Patent Publication Number: US-2016230702-A1

Title: Extended thrust reverser cascade

Description:
BACKGROUND 
     The present invention relates generally to gas turbine engines and, more particularly, to a cascade type thrust reverser for a gas turbine engine. 
     Modern aircraft turbofan engines have a nacelle or shroud surrounding the engine, spaced outwardly from a core engine cowl to define an annular passage or duct for flow of air rearwardly from the outer portion of a large fan or axial flow compressor. In this type of engine, a large proportion of the total thrust is developed by the reaction to the air driven rearward by the fan. The balance of the thrust results from ejection of the exhaust gas stream from the core engine. 
     Aircraft using gas turbine engines tend to have high landing speeds, placing great stress on wheel braking systems and requiring very long runways. Thrust reversers have been deployed in gas turbine engines to reduce braking stress and permit the use of shorter runways. 
     One type of thrust reverser is a cascade type thrust reverser. Gas turbine engines equipped with a cascade type thrust reverser utilize sets of cascade turning vanes in the sidewalls of the engine nacelle. A translating sleeve or cowl surrounds the cascade sets and forms a rearward outer wall portion of a bypass duct where bypass air flows between the nacelle and the core engine cowl. Upon deployment of the thrust reverser, the translatable sleeve moves rearwardly and blocking doors hinge radially inwardly to block the bypass duct and redirect bypass air flow through the cascade sets to an outlet. The direction of bypass air flowing through the cascade sets is substantially reversed, thereby slowing the aircraft&#39;s forward velocity. Bypass air is substantially reversed by contacting the turning vanes which comprise the cascade set. Normally each turning vane has the same surface area. Movement of the translating sleeve between a stowed forward position and a deployed rearward position may be provided by one or more actuators that extend between the nacelle and the translatable sleeve. 
     To contact the forward most turning vanes of the cascade set bypass air must make a very sharp turn. It is difficult to enable bypass air to turn sharp enough to contact the forward most turning vanes. As a result, a substantial amount of bypass air does not contact the forward most turning vanes of the cascade sets, and the thrust reverser operates less efficiently than it could. Accordingly, brake stress is increased and longer runways are required. In view of the foregoing problems, there is a need for improved cascade type thrust reversers that will operate more efficiently and help to create a sufficient amount of drag to slow an airplane. 
     SUMMARY 
     An aircraft turbofan engine includes an engine nacelle that circumscribes an airflow duct, and a translating cowl that forms an aft portion of the engine nacelle. A cascade set is positioned within a gap between the translating cowl and the nacelle and has a plurality of vanes. Vanes disposed upstream relative to the flow of air have a greater surface area than vanes disposed downstream relative to the flow of air. The aircraft turbo fan engine also includes blocker doors that cover the cascade set when the translating cowl is in a stowed position, and blocks a portion of the airflow duct when the translating cowl is in a deployed position. Movement of the translating cowl to the deployed position rotates blocker doors, causing air to travel through the cascade set. 
     In another aspect, a thrust reverser system for an aircraft engine is disclosed. The system includes a translating cowl that has a stowed and a deployed position. A cascade set is positioned to be blocked when the translating cowl is in the stowed position and open when the translating cowl is in the open position. The cascade set has a plurality of vanes. Vanes disposed upstream relative to the flow of air have a greater surface area than vanes disposed downstream relative to the flow of air. The thrust reverser system also includes blocker doors that cover the cascade set when the translating cowl is in a stowed position. This causes air to bypass the cascade set. 
     In yet a further aspect, of the current invention a cascade set for creating sufficient drag to slow an aircraft is disclosed. The cascade set includes one or more supporting vanes. A plurality of turning vanes are connected to the supporting vanes, and the turning vanes include forward and aft turning vanes. The forward turning vanes generally have a larger surface area than the aft turning vanes. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1A  is a partial cross-sectional view of a gas turbine engine in cruising mode, e.g. during flight. 
         FIG. 1B  is a partial cross-sectional view of the gas turbine engine in reverse thrust mode, e.g. during landing. 
         FIG. 2A  is a partial cross-sectional view of annular thrust reverser duct of the engine of  FIG. 1A  shown in cruising mode. 
         FIG. 2B  is a partial cross-sectional view of the annular thrust reverser duct of  FIG. 2A  shown in reverse thrust mode. 
     
    
    
     DETAILED DESCRIPTION 
       FIGS. 1A and 1B  are partial cross-sectional views of gas turbine engine  10  which can be mounted to an aircraft.  FIGS. 1A and 1B  show gas turbine engine  10  in cruising mode and thrust reversing mode, respectively. Gas turbine engine  10  includes fan  12 , multistage axial compressor  14 , combustor  16 , high pressure turbine  18 , low pressure turbine  20 , segmented cowl  22 , nacelle body  24 , engine core  26 , inner fixed structure  28 , core exhaust nozzle  30 , bypass duct  32 , blocker door  34 , annular thrust reverser duct  36 , cascade  38 , drag link  40 , translating cowl  42 , and translating sleeve  44 . 
