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US 10,378,453 B2Granted
Method and assembly for reducing secondary heat in a gas turbine engine
AssigneeUnited Technologies Corporation
InventorsGabriel L. Suciu, Julian Partyka, David R. Pack, James D. Hill
Life of this patent
Estimating the term…
Abstract
A turbine section for a gas turbine engine includes a first rotor assembly with a first rotor assembly bleed air source and an aft cavity that is in fluid communication with the first rotor assembly bleed air source. A second rotor assembly includes a forward cavity. A vane bleed air source is in fluid communication with the forward cavity. A seal extends between the first rotor assembly and the second rotor assembly.
Claims16
A turbine section for a gas turbine engine comprising:
- a first rotor assembly;
- a second rotor assembly located downstream of the first rotor assembly;
- a vane including a cooling passage having an outlet on a radial inner end of the vane;
- and
- a seal extending between the first rotor assembly and the second rotor assembly, the seal including a first plurality of openings in fluid communication with the cooling passage and a second plurality of opening directed at the second rotor assembly and spaced radially inward from the first plurality of openings, wherein the seal and the first rotor assembly at least partially define a first stage aft cavity and the seal and the second rotor assembly at least partially define a second stage forward cavity and the seal is configured to rotate with the first and second rotor assemblies.
2 more elements
A gas turbine engine comprising:
- a turbine section including:
- a first rotor assembly;
- a second rotor assembly located downstream of the first rotor assembly;
- a vane including a cooling passage on a radial inner end;
- and
- a seal extending between the first rotor assembly and the second rotor assembly, the seal including a first plurality of openings in fluid communication with the cooling passage and a second plurality of opening directed at the second rotor assembly and spaced radially inward from the first plurality of openings, wherein the seal and the first rotor assembly at least partially define a first stage aft cavity and the seal and the second rotor assembly at least partially define a second stage forward cavity and the seal is configured to rotate with the first and second rotor assemblies.
3 more elements
A method of operating a gas turbine engine comprising:
- directing bleed air from a first rotor assembly at a first pressure to purge a first stage aft cavity;
- and
- directing bleed air through a radially inner end of a vane at a second pressure to purge a second stage forward cavity and through a seal, wherein the first pressure is greater than the second pressure and the seal extends between a first rotor assembly and a second rotor assembly, the seal includes a first plurality of openings in fluid communication with a cooling passage in the vane and a second plurality of opening directed at the second rotor assembly and spaced radially inward from the first plurality of openings, wherein the seal and the first rotor assembly at least partially define the first stage aft cavity and the seal and the second rotor assembly at least partially define the second stage forward cavity and the seal is configured to rotate with the first and second rotor assembly.
Dependencies are inferred from the claim text.