Cascade thrust reverser assembly for a gas turbine engine

Abstract

A cascade thrust reverser assembly for a gas turbine engine includes a nacelle assembly defining a bypass passage. The cascade thrust reverser assembly includes a cascade assembly configured to be at least partially enclosed by the nacelle assembly, the cascade assembly including one or more cascade members movable between: a stowed configuration in which the one or more cascade members define a first radial extent; and a deployed configuration in which the one or more cascade members define a second radial extent, wherein the one or more cascade members form a cascade segment in the deployed configuration, and wherein the second radial extent is greater than the first radial extent.

Description

Technical Field

[0001] This disclosure generally relates to a gas turbine engine with a blade stack thrust reverser assembly. Background Technology

[0002] A turbofan engine typically comprises a fan and a turbine that are in fluid communication with each other. The turbine of a turbofan engine typically includes, in series flow sequence, a compression section, a combustion section, a turbine section, and an exhaust section. During operation, air supplied to the turbine flows through the compression section, where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel mixes with the compressed air and burns within the combustion section to provide combustion gases. These combustion gases are then directed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then directed through the exhaust section, for example, into the atmosphere.

[0003] Thrust reverser systems can be used in turbofan engines to reduce the landing distance of an aircraft or the load on its braking system. Thrust reverser systems can be bulky, increasing the size and weight of the engine. For example, some known thrust reverser systems require relatively large nacelle lengths, thereby increasing the weight and drag of the propulsion system and reducing engine efficiency. Furthermore, as fan sizes increase to improve propulsion efficiency, there is an increasing need to reduce nacelle thickness to decrease drag. Both of these effects present challenges to developing thrust reverser systems that are more compact than those traditionally used to achieve overall efficiency benefits for in-flight propulsion systems. Therefore, the inventors of this disclosure have discovered a need for one or more features to reduce the weight and / or overall size of turbofan engines. Attached Figure Description

[0004] Embodiments of this disclosure are illustrated in conjunction with the accompanying drawings, which are included and form part of this specification, and together with the description serve to explain the principles of this disclosure.

[0005] Figure 1 This is a schematic cross-sectional view of an exemplary gas turbine engine according to various embodiments of the present disclosure;

[0006] Figure 2 This is a perspective view of an exemplary turbofan engine according to an exemplary embodiment of the present disclosure, the exemplary turbofan engine having a thrust reverser assembly in a deployed configuration;

[0007] Figure 3 yes Figure 2 An axial side cross-sectional view of an exemplary turbofan engine depicts the thrust reverser assembly in a retracted configuration and an extended configuration in the upper and lower halves of the view, respectively.

[0008] Figure 4This is an axial side cross-sectional view of a turbofan engine according to another exemplary embodiment of the present disclosure, depicting the thrust reverser assembly in a retracted configuration and an extended configuration in the upper and lower halves of the view, respectively.

[0009] Figure 5 This is a perspective view of an exemplary leaf cascade assembly;

[0010] Figure 6 yes Figure 5 A cross-sectional side view of the blade assembly;

[0011] Figure 7 This is a schematic side view of a component of the leaf cascade assembly in a retracted configuration;

[0012] Figure 8 yes Figure 7 The components are shown in a schematic side view of their unfolded configuration;

[0013] Figure 9 This is a schematic side view of another embodiment of a component of a leaf cascade assembly in a retracted configuration;

[0014] Figure 10 yes Figure 9 The components are shown in a schematic side view of their unfolded configuration;

[0015] Figure 11 This is a schematic side view of another embodiment of a component of a leaf cascade assembly in a retracted configuration;

[0016] Figure 12 yes Figure 11 The components are shown in a schematic side view of their unfolded configuration;

[0017] Figure 13 This is a schematic perspective view of an embodiment of a component of a retracted blade assembly;

[0018] Figure 14 yes Figure 13 A schematic perspective view of the components in their unfolded configuration;

[0019] Figure 15 This is a schematic side view of another embodiment of a component of the leaf cascade assembly in a first retracted configuration;

[0020] Figure 16 yes Figure 15 The component is in a schematic side view of the second collapsed configuration;

[0021] Figure 17 yes Figure 15 The components are shown in a schematic side view of their unfolded configuration;

[0022] Figure 18This is a schematic side view of the leaf cascade assembly in its collapsed configuration;

[0023] Figure 19 yes Figure 18 The leaf cascade assembly is shown in a schematic side view of its unfolded configuration;

[0024] Figure 20 This is a schematic side view of another embodiment of the leaf cascade assembly in a retracted configuration;

[0025] Figure 21 yes Figure 20 The leaf cascade assembly is shown in a schematic side view of its unfolded configuration;

[0026] Figure 22 yes Figure 20 The leaf cascade assembly is shown in a schematic side view with its configuration extended and lowered.

[0027] Figure 23 This is a schematic side view of multiple cascade components in an extended configuration;

[0028] Figure 24 yes Figure 23 A schematic side view of multiple leaf cascade components in a nested configuration;

[0029] Figure 25 This is a schematic side view of a blade cascade assembly in an deployed configuration, the assembly having multiple blade cascade members, such as... Figure 23 and 24 As shown;

[0030] Figure 26 This is a schematic side view of another embodiment of multiple cascade components;

[0031] Figure 27 According to the first embodiment, Figure 25 The blade assembly shown is in a schematic side view in a retracted configuration;

[0032] Figure 28 According to the second embodiment, Figure 25 The blade assembly shown is in a schematic side view in a retracted configuration;

[0033] Figure 29 According to the third embodiment, Figure 25 The blade assembly shown is in a schematic side view in a retracted configuration;

[0034] Figure 30 This is a schematic side view of a plurality of cascade members in an extended configuration according to another embodiment;

[0035] Figure 31 This is a schematic side view of a plurality of cascade members in an extended configuration according to yet another embodiment;

[0036] Figure 32 yes Figure 30 A close-up and schematic perspective view of one of the multiple cascade components; and

[0037] Figure 33 This is a flowchart illustrating a method for operating a blade thrust reverser assembly for a gas turbine engine.

[0038] Other aspects and advantages of the embodiments disclosed herein will become apparent from consideration of the following detailed description, wherein similar or identical structures may have similar or identical reference characters. Detailed Implementation

[0039] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to denote features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to denote similar or analogous portions of this disclosure.

[0040] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0041] The terms "front" and "rear" refer to relative positions within a gas turbine engine, with "front" referring to the position closer to the engine inlet and "rear" referring to the position closer to the engine nozzle or exhaust port.

[0042] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction from which the fluid flows.

[0043] Unless the context clearly indicates otherwise, the singular forms “a,” “one,” and “the” include plural references.

[0044] As used throughout this specification and claims, approximate language is applied to modify any quantitative expression that allows for variation without altering its underlying function. Therefore, values ​​modified by one or more terms such as “about,” “approximate,” and “substantially” are not limited to specified exact values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the part and / or system. For example, in some contexts, approximate language may refer to a margin of 10%.

[0045] In this specification and throughout the claims, scope limitations may be combined and interchanged, and unless the context or language otherwise indicates otherwise, such scopes are identified and include all subscopes contained herein. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0046] The term “turbine” or “turbomachinery” refers to a machine that includes one or more compressors, a heating section (e.g., a combustion section), and one or more turbines that together produce torque output.

[0047] The term "gas turbine engine" refers to an engine that has a turbine as its power source, in whole or in part. Examples of gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and hybrid electric versions of one or more of these engines.

[0048] The term "combustion section" refers to any heat addition system used in a turbine. For example, the term combustion section can refer to a section including one or more of a knock combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assembly. In some example embodiments, the combustion section may include an annular burner, a can-shaped burner, a can-annular burner, a vortex burner (TVC), or other suitable combustion systems, or combinations thereof.

[0049] In some embodiments, one or more components of the gas turbine engine described below may be manufactured or formed using any suitable process, such as additive manufacturing processes, or 3D printing processes. Using such processes allows such components to be formed integrally, either as a single monolithic part or as any suitable number of sub-parts. In particular, additive manufacturing processes can allow such components to be formed integrally and include various features that are impossible using existing manufacturing methods. For example, the additive manufacturing methods described herein are capable of producing airfoils, turbine blades and rotors, compressor blades and rotors, and / or fan blades. Such components may have unique features, configurations, thicknesses, materials, densities, fluid passages, manifolds, and mounting structures that may be impossible or impractical using existing manufacturing methods. Some of these features are described herein.

[0050] As used herein, the term "additively manufactured" or "additive manufacturing technology or process" generally refers to a manufacturing process in which consecutive layers of material are provided on top of each other to "build" a three-dimensional part layer by layer. Consecutive layers are often fused together to form a monolithic part that may have multiple integral sub-parts. Although additive manufacturing technology is described herein as capable of manufacturing complex objects by building them point-by-point, layer-by-layer, typically in a vertical direction, other manufacturing methods are possible and within the scope of this subject matter. For example, although the discussion herein involves adding material to form consecutive layers, those skilled in the art will understand that the methods and structures disclosed herein can be practiced with any additive manufacturing technology or manufacturing technique. For example, embodiments of this disclosure may use layer-addition processes, layer-subtraction processes, or hybrid processes.

