Dual function cascade integrated variable area fan nozzle and thrust reverser

Patent metadata

Field	Value
Publication number	US-10677192-B2
Application number	US-201715674576-A
Country	US
Kind code	B2
Filing date	Aug 11, 2017
Priority date	Oct 12, 2006
Publication date	Jun 9, 2020
Grant date	Jun 9, 2020

Abstract

A gas turbine engine system according to an exemplary aspect of the present disclosure may include a fan nacelle that extends circumferentially about a fan, and at least one integrated mechanism coupled to the fan nacelle. The at least one integrated mechanism includes a variable fan nozzle and a thrust reverser, with the thrust reverser and the variable fan nozzle having a common part.

First claim

What is claimed is:

1. A gas turbine engine comprising:
a fan nacelle that extends circumferentially about a fan;
a bypass ratio of greater than 10;
a low pressure compressor and a high pressure compressor;
a gear train driving the fan, the gear train having a gear reduction ratio of greater than 2.5;
a combustion section in communication with the low pressure compressor and the high pressure compressor;
a low pressure turbine and a high pressure turbine, the low pressure turbine having a pressure ratio of greater than 5, with the low pressure turbine and high pressure turbine coupled for rotation with respective spools to rotationally drive the low pressure compressor, the high pressure compressor and the gear train, and the fan driven through the gear train;
at least one integrated mechanism coupled to the fan nacelle, the at least one integrated mechanism including a variable fan nozzle and a thrust reverser, the thrust reverser and the variable fan nozzle having a common part;
at least one actuator coupled to the at least one integrated mechanism; and
a controller that communicates with the at least one actuator in operation to selectively move the at least one integrated mechanism.

2. The gas turbine engine of claim 1 , wherein the at least one integrated mechanism includes at least two integrated mechanisms.

3. The gas turbine engine of claim 2 , wherein each of the two integrated mechanisms is coupled to a respective semi-circular portion of the fan nacelle.

4. The gas turbine engine of claim 3 , wherein the thrust reverser includes a blocker door having a stowed position and a fully deployed position.

5. The gas turbine engine of claim 4 , wherein the blocker door is pivotably connected to the common part.

6. The gas turbine engine of claim 5 , further comprising a link including a first portion slideably connected to the blocker door and a second portion connected to a support.

7. The gas turbine engine of claim 6 , wherein the support is an inner cowl, and the link is non-slideably connected to the inner cowl.

8. The gas turbine engine of claim 6 , wherein the blocker door includes a slot that receives and retains the first portion of the link.

9. The gas turbine engine of claim 8 , wherein the slot has a t-shaped cross section.

10. The gas turbine engine of claim 1 , wherein the engine has an engine control, and the controller is integrated with the engine control.

11. The gas turbine engine of claim 10 , wherein the controller communicates with the at least one actuator to selectively axially move the common part.

12. The gas turbine engine of claim 1 , wherein the thrust reverser includes a cascade section, the cascade section including a first set of apertures angled in a first direction and a second set of apertures angled in a second, different direction.

13. The gas turbine engine of claim 12 , wherein the first direction is a forward direction, and the second direction is an aft direction.

14. The gas turbine engine of claim 13 , wherein the first set of apertures and the second set of apertures are arranged in circumferential rows about the cascade section such that there are a larger number of circumferential rows in the second set of apertures than in the first set of apertures.

15. The gas turbine engine of claim 14 , wherein the first set of apertures are arranged in two circumferential rows.

16. The gas turbine engine of claim 15 , further comprising airfoil shaped vanes between the apertures.

17. A gas turbine engine comprising:
a fan nacelle that extends circumferentially about a fan, with a bypass passage extending between the fan nacelle and an inner housing;
a bypass ratio of greater than 10;
a low pressure compressor and a high pressure compressor;
a gear train defining a gear reduction ratio of greater than 2.5;
a combustion section in communication with the low pressure compressor and the high pressure compressor;
a low pressure turbine and a high pressure turbine, the low pressure turbine having a pressure ratio of greater than 5, with the low pressure turbine and the high pressure turbine coupled for rotation with respective spools to rotationally drive the low pressure compressor, the high pressure compressor and the gear train, and the fan driven through the gear train;
at least one integrated mechanism coupled to the fan nacelle, the at least one integrated mechanism including a variable area nozzle and a thrust reverser, the thrust reverser and the variable area nozzle having a common part; and
a plurality of actuators coupled to the at least one integrated mechanism, each one of the plurality of actuators moving both the variable area nozzle and the thrust reverser in operation.

18. The gas turbine engine of claim 17 , further comprising a controller, the controller communicating with the plurality of actuators to selectively move the common part between a plurality of axial positions in operation with respect to a centerline axis of the engine.

19. The gas turbine engine of claim 18 , wherein the thrust reverser includes a cascade section, the cascade section including a first set of apertures angled to discharge airflow in a forward direction and a second set of apertures angled to discharge airflow in an aft direction with respect to the centerline axis of the engine.

20. The gas turbine engine of claim 19 , wherein the first set of apertures and the second set of apertures are arranged in circumferential rows about the cascade section such that there are a larger number of circumferential rows in the second set of apertures than in the first set of apertures, and the first set of apertures are arranged in two or more circumferential rows.

21. The gas turbine engine of claim 19 , wherein:
the common part includes a hollow sleeve that extends about the cascade section;
the second set of apertures are axially aft of the first set of apertures with respect to the centerline axis of the engine;
the plurality of axial positions include a stowed position, an intermediation position, and a deployed position;
the first set of apertures are exposed in the intermediate position and the deployed position, but not the stowed position; and
the second set of apertures are exposed in the deployed position, but not in the stowed or intermediate positions.

22. The gas turbine engine of claim 21 , wherein:
the thrust reverser includes a blocker door pivotably connected to the common part; and
the blocker door is pivotable into the bypass passage when in the thrust reverser is in the deployed position, but not in the stowed or intermediate positions.

23. A gas turbine engine comprising:
a fan nacelle that extends circumferentially about a fan, with a bypass passage extending between the fan nacelle and an inner housing;
a bypass ratio greater than 10;
a low pressure compressor and a high pressure compressor;
a gear train defining a gear reduction ratio of greater than 2.5;
a combustion section in communication with the low pressure compressor and the high pressure compressor;
a low pressure turbine and a high pressure turbine, the low pressure turbine having a pressure ratio of greater than 5, with the low and high pressure turbines coupled for rotation with respective spools to rotationally drive the low pressure compressor, the high pressure compressor and the gear train, and the fan driven through the gear train;
at least one integrated mechanism coupled to the fan nacelle, the at least one int

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Marshall Richard M

Classifications

• F02K1/42

  the means being movable into an inoperative position · CPC title

• F02K1/09 Primary

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

• Y02T50/671

  Cross-Sectional Technologies · mapped topic

• F02K1/72

  the aft end of the fan housing being movable to uncover openings in the fan housing for the reversed flow · CPC title

• Y02T50/60

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