Rim seal for gas turbine engine

Patent metadata

Field	Value
Publication number	US-10533445-B2
Application number	US-201615244203-A
Country	US
Kind code	B2
Filing date	Aug 23, 2016
Priority date	Aug 23, 2016
Publication date	Jan 14, 2020
Grant date	Jan 14, 2020

Abstract

A rim seal for a rotor of a gas turbine engine includes a seal portion extending circumferentially across a rim cavity of a rotor, the sealing portion configured to seal the rim cavity and a first foot portion extending radially inwardly from a first end of the sealing portion. A rotor assembly for a gas turbine engine includes a rotor disc and a plurality of rotor blades secured to the rotor disc defining a rim cavity between the rotor disc and a rim portion of the plurality of rotor blades. A rim seal is located in the rim cavity and includes a seal portion extending circumferentially across the rim cavity, the sealing portion configured to seal the rim cavity. The seal portion has an increasing radial thickness with increasing distance from a first end of the rim seal and from a second end opposite the first end.

First claim

The invention claimed is:

1. A rim seal for a rotor of a gas turbine engine, the rotor including a rotor disc, a plurality of rotor blades secured to the rotor disc each including a blade platform with a blade platform top and a blade platform bottom, a retention portion extending radially inboard from the blade platform, and a rim portion extending circumferentially from the retention portion spaced inboard from the blade platform bottom, adjacent rim portions defining a rim cavity between the rim portions and the rotor disc, the seal comprising:
a seal portion extending circumferentially across the rim cavity, the sealing portion configured to seal the rim cavity, the seal portion extending radially outwardly to a seal peak disposed in the rim cavity and midway between a first circumferential seal end and a second circumferential seal end; and
a first foot portion extending radially inwardly from the first circumferential seal end of the sealing portion.

2. The rim seal of claim 1 , further comprising a second foot portion extending radially inwardly from the second circumferential seal end of the sealing portion opposite the first circumferential seal end.

3. The rim seal of claim 1 , wherein the seal portion has an increasing radial thickness with increasing distance from a circumferentially inboard end of the first foot.

4. The rim seal of claim 1 , wherein the seal portion includes a seal surface configured to abut the rim portion of the plurality of rotor blades.

5. The rim seal of claim 4 , wherein the seal surface has a cross-sectional shape such that a portion of the seal surface cross-sectional shape matches a cross-sectional shape of the rim portion.

6. A rotor assembly for a gas turbine engine comprising:
a rotor disc;
a plurality of rotor blades secured to the rotor disc each rotor blade including:
a blade platform with a blade platform top and a blade platform bottom;
a retention portion extending radially inboard from the blade platform; and
a rim portion extending circumferentially from the retention portion spaced inboard from the blade platform bottom, adjacent rim portions defining a rim cavity between the rotor disc and the rim portions; and

a rim seal disposed in the rim cavity including a seal portion extending circumferentially across the rim cavity, the sealing portion configured to seal the rim cavity;
wherein the seal portion has an increasing radial thickness with increasing distance from a first circumferential seal end of the rim seal and from a second circumferential seal end opposite the first end, the seal portion extending radially outwardly to a seal peak disposed in the rim cavity and midway between the first circumferential seal end and the second circumferential seal end.

7. The rotor assembly of claim 6 , further comprising a first foot portion extending radially inwardly from the first circumferential seal end of the rim seal, the seal portion having an increasing radial thickness with increasing distance from a circumferentially inboard end of the first foot.

8. The rotor assembly of claim 7 , further comprising a second foot portion extending radially inwardly from the second circumferential seal end of the rim seal opposite the first circumferential seal end.

9. The rotor assembly of claim 6 , wherein the seal portion includes a seal surface configured to abut the rim portion.

10. The rotor assembly of claim 9 , wherein the seal surface has a cross-sectional shape such that a portion of the seal surface cross-sectional shape matches a cross-sectional shape of the rim portion.

11. The rotor assembly of claim 6 , wherein the rim cavity is defined between circumferentially adjacent rotor blades of the plurality of rotor blades.

12. The rotor assembly of claim 11 , wherein the rim seal prevents leakage through a circumferential gap between the circumferentially adjacent rim portions.

13. The rotor assembly of claim 6 , wherein the rotor assembly is a turbine rotor assembly.

14. A gas turbine engine comprising:
a combustor; and
a rotor assembly in fluid communication with the combustor including:
a rotor disc;
a plurality of rotor blades secured to the rotor disc each rotor blade including:
a blade platform with a blade platform top and a blade platform bottom;
a retention portion extending radially inboard from the blade platform; and
a rim portion extending circumferentially from the retention portion spaced inboard from the blade platform bottom, adjacent rim portions defining a rim cavity between the rotor disc and the rim portions; and
a rim seal disposed in the rim cavity including:
a seal portion extending circumferentially across the rim cavity, the sealing portion configured to seal the rim cavity, the seal portion extending radially outwardly to a seal peak disposed in the rim cavity and midway between a first circumferential seal end and a second circumferential seal end; and
a first foot portion extending radially inwardly from the first circumferential seal end of the sealing portion.

15. The gas turbine engine of claim 14 , further comprising a second foot portion extending radially inwardly from the second circumferential seal end of the sealing portion opposite the first end.

16. The gas turbine engine of claim 14 , wherein the seal portion has an increasing radial thickness with increasing distance a circumferentially inboard end of the first foot.

17. The gas turbine engine of claim 14 , wherein the seal portion includes a seal surface configured to abut the rim portion.

18. The gas turbine engine of claim 17 , wherein the seal surface has a cross-sectional shape such that a portion of the seal surface cross-sectional shape matches a cross-sectional shape of the rim portion.

19. The gas turbine engine of claim 14 , wherein the rim cavity is defined between circumferentially adjacent rotor blades of the plurality of rotor blades.

20. The gas turbine engine of claim 19 , wherein the rim seal prevents leakage through a circumferential gap between the circumferentially adjacent rim portions.

Assignees

• United Technologies Corp

Inventors

• Banks Anton G
• Corcoran Christopher
• Dube Bryan P

Classifications

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F04D29/322

  Blade mountings · CPC title

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F01D11/00

  Preventing or minimising internal leakage of working-fluid, e.g. between stages (sealings in general F16J {; sealing arrangements for transition ducts of combustor cans F01D9/023}) · CPC title

• F04D29/083

  especially adapted for elastic fluid pumps · CPC title