[0051] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to denote features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to denote similar or analogous portions of this disclosure.

[0052] This application generally relates to a gas turbine engine including an extended cascade thrust reverser assembly (e.g., retracted in a retracted configuration within the nacelle of the gas turbine engine and deployed to an extended configuration). In particular, this disclosure relates to a cascade thrust reverser assembly for a gas turbine engine having a nacelle assembly, the cascade thrust reverser assembly including a cascade assembly configured to be at least partially surrounded by the nacelle assembly. The cascade assembly may include one or more cascade members forming a cascade segment, wherein the one or more cascade members are movable between a retracted configuration and a deployed configuration. In the retracted configuration, the one or more cascade members forming the cascade segment define a first radial range, and in the deployed configuration, one or more cascade members forming the cascade segment define a second radial range. The second radial range is greater than the first radial range.

[0053] The blade segment may be formed from a single blade member that rotates between a retracted configuration and a deployed configuration to increase its radial range. Alternatively, the blade segment may be formed from two or more blade members that move relative to each other, for example, between a retracted configuration and a deployed configuration to increase the radial range of the blade segment. The blade assembly may include multiple blade segments to provide desired airflow redirection in the deployed configuration, thereby generating reverse thrust.

[0054] The authors of this disclosure have discovered that this configuration allows for a more compact nacelle and potentially allows the engine to be mounted on the aircraft wing in a manner that makes the engine lighter overall and reduces drag during operation. In particular, the authors of this disclosure have discovered that providing a cascade thrust reverser assembly with radially extended cascade segments can achieve a more efficient thrust reverser and a more compact nacelle.

[0055] Now refer to the attached diagram, Figure 1 This is a schematic cross-sectional view of a gas turbine engine according to exemplary embodiments of the present disclosure. More specifically, for Figure 1 In one embodiment, the gas turbine engine is a high-bypass turbofan jet engine 10, referred to herein as "turbofan engine 10". Figure 1 As shown, the turbofan engine 10 defines an axial direction A (extending parallel to the longitudinal centerline 12 of the turbofan engine 10 for reference) and a radial direction R. An airflow direction F is also defined, which describes the general direction of airflow through the turbofan engine 10 during normal operation, such as takeoff and cruise. The turbofan engine 10 may also define a circumferential direction C extending circumferentially about the axial direction A (see...). Figure 2 Typically, the turbofan engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14.

[0056] The depicted exemplary turbine 16 is generally enclosed within a substantially tubular housing 18, which defines an annular inlet 20 and an annular outlet 21. The housing 18 surrounds, in a series flow relationship, a compressor section comprising a supercharger or low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; a turbine section comprising a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30; and an exhaust nozzle section 32. A high-pressure (HP) shaft or spool 34 drives the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft or spool 36 drives the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and nozzle section 32 together define a core airflow path 37 through which air flows.

[0057] In the depicted embodiment, fan section 14 includes a fixed-pitch fan 38 having a plurality of fan blades 40. The fan blades 40 may, for example, each be attached to a disk, wherein the fan blades 40 and the disk together can be rotated about a longitudinal centerline 12 via an LP shaft 36. In the depicted embodiment, the turbofan engine 10 is a direct-drive turbofan engine, such that the LP shaft 36 directly drives the fan 38 of fan section 14 without using a reduction gearbox. However, in other exemplary embodiments of this disclosure, the turbofan engine 10 may include a reduction gearbox, in which case the LP shaft 36 may drive the fan 38 of fan section 14 across the gearbox.

[0058] Still referencing Figure 1In an exemplary embodiment, the disc 42 is covered by a rotatable front hub 48, which is aerodynamically shaped to facilitate airflow through a plurality of fan blades 40. Furthermore, the exemplary turbofan engine 10 includes an annular nacelle assembly 50 circumferentially surrounding at least a portion of the fan 38 and / or turbine 16. In the depicted embodiment, the nacelle assembly 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide blades 52. Additionally, a downstream section 54 of the nacelle assembly 50 extends over the outer portion of the housing 18 to define a bypass airflow passage 56 therebetween. The ratio between a first portion of air flowing through the bypass airflow passage 56 and a second portion of air flowing through the inlet 20 of the turbine 16 and through the core airflow path 37 is generally referred to as the bypass ratio.

[0059] Furthermore, as will be discussed in more detail with reference to the following exemplary embodiments, the nacelle assembly 50 includes a thrust reverser assembly 100, which is depicted as a cascade thrust reverser assembly in a retracted configuration.

[0060] However, it should be understood that Figure 1 The exemplary turbofan engine 10 depicted herein is merely an example, and in other exemplary embodiments, the turbofan engine 10 may have any other suitable configuration, including, for example, any other suitable number of shafts or spools, a pitch mechanism for the fan 38 (no longer a fixed-pitch fan), a reduction gearbox between the LP shaft 36 and the fan 38, etc.

[0061] Now for reference Figure 2 and Figure 3 A turbofan engine 10 is provided according to another exemplary embodiment of the present disclosure. Figure 2 and Figure 3 The depicted exemplary turbofan engine 10 includes a thrust reverser assembly 100 according to an exemplary embodiment of the present disclosure. Specifically, Figure 2 A perspective view of an exemplary turbofan engine 10 is provided, wherein the thrust reverser assembly 100 is in an deployed configuration; and Figure 3 A schematic cross-sectional view of an exemplary turbofan engine 10 along the axial direction A is provided, the upper half of which depicts a thrust reverser assembly 100 in a retracted configuration and the lower half of which depicts a thrust reverser assembly 100 in an extended configuration. Figure 2 and Figure 3 The exemplary turbofan engine 10 can be coupled with Figure 1 The exemplary turbofan engine 10 is configured in essentially the same way. Therefore, the same designation can refer to the same or functionally equivalent parts.

[0062] As depicted, the nacelle assembly 50 of the turbofan engine 10 typically includes an inlet assembly 102, a fan shroud 104, and a thrust reverser assembly 100. The inlet assembly 102 is located at the front end of the nacelle assembly 50, and the fan shroud 104 is located at the rear of the inlet assembly 102 and at least partially surrounds the fan 38. When in a retracted configuration, the thrust reverser assembly 100 may again be located at least partially or substantially completely within the fan shroud 104. As depicted, the outer casing 18 of the turbine 16 defines a radially inward boundary of a bypass airflow passage 56, and the nacelle assembly 50 defines a radially outward boundary of the bypass airflow passage 56. During certain operations, bypass air from the engine 10 passes through the bypass airflow passage 56 and exits through the fan outlet nozzle 58.

[0063] Figure 2 and Figure 3 The thrust reverser assembly 100 may include a translation shroud 106 and an blade assembly 108 slidably mounted to the fan shroud 104. Figure 2 As can be clearly seen, the translation cowling 106 is the final section of the nacelle assembly 50, located behind the fan cowling 104 and surrounding the outer casing 18 of the turbine 16. When in the deployed configuration (see...), Figure 2 and Figure 3 The blade assembly 108 is also located at least partially at the rear of the fan shroud 104 and surrounds the outer casing 18 of the turbine 16. Conversely, when in the retracted configuration (see...), Figure 3 The top portion of the blade array 108 is substantially retracted within the fan shroud 104. It is worth noting that since the blade array 108 is substantially retracted within the fan shroud 104 when in the retracted configuration (and slides / translates approximately along the axial direction A to the deployed configuration), the retractable configuration including the blade array 108 allows for a reduced overall size of the nacelle assembly 50.

[0064] The blade assembly 108 can also be provided in an axially fixed configuration. For example, the blade assembly 108 can be fixed relative to the fan shroud 104 in the axial direction A. In this example, the translation shield 106 can be rearwardly translated relative to the fan shroud 104 and the blade assembly 108 in the axial direction A to expose... Figure 2 as well as Figure 3 and Figure 4 The lower half of the image shows the leaf cascade assembly 108 in its unfolded configuration.

[0065] It should be understood that axial movement of the blade assembly 108 can be used in various embodiments, independent of other movements or transformations of the blade assembly. For example, the blade assembly 108 may be lowered in the radial direction R, such as... Figure 3 The lower half is depicted without any initial axial movement or sliding within the turbofan engine 10. Furthermore, it should be understood that, as referenced... Figure 5-32 As described, the expansion of the blade assembly 108 in the radial direction R can also be independently combined with the axial sliding and / or radial descent of the blade assembly 108.

