Guide vane assembly with fixed and variable pitch inlet guide vanes

Abstract

A pilot vane assembly for a nacelle of a gas turbine engine includes a front vane and a rear vane. The rear wheel blade is located behind the front wheel blade and in front of the plurality of fan blades. The front blade defines a fixed pitch angle and the rear blade is movable between a first pitch angle and a second pitch angle.

Description

Technical Field

[0001] The present disclosure relates to a guide vane assembly, and more particularly, to a guide vane assembly for a gas turbine engine, the guide vane assembly being configured to guide airflow at an inlet of a nacelle. Background Art

[0002] A turbofan engine typically includes a fan and a turbine having a plurality of fan blades arranged to flow in communication with each other. In addition, the turbine of a turbofan engine typically includes a compressor section, a combustion section, a turbine section, and an exhaust section in a series order. In operation, air is provided from a fan to the inlet of the compressor section, wherein one or more axial compressors gradually compress the air until the compressed air reaches the combustion section. Fuel is mixed with the compressed air and burned in the combustion section to provide combustion gases. The combustion gases are directed from the combustion section to the turbine section. The combustion gases flow through the turbine section to drive the turbine section and are then directed through the exhaust section to the atmosphere. Fan efficiency losses may result in reduced efficiency of the turbofan engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] A complete and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0004] Figure 1 is a cross-sectional view of an exemplary gas turbine engine.

[0005] Figure 2 yes Figure 1 An enlarged view of the front end of an exemplary gas turbine engine showing the guide vane assembly.

[0006] Figure 3 yes Figure 1 An axial view of an inlet of an exemplary gas turbine engine having a plurality of evenly spaced guide vane assemblies.

[0007] Figure 4 is an axial view of an inlet of a gas turbine engine having a plurality of non-uniformly spaced guide vane assemblies according to another exemplary embodiment.

[0008] Figure 5 It is along line 5-5 Figure 1 A cross-sectional view of an exemplary gas turbine engine guide vane assembly with aft vanes disposed at a first angle.

[0009] Figure 6 yes Figure 1 A cross-sectional view of an exemplary gas turbine engine guide vane blade assembly with the aft vanes arranged at a second angle.

[0010] Figure 7 yes Figure 1 An enlarged view of an exemplary gas turbine engine showing a variable pitch mechanism. DETAILED DESCRIPTION

[0011] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. The same or similar names in the drawings and description have been used to refer to the same or similar parts of the present disclosure.

[0012] As used herein, the word "exemplary" means "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0013] For the purposes of this description, the terms "upper", "lower", "right", "left", "vertical", "horizontal", "top", "bottom", "lateral", "longitudinal" and their derivatives shall relate to the orientations disclosed in the accompanying drawings.

[0014] As used herein, the terms “first,” “second,” and “third,” and other ordinal numbers are used to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0015] The terms "fore" and "aft" refer to relative positions within a gas turbine engine, with "fore" referring to a position closer to the engine inlet and "aft" referring to a position closer to the engine nozzle or exhaust.

[0016] The terms "upstream" and "downstream" refer to relative directions relative to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, while "downstream" refers to the direction to which the fluid is flowing.

[0017] The term "attached" refers to two components being directly connected to one another. The term "integrated" refers to two components being formed simultaneously as a single piece, or two components being formed separately and then affixed to one another. The term "unitary structure" refers to an integrally formed, one-piece structure whose components are formed simultaneously.

[0018] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0019] For example, the term “at least one” in the context of “at least one of A, B, and C” means only A, only B, only C, or any combination of A, B, and C.

[0020] The phrases "from X to Y" and "between X and Y" each refer to a range of values ​​including the endpoints (ie, to a range of values ​​that includes X and Y).

[0021] The present disclosure generally relates to an inlet pre-swirl feature configured as a plurality of guide vane assemblies for an inlet of a gas turbine engine and a control system in communication with components of the gas turbine engine.

