Intercooled cooling air with plural heat exchangers

• US 10,480,419 B2
• Filed: 06/22/2015
• Issued: 11/19/2019
• Est. Priority Date: 04/24/2015
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Suciu, Gabriel L., Chandler, Jesse M., Staubach, Joseph Brent, Merry, Brian D.
• Primary Examiner(s): Verdier, Christopher
• Application Number:
  US14/745,567
• Publication Number:
  US 20160312711A1
• Time in Patent Office: 1,611 Days
• CPC Class Codes: F01D 25/12 Cooling
• CPC Class Codes: F02C 3/13 having variable working fluid interconnections between turbines or compressors or stages of different rotors controlling flow ratio between different flows of multi-flow jet-propulsion plant, e.g. ducted fan F02K3/075
• CPC Class Codes: F02C 6/08 the gas being bled from the gas-turbine compressor
• CPC Class Codes: F02C 7/143 before or between the compressor stages
• CPC Class Codes: F02C 7/185 Cooling means for reducing the temperature of the cooling air or gas
• CPC Class Codes: F02C 9/18 by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages F02C3/113 takes precedence
• CPC Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F04D 29/5826 Cooling at least part of the working fluid in a heat exchanger
• CPC Class Codes: F04D 29/5833 flow schemes and regulation thereto
• CPC Class Codes: F05D 2220/3212 the first stage of a turbine
• CPC Class Codes: F05D 2220/3219 for the last stage of a compressor or a high pressure compressor
• CPC Class Codes: F05D 2260/211 by intercooling, e.g. during a compression cycle
• CPC Class Codes: F05D 2260/213 by the provision of a heat exchanger within the cooling circuit
• CPC Class Codes: F05D 2260/606 Bypassing the fluid
• CPC Class Codes: F05D 2270/3062 of the auxiliary fluid for heating or cooling purposes
• CPC Class Codes: Y02T 50/60 Efficient propulsion technologies, e.g. for aircraft

Abstract

A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main compressor section, passing the tapped air through a first heat exchanger and then to a cooling compressor. A second tap taps air from a location closer to the downstream most end than the location of the first tap, and the first and second taps mix together and are delivered into the high pressure turbine. The cooling compressor is positioned downstream of the first heat exchanger, and upstream of a location where air from the first and second taps mix together.

Claims

1. A gas turbine engine comprising;

a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations;

a turbine section having a high pressure turbine;

a first tap tapping air from at least one of said more upstream locations in said main compressor section, passing said tapped air through a first heat exchanger and then to a cooling compressor;

a second tap tapping air from a location in said main compressor section closer to said downstream most end than the location of said first tap, and said first and second taps mixing together and being delivered into said high pressure turbine;

said cooling compressor being positioned downstream of said first heat exchanger, and upstream of a location where air from said first and second taps mix together;

wherein there is a second heat exchanger downstream of said cooling compressor, and air passes through said second heat exchanger before being delivered to said high pressure turbine;

wherein air is tapped from a third tap downstream of said second heat exchanger, and delivered to a component on the gas turbine engine to be cooled; and

wherein the location of said third tap is upstream of a location where said first and second taps mix together.

2. The gas turbine engine as set forth in claim 1, wherein said component to be cooled is a bearing compartment.

3. A gas turbine engine comprising;

a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations;

a turbine section having a high pressure turbine;

a first tap tapping air from at least one of said more upstream locations in said main compressor section, passing said tapped air through a first heat exchanger and then to a cooling compressor, said cooling compressor compressing air downstream of said first heat exchanger;

a second heat exchanger downstream of said cooling compressor;

a second tap tapping air from a location in said main compressor section closer to said downstream most end than the location of said first tap, wherein air from said first and second taps mixing together and being delivered into said high pressure turbine;

air being tapped from a third tap downstream of said second heat exchanger, and delivered to a component on the gas turbine engine to be cooled; and

wherein the location of said third tap is upstream of a location where said first and second taps mix together.

4. The gas turbine engine as set forth in claim 3, wherein a main fan delivers bypass air into a bypass duct and into said main compressor section and said first and second heat exchangers are positioned within said bypass duct to be cooled by bypass air.

5. The gas turbine engine as set forth in claim 4, wherein air temperatures at said downstream most location of said high pressure compressor are greater than or equal to 1350°

F.

6. The gas turbine engine as set forth in claim 4, wherein the second tap is at said downstream most end.

7. The gas turbine engine as set forth in claim 4, wherein air from said first tap mixes with air from said second tap at a location downstream of said cooling compressor.

8. The gas turbine engine as set forth in claim 3, wherein said component to be cooled is a bearing compartment.

9. The gas turbine engine as set forth in claim 8, wherein said bearing compartment supports a shaft adjacent to a combustor within said gas turbine engine.

10. The gas turbine engine as set forth in claim 9, wherein air temperatures at said downstream most location of said high pressure compressor are greater than or equal to 1350°

F.

11. The gas turbine engine as set forth in claim 3, wherein the second tap is at said downstream most end.

12. The gas turbine engine as set forth in claim 3, wherein air temperatures at said downstream most location of said high pressure compressor are greater than or equal to 1350°

F.