     Engine core  26  and fan  12  are circumscribed by segmented cowl  22 . Segmented cowl  22  includes nacelle body  24  and translating cowl  42 , which is capable of rearward translation along the longitudinal axis of gas turbine engine  10 . Axial movement of translating cowl  42  may be provided, for example by linear actuators (not shown). Disposed internally of segmented cowl  22  is translating sleeve  44  connected for movement with translating cowl  42 . Located closer to the engine centerline is inner fixed structure (IFS)  28 . IFS  28  is an outer surface of engine core  26 . Bypass duct  32  is located between translating sleeve  44  and IFS  28  and through which air is forced by fan  12  for operation of gas turbine engine  10 . 
     During operation, air A is pressurized in compressor  14  and mixed with fuel in combustor  16  for generating hot combustion gases  46  which flow through high and low pressure turbines  18 ,  20 , respectively, that extract energy therefrom. High pressure turbine  18  powers compressor  14  through high pressure shaft (HPS) there between and low pressure turbine  20  powers fan  12  through low pressure shaft (LPS) there between. 
     Gas turbine engine  10  illustrated in  FIGS. 1A and 1B  is a high bypass ratio engine whereby most of the air pressurized by fan  12  is discharged from engine  10  through bypass duct  32 , defined radially between IFS  28  of engine core  26  and nacelle  24  surrounding fan  12 . Core exhaust gases  46  are discharged from engine core  26  through core exhaust nozzle  30 . 
     Drag link  40  is primarily responsible for control in the deployment of blocker door  34  and is disposed within bypass duct  32 . Drag link  40  is secured at one end to blocker door  34  and to IFS  28  at another end. Drag link  40  can be pinned to blocker door  34  or attached in any other suitable manner. Drag link  40  can be configured to slide along IFS  28 . Drag link  40  can be shaped or contoured in such a way that when blocker door  34  moves from the stowed position shown in  FIG. 1A  to the deployed position shown in  FIG. 1B , it adheres to the contour of IFS  28  or has clearance thereto. Drag link  40  can have a number of possible geometric configurations. Drag link  40  can be a smooth curve, bent, have multiple bends, or be straight. Drag link  40  being a smooth curve can be especially desirable as it can help reduce air drag through bypass duct  32  during cruising mode. 
     Annular thrust reverser duct  36  is disposed circumferentially adjacent and radially outward of bypass duct  32 , defined between translating cowl  42  and translating sleeve  44 . In cruising mode, e.g. during flight, as depicted in  FIG. 1A , blocker door  34  lies generally contiguous with the surface of translating sleeve  44  and functions as a continuous extension thereof. Blocker door  34  is configured to mate and cooperate with a plurality of like blocker doors. When blocker door  34  is disposed in a mating engagement with like blocking doors, an annular ring is formed having a radius generally corresponding to the curvature of translating sleeve  44 . In this orientation annular thrust reverser duct  36  is not in fluid communication with air flow A. 
     In reverse thrust mode, e.g. during landing, after touchdown, as depicted in  FIG. 1B , annular thrust reverser duct  36  is in fluid communication with air flow A. To go from stowed to deployed, i.e., reverse thrust mode, translating cowl  42  translates axially rearward. Translating cowl  42  is usually moved using one or more suitable actuators (not shown) that can be a ball-screw actuator, hydraulic actuator, or any other actuator known in the art. As described above, translating sleeve  44  is connected to translating cowl  42  and also moves axially rearward. Blocking door  34  and drag link  40  are also responsive to the translation of translating cowl  42  and are moved into a deployed position. The movement of blocking door  34  is facilitated by the translation of drag link  40 . Drag link  40  translates along IFS  28  as blocking door  34  pivots into bypass duct  32 . Accordingly, bypass duct  32  is substantially blocked by a ring of blocker doors  34  interposed within bypass duct  32 . The rearward translation of translating cowl  42  and translating sleeve  44  puts annular thrust reverser duct  36  in fluid communication with bypass air A. Therefore, bypass air A is effectively diverted to annular thrust reverser duct  36 . 
       FIGS. 2A and 2B  are partial cross-sectional views of annular thrust reverser duct  36  of  FIGS. 1A and 1B  shown in stowed mode and deployed, i.e., thrust reversing mode, respectively. Annular thrust reverser duct  36  includes cascade  38  having forward end  48 , aft end  50 , turning vanes  52 , and support vane  54 . Annular thrust reverser duct  36  further includes bullnose  56 , aft cascade support ring  58 . Blocker door  34  is also shown and includes forward edge  60  and back edge  62 , drag link  40 , translating cowl  42  having forward edge  64 , and translating sleeve  46  having forward edge  66 . Nacelle body  24 , IFS  28 , and bypass duct  32  are also shown. 