[0022] The guide vane assemblies each include a front vane and a rear vane. The front vane is positioned at a fixed angle relative to the longitudinal axis of the gas turbine engine. The rear fan can move between a range of angles relative to the longitudinal axis of the gas turbine engine. In other words, the angle of the rear vane can be changed during operation of the gas turbine engine.

[0023] The angle of the front blades is fixed to provide rigidity for, for example, deflecting incoming debris. The angle of the rear blades is variable so as to match the swirl applied to the incoming air with the airspeed of the aircraft and the rotational speed of the fan so that the angular velocity of the air as it approaches the fan blades corresponds to the angular velocity of the fan blades in a desired manner. The rear blades are configured to pre-swirl an airflow provided through an inlet of an outer nacelle that is located upstream of a plurality of fan blades of the fan. As described herein, pre-swirl of the airflow provided through the inlet of the outer nacelle before it reaches the plurality of fan blades of the fan can reduce separation losses and / or shock losses, thereby allowing the fan to operate at a relatively high fan tip speed while reducing efficiency losses. The guide vane assembly has fixed and variable portions that provide rigid protection against incoming debris and pre-swirl of the incoming air, thereby improving the operation of the gas turbine engine.

[0024] Referring now to the drawings, wherein like numerals refer to like elements throughout the several views, Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In the embodiment of the present invention, the gas turbine engine is an aviation turbofan jet engine 10, referred to herein as a "turbofan engine 10". The turbofan engine 10 is configured to be mounted on an aircraft, such as in an underwing configuration or a tail mounted configuration. Figure 1 As shown, the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A). The longitudinal centerline defines a longitudinal axis 12 of the turbofan engine 10. Generally speaking, the turbofan engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14 (the turbine 16 is sometimes also or alternatively referred to as a "core turbine engine").

[0025] The exemplary turbine 16 shown generally includes a generally tubular casing 18 that defines an annular inlet 20. The casing 18 surrounds a compressor section, including a first supercharger or low pressure (LP) compressor 22 and a second high pressure (HP) compressor 24, in series flow relationship; a combustion section 26; a turbine section, including a first high pressure (HP) turbine 28 and a second low pressure (LP) turbine 30; and an ejection exhaust nozzle section 32. A high pressure (HP) shaft drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and ejection exhaust nozzle section 32 are arranged in a series flow sequence and together define a core air flow path 37 through the turbine 16. It is also contemplated that the present disclosure is compatible with engines having intermediate pressure turbines (e.g., engines having three spools).

[0026] Still refer to Figure 1 In the embodiment of the present invention, the fan section 14 includes a variable pitch, single-stage fan 38, and the turbine 16 can be operably coupled to the fan 38 to drive the fan 38. The fan 38 includes a plurality of rotatable fan blades 40 coupled to a disk 42 in a spaced manner. As shown, the fan blades 40 extend outwardly from the disk 42 generally in a radial direction R. With the fan blades 40 being operably coupled to an appropriate actuating member 44, the actuating member is configured to jointly change the pitch of the fan blades 40, and each fan blade 40 can rotate relative to the disk 42 about the pitch axis P, for example, in unison. The fan blades 40, the disk 42, and the actuating member 44 can rotate together around the longitudinal centerline 12 through the LP shaft 36 across the power gearbox 46. The power gearbox 46 includes a plurality of gears for reducing the rotational speed of the LP shaft 36 to a more efficient fan speed. Therefore, for the embodiment shown, the turbine 16 is operably coupled to the fan 38 through the power gearbox 46.

[0027] In an exemplary embodiment, fan section 14 includes twenty-two (22) or fewer fan blades 40. In other exemplary embodiments, fan section 14 includes a different number of fan blades 40, such as twenty (20), eighteen (18), sixteen (16), or another number of fan blades 40.

