustor casing of the gas turbine engine;
cooling the airflow received from the diffuser cavity with a heat exchanger;
providing a first portion of the airflow from the heat exchanger to a first cooling duct at least partially defined by a first pipe extending within the diffuser cavity and providing a second portion of the airflow from the heat exchanger to a second cooling duct at least partially defined by a second pipe extending within the diffuser cavity, the first cooling duct including a manifold within the diffuser cavity and extending in a circumferential direction from the first pipe about a longitudinal centerline of the gas turbine engine; and
providing the first portion of the airflow from the first cooling duct to an aft cavity defined at least in part by:
the inner combustor casing,
a seal that extends from a high-pressure shaft of the gas turbine engine to the inner combustor casing, and
a compressor section of the gas turbine engine; and

providing the second portion of the airflow from the second cooling duct to a turbine section of the gas turbine engine,
wherein providing the first portion of the airflow from the first cooling duct to the aft cavity comprises operating at least one valve located downstream of the manifold and upstream from the aft cavity and in fluid communication with the first cooling duct.

15. The method of claim 14 , further comprising: receiving data indicative of an operative condition of the gas turbine engine, wherein providing the first portion of the airflow from the first cooling duct to the aft cavity comprises providing the first portion of the airflow from the first cooling duct to the aft cavity in response to receiving the data indicative of the operative condition of the gas turbine engine.

16. The method of claim 15 , wherein the operative condition comprises at least one of: a high operating temperature condition; a high-pressure condition; a supersonic cruise condition; or a takeoff condition.

17. The method of claim 15 , wherein receiving the data indicati

Assignees

• Gen Electric

Inventors

• Jagata Lakshmi Narasimharao
• Jalan Prateek
• Vanapalli Veeraraju
• Saha Atanu
• Schimmels Scott Alan

Classifications

• F01D5/081

  Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence) · CPC title

• F02C7/185

  Cooling means for reducing the temperature of the cooling air or gas · CPC title

• F04D29/5833

  flow schemes and regulation thereto · CPC title

• F04D29/5826

  Cooling at least part of the working fluid in a heat exchanger · CPC title

• F05D2260/213

  by the provision of a heat exchanger within the cooling circuit · CPC title