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EP-4741634-A1 - AIRCRAFT PROPULSION SYSTEM WITH TURBOCOMPRESSOR UNIT

EP4741634A1EP 4741634 A1EP4741634 A1EP 4741634A1EP-4741634-A1

Abstract

An aircraft propulsion system (20) includes a turbocompressor unit (40), an engine exhaust assembly (32), and a tube assembly (118). The turbocompressor unit (40) includes a compressor (44), a turbine (42), a turbocompressor rotational assembly (46), and a static structure (48). The static structure (48) forms a first cavity (72) and a second cavity (84). The engine exhaust assembly (32) includes an annular scroll (38) disposed between the first cavity (72) and the second cavity (84). The annular scroll (38) includes a scroll body (86) forming a flow channel (90), and a plurality of hollow stator vanes (94). The plurality of hollow stator vanes (94) is disposed within the flow channel (90). The tube assembly (118) includes a tubular body (120). The tubular body (120) extends through one of the plurality of hollow stator vanes (94) between and to the first cavity (72) and the second cavity (84). The tubular body (120) forms an internal passage (112) extending from the first axial tube end (130) to the second axial tube end (132). The internal passage (112) connects the first cavity (72) in fluid communication with the second cavity (84).

Inventors

  • LEFEBVRE, GUY
  • SYNNOTT, REMY

Assignees

  • PRATT & WHITNEY CANADA CORP.

Dates

Publication Date
20260513
Application Date
20251110

Claims (15)

  1. An aircraft propulsion system (20) comprising: a turbocompressor unit (40) including a compressor (44), a turbine (42), a rotational assembly (46), and a static structure (48), the rotational assembly (46) rotatable about a rotational axis (56), the rotational assembly (46) including a bladed compressor rotor (54) of the compressor (44), a bladed turbine rotor (52) of the turbine (42), and a shaft (50), the shaft (50) interconnecting the bladed compressor rotor (54) and the bladed turbine rotor (52), the static structure (48) forming a first cavity (68) and a second cavity (84), the first cavity (68) connected in fluid communication with the compressor (44), the second cavity (84) disposed at the turbine (42); an engine exhaust assembly (32) including an annular scroll (38) disposed between the first cavity (68) and the second cavity (84), the annular scroll (38) including a scroll body (86) and a plurality of hollow stator vanes (94), the scroll body (86) forming a flow channel (90) connected in fluid communication with the turbine (42), the plurality of hollow stator vanes (94) disposed within the flow channel (90) and circumferentially distributed about the rotational axis (56); and a tube assembly (118) including a tubular body (120), the tubular body (120) extending along a tube axis (128) between and to a first axial tube end (130) and a second axial tube end (132), the tubular body (120) extending through one of the plurality of hollow stator vanes (94) between and to the first cavity (68) and the second cavity (84), the tubular body (120) forming an internal passage (112) extending along the tube axis (128) from the first axial tube end (130) to the second axial tube end (132), the internal passage (112) connecting the first cavity (68) in fluid communication with the second cavity (84).
  2. The aircraft propulsion system (20) of claim 1, wherein the scroll body (86) extends axially between and to a first axial scroll end (96) and a second axial scroll end (98), and the scroll body (86) forms a portion of the second cavity (84) at the second axial scroll end (98).
  3. The aircraft propulsion system (20) of claim 2, wherein the tubular body (120) is mounted to the scroll body (86) at the second axial scroll end (98).
  4. The aircraft propulsion system (20) of any preceding claim, wherein the scroll body (86) forms a third cavity (78), and the tubular body (120) extends through the third cavity (78) to the first cavity (68).
  5. The aircraft propulsion system (20) of claim 4 when dependent on claims 2 or 3, wherein the third cavity (78) is formed at the first axial scroll end (96).
  6. The aircraft propulsion system (20) of claim 4 or 5, wherein the third cavity (78) is a dead cavity.
  7. The aircraft propulsion system (20) of any preceding claim, wherein the tube assembly (118) further includes a sealing body (122) mounted on the static structure (48) at the first cavity (68), the sealing body (122) includes an inner sealing surface (144) extending circumferentially about the tube axis (128), and the first axial tube end (130) is disposed within the sealing body (122) at the inner sealing surface (144).
  8. The aircraft propulsion system (20) of claim 7, wherein the tube assembly (118) further includes a seal (124) disposed at the first axial tube end (130), and the seal (124) is disposed in sealing engagement with the inner sealing surface (144).
  9. The aircraft propulsion system (20) of claim 7 or 8, wherein the first axial tube end (130) is axially translatable relative to the inner sealing surface (144) along the tube axis (128).
  10. The aircraft propulsion system (20) of any preceding claim, further comprising an engine (26) including an exhaust port (30), and the flow channel (90) forms a portion of a combustion gas flow path (58) from the exhaust port (30) to the turbine (42).
  11. The aircraft propulsion system (20) of any preceding claim, wherein the static structure (48) further includes a bearing assembly mounted to rotationally support the shaft (50), and the bearing assembly is connected in fluid communication with the first cavity (68).
  12. The aircraft propulsion system (20) of any preceding claim, wherein the static structure (48) includes a turbine case (66) circumscribing the bladed turbine rotor (52), and the second cavity (84) is disposed radially outward of the turbine case (66).
  13. The aircraft propulsion system (20) of any preceding claim, wherein the tubular body (120) is mounted to the scroll body (86).
  14. The aircraft propulsion system (20) of any preceding claim, wherein: the annular scroll (38) is axially disposed between the compressor (44) and the turbine (42); the scroll body (86) includes a first axial side wall (96), a second axial side wall (98), and an outer wall (108) forming the flow channel (90) of the annular scroll (38); the plurality of hollow stator vanes (94) extends between and to the first axial side wall (96) and the second axial side wall (98); and the second axial tube end (132) is mounted to the second axial side wall (98).
  15. The aircraft propulsion system (20) of any of claim 8 to 14, wherein the tubular body (120) forms a seal groove (136) at the first axial tube end (130), the seal (124) is disposed in the seal groove (136), the seal (124) is sealingly engaged with the sealing body (122), and the first axial tube end (130) is moveable within the sealing body (122).