[0028] Still refer to Figure 1 In the exemplary embodiment, disk 42 is covered by a rotatable forward nacelle or hub 48 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. In addition, exemplary fan section 14 includes an annular fan case or outer nacelle 50 that at least partially (circumferentially for the illustrated embodiment) surrounds fan 38 and at least a portion of turbine 16.

[0029] More specifically, the outer nacelle 50 includes an inner wall 52, and a downstream section 54 of the inner wall 52 of the outer nacelle 50 extends above an outer portion of the turbine 16, thereby defining a bypass airflow passage 56 therebetween. Additionally, for the illustrated embodiment, the outer nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 55. The outer nacelle 50 includes an inlet 60 located at a leading edge 61 of the outer nacelle 50.

[0030] During operation of turbofan engine 10, a volume of air 58 enters turbofan engine 10 through inlet 60 of outer nacelle 50 and / or fan section 14. As volume of air 58 passes through fan blades 40, a first portion of air 58 is directed or directed into bypass airflow passage 56 as indicated by arrow 62, while a second portion of air 58 is directed or directed into core air flow path 37 as indicated by arrow 64. The pressure of the second portion of air indicated by arrow 64 then increases as it is directed through HP compressor 24 into combustion section 26, where the second portion of air is mixed with fuel and combusted to provide combustion gases 66. Combustion gases 66 are directed from combustion section 26 through HP turbine 28. In HP turbine 28, a portion of the thermal and / or kinetic energy in combustion gases 66 is extracted via successive stages of HP turbine stator blades 68 coupled to casing 18 and HP turbine rotor blades 70 coupled to high pressure (HP) shaft 34, thereby rotating HP shaft 34 to support operation of HP compressor 24. The combustion gases 66 are then directed through the LP turbine 30 wherein a second portion of thermal and / or kinetic energy is extracted from the combustion gases 66 via successive stages of LP turbine stator blades 72 coupled to the casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36 , thereby rotating the LP shaft 36 to support operation of the LP compressor 22 and / or rotation of the fan 38 .

[0031] The combustion gases 66 are then directed through the jet exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of the air 62 is significantly increased after being directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the turbofan engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbine 16.

[0032] Still see Figure 1 , the turbofan engine 10 of the present disclosure also provides pre-swirl flow ahead of the tips of the fan blades 40 as described herein. For example, the turbofan engine 10 also includes one or more inlet guide vane assemblies 100, as described in more detail below.

[0033] In some exemplary embodiments, the exemplary turbofan engine 10 of the present disclosure may be a relatively high-power class turbofan engine 10. Therefore, when operating at rated speed, the turbofan engine 10 may be configured to generate relatively large thrust. More specifically, when operating at rated speed, the turbofan engine 10 may be configured to generate at least 20,000 pounds of thrust, such as at least about 25,000, 30,000, and up to, for example, 150,000 pounds of thrust. Therefore, the turbofan engine 10 may be referred to as a relatively high-power class gas turbine engine.

[0034] also, Figure 1 The exemplary turbofan engine 10 shown in FIG. 1 is for example only, and in other exemplary embodiments, the turbofan engine 10 may have any other suitable configuration. For example, in certain exemplary embodiments, the fan may not be a variable pitch fan, the engine may not include a reduction gearbox (e.g., power gearbox 46) to drive the fan, and may include any other suitable number or arrangement of shafts, spools, compressors, turbines, etc.

[0035] See now Figure 2 , which provides an enlarged view of the front end of the fan section 14 and the turbine 16. The guide vane assembly 100 is located in front of the plurality of fan blades 40 in the axial direction A, i.e., upstream of the fan blades 40. The guide vane assembly 100 may be attached to or integrated into the outer nacelle 50, for example, formed separately and then attached to each other or formed simultaneously as an integral structure.