[0066] Still referencing Figure 2 and Figure 3 The depicted blade array assembly 108 is formed by and includes a plurality of individual blade segments 110, which are circumferentially spaced around the nacelle assembly 50. Figure 3 It can be clearly seen that the blade section 110 of the blade assembly 108 can be adapted from... Figure 3 The axially retracted configuration shown in the upper half unfolds to Figure 3 The lower half of the diagram shows the axially deployed configuration. For the depicted embodiment, when the thrust reverser assembly 100 moves from a retracted configuration to a deployed configuration (i.e., deployed), the translation shroud 106 and the blade assembly 108 are adapted to translate uniformly in the rearward direction of the turbofan engine 10, typically along the axial direction A. More specifically, in order to deploy the blade assembly 108 into the bypass airflow passage 56, the translation shroud 106 typically moves rearward from the fan shroud 104 along the axial direction A, and the blade assembly 108 translates and pivots, causing the bypass airflow within the bypass airflow passage 56 to be diverted through the deployed blade assembly 108 to provide a thrust reversing effect. For this purpose, Figure 2 and Figure 3 The blade segment 110 is shown as being pivotally coupled to the nacelle assembly 50 via a corresponding actuator 112 mounted to the nacelle assembly 50. The actuator 112 is configured to move the thrust reverser assembly 100 from a retracted configuration to a deployed configuration. The actuator 112 can be of any suitable type and can be driven by, for example, a pneumatic motor, a hydraulic motor, or an electric motor. Furthermore, the blade segment 110 is depicted as a fixed structure coupled to the nacelle assembly 50 via a guide connection 125. Figure 3 The blade section 110 is represented as being pivotally connected to the outer casing 18 of the turbine 16 via a towing link 114, and the translation shield 106 is represented as being pivotally connected to the blade section 110 via a towing link 116 to translate together with it.

[0067] The translation of the blade cascade assembly 108 and the translation shield 106 in the rearward direction along the axial direction A causes the blade cascade segment 110 to... Figure 3 It unfolds into the bypass airflow channel 56 as shown. From Figure 3It is understood that when retracted, the blade segment 110 can be closed and completely concealed between the inner engine fan housing 118 and the outer engine fan housing 120 of the fan shroud 104, and in the depicted embodiment, the blade segment 110 can be closed and completely concealed between the inner wall 122 and the outer wall 124 of the translation shroud 106. Therefore, when the thrust reverser assembly 100 is retracted, the inner engine fan housing 118 and the inner wall 122 of the translation shroud 106 define a portion of the radially outward flow surface of the bypass airflow passage 56, and completely separate the blade assembly 108 from the bypass airflow passage 56.

[0068] Conversely, when moved to the deployed configuration, the blade segment 110 of the thrust reverser assembly 100 may, but does not need to, extend fully across the radial width of the bypass airflow passage 56, thereby allowing the rear end 126 of the thrust reverser assembly 100 to contact or nearly contact the outer casing 18 of the turbine 16. Figure 3 As shown, as the bypass air in the bypass airflow passage 56 encounters the blade assembly 108, the air is redirected by the opening in the blade section 110 and discharged through the circumferential opening 128 defined between the inner engine fan housing 118 and the outer engine fan housing 120, as well as the inner wall 122 and outer wall 124 of the translation shroud 106. Figure 3 As depicted, each blade segment 110 may be equipped with an extension 127 that facilitates the capture of air flowing along the outer surface of the casing 18 of the turbine 16.

[0069] As can be clearly seen from the above, the depicted embodiments incorporate the function of a blocking gate into the blade assembly 108 to some extent, and achieve this through the rotation of the blade assembly 108. In order to perform this function, each blade segment 110 must be of sufficient length and angled downwards sufficiently when in the deployed configuration to completely or substantially completely block the fan bypass airflow passage 56 in some embodiments.

[0070] However, it should be understood that the exemplary thrust reverser assembly 100 depicted is provided merely as an example, and in other exemplary embodiments, the thrust reverser assembly 100 may have any other suitable configuration. For example, although Figure 2 and Figure 3 The embodiments depict each blade segment 110 as equipped with two different towing links 114 and 116, which are rotatably coupled near the rear end 126 of each blade segment 110 to impart and control rotational motion of the blade segment 110 during deployment. However, in other exemplary embodiments, towing links 114 may be eliminated in order to further reduce aerodynamic drag and other flow disturbances that may lead to aerodynamic or acoustic inefficiencies.

[0071] Furthermore, it should be understood that, apart from the requirement that the blade assembly 108 can redirect airflow within the bypass airflow passage 56, the translational and / or rotational motion of the blade segment 110 does not depend on any particular type of blade design. For example, in other embodiments, the thrust reverser assembly 100 may not include any of the towing links 114, 116 shown, but may instead rely, for example, on the geometry and kinematic deployment system of the blade assembly 108.

[0072] Furthermore, although for Figure 2 and Figure 3 In one embodiment, the plurality of blade segments 110 are configured to translate generally along the axial direction A and further rotate inward along the radial direction R as they move from a retracted configuration to an extended configuration, but in other exemplary embodiments of this disclosure, the blade assembly 108 may not be configured to pivot or rotate.

[0073] For example, now refer to Figure 4 A cross-sectional schematic diagram of a turbofan engine 10 according to another exemplary embodiment is provided. Specifically, Figure 4 An exemplary embodiment depicts a thrust reverser assembly 100 according to another exemplary embodiment in a collapsed configuration in the upper half, and an exemplary thrust reverser assembly 100 in an deployed configuration in the lower half. The turbofan engine 10 and the thrust reverser assembly 100 can be coupled with... Figure 3 The turbofan engine 10 and the thrust reverser assembly 100 are configured in essentially the same way. However, for Figure 4 In some embodiments, the plurality of blade segments 110 may alternatively simply translate generally along the axial direction A between a retracted configuration and an extended configuration, and deploy one or more blocking doors 111 within the bypass airflow passage 56, or in other embodiments, certain other structures (not shown), to deflect airflow from the bypass airflow passage 56 through the blade segments 110, thereby achieving the desired thrust reversal effect. As described above, regarding Figure 4 As will be understood from the embodiments, the blade assembly 108 may also be axially fixed relative to the fan shroud 104. For example, a translation shroud 106 with a large axial and / or radial dimension may be provided to cover the fixed configuration of the blade assembly 108 in a retracted configuration.

[0074] Now for reference Figure 5A perspective view of an exemplary blade cascade assembly 108 removed from a turbofan engine 10 is shown. As depicted, a blade cascade segment 110 may be supported by at least one blade cascade support 113. For example, a plurality of blade cascade supports 113 may be arranged at regular intervals between blade cascade segments 110. The blade cascade supports 113 may be configured to facilitate movement of the blade cascade segment 110 relative to other components. For example, as described in more detail below, the blade cascade supports 113 may facilitate rotation of at least a portion of the blade cascade segment 110. In an embodiment, the blade cascade segment 110 provides at least one anchor point for an actuation assembly, such as... Figure 7-17 As shown.

[0075] like Figure 5 and Figure 2 As shown, multiple blade support members 113 may be radially spaced around the turbofan engine 10. The thrust reverser assembly 100 may further include multiple blade assemblies 108. For example, multiple blade assemblies 108 may be provided for ease of manufacture, repair, or replacement. Figure 5 The individual blade cascade assemblies 108 shown can be separated by blade cascade boundaries 115. In an embodiment, the blade cascade boundary 115 may be provided with one or more mounting features (not shown) for attachment to the turbofan engine 10.

[0076] Figure 6 yes Figure 5 A cross-sectional side view of the blade assembly. (See image.) Figure 6 As shown, a plurality of blade segments 110 may be positioned by at least one blade support 113, for example, between a pair of blade supports 113. For example, the plurality of blade segments 110 may be displaceable and / or rotatable relative to the blade support 113, as referenced below. Figure 7-25 A more detailed description.

[0077] Now for reference Figure 7 This illustration depicts embodiments of one or more blade members 209 of a blade array assembly (such as blade array assembly 108 as described above) of a thrust reverser assembly (such as thrust reverser assembly 100 as described above). Blade array assembly 108 includes one or more blade members 209, and more specifically, in the illustrated embodiments, includes multiple blade members 209. As depicted, the multiple blade members 209 can be configured in a retractable manner, allowing for a reduction in the size of the nacelle assembly 50. It should be understood that such a configuration of blade members 209 can be used in various embodiments, including axially translated, axially fixed, tilted, and / or radially lowered blade array 108.

[0078] exist Figure 7In the illustration, multiple blade cascade members 209 are depicted in a retracted configuration. As depicted, the multiple blade cascade members 209 may be described as stacked. The depicted multiple blade cascade members 209 can be deployed to form a blade segment 210. More specifically, when in a deployed configuration, the multiple blade cascade members 209 form a blade segment 210. It should be understood that, as used herein, "blade cascade member" generally refers to any structure, alone or in combination with other blade cascade members, that forms a blade segment for redirecting airflow. A blade cascade member may be an airfoil, or may cooperate with other blade cascade members to form an airfoil.

[0079] By moving between a collapsed configuration and an expanded configuration, a relatively compact but efficient leaf cascade assembly 108 can be provided.

[0080] For example, when in a retracted configuration, multiple blade members 209 define a first radial range R1, such as Figure 7 As shown. The term "radial range" relative to one or more blade members (e.g., blade member 209) forming a blade segment (e.g., blade segment 210) refers to the total length of blade member 209 in the radial direction R. In some embodiments, the radial range R1 may refer to a measurement along the radial direction R from the leading edge to the trailing edge of the blade segment (e.g., blade segment 210). In other embodiments, the radial range R1 may refer to the maximum measured value along the radial direction R of blade segment 210 at a given axial position along the longitudinal centerline 12. Multiple blade members 209 can then be deployed to form blade segment 210, as... Figure 8 As shown. In Figure 8 In the process, multiple blade members 209 define a second radial range R2, which is larger than a first radial range R1. The multiple blade members 209 of the blade segment 210... Figure 7 Collapse configuration in Figure 8 The movement of the unfolded configuration can therefore be described as increasing from the first radial range R1 to the second radial range R2 or increasing the length of the blade segment 210.

