WO1996034193A1

WO1996034193A1 - Thrust reverser mechanism for reducing fan distortion

Info

Publication number: WO1996034193A1
Application number: PCT/US1996/005890
Authority: WO
Inventors: Michael O. Lee
Original assignee: United Technologies Corporation
Priority date: 1995-04-28
Filing date: 1996-04-25
Publication date: 1996-10-31

Abstract

A thrust reverser (30) for a gas turbine engine (10) includes a plurality of cascades (38), a translating sleeve, and a blocker door (36). The translating sleeve and the blocker door (36) have a stowed position and a deployed position. In the stowed position, the translating sleeve encloses the cascades. In the deployed position, the translating sleeve exposes the cascades. The blocker door in the deployed position is disposed substantially downstream from the trailing edge of the cascades. The additional spacing between the fan and the deployed blocker door allows back pressure distortion to attenuate.

Description

Thrust Reverser Mechanism for Reducing Fan Distortion

Technical Field This invention relates to gas turbine engines and, more particularly, to thrust reversers therefor.

Background of the Invention

Conventional gas turbine engines include a fan section and a core engine with the fan section having a larger outer diameter than that of die core engine. The fan section and the core engine are disposed sequentially about a longitudinal axis and are enclosed in a nacelle. An annular path of primary airflow passes through the fan section and the core engine to generate primary thrust. An annular path of duct or fan flow, disposed radially outward of the primary airflow path, passes through me fan section, bypassing me core engine, and exits through a fan nozzle to generate fan thrust.

After touch down, a thrust reverser is activated to slow down the aircraft. The thrust reverser slows down the aircraft by preventing the gas turbine engine from generating forward fan thrust and by generating reverse thrust to counteract the forward primary thrust. The thrust reverser is disposed in the downstream portion of the nacelle and comprises a translating sleeve and a blocker door, each having stowed and deployed positions. In the stowed position, the blocker door is typically disposed in a substantially parallel relationship to the longitudinal axis of the engine. Upon actuation, the blocker door swings into the deployed position to block the fan flow path, thereby preventing the fan flow from generating the forward fan thrust. Upon actuation, the translating sleeve slides axially downstream into the deployed position to expose a plurality of cascades that are disposed around the circumference of the gas turbine engine. Each cascade includes a plurality of turning vanes to guide the fan flow. Since the blocker door blocks the path of the fan flow, substantially the entire fan flow is diverted through the cascades. The turning vanes turn the fan flow to generate reverse thrust that counteracts the forward primary thrust.

The function of the thrust reversers during ground operations in the newer generation of gas turbine engines has expanded. For example, the thrust reverser in a C- 17 transport plane manufactured by McDonnell Douglas Aerospace of St. Louis, Missouri, generates reverse thrust to back up the plane on the ground. A conventional cascade thrust reverser that generates reverse thrust around the circumference of the engine is not suitable for backing up operations for a number of reasons. First, a portion of the reverser flow that is directed towards the ground kicks up ground debris. As the plane backs up, the engine reingests the reverser flow. The debris that enter the engine with the reingested flow may cause severe damage to the engine. Second, during the landing, a portion of the reverser flow that is directed toward the ground also passes under the airplane wing resulting in increased airplane lift. Such interference with the wing aerodynamics leads to a reduction in the weight on the wheels of an airplane and results in reduced brake effectiveness thereof. Thus, conventional cascade thrust reversers are not well suited for the ground maneuvers of the newer gas turbine engines. The thrust reversers of the newer generation of the gas turbine engines include greater directional control of the flow exiting the thrust reverser to avoid adverse effects of the conventional thrust reversers during the ground operations. This directional control is referred to as "targeting." Targeting does not allow the flow to exit uniformly from the cascades but rather blocks and/or reduces the flow exiting in certain areas while allowing an increase in the flow in other areas so that the total system flow remains constant. Specifically, in the C- 17 application, the reverser flow directed toward the ground is blocked. However, the reverser flow non-uniformity or asymmetry results in a static pressure distortion within the engine fan section. High distortion is known to cause dangerous vibratory stresses and reduce efficiency in gas turbines in general and in the fan stage in particular.