[0036] The guide vane assembly 100 includes a front vane 102 and a rear vane 104. The front vane 102 provides rigidity and debris protection for the guide vane assembly 100. Specifically, the front vane 102 is configured to deflect objects and other debris entering the nacelle 50, and the front vane 102 may be formed of a material having a specified stiffness or modulus so as to deflect the object while suppressing deformation. The material of the front vane 102 may have a higher stiffness or modulus than the material of the rear vane 104. For example, the front vane 102 may be a metal, such as steel or titanium, and the rear vane 104 may be a composite material, such as a carbon fiber polymer. As another example, both the front vane 102 and the rear vane 104 may be metal, or both may be composite materials, wherein the stiffness or modulus of the front vane 102 is greater than the stiffness or modulus of the rear vane 104. When the guide vane assembly 100 is positioned in the nacelle 50, the front vane 102 may be attached to the nacelle 50 or integrated with the nacelle 50.

[0037] The rear blades 104 provide a swirl control feature to control the swirl of the air 58 flowing through the guide vane assembly 100 toward the fan blades 40. Specifically, as will be explained in more detail below, the rear blades 104 are movable relative to the front blades 102 to control the air flowing toward the fan blades 40. The rear blades 104 are located behind the front blades 102 and in front of the fan blades 40. When the guide vane assembly 100 is positioned in the nacelle 50, at least a portion of the rear blades 104 are attached to or integrated with the nacelle 50. That is, while the rear blades 104 are generally movable relative to the front blades 102, a portion of the rear blades 104 may be attached to or integrated with the nacelle 50 to secure the rear blades 104 in place. The fixed portion may be, for example, a rod about which the movable portion of the rear blades 104 rotates (e.g., a rod). Figure 5-6 ), and the fixed portion may extend into the nacelle 50 to a suitable attachment point or may be integrated into the nacelle 50 .

[0038] The front blade 102 generally extends from an outer end 106 to an inner end 108 in a radial direction R, and the rear blade 104 generally extends from an outer end 110 to an inner end 112 in a radial direction R. The "span" of the front blade 102 or the rear blade 104 is the length from the outer end 106, 110 to the inner end 108, 112 in the radial direction R. The span of the front blade 102 is the "front blade span" 114, and the span of the rear blade is the "rear blade span" 116. The inner ends 108, 112 of the front blade 102 and the rear blade 104 extend freely and do not include any intermediate connection members, such as connection rings, struts, etc., at the inner ends 108, 112. More specifically, the front blade 102 and the rear blade 104 are completely supported by the connection to the nacelle 50 at the outer ends 106, 110, and are not supported by any structure extending between adjacent guide vane assemblies 100, for example. In the exemplary embodiment, front blade span 114 is less than rear blade span 116 , and it will be appreciated that the larger of front blade span 114 and rear blade span 116 may define the overall span of guide vane assembly 100 , ie, the “guide vane assembly span.”

[0039] It will be appreciated that each of the plurality of fan blades 40 defines a fan blade span 118. In the exemplary embodiment, the front blade span 114 and the rear blade span 116 (as well as the guide vane assembly span) are 5% to 50% of the fan blade span, for example, 5% to 10%. In general, the guide vane assembly span may be expressed as a percentage of the fan blade span 118.

[0040] Front bucket 102 generally extends in axial direction A from leading edge 120 to trailing edge 122 , defining a “front axial length”, while rear bucket 104 generally extends in axial direction A from leading edge 124 to trailing edge 126 , defining a “rear axial length”.

[0041] More specifically, the trailing edge 122 of the front blade 102 abuts the leading edge 124 of the rear blade 104. Alternatively, the trailing edge 122 of the front blade 102 may be spaced or separated from the leading edge 124 of the rear blade 104, thereby defining a gap therebetween. In this embodiment, at a given radial position within the nacelle 50, the front blade 104 defines a front axial length L between the leading edge 120 and the trailing edge 122. f , and the rear blade 104 defines a rear axial length L between the leading edge 124 and the trailing edge 126 a In the exemplary embodiment shown, the front axial length L f and rear axial length L a are measured at a common position along the radial direction R and along the axial direction A. Specifically, for the embodiment shown, the front axial length L f and rear axial length L a Both are measured at a position along the radial direction R corresponding to 50% of the span of the front blade 102 .