[0081] like Figure 8 As depicted, the plurality of blade members 209 are configured to form a composite aerodynamic shape of the blade segment 210. It should be understood that the arrangement and configuration of the plurality of blade members 209 are adjustable when in a deployed configuration to achieve desired performance. For example, the plurality of blade members 209 may be arranged in a high-lift configuration using independent aerodynamic surfaces (e.g., slats, slots, flaps, and / or flange configurations).

[0082] like Figure 7 and Figure 8 As depicted, the blade actuation assembly 212 is configured for... Figure 7 Collapse configuration and Figure 8 The blades move between the unfolded configurations to form the cascade assembly 108 (see Figure 1-4 A plurality of blade members 209 of a blade segment 210 (at least in the deployed configuration). A blade actuation assembly 212 is operatively connected to the plurality of blade members 209 to control their movement. For example, the blade actuation assembly 212 is configured to control the radial expansion of the plurality of blade members 209 and may be further configured to control the rotation and / or alignment of the plurality of blade members 209. See reference Figure 5 and Figure 6 It should be understood that the blade actuation assembly 212 may be further operably connected to at least one blade support 113.

[0083] Figure 7 and Figure 8 The embodiments include a curved blade cascade actuation assembly 212. As depicted, one or more actuation points 214 can be actuated by pushing, pulling, pressing, and / or inducing strain. With such actuation, the deployment of the blade cascade segment 210 of the blade cascade assembly 108 can be achieved by changing the length and / or orientation of the blade cascade actuation assembly 212. Figure 7 and Figure 8 The described configuration can also be referred to as the bellows configuration of the blade actuation assembly 212.

[0084] Now for reference Figure 9 and Figure 10 The image depicts one or more blade members 309 of a blade segment 310 forming a blade assembly (e.g., blade assembly 108 described above) in an deployed configuration according to another embodiment. Figure 7 and Figure 8 The implementation is the same as the previous one. Figure 9 and Figure 10 The embodiments similarly include a cascade actuation assembly 312. However, compared with... Figure 7 and Figure 8 Compared to the embodiments described above, the blade actuation assembly 312 is configured in a linked manner, operable to move one or more blade members 309 forming the blade segment 310 between a retracted configuration and an extended configuration. The depicted blade actuation assembly 312 can be actuated at one or more actuation points 314 by rotational, linear, and / or flexural motion. For example, the blade actuation assembly 312 can be a hinge mechanism, a sliding crank, an n-bar linkage, a cam mechanism, or any combination thereof. In embodiments, one or more actuation points 314 can be controlled by one or more parallel motion devices. Actuation at one or more actuation points 314 can be aligned with one or more blade members 309, or more precisely, in the illustrated embodiment, with multiple blade members 309 to form the blade segment 310.

[0085] Figure 7-10The actuation motion in the embodiments can be achieved by any suitable prime mover and can be driven by, for example, a pneumatic, hydraulic, or electric motor. For example, as referenced... Figure 3 and Figure 4 One or more actuators 112 described herein may provide prime mover to move the blade assembly 108 between a retracted configuration and an extended configuration. In embodiments, the same components or groups of components may provide for movement of the translation shroud 106 in the axial direction A (see [link to documentation]). Figure 2-4 ), blade assembly 108 (see Figure 2-4 ) or the movement of the blade assembly 309 in the radial direction R (see Figure 3 ), movement of the blade assembly 108 in the axial direction A (see Figure 3 and Figure 4 ), the blade assembly 108, for example, is deployed in the radial direction R (see Figure 7-10 ) and the movement of blocking door 111 or other suitable thrust redirection mechanism (see Figure 3 and Figure 4 The driving force of ). For example, Figure 7-10 The blade actuation assemblies 212 and 312 can be operably connected to including Figure 4 The blocking gate component of blocking gate 111.

[0086] Still referencing Figure 7-10 In the embodiments described, it should be understood that the retracted and expanded configurations are merely guidelines and not a limitation on the total number of configurations for the leaf segments 210, 310. For example, one or more intermediate configurations between the retracted and expanded configurations may also be provided. It is also conceivable that variable control of the multiple leaf elements 209, 309 may be used independently or as grouped leaf segments 210, 310, for example, to rotate the multiple leaf elements 209, 309.

[0087] Now for reference Figure 11 and Figure 12 Another embodiment of the blade cascade segment 1310 is shown. The depicted embodiment includes... Figure 11 Collapse and as Figure 12 The diagram shows multiple blade members 1309 that can be unfolded to form multiple blade segments 1310. Figure 11 and Figure 12 Implementation examples and Figure 7-10 The difference in the embodiments is at least that the blade actuation assembly 1312 is configured such that its actuation point 1314 overlaps with the blade member 1309. Figure 11 and Figure 12 The configuration depicted herein can provide a particularly compact blade section 1310 in the axial direction A.

[0088] like Figure 11 and Figure 12As further depicted, individual blade segments 1310 may include one or more actuation constraints 1319. The depicted actuation constraints 1319 differ from the actuation point 1314, but it should be understood that these features may also be combined. As depicted, one or more actuation tracks 1317 may be further provided to position the one or more actuation constraints 1319. In some embodiments, the actuation constraints 1319 and actuation tracks 1317 may be configured to control the relative movement of the respective blade members 1309 relative to each other. For example, a pair of actuation tracks 1317 may be provided to facilitate rotation during a portion of the actuated movement and maintain the rotated position during another portion of the actuated movement.

[0089] Now for reference Figure 13 and Figure 14 The illustration shows a configuration of a leaf cascade assembly 1408 according to another embodiment. The depicted embodiment includes... Figure 13 Collapse and as Figure 14 The diagram shows a plurality of blade members 1409 that can be unfolded to form a plurality of blade segments 1410. The blade members 1409 are arranged in pairs, such that a blade member 1409 arranged on a first blade actuation assembly 1412 cooperates with a corresponding blade member 1409 arranged on a second blade actuation assembly 1412 to form a blade segment 1410. Figure 13 and Figure 14 As shown, the relative movement between the individual actuation components 1412 can be sufficient to form a plurality of blade segments 1410. However, it should be understood that various other configurations can be used for further actuation as described herein, such as further control of rotation. Each blade actuation component 1412 is provided with at least one actuation point 1414 for, for example, relative to a reference point. Figure 5 and Figure 6 The blade support 113 is controlled to actuate.

[0090] Now for reference Figure 15-17 A further embodiment of the blade cascade assembly 1508 is depicted. The blade cascade assembly 1508 shown includes a plurality of blade cascade members 1509 that cooperate to form a plurality of blade cascade segments 1510. One or more blade cascade actuation components 1512 can be used to control and / or actuate a first blade cascade segment 1510 relative to another blade cascade segment 1510. In the depicted exemplary embodiment, a pair of blade cascade actuation components 1512 are provided with a pair of blade cascade segments 1510. However, it should be understood that the number of blade cascade actuation components 1512 and the number of blade cascade segments 1510 do not need to correspond to each other. For example, one blade cascade actuation component 1512 may be configured to actuate a plurality of blade cascade segments 1510.

[0091] Figure 15The image depicts a blade assembly 1508 in a first retracted configuration, which may also be referred to as a retracted and rotated configuration. As shown, both the first and second blade segments 1510 include a plurality of nested blade members 1509. The first and second blade segments 1510 retract adjacent to each other. In some embodiments, the first and second blade segments 1510 may retract adjacent to each other.

[0092] The movement between the configurations of the depicted blade cascade assemblies 1508 can be controlled by one or more blade cascade actuation assemblies 1512 as described above. Figure 15-17 The depicted pair of blade actuation assemblies 1512 control Figure 15 First collapse configuration, Figure 16 The second collapse configuration and Figure 17 Movement between unfolded configurations. As depicted, the cascade actuation assembly 1512 includes multiple control features: an actuation point 1514, an actuation track 1517, and an actuation constraint 1519. It should be understood that these features can be combined or omitted to achieve various actuations between configurations. The illustrated embodiment provides an actuation point 1514 for attaching the cascade actuation assembly 1512 to a first cascade segment 1510. The actuation point 1514 may be completely fixed to the first cascade segment 1510, or may have some degrees of freedom relative to the first cascade segment 1520, such as rotation. In some embodiments, the actuation point 1514 may be connected from, for example, to, as referenced in the figure Figure 5 and Figure 6 The actuator 112 of the blade support 113 transmits force.

[0093] Turning Figure 16 The second retracted configuration shown represents rotation of the blade assembly 1508. This rotation can be used to increase the radial extension of the blade segment 1510, thereby increasing the radial extension of the blade assembly 1508 (as shown in the reference). Figure 7-10 Rotation can be achieved relative to another component (e.g., a reference). Figure 5 and Figure 6 The actuation of the described blade support 113 is used for control.