     In stowed mode, e.g., when cruising, bypass air A does not enter annular thrust reverser duct  36 . As shown by  FIG. 2A , forward edge  64  of translating cowl  42  is in contact with nacelle body  24 , which would substantially block bypass air A from leaving annular thrust reverser duct  36 . Forward edge  60  of blocking door  34  is substantially in contact with bullnose  56 , substantially preventing bypass air A from entering annular thrust reverser duct  36 . Because annular thrust reverser duct  36  is substantially blocked from bypass air A flow in cruising mode, bypass air A flows through bypass duct  32  and exits gas turbine engine  10  creating forward thrust. 
     In reverse thrust mode as shown in  FIG. 2B , forward edge  64  of translating cowl  42  and forward edge  66  of translating sleeve  44  are translated axially rearward by an actuator (not shown) towards aft end  50  of cascade  38 . As translating cowl  42  and translating sleeve  44  translate axially rearward, blocking door  34  pivots into bypass duct  32 , the pivoting motion facilitated in part by drag link  40 . After translation, forward edge  60  of blocking door  34  is disposed near aft cascade support ring  58 . Back edge  62  of blocking door  34  extends radially towards and contacts IFS  28  effectively blocking bypass air A from flowing through bypass duct  32 . As a result of the movement of translating cowl  42 , translating sleeve  46 , blocking door  34 , and drag link  40 , annular thrust reverser duct  36  is in fluid communication with bypass air A. 
     Cascade  38  is shown disposed within annular thrust reverser duct  36 . Cascade  38  is disposed extending axially between bullnose  56  and aft cascade support ring  58 . Bullnose  56  is fixed to nacelle body  24  and can be attached to the forward most turning vane  52  of cascade  38 . Bullnose  56  can be aerodynamically configured to turn bypass air A toward turning vanes  52  disposed near forward end  48  of cascade  38 . The configuration of bullnose  56  can also help direct bypass air A toward turning vanes  52  disposed near aft end  50  of cascade  38 . 
     Aft cascade support ring  58  is fixed to nacelle body  24  and is attached to the aft portion of cascade  38 . When in cruising mode as depicted by  FIG. 2A , cascade  38  is disposed circumferentially adjacent and radially outward from translating sleeve  44 , which is attached to blocker door  34  and connected for movement with translating cowl  42 . Cascade  38  is also disposed circumferentially adjacent and radially inward from translating cowl  42 . In reverse thrust mode as depicted in  FIG. 2B , cascade  38  is not circumvented by either translating cowl  42  nor translating sleeve  44 , and bypass air A can flow through cascade  38 . 
     Cascade  38  can be made from a carbon composite or any other suitable material. Cascade  38  includes a plurality of vanes arranged as a matrix of turning vanes  52  and support vanes  54 . Turning vanes  52  are disposed substantially perpendicular to the centerline of gas turbine engine  10  and support vanes  54  are disposed substantially parallel to the centerline of gas turbine engine  10 . Turning vanes  52  can be curved with a forward aspect to divert air in a direction substantially reversed from its rearward flow through bypass duct  32 . 
     Turning vanes  52  disposed toward forward end  48  of cascade  38  generally have a larger surface area than turning vanes  52  disposed toward aft end  50  of cascade  38 . The difference in surface area can be the result of turning vanes  52 , disposed toward forward end  48  extending radially longer than turning vanes  52  disposed toward aft end  50 . The length of turning vanes  52  is limited by the distance between supporting vane  54  and translating sleeve  44 . Although sixteen turning vanes  52  are depicted, more or fewer turning vanes can be employed in further embodiments without departing from the scope of the invention. Cascade  38  can be one of many cascade  38  matrices disposed within annular thrust reverser duct  36  circumferentially around gas turbine engine  10 . 
     Turning vanes  52  disposed at forward end  48  of cascade  38  generally have a larger surface area than turning vanes  52  disposed at aft end  50  of cascade  38 . The larger surface area of turning vanes  52  disposed at forward end  48  of cascade  38  can result in those turning vanes  52  being disposed closer to bypass air A than they would be if they had the same surface area as those turning vanes  52  disposed at aft end  50  of cascade  38 . Accordingly, the generally larger surface area helps forward turning vanes  52  engage more bypass air A directed towards cascade  38  by bullnose  56 . As turning vanes  52  engage bypass air A the direction of bypass air A is substantially reversed from it rearward path. Thus, drag sufficient to help slow an aircraft&#39;s forward velocity is created. 
