Compressors with integrated secondary air flow systems

Patent metadata

Field	Value
Publication number	US-10072522-B2
Application number	US-201113182976-A
Country	US
Kind code	B2
Filing date	Jul 14, 2011
Priority date	Jul 14, 2011
Publication date	Sep 11, 2018
Grant date	Sep 11, 2018

Abstract

A compressor includes a rotor platform; a rotor blade; and a casing having an inner surface surrounding the tip and spaced radially outwardly from the tip to define a gap. A secondary air flow system includes a bleed inlet configured to remove secondary air flow from the primary air flow; an injection opening disposed in the inner surface of the casing upstream of the bleed inlet; an accessory conduit; and a plenum fluidly coupled to the bleed inlet, the injection opening, and the accessory conduit. The bleed inlet and plenum at least partially define a secondary air flow path such that a first portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the injection opening and a second portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the accessory conduit.

First claim

What is claimed is:

1. A compressor for a gas turbine engine, comprising:
a first stator;
a rotor platform positioned downstream of the first stator;
an exit stator positioned downstream of the rotor platform;
a rotor blade extending radially outwardly from the rotor platform, the rotor blade including a pressure sidewall and a circumferentially opposing suction sidewall extending in a radial direction between a root and a tip and in an axial direction between a leading edge and a trailing edge;
a casing having an inner surface surrounding the tip and spaced radially outwardly from the tip to define a gap between the casing and the tip, the casing and the rotor platform defining a primary air flow path between the casing and the rotor platform to direct a primary air flow in a direct air in a downstream direction;
a secondary air flow system comprising
a bleed inlet disposed in the inner surface of the casing downstream of the rotor blade and configured to remove secondary air flow from the primary air flow;
an injection opening disposed in the inner surface of the casing upstream of the bleed inlet, wherein the injection opening is disposed in the casing upstream of the rotor blade;
an accessory conduit configured to be fluidly coupled to an accessory system of the gas turbine engine; and
a plenum fluidly coupled to the bleed inlet, the injection opening, and the accessory conduit,
wherein the bleed inlet and plenum at least partially define a secondary air flow path for the secondary air flow such that a first portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the injection opening and a second portion of the secondary air flow is directed in through the bleed inlet, through the plenum, and out through the accessory conduit; and
an impeller positioned downstream of the rotor blade and the casing and immediately downstream of the exit stator, the bleed inlet being arranged between the exit stator and the impeller.

2. The compressor of claim 1 , wherein the injection opening is configured to direct the first portion of secondary air flow into a gap between the rotor blade and the casing.

3. The compressor of claim 1 , wherein the bleed inlet is sized to remove the secondary air flow at a flow rate of less than 0.15 Mach at a maximum flow rate of the primary air flow.

4. The compressor of claim 1 , wherein the plenum is sized such that the secondary air flow has a flow rate of less than 0.05 Mach within the plenum.

5. The compressor of claim 1 , wherein the plenum is sized such that a distance between the accessory conduit and the bleed inlet is greater than twice a width of the bleed inlet.

6. The compressor of claim 1 , wherein the accessory conduit is a first accessory conduit, and wherein secondary air flow system further comprises a second accessory conduit fluidly coupled to the plenum for receiving a third portion of the secondary air flow.

7. The compressor of claim 6 , wherein the first accessory conduit has a first diameter and the second accessory conduit has a second diameter, the first accessory conduit being spaced from the second accessory conduit at a distance at least three times greater than the first diameter.

8. The compressor of claim 7 , wherein the distance is at least three times greater than the second diameter.

9. The compressor of claim 1 , wherein the gas turbine engine is arranged on an aircraft, and wherein the secondary air flow system further comprises an exhaust conduit in fluid communication with the plenum and configured to direct a third portion of the secondary air flow from the plenum to an exhaust system that dumps the third portion of the secondary air flow overboard the aircraft.

10. The compressor of claim 9 , wherein the plenum is sized such that a distance between the exhaust conduit and the bleed inlet is greater than twice a width of the bleed inlet.

11. The compressor of claim 9 , wherein the accessory conduit has a first diameter and the exhaust conduit has a second diameter, the accessory conduit being spaced from the exhaust conduit at a distance at least three times greater than the first diameter.

12. The compressor of claim 11 , wherein the distance is at least three times greater than the second diameter.

13. A compressor for a gas turbine engine, comprising:
a first stator;
a rotor platform positioned downstream of the first stator;
an exit stator positioned downstream of the rotor platform;
a rotor blade extending radially outwardly from the rotor platform, the rotor blade including a pressure sidewall and a circumferentially opposing suction sidewall extending in a radial direction between a root and a tip and in an axial direction between a leading edge and a trailing edge;
a casing having an inner surface surrounding the tip and spaced radially outwardly from the tip to define a gap between the casing and the tip, the casing and the rotor platform defining a primary air flow path between the casing and the rotor platform to direct a primary air flow in a direct air in a downstream direction;
a secondary air flow system comprising
a bleed inlet disposed in the inner surface of the casing downstream of the rotor blade and configured to remove secondary air flow from the primary air flow, the bleed inlet being sized to remove the secondary air flow at a flow rate of less than 0.15 Mach at a maximum flow rate of the primary air flow;
a plenum fluidly receiving the secondary air flow from the bleed inlet, the plenum being sized such that the secondary air flow has a flow rate of less than 0.05 Mach within the plenum;
an injection opening disposed in the inner surface of the casing upstream of the bleed inlet and fluidly coupled to the plenum for directing a first portion of secondary air flow from the plenum into a gap between the rotor blade and the casing;
an accessory conduit fluidly coupled to the plenum for directing a second portion of secondary air flow from the plenum to an accessory system of the gas turbine engine, wherein the plenum is sized such that a distance between the accessory conduit and the bleed inlet is greater than twice a width of the bleed inlet; and
an exhaust conduit fluidly coupled to the plenum for directing a third portion of secondary air flow from the plenum to an exhaust system of the gas turbine engine, wherein the plenum is sized such that a distance between the exhaust conduit and the bleed inlet is greater than twice a width of the bleed inlet; and

an impeller positioned downstream of the rotor blade and the casing and immediately downstream of the exit stator, the bleed inlet being arranged between the exit stator and the impeller.

14. The compressor of claim 9 , wherein the plenum is a common plenum directly connected to the bleed inlet, the injection opening, the accessory conduit, and the exhaust conduit.

15. The compressor of claim 1 , wherein the bleed inlet has a leading inlet edge that is angled in an upstream direction.

16. A gas turbine engine assembly, comprising: a compressor section comprising:
a rotor platform;
a rotor blade extending radially outwardly from the rotor platform, the rotor blade including a pressure sidewall and a circumferentially opposing suction sidewall extending in a radial direction between a root and a tip and in an axial direction between a leading edge and a trailing edge;
a casing having an inner surface surrounding the tip and spaced radially outwardly from the tip to define a gap between the casing and the tip, the casing and the rotor platform defining a primary air flow path between

Assignees

• Nolcheff Nick
• Lee Jong
• Honeywell Int Inc

Inventors

• Nolcheff Nick
• Lee Jong

Classifications

• F02C9/18

  by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} · CPC title

• F04D27/023

  Details or means for fluid extraction · CPC title

• F04D29/545

  Ducts · CPC title

• F04D27/0238

  Details or means for fluid reinjection · CPC title

• F04D27/0215

  Arrangements therefor, e.g. bleed or by-pass valves · CPC title