[0094] Figure 15-17 The actuation track 1517 of the blade actuation assembly 1512 is further depicted. (Steering) Figure 17 The actuation track 1517 can facilitate movement into the depicted deployment configuration. As shown, the actuation constraint 1519 is configured to control the relative movement of the first leaf segment 1510 into the deployment configuration. However, it should be understood that... Figure 15 , 16 The moves between 1 and 17 can be performed step by step or in combination in any order.

[0095] In some embodiments, Figure 17The unfolded configuration can represent a first unfolded configuration, where further unfolding is possible. For example, multiple cascade members 1509 can be spaced apart, for example, as referenced. Figure 23-26 A more detailed description.

[0096] Turn now Figure 18 An embodiment of the thrust reverser assembly 400 having a blade array 408 is schematically shown as having a plurality of blade array members 409 which are configured in a retracted configuration within the nacelle assembly 450 to form a plurality of blade array segments 410.

[0097] As depicted, cabin assembly 450 includes a first cabin component 404 (e.g., reference 404). Figure 2-4 The aforementioned fan shroud and the second cabin component 406 (e.g., referenced to...) Figure 2-4 (Described translation cowling). Multiple blade cascade members 409 are at least partially contained within the first nacelle component 404 and the second nacelle component 406 in the depicted retracted configuration. It should be understood that, in the retracted configuration, the multiple blade cascade members 409 arising together from the blade cascade assembly 408 may be entirely contained within a single component (e.g., the second nacelle component 406 or the translation cowling). Alternatively, as in the reference... Figure 2-4 In the case of the axial translational blade assembly 108, the plurality of blade members 409 can be completely contained within the first nacelle component 404 or the fan shroud.

[0098] Figure 19 Depicting Figure 18 An embodiment of the blade cascade assembly 408 is shown, wherein a plurality of blade cascade members 409 form corresponding plurality of blade cascade segments 410 in an deployed configuration. In the deployed configuration, a first nacelle member 404 and a second nacelle member 406 are displaced relative to each other to form a blade cascade opening 428. The blade cascade opening 428 exposes the blade cascade assembly 408 in its deployed configuration. As described above, the blade cascade opening 428 can be provided by moving the second nacelle member 406 or a translational cover rearward relative to the first nacelle member 404 in the axial direction A. In this case, the second nacelle member 406 may also be referred to as a blade cascade cover, which is movable to expose the blade cascade assembly 408 in an open position and to at least partially cover the blade cascade assembly 408 in a closed position. Although Figure 18 and Figure 19 The second nacelle component 406 or translation cowl depicted is described as forming a blade cover in the depicted embodiment; however, it should be understood that any number of existing components of the nacelle assembly 50 or other components of the turbofan engine 10 may be configured to form a blade cover.

[0099] With the blade cascade opening 428 closed by the blade cascade cover, airflow through the blade cascade assembly 408 can be suppressed, as described above with reference to the second nacelle component 406. For example, air can flow substantially uninterruptedly through the bypass airflow passage 56 in the airflow direction F (see...). Figure 1 For example, when the blade cascade opening 428 is closed by the blade cascade cover, less than 10% of the mass flow rate of air through the bypass airflow passage 56 can flow through the blade cascade opening 428. Figure 19 In the case of the exposed blade opening 428, as indicated by the unlabeled flow arrows, flow is facilitated through the blade assembly 408. Airflow from within the turbofan engine 10, such as from the bypass airflow passage 56 (see...) Figure 1 The airflow can be redirected from the airflow direction F. In various embodiments, the cascade assembly 408 is configured to redirect the airflow to a direction that is generally opposite to the airflow direction F. Although the airflow through the cascade assembly 408 and the cascade opening 428 may not leave in exactly the opposite direction to the airflow direction F, airflow with any velocity component opposite to the airflow direction F can advantageously provide reverse thrust.

[0100] For comparison, the radial ranges R1 and R2 of the blade cascade assemblies 108, 208, 308, and 408 described herein can be based on their distance from the longitudinal centerline 12 (see, for example, see...). Figure 3 and Figure 4 The common axial position A1 of the nacelle is related to the radial dimensions or ranges of other components. For example, the nacelle thickness N can be defined at a fixed position in the axial direction A relative to, for example, the longitudinal centerline 12 of the turbine 16 or the aircraft (not shown). At this fixed position, the nacelle thickness N can be compared with a first radial range R1 or a second radial range R2 defined at the same axial position. For example, as Figure 18 The depicted nacelle thickness N can be compared in this way with a first radial range R1 in the corresponding axial position A1 of the retracted blade assembly 408. For example... Figure 19 As shown, the nacelle thickness N can be compared with a second radial range R2 at the corresponding axial position A1 of the blade assembly 408 in the deployed configuration. It should be understood that the axial position A1 is provided only for reference anywhere within the range of movement of the measured component. As depicted, reference measurements of the first radial range R1 and the second radial range R2 are performed at axial position A1 in both the retracted and deployed configurations, while the radial thickness N is measured in the retracted configuration.

[0101] In the depicted embodiment, at a given axial position, the nacelle thickness N is greater than a first radial range R1, such that in a retracted configuration, the nacelle assembly 50 can close the blade cascade assembly 408. In this embodiment, for example in a deployed configuration of the blade cascade assembly 408, a second radial range R2 is greater than the nacelle thickness N. In this way, a larger radially more efficient aerodynamic surface or combination of aerodynamic surfaces can be provided as the blade cascade segment 410. Therefore, as described herein, a nacelle thickness N less than the second radial range R2 can promote more efficient reverse thrust for the same compactness of the turbofan engine 10, promote a more compact turbofan engine 10 for the same reverse thrust capability, or promote a combination of more efficient reverse thrust and a more compact turbofan engine 10.

[0102] Figure 18 and Figure 19 Implementation examples and Figure 7-18 The difference in the illustrated embodiment is that each blade segment 410 is formed by only one blade member 409. To achieve the relevant dimensions N, R1, R2 described above, each blade member 409 can be repositioned to a more compact size relative to its unfolded configuration. For example, each blade member 409 can be configured to rotate at least 30 degrees between a collapsed configuration and an unfolded configuration, such as at least 45 degrees, at least 60 degrees, up to 120 degrees, or approximately 90 degrees (90°). Figure 18 and Figure 19 As depicted in the diagram. Similarly, as depicted, multiple leaf cascade members 409 can be configured to nest within each other to further increase the compactness of the collapsed configuration. Although for clarity... Figure 18 and Figure 19 Not shown in the schematic diagram, but can be achieved using the blade cascade actuation assembly 212 (see...). Figure 7-18 To achieve various configurations Figure 18 and Figure 19 The expected rotation described between.

[0103] Now for reference Figure 20 and Figure 21 Another embodiment of the blade assembly 508 is shown schematically. Figure 20 The image depicts a blade array assembly 508 with multiple blade elements 509 in a retracted configuration within the nacelle assembly 550, and... Figure 21 A blade array assembly 508 with multiple blade elements 509 is depicted in a deployed configuration. As described above, the deployed configuration provides a second cabin component 506 spaced apart from the first cabin component 504.

[0104] Figure 20 and Figure 21 The embodiments described herein are similar to Figure 18 and Figure 19 The difference between the embodiments described herein and those described is that the plurality of blade segments 510 of the blade assembly 508 are each formed by a corresponding plurality of blade members 509, such as Figure 21 As shown. As depicted, three of the depicted multiple blade members 509 are configured to be aligned to form corresponding blade segments 510 (similar to...). Figure 7 and Figure 8 (Example). Figure 20 The first radial range R1 depicted can be measured as the total radial dimension of a plurality of blade members 509 that cooperate to form a single blade segment 510 when in an unfolded configuration at a given axial position A1.

[0105] Now for special reference Figure 21 This embodiment is illustrated, wherein each of the plurality of blade members 509 forms a corresponding plurality of blade segments 510 in an unfolded configuration, with blade openings 528 exposing the blade assembly 508. As described above, one or more blade segments 510 at corresponding axial positions A1 in the unfolded configuration define a second radial range R2, which substantially corresponds to a reference... Figure 18 and Figure 19 The defined radial range. It should be understood that... Figure 20-22 The configuration can achieve a radial range R2 that is even greater than the nacelle thickness N.

[0106] Figure 21 The blocking component 511 is also described, including, for example, a blocking gate, as referenced above. Figure 4 The aforementioned blocking gate 111. Although this embodiment is depicted as having a separate blocking component 511, the blocking component 511 can also be combined with another component (e.g., as mentioned above). Figure 3 The blade cascade assembly 508 is combined with the blade cascade assembly 508. The blocking assembly 511 is configured to redirect airflow in the turbofan engine 10 through the blade cascade assembly 508 out of the blade cascade opening 528.