[0042] As part of the design of the guide vane assembly 100 of the present disclosure, the inventors of the present disclosure designed several iterations to meet the identified design requirements. Specifically, the inventors of the present disclosure designed several iterations that had sufficient stiffness to withstand contact with debris while having sufficient variability to provide the desired aerodynamic advantage under various flight conditions. These iterations included inlet guide vanes of various spans relative to the fan blades. As part of these design iterations, the inventors of the present disclosure discovered that the front axial length L f and rear axial length L a to meet these design requirements.

[0043] Specifically, the inventors of the present disclosure have discovered that as the span of the guide vane assembly increases, more debris contacts the front vane 102 during flight operations and generally requires an increase in the front axial length L f This configuration increases the overall rigidity of the guide vane assembly 100. As a result, the rear axial length L a can be reduced to keep the axial length of the guide vane assembly 100 constant, as shown in Table 1. That is, as the span of the guide vane assembly increases, the rear axial length L a Can be with front axial length L f Inversely proportional.

[0044] In this exemplary embodiment, the rear axial length L a With front axial length L fThe ratio of is 0.2 to 5.0. More specifically, the guide vane assembly 100 is designed so that the front axial length L is determined based on the guide vane assembly span. f and rear axial length L a , such that the front vanes 102 provide debris protection and the rear vanes 104 provide swirl control. Table 1 below shows an example front axial length L of an example guide vane assembly 100 having a specified guide vane assembly span f and rear axial length L a Table 1 also shows the minimum rear axial length of the rear blade 104 of the exemplary embodiment (ie, 0.2L f ) and the maximum rear axial length of the rear impeller 104 (ie, 5.0L f ).

[0045]

[0046] Table 1: Axial lengths of front and rear blades

[0047] like Figure 3 , an axial view of the inlet 60 of the turbofan engine 10 is shown. In this embodiment, the plurality of guide vane assemblies 100 include a relatively large number of guide vane assemblies 100 arranged circumferentially around the nacelle 50. More specifically, Figure 3 The plurality of guide vane assemblies 100 shown in the drawings are about ten guide vane assemblies 100 to about fifty guide vane assemblies 100. The plurality of guide vane assemblies 100 are substantially evenly arranged along the circumferential direction C. More specifically, each guide vane assembly 100 of the plurality of guide vane assemblies 100 defines a circumferential spacing 128 with an adjacent guide vane assembly 100, and each circumferential spacing 128 is substantially equal to one another.

[0048] Or, if Figure 4 As shown, the circumferential spacing 128 between the guide vane assemblies 100 can be different. That is, at least one circumferential spacing 128 can be different from at least one other circumferential spacing 128. For example, a first circumferential spacing 128A defined between two adjacent guide vane assemblies 100A, 100B is different from a second circumferential spacing 128B between another two adjacent guide vane assemblies 100B, 100C. In this example, the first circumferential spacing 128A is at least about 20% greater than the second circumferential spacing 128B, such as at least about 25% greater and at most about 200% greater. The non-uniform circumferential spacing 128 can, for example, be offset by a structure upstream of the guide vane assembly 100.

[0049] Reference now Figure 5-6 , showing Figure 2FIG. 1 is a cross-sectional view of a partial span inlet guide vane of guide vane assembly 100. As shown, guide vane assembly 100 is generally configured as an airfoil having a pressure side 130 and a suction side 132 and extends from a leading edge 120 of a forward vane 102 to a trailing edge 126 of a trailing vane 104. For the exemplary embodiment, the direction of airflow 58 is substantially parallel to the axial direction A and the longitudinal axis 12 of the turbofan engine 10.

