Turbine vane assembly with integrally cast cooling fluid nozzle

• US 5,217,348 A
• Filed: 09/24/1992
• Issued: 06/08/1993
• Est. Priority Date: 09/24/1992
• Status: Expired due to Term

• Current Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: United Technologies Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: Rup, John J. Jr., Nikkanen, John P.
• Primary Examiner(s): KWON, JOHN
• Application Number:
  US07/949,968
• Time in Patent Office: 257 Days
• Field of Search: 415/115, 415/116
• US Class Current: 415/115
• CPC Class Codes: F01D 11/001 for sealing space between stator blade and rotor
• CPC Class Codes: F01D 25/12 Cooling

Abstract

A gas turbine engine having a turbine vane assembly including an integrally cast cooling fluid nozzle is disclosed. Various construction details are developed which disclose a cooling fluid nozzle including a flow passage having an exit and a wall. In one embodiment, the wall includes an angled leading edge which mates with a circumferentially adjacent trailing edge of an adjacent wall. The leading edge is tapered such that in a most open position of the turbine vane assembly the leading edge and the trailing edge circumferentially align. In a most closed position of the turbine vane assembly, the angled leading edge aligns with the trailing edge such that a step down is created in the circumferentially directed flow of the sealed cavity. The plurality of wall thereby produce a waterfall arrangement within the sealed cavity to reduce the windage losses.

Claims

1. A gas turbine engine disposed about a longitudinal axis and including an axially extending flow path, a turbine section, and means to conduct cooling fluid into the turbine section, the turbine section including a rotor assembly disposed circumferentially about the axis, sealing means disposed axially downstream of the rotor assembly, the rotor assembly and sealing means adapted to rotate about the axis in an operational condition, and a turbine vane assembly disposed axially downstream of the rotor assembly and radially outward of the sealing means, wherein an annular seal cavity is defined in part by the separation between the rotor assembly, the turbine vane assembly and the sealing means, the turbine vane assembly including a plurality of vanes, and a sealing shroud adapted to engage the sealing means to block the passage of fluid from the cavity, wherein each vane has a pitch angle γ

with γ

≧

γ

₁ wherein γ

₁ is the most open pitch angle, each vane including an airfoil portion, a platform, and an integrally cast nozzle, the airfoil portion being hollow and in fluid communication with the means to conduct cooling fluid, the nozzle including wall means having a flow surface facing the cavity, and a smooth, continuous flow passage in fluid communication with the airfoil portion and adapted to direct cooling fluid into the cavity, the wall means extending circumferentially between adjacent wall means and extending radially inward from the platform, each wall means having a trailing edge and a leading edge, the leading edge adapted to circumferentially align with the trailing edge of an adjacent wall means with γ

=γ

₁, and the leading edge adapted to form a step down relative to the direction of flow within the cavity with γ

>

γ

₁.

2. The gas turbine engine according to claim 1, wherein the nozzle is adapted to direct cooling fluid into an axially forward, radially inward portion of the cavity.

3. The gas turbine engine according to claim 1, wherein the nozzle further includes a throat portion adapted to meter the cooling fluid entering the cavity.

4. A turbine vane assembly for a turbomachine disposed about a longitudinal axis and including an axially extending flow path, a turbine section, and means to conduct cooling fluid into the turbine section, the turbine section including a rotor assembly disposed circumferentially about the axis, sealing means disposed axially downstream of the rotor assembly, the rotor assembly and sealing means adapted to rotate about the axis in an operational condition, and a turbine vane assembly disposed axially downstream of the rotor assembly and radially outward of the sealing means, wherein an annular seal cavity is defined in part by the separation between the rotor assembly, the turbine vane assembly and the sealing means, the seal cavity having effectively continuous flow surfaces, the turbine vane assembly including a plurality of vanes and a sealing shroud adapted to engage the sealing means to block the passage of fluid from the cavity, wherein each vane has a pitch angle γ

with γ

≧

γ

₁ wherein γ

₁ is the most open pitch angle, each vane including an airfoil portion, a platform, and an integrally cast nozzle, the airfoil portion being hollow and in fluid communication with the means to conduct cooling fluid, the nozzle including wall means having a flow surface facing the cavity, and a smooth, continuous flow passage in fluid communication with the airfoil portion and adapted to direct cooling fluid into the cavity, the wall means extending circumferentially between adjacent wall means and extending radially inward from the platform, each wall means having a trailing edge and a leading edge, the leading edge adapted to circumferentially align with the trailing edge of an adjacent wall means with γ

=γ

₁, and the leading edge adapted to form a step down relative to the direction of flow within the cavity with γ

>

γ

₁.

5. The turbine vane assembly according to claim 4, wherein the nozzle is adapted to inject cooling fluid in a tangential direction relative to the flow within the seal cavity and to direct cooling fluid into an axially forward, radially inward portion of the cavity.

6. The turbine vane assembly according to claim 4, wherein the nozzle further includes a throat portion adapted to meter the cooling fluid entering the cavity.