[0107] Figure 22 Depicting Figure 20 and Figure 21 Possible configurations of embodiments of the leaf cascade assembly 508. Figure 22 A blade cascade assembly 508 in a radially lowered configuration is depicted. For example, the blade cascade assembly 508 shown can be lowered in the radial direction R to a turbofan engine (e.g., bypass airflow passage 56, see...). Figure 1Within the airflow. Lowering of the blade cascade assembly 508 can be achieved by translating and / or rotating the blade cascade assembly 508. For example, the blade cascade assembly 508 can be translated relative to the first nacelle component 504 and / or rotated relative to the airflow direction F. As depicted, the blade cascade assembly is disposed on a blade cascade support 513, which is configured to support and / or control individual blade cascade segments in the blade cascade segment 510 and / or individual blade cascade members in the blade cascade member 509. As depicted, the blade cascade support 13 is configured to pivot at or near the rear end of the first nacelle component 504 or the fan shroud. The blade cascade support 513 can also be used for, for example, blade cascade actuation assemblies (see...). Figure 7-17 Other features of the blocking component 511 provide support and / or control.

[0108] like Figure 22 As further shown, the third radial range R3 is defined as a result of the radially reduced configuration of the blade assembly 508. The third radial range R3 may differ from the second radial range. For example, some embodiments of the blade assembly 508 may be configured such that the third radial range R3 is greater than the second radial range R2.

[0109] Figure 20-22 The blade support 513, depicted schematically, also represents the longitudinal centerline 512 of the blade assembly 508. The longitudinal centerline 512 of the blade assembly 508 can represent or approximate a line of symmetry of the blade assembly 508 from its front end 524 to its rear end 526. For example, the longitudinal centerline 512 of the blade assembly 508 can be a line defined such that the blade assembly 508 extends perpendicularly to its longitudinal centerline 512 at approximately equal dimensions on opposite sides of the longitudinal centerline 512 at its front end 524 and rear end 526. In an embodiment, the blade assembly 508 extends perpendicularly to its longitudinal centerline 512 at approximately equal dimensions along its entire length from its front end 524 to its rear end 526 on opposite sides of the longitudinal centerline 512.

[0110] The longitudinal centerline 512 of the blade assembly 508 can be used to define the thickness of the blade assembly 508 at one or more locations. For example... Figure 20 As depicted, the first blade thickness T1 is defined by a plurality of blade members 509 in a retracted configuration within the nacelle assembly 550. As depicted, the first blade thickness T1 is measured as the maximum vertical range of a blade segment 510 or one or more blade members 509 that may form a blade segment 510 relative to the longitudinal centerline 512 of the blade assembly 508, associated with a second axial position A2.

[0111] like Figure 21As shown, the second blade thickness T2 can be measured as the maximum vertical range of the blade segment 510 or one or more blade members 509 forming the blade segment 510 relative to the longitudinal centerline 512 of the blade assembly 508, associated with the second axial position A2. It should be understood that... Figure 20 The thickness T1 of the first leaf cascade in the collapsed configuration and Figure 21 Comparing the second blade thickness T2 in the unfolded configuration can produce a relative change in the blade thickness between the collapsed and unfolded configurations of the blade assembly 508.

[0112] It should be understood that the first cascade thickness T1 and the second cascade thickness T2 are generally related to the first radial range R1 and the second radial range R2. Similar to the first radial range R1 and the second radial range R2, the first cascade thickness T1 and the second cascade thickness T2 can be related to the nacelle thickness N. For example, the first cascade thickness T1 can be less than the nacelle thickness N, and the second cascade thickness T2 can be greater than the nacelle thickness N.

[0113] Refer again Figure 22 The second blade thickness T2 can still be used to describe the thickness of a given blade segment 510 relative to the longitudinal centerline 512 of the blade assembly 508. It should be understood that this measurement of the first blade thickness T1 and the second blade thickness T2 does not simply change with the tilt or angle of the blade assembly 508, because it is relative to the longitudinal centerline 512 of the blade assembly 508, not relative to a reference. Figure 1 The longitudinal centerline 12 of the described turbofan engine 10 is traversed.

[0114] Figure 20-22 The illustrated embodiment depicts a blade cascade segment 510 having multiple aerodynamic surfaces, each blade cascade segment 510 being formed by a separate blade cascade member 509. One or more of the depicted blade cascade segments 510 can be configured with high-lift aerodynamic characteristics to provide enhanced control over airflow through the blade cascade assembly 508. For example, the blade cascade segment 510 can be designed as or include slats, slots, flaps, fencing, and / or vortex generators. It should be understood that providing multiple blade cascade members 509 in a given blade cascade segment 510 provides the opportunity to manage aerodynamic discontinuities to create desired flow properties, such as increased lift, reduced drag, and / or increased flow redirection.

[0115] Turn now Figure 23 The diagram depicts a plurality of cascade members 609 in an extended configuration. The depicted extended configuration can typically be coupled with, for example... Figure 25The unfolded configuration of the blade cascade assembly 608 shown and described below corresponds to this. Multiple three-blade cascade members 609 cooperate to form a blade cascade segment 610 together. It should be understood that any number of blade cascade members 609 can cooperate or be aligned to form a blade cascade segment 610, such as one, two, three, four, five, six, or more blade cascade members 609.

[0116] The extended cascade member 609 may include one or more nested features, such as the depicted first nested feature 660 and a second nested feature 662 complementary to the first nested feature 660. The size and shape of the first nested feature 660 may be configured to be compactly nested, stowed, or folded with the second nested feature 662 of appropriate size and shape. It should be understood that various configurations of the complementary first nested feature 660 and the second nested feature 662 may be provided, and the design of these nested features may also be tailored for aerodynamic adaptability. For example, features may be added to avoid turbulence and / or provide damping.

[0117] Turn now Figure 24 It describes a nested configuration. Figure 23 A schematic diagram of the plurality of blade members 609 shown. The depicted nested configuration can generally correspond to a retracted configuration of the blade members 609 as shown and described with reference to previous embodiments. It should be understood that the complementary first nesting feature 660 and the second nesting feature 662 facilitate the compact retraction of the plurality of blade members 609 within, for example, a nacelle assembly 650 in a retracted configuration.

[0118] As described above, the dimensions and shapes of the complementary first nesting feature 660 and the second nesting feature 662 can be configured to reduce the size of the nested configuration of the plurality of leaf cascade members 609 by compact nesting. Exemplary first nesting feature 660 and second nesting feature 662 include recesses corresponding to protrusions, for example, as... Figure 23 As shown, it has a generally circular configuration of recesses and protrusions. In this configuration, the airfoil of the leading blade member among the plurality of blade members 609 can be extended with a first nesting feature 660, while the effective leading edge can be provided with a second nesting feature 662 on the next trailing blade member among the plurality of blade members 609. It should also be understood that the blade segment 610 may have one or more blade members 609 having one or more of the first nesting features 660 and the second nesting features 662, and may also have one or more blade members 609 without the first nesting features 660 and the second nesting features 662. Figure 24 As shown by the dashed lines, it should also be understood that various embodiments of the first nested feature 660 and the second nested feature 662 may be provided, for example, to add or remove all or part of the slat recess configuration from one or more of the first nested feature 660 and the second nested feature 662.

[0119] Turn now Figure 25 A schematic side view of the blade cascade assembly 608 in an deployed configuration is shown, wherein the blade cascade assembly has features configured for, as Figure 23 and Figure 24 The nested multiple blade members 609 forming the blade segment 610 are shown. Figure 25 The configuration of the blade support 613 is also depicted, wherein the blade support 613 can generally conform to the available space in the nacelle assembly 650. As shown, the second nacelle assembly 606 is deployable from the first nacelle assembly 604 to expose the blade support 613 and the associated blade assembly 608 through the blade opening 628. The second nacelle assembly 606 can be operatively connected to a blocking assembly 611 configured to redirect airflow as indicated by an unlabeled arrow. Movement of the second nacelle assembly 606 can also trigger or otherwise facilitate the movement of multiple blade segments 610 into the depicted deployment configuration.

[0120] Turn now Figure 26 Further embodiments of the cascade segment 1610 may be provided to, for example, Figure 25 It is used in the blade assembly 608 shown. Figure 26 An exemplary embodiment of a blade section 1610 comprising a plurality of actuation tracks 1617 configured to facilitate relative movement between a plurality of blade members 1609 is depicted. The depicted embodiment provides pairs of actuation tracks 1617 that can be arranged parallel to each other to control movement of lower and upper blade members 1609. Various other embodiments may provide fewer or more actuation tracks 1617. The actuation tracks 1617 may be anchored to a blade support, such as referenced in [reference 1617]. Figure 5 and Figure 6 The blade support 113 may be anchored to another, for example, such as Figure 26 The central blade assembly in the blade assembly 1609 shown.

[0121] Figure 26 The actuation tracks 1617 depicted are all configured to control at least one corresponding actuation constraint 1619. As shown, each of the actuation constraints 1619 is included in a corresponding... Figure 26 The first limit position of the expanded configuration is provided, wherein the first nested feature 1660 is spaced apart from the corresponding second nested feature 1662. A second position 1623 is also provided for each of the actuation constraints 1619, the second position 1623 corresponding to the collapsed configuration, wherein the first nested feature 1660 is adjacent to the corresponding second nested feature 1662, and in some embodiments adjacent to the corresponding second nested feature 1662. Figure 26 The embodiments can represent Figure 17A second expansion configuration in addition to the expansion configuration, and it should be understood that these configurations and features can be optionally combined.

