Coolant flow redirection component

Patent metadata

Field	Value
Publication number	US-9890645-B2
Application number	US-201514841265-A
Country	US
Kind code	B2
Filing date	Aug 31, 2015
Priority date	Sep 4, 2014
Publication date	Feb 13, 2018
Grant date	Feb 13, 2018

Abstract

A gas turbine engine includes a compressor section, a combustor fluidly connected to the compressor section, and a turbine section fluidly connected to the combustor and mechanically connected to the compressor section via a shaft. Multiple rotors are disposed in one of the compressor section and the turbine section. Each of the rotors includes a rotor disk portion having a radially inward bore, and is static relative to the shaft. Each rotor is axially adjacent at least one other rotor and a gap is defined between each rotor and an adjacent rotor. A cooling passage for a cooling flow is defined between the shaft and the rotors, and a cooling flow redirection component is disposed at the gap and is operable to redirect the cooling flow in the cooling passage into the gap.

First claim

The invention claimed is:

1. A gas turbine engine comprising:
a compressor section;
a combustor fluidly connected to the compressor section;
a turbine section fluidly connected to the combustor and mechanically connected to said compressor section via a shaft;
a plurality of rotors disposed in one of said compressor section and said turbine section, each of said rotors including a rotor disk portion including a radially inward bore, and each of said rotors being static relative to said shaft;
each rotor in said plurality of rotors being axially adjacent at least one other of said rotors in said plurality of rotors and defining a gap between each of said rotors and said axially adjacent rotors, wherein each of said rotors includes a joint arm partially crossing said gap;
a cooling passage for a cooling flow defined between said shaft and said rotors; and
a cooling flow redirection component disposed at said gap and operable to redirect said cooling flow in said cooling passage into said gap, wherein the cooling flow redirection component is distinct from said rotors in said plurality of rotors, wherein the cooling flow redirection component comprises a cooling flow redirection portion, said cooling flow redirection portion being angled relative to flow through said cooling passage such that coolant flow is redirected into said gap, and a radially aligned portion including a tapered through hole having an inlet with a first cross sectional area, and an outlet with a second cross sectional area, the radially aligned portion being radially outward from said cooling flow redirection portion, relative to a radius of the engine.

2. The gas turbine engine of claim 1 , wherein said gap is defined between radially aligned surfaces of adjacent rotor disks.

3. The gas turbine engine of claim 1 , wherein said cooling flow redirection component includes a radially outward protrusion from said shaft.

4. The gas turbine engine of claim 3 , wherein cooling flow through said cooling passage contacts a planar surface of said cooling flow redirection component.

5. The gas turbine engine of claim 1 , wherein the second cross sectional area is smaller than the first cross sectional area.

6. The gas turbine engine of claim 1 , wherein said cooling flow redirection component is integral to one of said rotors defining said gap.

7. The gas turbine engine of claim 1 , further comprising a secondary cooling flow inlet positioned at a radially outward edge of said gap.

8. The gas turbine engine of claim 1 , wherein said cooling flow redirection component includes at least one cooling flow acceleration component positioned on an upstream surface of said cooling flow redirection component, relative to cooling flow through said cooling flow passage.

9. The gas turbine engine of claim 1 , wherein each of said joint arms contacts a joint arm of an adjacent rotor at a joint contact.

10. The gas turbine engine of claim 9 , wherein the cooling flow redirection component includes a portion disposed at the joint contact.

11. The gas turbine engine of claim 10 , wherein the joint contact is configured to retain the portion of the cooling flow redirection component disposed at the joint contact.

12. The gas turbine engine of claim 1 , wherein each of said joint arms is integral to the corresponding rotor.

13. The gas turbine engine of claim 1 , wherein the tapered through hole defines a through hole axis, and wherein the through hole axis is aligned with an axis defined by the shaft.

14. The gas turbine engine of claim 1 , wherein the flow redirection portion is a planar surface, and wherein the cooling flow redirection portion is angled relative to an axis defined by the shaft.

Assignees

• United Technologies Corp

Inventors

• Mccaffrey Michael G
• Vinson Elizabeth F

Classifications

• F05D2260/201

  by impingement of a fluid · CPC title

• F04D29/053

  Shafts · CPC title

• F05D2240/60

  Shafts · CPC title

• F04D29/321

  for axial flow compressors · CPC title

• F04D29/584

  cooling or heating the machine (F04D29/5846, F04D29/5853 take precedence) · CPC title