If the distortion is not too great, the thrust reverser in general or cascades in particular, can be oversized. Although the oversized thrust reverser can result in an efficiency penalty, it is usually acceptable for lightly targeted applications where the reverser is used to assist breaking only during the initial part of landing. Oversized thrust reversers are suitable for most commercial aircraft. In highly targeted reversers, such as the C-17, the distortion is so high that oversizing the reverser cascades does not provide a practical solution. Presently, this problem is addressed by locating the cascades further downstream (i.e. increasing the spacing between the fan and the cascades) so that the static pressure distortion produced by the reverser is attenuated before it reaches the fan. The disadvantage of this approach is that it requires a much longer translating sleeve for stowing the cascades during forward operation, and leads to a heavier, more costly system. The current C-17 thrust reverser rmnimized the increase in translating sleeve length by incorporating a translating cascade system. During forward operations, the cascade stows within the nacelle. For reverse operations, the cascade moves on rails downstream with the translating sleeve so that a large fan to cascade spacing is maintained. The disadvantage of this approach is that the translating cascade is prone to reliability and maintainability problems as well as being heavy and costly to produce. It should be noted that weight is always a critical concern in aircraft design, since excessive weight in the components limits the useful load and flight range capability (fuel weight) of the aircraft.

Disclosure of the Invention

According to the present invention, a thrust reverser for a gas turbine engine having a longitudinal axis and an annular fan flow passing through the engine in a substantially parallel relationship to the longitudinal axis includes a blocker door having a stowed position and a deployed position and being disposed in the deployed position substantially downstream from a trailing edge of a cascade. The blocker door in the deployed position is disposed in a substantially pe endicular relationship to the longitudinal axis of the engine and obstructs the fan flow path. In the stowed position, the blocker door is disposed in a substantially parallel relationship to the longitudinal axis of the engine without obstructing the fan flow. The thrust reverser also includes a translating sleeve that blocks the fan flow from passing through the cascade in the stowed position and exposes the cascade to allow fan flow therethrough in the deployed position, when the blocker door obstructs the fan flow from generating forward thrust. Placing the deployed blocker door substantially downstream from the trailing edge of the cascade provides a plenum between the trailing edge of the cascades and the deployed blocker door. The plenum allows the fan flow to recirculate before exiting through the cascade, thereby enhancing the back pressure distortion decay rate. The resulting recirculation of the fan flow reduces fan back pressure distortion in gas turbine engines. The present invention is particularly beneficial for high by-pass ratio gas turbine engines employing targeted or asymmetric thrust reverser mechanisms. One major advantage of the present invention is that fan distortion can be minimized without adding substantial weight, expense, or complexity to the gas turbine engine. The blocker door can be relatively easily located downstream from the cascades, without lengthening or relocating the cascades. The foregoing and other advantages of the present invention become more apparent in light of the following detailed description of the exemplary embodiments thereof, as illustrated in the accompanying drawings.

Brief Description of the Drawings FIG. 1 is a simplified, partially broken away representation of a gas turbine engine with a thrust reverser shown in a stowed position; and

FIG. 2 is an enlarged, simplified representation of the thrust reverser of FIG. 1 in a deployed position, according to the present invention.

Best Mode for Carrying Out the Invention Referring to FIG. 1, a gas turbine engine 10 includes a fan section 12 and a core engine 14 disposed sequentially about a longitudinal axis 16. The fan section 12 and the core engine 14 are enclosed in a nacelle 18. An annular path of primary airflow 20 passes through the fan section 12 and the core engine 14 generating primary thrust 22. An annular path of fan flow 24, disposed radially outward of the path of the primary airflow 20, bypasses the core engine 14, flows through the fan section 12, and exits through a fan nozzle 25 to generate fan thrust 26.

A thrust reverser mechanism 30, shown in greater detail in FIG. 2, is disposed in the downstream portion of the nacelle 18 with a trailing edge 32 of the thrust reverser mechanism 30 defining the fan exit nozzle 25. The thrust reverser mechanism 30 comprises a plurality of blocker doors 36, a plurality of cascades 38, a plurality of actuators 40, a plurality of tracks (not shown), and a translating sleeve 42. The translating sleeve 42 includes an aerodynamically shaped body 46 with a recess 48 that accommodates the cascades 38 therein. The translating sleeve 42 has a stowed position and a deployed position. In the stowed position, shown in FIG. 1, the translating sleeve 42 encloses the cascades 38 within the recess 48. In the deployed position, shown in FIG. 2, the translating sleeve 42 moves axially downstream to expose the cascades 38. Each cascade 38 has a trailing edge 49 and comprises a plurality of tinning vanes 50.

Each actuator 40 includes a cylinder 52 and a movable rod 54. Each cylinder 52 is attached to a torque box 56 and a support beam 58. Each rod 54 is secured onto the translating sleeve 42. Hydraulic pressure to the actuators 40 is provided through tubing 60.