[0122] Figure 27 , 28 29 describes the use of cascade segments (e.g., according to...) Figure 23-26 An exemplary embodiment of the leaf cascade segment 610 in the embodiment. However, it should be understood that, Figure 27 , 28 The configuration of 29 can also be applied to any other embodiment of the leaf cascade segments 110, 210, 310, 410, 510, and 610 described herein. Therefore, reference will be made to... Figure 27 , 28 29, using the general configuration of the exemplary leaf segment (e.g., leaf segment 610), while understanding that any configuration or other related features of the leaf segment may also be adopted.

[0123] Figure 27 According to the first embodiment, Figure 25 The illustrated blade cascade assembly 708 is a schematic side view in a retracted configuration. In this first embodiment, a plurality of blade segments 710 are depicted in both a retracted and nested configuration. As depicted, the plurality of blade segments 710 are generally configured in an axial position of the blade cascade assembly 708 in an extended configuration. The relatively compact configuration of the nacelle assembly 750 can be achieved by one or more techniques as described above, such as rotation and nesting of the plurality of blade members 709.

[0124] Figure 28 The second embodiment is described, which is consistent with... Figure 27 The difference in the embodiments is that the plurality of blade members 809 or blade segments 810 of the blade assembly 808 do not rotate, but rather, as in reference... Figure 3 and Figure 4 The retracted configuration involves axial translation. According to this embodiment, multiple blade cascade members 809 can still be nested in groups according to their corresponding blade cascade segments 810. By axially translating the blade cascade assembly 808 as described above, the nacelle assembly 850 can still be advantageously configured by reducing the radial dimension of the nacelle assembly 850 near its trailing edge.

[0125] from Figure 28 It should be understood that the axial dimensions of the blade assembly 808 can also be defined. For example, multiple blade members 809 can define the rear end X and the front end Y in both a retracted configuration and an extended configuration. Figure 28 The collapse configuration described above is illustrated, where the backend X and frontend Y define the first end-to-end distance L1. In contrast, Figure 25 The relevant deployment configuration is described, where the backend X and frontend Y define the second end-to-end distance L2. It should be understood that from... Figure 28Move the collapse configuration to Figure 25 The expanded configuration can result in an increased end-to-end distance L2. Various actuation mechanisms (e.g., reference) can be used when moving from the collapsed configuration to the expanded configuration. Figure 7-17 The described actuation mechanism is used to increase the end-to-end distance, for example, from a first end-to-end distance L1 to a second end-to-end distance L2. In this way, this axially extended configuration of the blade assembly 808 can further reduce the size requirements of the nacelle assembly 850.

[0126] Figure 29 The third embodiment is described, which is related to Figure 27 and Figure 28 The difference in the embodiments is that the plurality of blade members 909 or blade segments 910 of the blade assembly 908 rotate and translate axially relative to their respective positions in the deployed configuration in the retracted configuration. As depicted by the dashed nacelle boundary line 951, Figure 29 This configuration can further increase the possible compactness of the nacelle assembly 950, which is configured to cover the cascade assembly 908 in a retracted configuration.

[0127] Turn now Figure 30 The illustration depicts an embodiment of a plurality of blade members 1009 aligned to form blade segments 1010 in an unfolded configuration of a corresponding blade assembly (not shown). Figure 30 The embodiment differs from the previous embodiment in that one or more flow control features 1064 are disposed on, in, or defined by at least one of the blade members 1009. The flow control features 1064 can further facilitate aerodynamic interactions between adjacent blade members 1009 of the blade segment 1010 as described above. For example, allowing controlled flow through multiple through-holes that define the flow control features 1064 in the leading blade member of the blade member 1009 can promote flow adhesion and reduce the chance of flow separation in use.

[0128] Turn now Figure 31 The illustration depicts an embodiment of a plurality of blade members 1109 aligned to form blade segments 1110 in an unfolded configuration of a corresponding blade assembly (not shown). Figure 31 Implementation examples and Figure 30 The difference in the embodiments is that, as referenced above... Figure 23 and Figure 24As generally described, corresponding first nested feature 1160 and second nested feature 1162 are provided. The corresponding first nested feature 1160 and second nested feature 1162 can be further configured to aerodynamically interact with one or more flow control features 1164. For example, one or more flow control features 1164 can be configured to control the flow around the nested features 1160, 1162 to promote flow adhesion, reduce drag, and / or increase lift.

[0129] Figure 32 yes Figure 30 The circle C32 depicts Figure 30 A close-up isometric view of the blade cascade component 1009. However, it should be understood that, with reference to... Figure 30 The flow control feature 1064 described in the embodiments can be used in other embodiments described herein. As depicted, a plurality of flow control features 1064 are disposed at the leading edge 1066 of the blade cascade member 1009, and another plurality of flow control features 1064 are disposed at the trailing edge 1068 of the blade cascade member 1009. As described above, the flow control feature 1064 can be a passive feature to enhance flow control and provide flow control with high lift separation. Various tilting and / or yaw jets can be employed as flow control features 1064. It should be understood that the flow control feature 1064 can be individually customized for a given location to avoid flow separation, increase aerodynamic loading capacity, and / or increase thrust reverse flow steering. Relatively complex configurations of the flow control feature 1064 can be provided using advanced and / or additive manufacturing techniques as described above.

[0130] Turn now Figure 33 A flowchart depicts the operation of the blade stack thrust reverser assembly 100 for a gas turbine engine 10. In the first process P1 of the depicted operation method, one or more blade stack members 209, 309, 409, 509, 609, 709, 809, 909, 1009, 1109 move from a collapsed configuration to an expanded configuration. This movement can be described with reference to the above. Figure 7-17 The blade cascade actuation assemblies 212 and 312 are implemented as described. During this movement, the blade cascade assemblies 108, 408, 508, 608, 708, 808, 908, 1408, and 1508 can transition from defining a first radial range R1 to defining a second radial range R2. For example, one or more blade cascade members 209, 309, 409, 509, 609, 709, 809, 909, 1009, and 1109 can be configured such that the second radial range R2 is greater than the first radial range R1, as referenced above. Figure 7As described above, it should be understood that one or more blade elements 209, 309, 409, 509, 609, 709, 809, 909, 1009, 1109 in this manner can form blade segments 110, 210, 310, 410, 510, 610, 710, 810, 910, 1010, 1110.

[0131] exist Figure 33 In the second process P2 of the described method, air flows through one or more blade members 209, 309, 409, 509, 609, 709, 809, 909, 1009, and 1109 in an deployed configuration. For example, as referenced above... Figure 2 and Figure 3 The bypass airflow within the bypass airflow channel 56 can be diverted through the deployed blade assemblies 108, 408, 508, 608, 708, 808, 908, 1408, and 1508 to provide a thrust reversal effect.

[0132] This written description uses examples to disclose preferred embodiments, including best practices, and to enable any person skilled in the art to practice this disclosure, including making and using any device or system and performing any combination of methods. The patentable scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0133] The following items provide further details:

[0134] A cascade thrust reverser assembly for a gas turbine engine, the gas turbine engine including a nacelle assembly defining a bypass passage, the cascade thrust reverser assembly comprising: a cascade assembly configured to be at least partially surrounded by the nacelle assembly, the cascade assembly including one or more cascade members movable between: a retracted configuration, wherein the one or more cascade members define a first radial range; and a deployed configuration, wherein the one or more cascade members define a second radial range, wherein the one or more cascade members form cascade segments in the deployed configuration, and wherein the second radial range is greater than the first radial range.

[0135] According to the blade stack thrust reverser assembly described in the preceding clause, wherein when the one or more blade stack members of the blade stack assembly are in the retracted configuration, the nacelle assembly defines a nacelle thickness in the radial direction at an axial position, wherein at the axial position, the nacelle thickness is less than the second radial range of the one or more blade stack members.

[0136] According to any of the foregoing clauses, the blade stack thrust reverser assembly, wherein the nacelle thickness is greater than the first radial range of the one or more blade stack members.

[0137] According to any of the foregoing clauses, the blade cascade thrust reverser assembly defines a front end and a rear end, wherein the distance between the front end and the rear end is increased by moving one or more blade cascade members from the retracted configuration to the deployed configuration.

[0138] According to any of the foregoing clauses, at least two of the one or more blade members are aligned to form the blade segment defining the second radial range.

[0139] The blade stack thrust reverser assembly according to any of the foregoing clauses further includes: a blade stack actuation assembly, wherein the blade stack actuation assembly is operatively connected to the one or more blade stack members and configured to move the one or more blade stack members between the retracted configuration and the deployed configuration.

[0140] The blade cascade thrust reverser assembly according to any of the foregoing clauses further includes: a blocking component configured to redirect airflow through the blade cascade assembly, wherein the blade cascade actuation component is operatively connected to the blocking component.