Description

TECHNICAL FIELD This disclosure relates generally to aircraft propulsion systems, and more particularly to air flow arrangements for turbocompressors. BACKGROUND OF THE ART Engines for aircraft may typically include rotational equipment configured for facilitating aircraft propulsion and/or other functions of aircraft propulsion system operation. In many cases, rotational equipment components and supporting static structures may require cooling, for example, using air from one or more compressed air sources. Various systems for distributing cooling air are known in the art. While these known systems may be useful for their intended purposes, there is always room in the art for improvement. SUMMARY According to one aspect of the present invention, there is provided an aircraft propulsion system including a turbocompressor unit, an engine exhaust assembly, and a tube assembly. The turbocompressor unit includes a compressor, a turbine, a rotational assembly, and a static structure. The rotational assembly is rotatable about a rotational axis. The rotational assembly includes a bladed compressor rotor of the compressor, a bladed turbine rotor of the turbine, and a shaft. The shaft interconnects the bladed compressor rotor and the bladed turbine rotor. The static structure forms a first cavity and a second cavity. The first cavity is connected in fluid communication with the compressor. The second cavity is disposed at the turbine. The engine exhaust assembly includes an annular scroll disposed between the first cavity and the second cavity. The annular scroll includes a scroll body and a plurality of hollow stator vanes. The scroll body forms a flow channel connected in fluid communication with the turbine. The plurality of hollow stator vanes is disposed within the flow channel and circumferentially distributed about the rotational axis. The tube assembly includes a tubular body. The tubular body extends along a tube axis between and to a first axial tube end and a second axial tube end. The tubular body extends through one of the plurality of hollow stator vanes between and to the first cavity and the second cavity. The tubular body forms an internal passage extending along the tube axis from the first axial tube end to the second axial tube end. The internal passage connects the first cavity in fluid communication with the second cavity. Optionally, and in accordance with the above, the scroll body may extend axially between and to a first axial scroll end and a second axial scroll end, and the scroll body may form a portion of the second cavity at the second axial scroll end. Optionally, and in accordance with any of the above, the tubular body may be mounted to the scroll body at the second axial scroll end. Optionally, and in accordance with any of the above, the scroll body may form a third cavity at the first axial scroll end, and the tubular body may extend through the third cavity to the first cavity. Optionally, and in accordance with any of the above, the third cavity may be a dead cavity. Optionally, and in accordance with any of the above, the tube assembly may further include a sealing body mounted on the static structure at the first cavity, the sealing body may include an inner sealing surface extending circumferentially about the tube axis, and the first axial tube end may be disposed within the sealing body at the inner sealing surface. Optionally, and in accordance with any of the above, the tube assembly may further include a seal disposed at the first axial tube end, and the seal may be disposed in sealing engagement with the inner sealing surface. Optionally, and in accordance with any of the above, the first axial tube end may be axially translatable relative to the inner sealing surface along the tube axis. Optionally, and in accordance with any of the above, the aircraft propulsion system may further include an engine including an exhaust port, and the flow channel may form a portion of a combustion gas flow path from the exhaust port to the turbine. Optionally, and in accordance with any of the above, the static structure may further include a bearing assembly mounted to rotationally support the shaft, and the bearing assembly may be connected in fluid communication with the first cavity. According to another aspect of the present invention, there is provided an aircraft propulsion system including a turbocompressor unit, an engine exhaust assembly, and a tube assembly. The turbocompressor unit includes a compressor, a turbine, a rotational assembly, and a static structure. The rotational assembly is rotatable about a rotational axis. The rotational assembly includes a bladed compressor rotor of the compressor, a bladed turbine rotor of the turbine, and a shaft, the shaft interconnecting the bladed compressor rotor and the bladed turbine rotor. The static structure forming a first cavity and a second cavity. The first cavity is connected in fluid communication with the compressor. The