[0050] The front blades 102 and the rear blades 104 define respective pitch angles 134, 136. In this context, the "pitch angle" is defined as the pitch angle of the turbofan engine 10 ( Figure 2 ) and a chord extending from a leading edge 120, 124 of the front blade 102 or the rear blade 104 to a trailing edge 122, 126. More specifically, a chord 138 of the front blade 104 is defined between a forward-most point of the front blade 102 and a rearward-most point of the front blade 102, and a chord 140 of the rear blade 104 is defined between a forward-most point of the rear blade 104 and a rearward-most point of the rear blade 104. The pitch angle 134 of the front blade 102 is fixed, such that the rigidity of the front blade 102 is increased, thereby improving debris deflection of the front blade 102. That is, by fixing the pitch angle 134 of the front blade 102, the front blade 102 may be more rigidly fixed to the nacelle 50, thereby increasing the stiffness of the front blade 102.

[0051] The pitch angle 136 of the rear blade 104 is movable to provide a controlled swirl of the airflow 58 passing through the guide vane assembly 100. As described above, the rear blade 104 may include a fixed portion 104A, such as a rod attached to or integrated with the nacelle 50, about which a movable portion 104B of the rear blade 104 rotates. The movable portion 104B is rotatable about a pitch axis P extending through the fixed portion 104A so that the rear blade 104 can be at a first pitch angle 136A (e.g., Figure 5 ) and a second pitch angle 136B (eg, Figure 6 By moving the rear vanes 104 between different pitch angles 136, the rear vanes 104 control the swirl of the airflow 58 flowing through the guide vane assembly 100, which can reduce the turbulence of the airflow 58 and / or provide a specified amount of pre-swirl at the radially outer end of the fan blade 40, where the fan blade 40 ( Figure 2 ) is maximum, thereby providing a desired reduction in flow separation and / or shock losses that may otherwise occur during operation of turbofan engine 10 due to the relatively high speed of the plurality of fan blades 40 at the fan tip. Figure 5 to Figure 6In the exemplary embodiment of the invention, aft blades 104 are movable between a minimum pitch angle of 5 degrees and a maximum pitch angle of 35 degrees.

[0052] The specified pitch angle 136 of the rear blades 104 may be determined so that the swirl imparted to the incoming airflow 58 is proportional to the airspeed of the aircraft and the fan 38 ( Figure 2 ) so that the angular velocity of the air as it approaches the fan blades 40 corresponds as closely as possible to the angular velocity of the fan blades 40. This minimizes the possibility of fan 38 surging or stalling. The faster the fan 38 rotates, the more swirl needs to be applied by the guide vane assembly 100. As the aircraft airspeed increases, the time required for the incoming air to pass from the guide vane assembly 100 to the leading edge of the fan 38 decreases, and thus the amount of swirl required is correspondingly reduced. Thus, the maximum swirl needs to be applied when the turbofan engine 10 is applying maximum thrust to a stationary aircraft, just prior to commencing the takeoff roll.

[0053] Reference now Figure 7 , showing an enlarged view of the turbofan engine 10. The turbofan engine 10 may include a variable pitch mechanism 142 that is operably coupled to the rear blades 104. The variable pitch mechanism 142 is configured to move the rear blades 104 about the pitch axis P to a specified pitch angle 136, for example, Figure 5-6 , shown in the figure, moves from a first angle 136A to a second angle 136B. It is contemplated that the variable pitch mechanism 142 may include, for example, a stepper motor, a torque motor, or a similar drive component. The controller 144 communicates with the variable pitch mechanism 142 and actuates the variable pitch mechanism 142 to move the rear blades 104. The controller 144 includes a processor and a memory, and the processor is configured to determine when to change the pitch angle 136 of the rear blades 104 based on data from one or more sensors (not shown), such as a side wind sensor, a pressure sensor, a blade passing sensor, a temperature sensor, etc. The processor then actuates the variable pitch mechanism 142 to move the rear blades 104 to the specified pitch angle 136.