[0141] According to any of the foregoing clauses, the blade cascade thrust reverser assembly, wherein the blade cascade actuator assembly is further configured to rotate the one or more blade cascade members.

[0142] According to any of the foregoing clauses, in a blade stack thrust reverser assembly, at least one of the one or more blade stack members includes a flow control feature configured to control flow separation through the one or more blade stack members.

[0143] The blade stack thrust reverser assembly according to any of the foregoing clauses further includes: a blade stack cover, the blade stack cover being movable between: an open position, wherein flow is facilitated through the blade stack assembly; and a closed position, wherein flow is inhibited through the blade stack assembly.

[0144] A method of operating a blade stack thrust reverser assembly for a gas turbine engine, the method comprising: moving one or more blade stack members of the blade stack assembly from a retracted configuration defining a first radial range to an extended configuration defining a second radial range, wherein the one or more blade stack members form blade stack segments in the extended configuration, and wherein the second radial range is greater than the first radial range; and allowing air to flow through the one or more blade stack members in the extended configuration.

[0145] The method according to the foregoing clause further includes: moving the blade cover between an open position and a closed position, wherein the open position promotes flow through the blade assembly, and the closed position inhibits flow through the blade assembly.

[0146] According to any of the preceding clauses of the method, wherein the cascade cover defines a nacelle thickness in the radial direction at an axial position of the cascade assembly, wherein the nacelle thickness is less than the second radial range of the cascade segment formed by the one or more cascade members.

[0147] According to any of the foregoing descriptions, the nacelle thickness is greater than the first radial range of the one or more blade array members.

[0148] The method according to any of the foregoing clauses further includes: aligning at least two of the one or more blade members to form the blade segment and define the second radial range.

[0149] The method according to any of the foregoing clauses further includes: actuating a blade actuation component operatively connected to the blade assembly, wherein actuating the blade actuation component moves the one or more blade members of the blade assembly forming the blade segment from the collapsed configuration to the deployed configuration.

[0150] The method according to any of the foregoing clauses further includes: moving the blocking component by actuating the cascade actuation component; and redirecting the airflow through the cascade component using the blocking component.

[0151] The method according to any of the foregoing clauses further includes: rotating the one or more blade members of the blade assembly forming the blade segment by actuating the blade actuation assembly.

[0152] A gas turbine engine including a blade stack thrust reverser assembly, the gas turbine engine comprising: a blade stack assembly including one or more blade stack members, the one or more blade stack members being movable between: a retracted configuration, wherein the one or more blade stack members define a first radial range; and a deployed configuration, wherein the one or more blade stack members define a second radial range, wherein the one or more blade stack members form blade stack segments in the deployed configuration, and wherein the second radial range is greater than the first radial range.

[0153] The gas turbine engine according to the foregoing clause further includes: a turbine; a nacelle at least partially surrounding the turbine; and a bypass passage defined between the turbine and the nacelle, wherein the plurality of blade members are configured to facilitate flow from the bypass passage through the blade members in the deployed configuration.

Claims

1. A blade-cascade thrust reverser assembly for a gas turbine engine, the gas turbine engine including a nacelle assembly, the nacelle assembly including a fan shroud that at least partially defines a bypass passage, characterized in that, The blade stack thrust reverser assembly includes: A blade array assembly configured to at least partially surround the nacelle assembly, the blade array assembly including a plurality of blade elements, the plurality of blade elements including a first blade element and a second blade element, the plurality of blade elements being movable between: Collapse configuration, wherein the plurality of cascade members define a first radial range; and The unfolded configuration, wherein the plurality of blade members define a second radial range, wherein the plurality of blade members form blade segments in the unfolded configuration, and wherein the second radial range is greater than the first radial range, and wherein at least one of the first blade member and the second blade member extends radially over the outer surface of the fan shroud and radially inward into the bypass channel; A leaf actuation assembly, wherein the leaf actuation assembly is operatively connected to the plurality of leaf members, the leaf actuation assembly is connected to a first actuation point of the first leaf member and to a second actuation point of the second leaf member, and the leaf actuation assembly is configured to actuate the first leaf member independently of the second leaf member, such that the plurality of leaf members move between the retracted configuration and the extended configuration.

2. The blade cascade thrust reverser assembly according to claim 1, characterized in that, When the plurality of blade members of the blade assembly are in the retracted configuration, the nacelle assembly defines a nacelle thickness in the radial direction at an axial position, wherein at the axial position, the nacelle thickness is less than the second radial range of the plurality of blade members.

3. The blade cascade thrust reverser assembly according to claim 2, characterized in that, The nacelle thickness is greater than the first radial range of the plurality of blade cascade members.

4. The blade cascade thrust reverser assembly according to claim 1, characterized in that, The leaf cascade assembly defines a front end and a rear end, wherein the distance between the front end and the rear end is increased by moving the plurality of leaf cascade members from the retracted configuration to the unfolded configuration.

5. The blade cascade thrust reverser assembly according to claim 1, characterized in that, The first and second blade members of the plurality of blade members are aligned to form the blade segment defining the second radial range.

6. The blade cascade thrust reverser assembly according to claim 1, characterized in that, Further includes: A blocking component configured to redirect airflow through the cascade assembly, wherein the cascade actuation component is operatively connected to the blocking component.

7. The blade cascade thrust reverser assembly according to claim 1, characterized in that, The blade actuator assembly is further configured to rotate the plurality of blade members.

8. The blade cascade thrust reverser assembly according to claim 1, characterized in that, At least one of the first and second blade members of the plurality of blade members includes a flow control feature configured to control flow separation through the plurality of blade members.

9. The blade cascade thrust reverser assembly according to claim 1, characterized in that, Further includes: Blade cover, the blade cover being movable between: Open position, wherein flow is facilitated through the cascade assembly; and The closed position suppresses flow through the cascade assembly.

10. A method of operating a blade thrust reverser assembly for a gas turbine engine, characterized in that, The method includes: A plurality of blade members, including a first blade member and a second blade member, are moved from a retracted configuration defining a first radial range to an extended configuration defining a second radial range via a blade actuation assembly. In the extended configuration, the plurality of blade members form blade segments, and the second radial range is greater than the first radial range. At least one of the plurality of blade members extends radially beyond the outer surface of the fan shroud and radially inward into a bypass channel at least partially formed by the fan shroud. The blade actuation assembly is operatively connected to the plurality of blade members, connected to a first actuation point of the first blade member and a second actuation point of the second blade member, and configured to actuate the first blade member independently of the second blade member, such that the plurality of blade members move between the retracted configuration and the extended configuration. Air is allowed to flow through the plurality of blade members in the deployed configuration.

11. The method according to claim 10, characterized in that, Further includes: The blade cover is moved between an open position and a closed position, wherein the open position promotes flow through the blade assembly and the closed position inhibits flow through the blade assembly.

12. The method according to claim 11, characterized in that, The blade cascade cover defines a nacelle thickness in the radial direction at an axial position of the blade cascade assembly, wherein the nacelle thickness is less than the second radial range of the blade cascade segment formed by the plurality of blade cascade members.

13. The method according to claim 12, characterized in that, The nacelle thickness is greater than the first radial range of the plurality of blade cascade members.

14. The method according to claim 10, characterized in that, Further includes: Align the first and second blade members of the plurality of blade members to form the blade segment and define the second radial range.

15. The method according to claim 10, characterized in that, Further includes: The actuation of the blade actuation component operably connected to the blade assembly, wherein actuation of the blade actuation component moves the plurality of blade members of the blade assembly forming the blade segment from the collapsed configuration to the deployed configuration.

16. The method according to claim 15, characterized in that, Further includes: The blocking component is moved by actuating the blade actuation component; and The blocking component is used to redirect the airflow through the cascade assembly.

17. The method according to claim 15, characterized in that, Further includes: The plurality of blade members of the blade assembly forming the blade segment are rotated by actuating the blade actuation assembly.

18. A gas turbine engine comprising a blade thrust reverser assembly and a fan shroud, the fan shroud having an outer surface and at least partially defining a bypass passage, characterized in that, The gas turbine engine includes: A blade cascade assembly includes a plurality of blade cascade members, the plurality of blade cascade members including a first blade cascade member and a second blade cascade member operably connected to a blade cascade actuation assembly, the blade cascade actuation assembly being connected to a first actuation point of the first blade cascade member and a second actuation point of the second blade cascade member, and the blade cascade actuation assembly being configured to actuate the first blade cascade member independently of the second blade cascade member, such that the plurality of blade cascade members are movable between: Collapse configuration, wherein the plurality of cascade members define a first radial range; and The unfolded configuration, wherein the plurality of blade members define a second radial range, wherein the plurality of blade members form blade segments in the unfolded configuration, wherein the second radial range is greater than the first radial range, and wherein at least one of the first blade member and the second blade member extends radially above the outer surface of the fan shroud and radially inward into the bypass channel.

19. The gas turbine engine according to claim 18, characterized in that, Further includes: Turbine; The nacelle that at least partially surrounds the turbine; and A bypass passage is defined between the turbine and the nacelle, wherein the plurality of blade cascade members are configured to facilitate flow from the bypass passage through the blade cascade assembly in the deployed configuration.

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