     In view of the entirety of the present disclosure, including the accompanying figures, persons of ordinary skill in the art will recognize that the present invention can provide numerous advantages and benefits. For example, the ability of cascade  38  to engage more bypass air A can make cascade  38  more efficient than traditional cascades where every turning vane has a generally equivalent surface area. Because cascade  38  can engage more bypass air A, the ability of cascade  38  to create drag can be increased. This can help reduce braking stress and allow the use of shorter runways because the airplane will be able to stop quicker, while relying less on its brakes. Because cascade  38  can help an airplane stop quicker overall flight safety can be increased. Also, disposing turning vanes  52  closer to bypass air A allows the design of cascade  38  to have a shorter axial length than traditional cascade sets because turning vanes  52  can engage more air. A further benefit of cascade  38  is that it can be retrofit into annular thrust reverser duct  36  of any gas turbine engine or be built into any new gas turbine engine. 
     Any relative terms of degree used herein, such as “substantially”, approximately”, “essentially”, “generally” and the like, should be interpreted in accordance with and subject to any applicable definitions or limits expressly stated herein. In all instances, and relative terms or terms of degree used herein should be interpreted to broadly encompass any relevant disclosed embodiments as well as such ranges or variations as would be understood by a person of ordinary skill in the art in view of the entirety of the present disclosure, such as to encompass ordinary manufacturing tolerance variations, incidental alignment variations, and the like. 
     While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims. 
     Discussion of Possible Embodiments 
     The following are non-exclusive descriptions of possible embodiments of the present invention. 
     An aircraft turbofan engine can include an engine nacelle that circumscribes an airflow duct, and a translating cowl that forms an aft portion of the engine nacelle. A cascade can be positioned within a gap between the translating cowl and the nacelle and has a plurality of vanes. Vanes disposed upstream relative to the flow of air can have a greater surface area than vanes disposed downstream relative to the flow of air. The aircraft turbofan engine can also include a blocker door that covers the cascade set when the translating cowl is in a stowed position and that blocks a portion of the airflow when the translating cowl is in a deployed position, such that flow of air travels through the cascade set. 
     The aircraft turbofan engine of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components. The vanes disposed upstream relative to the flow of air can be generally disposed closer to the air flow path than the vanes disposed downstream relative to the air flow path. Vanes of the plurality of the vanes can have different surface areas. The gas turbine engine can include a drag link that is connected to the blocker door, the drag link moving with the translating cowl as the translating cowl moves to the deployed position. The cascade set can be made of a composite carbon material. The vane disposed furthest upstream relative to the flow of air can be supported by a bullnose structure integrated to the engine nacelle. The cascade set can be a static structure disposed within the gap between the translating cowl and an engine nacelle. The plurality of vanes can engage the air flow and substantially reverse the generally rearward path of the air flow when the translating cowl is in a deployed position. 
     In another aspect, a thrust reverser system for an aircraft engine is disclosed. The system can include a translating cowl that can have a stored and a deployed position. A cascade set can be positioned to be blocked when the translating cowl is in the stowed position and open when the translating cowl is in the open position. The cascade set can have a plurality of vanes. Vanes disposed upstream relative to the flow of air can have a greater surface area than vanes disposed downstream relative to the flow of air. The thrust reverser system can also include a blocker door that covers the cascade set when the translating cowl is in a stowed position. The blocker door blocks a portion of the airflow duct when the translating cowl is in the deployed position such that the flow of air travels through the cascade set. 
     The thrust reverser system for an aircraft engine of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components. The vanes disposed upstream relative to the flow of air can be generally disposed closer to the air flow path than the vanes disposed downstream relative to the air flow path. Vanes of the plurality of the vanes can have different surface areas. The thrust reverser system can include a drag link that is connected to the blocker door, the drag link moving with the translating cowl when the translating cowl moves to the deployed position. The cascade set can be made of a composite carbon material. The vane disposed furthest upstream relative to the flow of air can be supported by a bullnose structure integrated to an engine cover. The cascade set can be a static structure disposed within the gap between the translating cowl and the engine. The plurality of vanes can engage the air flow and substantially reverse the generally rearward path of the air flow when the translating cowl is in a deployed position. 
     In yet another embodiment, a cascade set for creating sufficient drag to slow an aircraft can include the following features. The cascade set can include one or more supporting vanes. A plurality of turning vanes can be connected to the supporting vanes, and the turning vanes include forward and aft turning vanes. The forward turning vanes can generally have a larger surface area than the aft turning vanes. 
     The cascade set of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components. The turning vanes can have progressively smaller surface areas as they approach an aft end of the cascade set. The cascade set can be made of a composite carbon material. Finally, the forward most turning vane can be supported by a bullnose structure integrated to an engine nacelle.