[0054] Further aspects are provided by the subject matter of the following clauses:

[0055] A guide vane assembly for a nacelle of a gas turbine engine, the gas turbine engine defining a longitudinal axis and an axial direction, the nacelle circumferentially surrounding a plurality of fan blades of the gas turbine engine, the guide vane assembly comprising: a front vane; and a rear vane, when the guide vane assembly is positioned in the nacelle of the gas turbine engine, the rear vane is located behind the front vane and in front of the plurality of fan blades, when the guide vane assembly is positioned in the nacelle of the gas turbine engine, the front vane defines a fixed pitch angle and the rear vane is movable between a first angle relative to the longitudinal axis and a second angle relative to the longitudinal axis.

[0056] A guide vane assembly according to any preceding clause, wherein the leading vane defines a leading axial length at a radial position, wherein the trailing vane defines a trailing axial length at the radial position, wherein a ratio of the trailing axial length to the leading axial length is 0.2 to 5.0.

[0057] A guide vane assembly as claimed in any preceding clause, wherein the axial length of the aft vane is based on the span of the forward vane.

[0058] A guide vane assembly as claimed in any preceding clause, wherein the trailing edge of the leading vane abuts the leading edge of the trailing vane.

[0059] The guide vane assembly of any preceding clause, wherein the leading vane defines a leading vane span, wherein the plurality of fan blades each define a fan blade span, and wherein the leading vane span is 5% to 50% of the fan blade span of the plurality of fan blades.

[0060] A guide vane assembly as claimed in any preceding clause, wherein the front vane span of the front vane is between 5% and 10% of the fan blade span of the plurality of fan blades.

[0061] A guide vane assembly as claimed in any preceding clause, wherein the leading vane comprises a first material, wherein the trailing vane comprises a second material, and wherein the first material has a higher stiffness than the second material.

[0062] A guide vane assembly as claimed in any preceding clause, wherein the front vanes are configured to deflect objects entering the nacelle.

[0063] The guide vane assembly of any preceding clause, wherein when the guide vane assembly is positioned in the nacelle of the gas turbine engine, the front vane and the rear vane are each attached to or integrated with the nacelle.

[0064] A guide vane assembly according to any preceding clause, wherein when the guide vane assembly is positioned in the nacelle of the gas turbine engine, the front vane is fixed to a specified angle relative to the longitudinal axis of the nacelle.

[0065] The guide vane assembly of any preceding clause, wherein the first angle and the second angle are between 5 and 35 degrees.

[0066] A guide vane assembly as described in any preceding clause, further comprising a variable pitch mechanism operably coupled to the rear vane and configured to move the rear vane from the first angle to the second angle.

[0067] A guide vane assembly as claimed in any preceding clause, wherein a front vane span of the front vanes is less than a rear vane span of the rear vanes.

[0068] A turbofan engine, defining a longitudinal axis and an axial direction, comprises: a fan, the fan comprising a plurality of fan blades; a turbine, the turbine being operably coupled to the fan and configured to drive the fan; a nacelle, the nacelle surrounding and at least partially enclosing the fan; and a guide vane assembly, the guide vane assembly being located in front of the plurality of fan blades in the axial direction, the guide vane assembly comprising front blades and rear blades, when the guide vane assembly is positioned in the nacelle of the gas turbine engine, the rear blades being located behind the front blades and in front of the plurality of fan blades, when the guide vane assembly is positioned in the nacelle of the turbofan engine, the front blades defining a fixed pitch angle and the rear blades being movable between a first angle relative to the longitudinal axis and a second angle relative to the longitudinal axis.

[0069] A turbofan engine as described in any preceding clause, wherein the guide vane assembly is attached to or integrated with the nacelle.

[0070] The turbofan engine according to any of the preceding clauses, further comprising: a plurality of guide vane assemblies arranged circumferentially around the nacelle.

[0071] A turbofan engine as described in any preceding clause, wherein said plurality of guide vane assemblies defines a circumferential spacing between each pair of adjacent guide vane assemblies of said plurality of guide vane assemblies.