The blocker door 36 has a stowed position and a deployed position. In the stowed position, the blocker door 36 is in a substantially parallel relationship to ιe longitudinal axis 16, as shown in FIG. 1 and in phantom in FIG. 2. The blocker door 36 swings about a pivot point (not shown) into the deployed position upon actuation, as shown in FIG. 2. The blocker door 36 in the deployed position is disposed substantially downstream from the trailing edge 49 of the cascades 38 to define a plenum 62.

During takeoff, climb, cruise, and descent, the translating sleeve 42 and the blocker door 36 are in their respective stowed positions. Primary thrust 22 is generated by the primary airflow 20 that exits the core engine 14. Fan thrust 26 is generated by the fan flow 24 exiting through the fan exit nozzle 25. During those modes of operation, the blocker door 36 and the thrust reverser 42 do not interfere with the fan flow 24, as shown in FIG. 1.

After touch down and during reverse maneuvers, the translating sleeve body 46 is moved axially downstream into die deployed position when the hydraulic pressure builds up in the thrust reverser cylinders 52 and extends the movable rod 54 axially downstream, as shown in FIG. 2. The translating sleeve body 46 moves downstream along a set of tracks (not shown), thereby uncovering the cascades 38. The blocker door 36 swings radially inward into the deployed position to obstruct the flow path of the fan flow 24. A portion of the fan flow 24 exits through the cascades 38. Another portion of the fan flow enters the plenum 62. After circulating within the plenum 62, the flow also exits through the cascades 38. The present invention allows a portion of the fan flow to recirculate through the plenum 62 before exiting through the cascade 38. The recirculation significantly reduces the back pressure distortion otherwise produced on the fan 12. The plenum 62 minimizes distortion on the fan by allowing the fan flow to recirculate within the plenum and to attenuate to an acceptable level. Reduction of fan distortion reduces the risk of engine stall and reduces structural stress on the engine. The present invention is suitable even for highly targeted thrust reversers. The present invention is particularly advantageous because placing the blocker door downstream from the trailing edge of the cascades does not add significant weight to the engine and does not significantly impact the complexity of the gas turbine engine.

The position of the blocker door can be optimized for each particular application. For example, in a C- 17 transport plane manufactured by McDonnell Douglas Aerospace of St. Louis, Missouri, the blocker door would be disposed approximately thirty inches (30") downstream from the trailing edge of the cascades.

Although the invention has been shown and described with respect to exemplary embodiments thereof, it should be understood by those skilled in the art that various changes, omissions, and additions may be made thereto, without departing from the spirit and scope of the invention. For example, the blocker door may include a variety of actuating mechanisms. The particular actuating mechanism is not crucial to the present invention, as long as the blocker door is located substantially downstream from the trailing edge of the cascades.

Claims

1. A thrust reverser for a gas turbine engine includes a plurality of cascades, each said cascade having a trailing edge, said gas turbine engine having a longitudinal axis passing therethrough and an annular fan flow flowing substantially parallel to said longitudinal axis, said thrust reverser

5 characterized by: a blocker door having a stowed position and a deployed position, in said deployed position said blocker door substantially obstructing said fan flow, said deployed blocker door being disposed substantially downstream from said trailing edge of said cascade to form a plenum

!0 between said trailing edge of said cascade and said deployed blocker door.

2. A thrust reverser for a gas turbine engine includes a plurality of cascades, each said cascade having a trailing edge, said gas turbine engine having a longitudinal axis passing therethrough and an annular fan flow flowing substantially parallel to said longitudinal axis, said thrust reverser

5 characterized by: a blocker door having a stowed position and a deployed position, in said deployed position said blocker door substantially obstructing said fan flow, said deployed blocker door being disposed substantially downstream from said trailing edge of said cascade to allow attenuation of 10 back pressure distortion therein.

3. A gas turbine engine having a longitudinal axis passing therethrough and fan flow path passing in substantially parallel relationship to said longitudinal axis, said gas turbine engine including a ₁₉₃

thrust reverser, said thrust reverser having a translating sleeve and a plurality of cascades, said translating sleeve having a translating sleeve body with a recess formed therein to accommodate said plurality of cascades therein, said translating sleeve having a stowed position and a deployed position, said translating sleeve enclosing said plurality of cascades within said recess in said stowed position, said translating sleeve exposing said plurality of cascades in said deployed position to said exiting fan flow therethrough, each said cascade having a trailing edge, said gas turbine engine characterized by: a blocker door having a stowed position and a deployed position, in said deployed position said blocker door substantially obstructing said fan flow, said deployed blocker door being disposed substantially downstream from said trailing edge of said cascade to form a plenum between said trailing edge of said cascade and said deployed blocker door.