Dual-mode plug nozzle

Patent metadata

Field	Value
Publication number	US-11143052-B2
Application number	US-201916267978-A
Country	US
Kind code	B2
Filing date	Feb 5, 2019
Priority date	May 31, 2013
Publication date	Oct 12, 2021
Grant date	Oct 12, 2021

Abstract

A method for controlling flow through an exhaust nozzle includes: providing a centerbody including a maximum diameter section; providing an inner shroud surrounding the centerbody, including at least a middle section of decreased diameter and terminating at an aft edge; providing an outer shroud. wherein the centerbody and the inner shroud collectively define a throat, and the outer shroud and the centerbody collectively define an exit; selectively translating the inner shroud and outer shroud to vary the throat; and selectively translating the outer shroud to vary the ratio of the exit to the throat; wherein, when the inner shroud is in a forward position, its aft edge is forward of the maximum diameter section of the centerbody, such that the throat of the nozzle is formed between the aft edge of the inner shroud and the centerbody.

First claim

What is claimed is:

1. A method for controlling a fluid flow through an exhaust nozzle of a gas turbine engine for a supersonic aircraft, the method comprising:
providing a centerbody extending rearward along a longitudinal axis, the centerbody including a maximum diameter section relative to the remainder of the centerbody;
providing an inner shroud surrounding the centerbody, the inner shroud having an outer surface and an inner surface, the inner surface including at least a middle section of decreased diameter relative to the remainder of the inner surface, the inner shroud terminating at an aft edge;
providing an outer shroud surrounding the inner shroud, the outer shroud having a forward edge, an aft edge, and an inner surface extending from the forward edge to the aft edge, wherein the centerbody and the inner shroud collectively define a throat area of the exhaust nozzle, and the outer shroud and the centerbody collectively define an exit area of the exhaust nozzle;
selectively translating the inner shroud and the outer shroud to vary the throat area; and
selectively translating the outer shroud to vary a ratio of the exit area to the throat area,
wherein the inner shroud is movable between forward and aft positions, wherein, when the inner shroud is in the forward position, the aft edge of the inner shroud is forward of the maximum diameter section of the centerbody, such that the throat area of the exhaust nozzle is formed between the aft edge of the inner shroud and the centerbody, and
wherein the outer shroud is movable between forward and aft positions, wherein, when the outer shroud is in the forward position, the aft edge of the outer shroud is forward of the maximum diameter section of the centerbody, such that the exit area of the exhaust nozzle is formed between the aft edge of the outer shroud and the centerbody.

2. The method of claim 1 , wherein the centerbody includes, sequentially, a sloping forward section, the maximum diameter section, and an aft section.

3. The method of claim 2 , wherein the aft section of the centerbody tapers in diameter to form an aft-facing conical shape.

4. The method of claim 1 , wherein the throat area is at a minimum size when the inner shroud is at the aft position, and the throat area is at a maximum size for a converging-diverging nozzle when the inner shroud is in an intermediate position, and is at a maximum size for a converging nozzle when the inner shroud is in the forward position.

5. The method of claim 1 , further comprising independent translation of the centerbody with respect to the inner shroud and the outer shroud.

6. A gas turbine engine for a supersonic aircraft, the gas turbine engine having a nozzle for directing exhaust flow to atmosphere, the nozzle comprising:
a centerbody extending along a longitudinal axis and including, sequentially, a sloping forward section, a maximum diameter section, and an aft section;
an annular inner shroud having an outer surface and an inner surface, the inner surface including a middle section of decreased diameter relative to the remainder of the inner surface, the annular inner shroud terminating at an aft edge and being selectively moveable along the longitudinal axis between forward and aft positions relative to the centerbody;
an annular outer shroud having an inner surface and an outer surface, the annular outer shroud being selectively movable between forward and aft positions relative to the centerbody; and
actuators operable to independently translate the annular inner shroud and the annular outer shroud between the respective forward and aft positions of the annular inner shroud and the annular outer shroud,
wherein, when the annular inner shroud is in the forward position, the aft edge of the annular inner shroud is forward of the maximum diameter section of the centerbody, such that a throat area of the nozzle is formed between the aft edge of the annular inner shroud and the sloping forward section of the centerbody, and
wherein, when the annular outer shroud is in the forward position, an aft edge of the annular outer shroud is forward of the maximum diameter section of the centerbody, such that an exit area of the nozzle is formed between the aft edge of the annular outer shroud and the centerbody.

7. The gas turbine engine of claim 6 , wherein the inner surface of the annular outer shroud defines a substantially constant cross-sectional area from a forward edge thereof to the aft edge thereof.

8. The gas turbine engine of claim 6 , wherein the aft section of the centerbody tapers in diameter to form an aft-facing conical shape.

9. The gas turbine engine of claim 6 , wherein the centerbody, the annular inner shroud, and the annular outer shroud are bodies of revolution about the longitudinal axis.

10. The gas turbine engine of claim 6 , wherein the centerbody, the annular inner shroud, and the annular outer shroud are oval in cross-section.

11. A gas turbine engine for a supersonic aircraft, the gas turbine engine comprising:
a compressor, a combustor, and a turbine disposed in series flow sequence along a longitudinal axis;
a nozzle disposed downstream of the turbine, including:
a centerbody extending along the longitudinal axis and including, sequentially, a sloping forward section, a maximum diameter section, and an aft section;
an annular inner shroud having an outer surface and an inner surface, the inner surface including a middle section of decreased diameter relative to the remainder of the inner surface, the annular inner shroud terminating at an aft edge and being selectively moveable along the longitudinal axis between forward and aft positions relative to the centerbody;
an annular outer shroud having an inner surface and an outer surface, the annular outer shroud being selectively movable between forward and aft positions relative to the centerbody; and
actuators operable to independently translate the annular inner shroud and the annular outer shroud parallel to the longitudinal axis,
wherein, when the annular inner shroud is in the forward position, the aft edge of the annular inner shroud is forward of the maximum diameter section of the centerbody, such that a throat area of the nozzle is formed between the aft edge of the annular inner shroud and the sloping forward section of the centerbody, and
wherein, when the annular outer shroud is in the forward position, an aft edge of the annular outer shroud is forward of the maximum diameter section of the centerbody, such that an exit area of the nozzle is formed between the aft edge of the annular outer shroud and the centerbody.

12. The gas turbine engine of claim 11 , wherein the inner surface of the annular outer shroud defines a substantially constant cross-sectional area from a forward edge thereof to the aft edge thereof.

13. The gas turbine engine of claim 11 , wherein the aft section of the centerbody tapers in diameter to form an aft-facing conical shape.

14. The gas turbine engine of claim 11 , wherein the centerbody, the annular inner shroud, and the annular outer shroud are bodies of revolution about the longitudinal axis.

15. The gas turbine engine of claim 11 , wherein the centerbody, the annular inner shroud, and the annular outer shroud are oval in cross-section.

Assignees

• Gen Electric

Inventors

• Tersmette Trevor Andrew
• Lee Jason Adam

Classifications

• F02K1/08 Primary

  by axially moving or transversely deforming an internal member, e.g. the exhaust cone · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F01D9/02

  Nozzles; Nozzle boxes; Stator blades; Guide conduits {, e.g. individual nozzles (nozzle boxes F01D9/047)} · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F02K1/09

  by axially moving an external member, e.g. a shroud (F02K1/12 takes precedence) · CPC title