Jet engine cold air cooling system

Patent metadata

Field	Value
Publication number	US-10247100-B2
Application number	US-201514755398-A
Country	US
Kind code	B2
Filing date	Jun 30, 2015
Priority date	Jul 3, 2014
Publication date	Apr 2, 2019
Grant date	Apr 2, 2019

Abstract

Airplanes and jet engines are provided that includes an engine compressor; a combustor in flow communication with the engine compressor; an engine turbine in flow communication with the combustor to receive combustion products from the combustor; and a bleed air cooling system in fluid communication with bleed air from the engine compressor. The bleed air cooling system can include a first precooler in fluid communication with the bleed air from the engine compressor; a cooling system turbine in fluid communication with and downstream from the first precooler; and a discharge conduit from the cooling system turbine that is configured to be in fluid communication with at least one of an aircraft thermal management system and an aircraft environmental control system. Methods are also described for providing cooling fluid from a jet engine.

First claim

What is claimed is:

1. A jet engine comprising:
an engine compressor;
a combustor in fluid communication with the engine compressor;
an engine turbine in fluid communication with the combustor to receive combustion products from the combustor; and
a bleed air cooling system in fluid communication with bleed air from the engine compressor, wherein the bleed air cooling system comprises:
a first precooler in fluid communication with the bleed air from the engine compressor;
a cooling system turbine in fluid communication with and downstream from the first precooler;
a cooling system compressor in fluid communication with and downstream from the first precooler, the cooling system compressor in fluid communication with and downstream from the engine compressor;
a second precooler in fluid communication with and downstream from the cooling system compressor, the second precooler in fluid communication with and downstream from the engine compressor; and
a third precooler, the third precooler in fluid communication with and downstream from both the first precooler and the engine compressor, wherein each of the first, second, and third precoolers is cooled by an engine cooling stream passing through the jet engine,

wherein a discharge conduit from the cooling system turbine is configured to be in fluid communication with at least one of an aircraft thermal management system and an aircraft environmental control system.

2. The jet engine of claim 1 , wherein the bleed air cooling system further comprises a gearbox operatively coupled to and powered by the cooling system turbine.

3. The jet engine of claim 2 , wherein the cooling system compressor is operatively coupled to and powered by the cooling system turbine, and
wherein the cooling system turbine has a multiple position turbine nozzle.

4. The jet engine of claim 3 , wherein the cooling system compressor comprises a suction compressor, the jet engine further comprising:
a first bypass line, the first bypass line connected between an inlet of the cooling system turbine and an outlet of the cooling system turbine;
a first bypass control valve, the first bypass control valve in series with the first bypass line; and
a thermocouple, the thermocouple in thermal communication with air discharged from the cooling system turbine,
wherein based on temperature readings from the thermocouple, the first bypass line selectively allows air discharged from the second precooler to be directed to at least one of the aircraft thermal management system and the aircraft environmental control system, without traveling through the cooling system turbine.

5. The jet engine of claim 2 , further comprising:
a second bypass line, the second bypass line connected between an outlet of the third precooler and the discharge conduit, the discharge conduit downstream from an outlet of the cooling system turbine, and upstream of at least one of the aircraft thermal management system and the aircraft environmental control system, without traveling through the cooling system turbine; and
a second bypass control valve, the second bypass control valve downstream from and in fluid communication with the first precooler, the second bypass control valve upstream from and in fluid communication with both the second precooler and the third precooler,
wherein the second bypass control valve selectively controls air flow to both the second precooler and the third precooler.

6. The jet engine of claim 5 , wherein the cooling system turbine has a variable area turbine nozzle and is operative to generate power to drive a pump associated with the jet engine.

7. The jet engine of claim 6 , further comprising:
an engine thermal management system in two-way fluid communication with at least one of the aircraft thermal management system and the aircraft environmental control system of an aircraft,
wherein relative to a flow direction of thermal communication with an engine inlet airstream:
the engine thermal management system is upstream of the second precooler;
the second precooler is upstream of the third precooler; and
the third precooler is upstream of the first precooler.

8. A method of providing cooling fluid from a jet engine, the jet engine comprising: an engine compressor; a combustor in fluid communication with the engine compressor; an engine turbine in fluid communication with the combustor to receive combustion products from the combustor; and a bleed air cooling system in fluid communication with bleed air from the engine compressor, wherein the bleed air cooling system comprises: a first precooler in fluid communication with the bleed air from the engine compressor; a cooling system turbine in fluid communication with and downstream from the first precooler; a cooling system compressor in fluid communication with and downstream from the first precooler, the cooling system compressor in fluid communication with and downstream from the engine compressor; a second precooler in fluid communication with and downstream from the cooling system compressor, the second precooler in fluid communication with and downstream from the engine compressor; and a third precooler, the third precooler in fluid communication with and downstream from both the first precooler and the engine compressor, wherein each of the first, second and third precoolers is cooled by an engine cooling stream passing through the jet engine, wherein a discharge conduit from the cooling system turbine in configured to be in fluid communication with at least one of an aircraft thermal management system and an aircraft environmental control system; the method comprising extracting a first portion of an airflow flowing through the engine compressor of the jet engine at a first temperature; directing the first portion of the airflow into fluid communication with the first precooler to reduce a temperature of the first portion of the airflow to a second temperature; and flowing the first portion of the airflow downstream from the first precooler through a cooling system turbine to reduce the temperature of the first portion of the airflow to a third temperature, wherein the third temperature is less than a flow path temperature of air flowing into the jet engine.

9. The method of claim 8 , wherein flowing the first portion of the airflow through the turbine is operative to generate power to drive a gearbox associated with the jet engine.

10. The method of claim 8 , wherein flowing the first portion of the airflow through the turbine is operative to generate power to drive the cooling system compressor.

11. The method of claim 10 , wherein the cooling system compressor comprises a suction compressor.

12. The method of claim 8 , further comprising compressing the first portion of the airflow downstream from the first precooler from a first lower pressure to a second higher pressuring using the cooling system compressor, wherein compressing the first portion of the airflow increases the temperature of the first portion of the airflow to a fourth temperature greater than the second temperature and less than the first temperature.

13. The method of claim 12 , wherein directing the first portion of the airflow through the cooling system turbine is operative to power the cooling system compressor.

14. The method of claim 12 , further comprising directing the first portion of the airflow into fluid communication with the second precooler, downstream from the cooling system compressor and upstream from the cooling system turbine, to reduce the temperature of the first portion of the airflow to a fifth temperature, where the fifth temperature is less than the fo

Assignees

• Gen Electric

Inventors

• Leamy Kevin Richard
• Breig Benjamin Paul
• Epstein Michael Jay
• Parrilla Javier Armando
• Hurt Matthew John
• Brinson Thomas Edward
• Fleming Andrew James

Classifications

• F02C7/32

  Arrangement, mounting, or driving, of auxiliaries · CPC title

• F01D13/02

  Working-fluid interconnection of machines or engines · CPC title

• F05D2270/301

  Pressure · CPC title

• B64D13/006

  the air being used to cool structural parts of the aircraft · CPC title

• F05D2270/303

  Temperature · CPC title