[0072] A turbofan engine as described in any preceding clause, wherein said circumferential spacing between each pair of adjacent guide vane assemblies of said plurality of guide vane assemblies is equal.

[0073] A turbofan engine according to any preceding clause, wherein the circumferential spacing between at least one pair of adjacent guide vane assemblies of the plurality of guide vane assemblies is different from the circumferential spacing between at least one other pair of adjacent guide vane assemblies of the plurality of guide vane assemblies.

[0074] A turbofan engine as described in any preceding clause, wherein the circumferential spacing between each pair of adjacent guide vane assemblies of the plurality of guide vane assemblies is different from the circumferential spacing between each other pair of adjacent guide vane assemblies of the plurality of guide vane assemblies.

[0075] A turbofan engine according to any preceding clause, wherein the front blade defines a front axial length at a radial position, wherein the rear blade defines a rear axial length at the radial position, wherein a ratio of the rear axial length to the front axial length is 0.2 to 5.0.

[0076] A turbofan engine as described in any preceding clause, wherein said rear axial length is based on a span of said front blades.

[0077] A turbofan engine as described in any preceding clause, wherein the span of the front vane is 5% to 50% of the span of the plurality of fan blades.

[0078] A turbofan engine as described in any preceding clause, further comprising a variable pitch mechanism operably coupled to the rear blades and configured to move the rear blades from the first angle to the second angle.

[0079] A turbofan engine according to any preceding clause, further comprising a controller in communication with the variable pitch mechanism and configured to actuate the variable pitch mechanism to move the rear blades.

[0080] A turbofan engine according to any preceding clause, wherein the front blades are fixed to a specified pitch angle when the guide vane assembly is positioned in the nacelle of the turbofan engine.

[0081] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable those skilled in the art to practice the present disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims.

Claims

1. A guide vane assembly for a nacelle of a gas turbine engine, characterized in that: The gas turbine engine defines a longitudinal axis and an axial direction, the nacelle circumferentially surrounds a plurality of fan blades of the gas turbine engine, the guide vane assembly comprising: front blades; and A rear blade, when the guide blade assembly is positioned in the nacelle of the gas turbine engine, the rear blade is located behind the front blade and in front of the plurality of fan blades, when the guide blade assembly is positioned in the nacelle of the gas turbine engine, the front blade defines a fixed pitch angle and the rear blade is movable between a first pitch angle and a second pitch angle.

2. The guide vane assembly according to claim 1, characterized in that: The front blade defines a front axial length at a radial position, wherein the rear blade defines a rear axial length at the radial position, wherein a ratio of the rear axial length to the front axial length is 0.2 to 5.

0.

3. The guide vane assembly according to claim 2, characterized in that: Wherein the axial length of the rear blade is based on the span of the front blade.

4. The guide vane assembly according to claim 1, characterized in that: The trailing edge of the front blade abuts against the leading edge of the rear blade.

5. The guide vane assembly according to claim 1, characterized in that: Wherein the front blade defines a front blade span, wherein the plurality of fan blades each define a fan blade span, and wherein the front blade span is 5% to 50% of the fan blade span of the plurality of fan blades.

6. The guide vane assembly according to claim 5, characterized in that: The front wheel blade span is 5% to 10% of the fan blade span.

7. The guide vane assembly according to claim 1, characterized in that: Wherein the front blade comprises a first material, wherein the rear blade comprises a second material, and wherein the first material has a higher stiffness than the second material.

8. The guide vane assembly according to claim 1, characterized in that: The front blades are configured to deflect objects entering the nacelle.

9. The guide vane assembly according to claim 1, characterized in that: Wherein when the guide vane assembly is positioned in the nacelle of the gas turbine engine, the front vane and the rear vane are each attached to or integrated with the nacelle.

10. The guide vane assembly according to claim 1, characterized in that: Wherein when the guide vane assembly is positioned in the nacelle of the gas turbine engine, the front blades are fixed to a specified pitch angle.

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