Adaptive fan with cold turbine

• US 9,157,366 B2
• Filed: 05/30/2012
• Issued: 10/13/2015
• Est. Priority Date: 05/30/2012
• Status: Active Grant

• Current Assignee: Rtx Corporation
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Kupratis, Daniel Bernard
• Primary Examiner(s): SUTHERLAND, STEVEN M
• Application Number:
  US13/483,449
• Publication Number:
  US 20130318981A1
• Time in Patent Office: 1,231 Days
• Field of Search: 607/73, 602/261, 602/68, 607/92, 60/391.62, 602/62, 415/68, 415/69, 415/1, 415144-145, 415/159, 415/162, 416/129, 416/198.A
• US Class Current: 1/1
• CPC Class Codes: F02C 3/06 the compressor comprising only axial stages F02C3/10 takes precedence
• CPC Class Codes: F02C 3/10 with another turbine driving an output shaft but not driving the compressor
• CPC Class Codes: F02C 9/16 Control of working fluid flow F02C9/48 takes precedence; control of air-intake flow F02C7/057
• CPC Class Codes: F02K 3/06 with front fan
• CPC Class Codes: F02K 3/075 controlling flow ratio between flows

Abstract

A disclosed gas turbine engine includes a first fan section including a plurality of fan blades rotatable about an axis, a compressor in fluid communication with the first fan section, a combustor in fluid communication with the compressor and a first turbine section in fluid communication with the combustor. The first turbine section includes a low pressure turbine that drives the first fan section. A second fan section is supported between the first fan section and the compressor and is driven by a second turbine section disposed between the second fan section and the compressor for driving the second fan section.

Claims

1. A gas turbine engine comprising:

a first fan section including a plurality of fan blades rotatable about an axis;

a compressor in fluid communication with the first fan section;

a combustor in fluid communication with the compressor;

a first turbine section in fluid communication with the combustor, the first turbine section driving the first fan section and the compressor;

a core flow path directing airflow to the compressor;

a first bypass radially outward of the core flow path;

a first splitter fixed between the core flow path and the first bypass;

a second fan section disposed between the first fan section and the compressor directing airflow into the first bypass and the core flow path; and

a second turbine section disposed axially aft of a forward end of the first splitter within the core flow path between the second fan section and the compressor, the second turbine section coupled to the second fan section for driving the second fan section.

2. The gas turbine engine as recited in claim 1, including a variable vane disposed between the second fan section and the second turbine section for controlling airflow through the second turbine section.

3. The gas turbine engine as recited in claim 2, wherein the variable vane is movable between a first position for driving the second turbine at a first speed and a second position for driving the second turbine at a second speed slower than the first speed.

4. The gas turbine engine as recited in claim 2, including a fixed vane disposed between the second fan section and the variable vane.

5. The gas turbine engine as recited in claim 1, including a second bypass stream defined radially outward of the core flow path and the first bypass.

6. The gas turbine engine as recited in claim 5, wherein the second fan section directs airflow only into the first bypass and the core flow path.

7. The gas turbine engine as recited in claim 6, including a second splitter between the first bypass and the second bypass stream, with the second fan disposed axially aft of a forward most portion of the second splitter.

8. The gas turbine engine as recited in claim 1, wherein the second turbine and the second fan section are attached.

9. The gas turbine engine as recited in claim 1, including an augmenter aft of the first turbine section.

10. The gas turbine engine as recited in claim 1, wherein the second turbine is disposed only within the core flow path.

11. A gas turbine engine comprising:

a first fan section including a plurality of fan blades rotatable about an axis;

a core engine including a compressor in fluid communication with the first fan section, a combustor in fluid communication with the compressor and a first turbine section in fluid communication with the combustor, the turbine section driving the first fan section and the compressor;

a core flow passage providing air to the core engine;

an annular bypass passage disposed about the core engine;

a fixed splitter that includes a forward end that splits incoming airflow between the core flow passage and the annular bypass passage;

a second fan section disposed axially forward of the fixed splitter and between the first fan section and the compressor, the second fan section driving flow into the annular bypass passage; and

a second turbine section disposed within the core flow passage and axially aft of the forward end of the fixed splitter between the second fan section and the compressor, wherein the second turbine section is coupled to the second fan section.

12. The gas turbine engine as recited in claim 11, including a variable vane disposed between the second fan section and the second turbine section for controlling airflow through the second turbine section.

13. The gas turbine engine as recited in claim 12, wherein the variable vane is movable between a first position for driving the second turbine at a first speed and a second position for driving the second turbine at a second speed slower than the first speed.

14. The gas turbine engine as recited in claim 11, wherein the annular bypass passage comprises an inner annular bypass passage disposed about the core engine, an outer annular bypass passage radially outward of the inner bypass passage, and a splitter disposed between the inner and outer annular bypass passages.

15. The gas turbine engine as recited in claim 11, including an augmenter aft of the first turbine section.

16. The gas turbine engine as recited in claim 11, including at least one bearing assembly supporting rotation of the second fan section and the second turbine section.

17. A method of operating a gas turbine engine comprising

defining a core gas flow path through a core engine, where the core engine includes a high pressure compressor, a combustor in communication with the compressor and a first turbine driven by gas flow generated by the combustor;

adjusting a vane disposed within the core gas flow path to a first position for driving a second turbine disposed forward of the high pressure compressor and only within in the core gas flow path to provide a first condition of core flow to the high pressure compressor; and

adjusting the vane to a second position for driving the second turbine to provide a second condition of core flow to the high pressure compressor.

18. The method as recited in claim 17, wherein the first condition comprises a first temperature of flow through the high pressure compressor and the second condition comprises a second temperature of flow through the high pressure compressor.

19. The method as recited in claim 18, including driving a fan blade forward of the core gas flow path with the second turbine for controlling flow into the gas flow path.

20. The method as recited in claim 17, including providing a maximum level of de-supercharging of flow into the core gas flow path at a first flight condition and providing a minimum level of de-supercharging of flow into the core gas flow path at a second flight condition.