CA2697292A1 - Method and apparatus for a structural outlet guide vane - Google Patents
Method and apparatus for a structural outlet guide vane Download PDFInfo
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- CA2697292A1 CA2697292A1 CA2697292A CA2697292A CA2697292A1 CA 2697292 A1 CA2697292 A1 CA 2697292A1 CA 2697292 A CA2697292 A CA 2697292A CA 2697292 A CA2697292 A CA 2697292A CA 2697292 A1 CA2697292 A1 CA 2697292A1
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- Prior art keywords
- outlet guide
- guide vane
- assembly
- fabricated
- gas turbine
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
- F01D5/288—Protective coatings for blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Architecture (AREA)
- Chemical & Material Sciences (AREA)
- Materials Engineering (AREA)
- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
An outlet guide vane (62,100) for a gas turbine engine is provided. The outlet guide vane (62,100) includes a first flange (64), a second flange (66) positioned radially outwardly from the first flange, an airfoil (102) extending between the first and second flanges and coupled thereto, the airfoil including a leading edge portion (130) and a trailing edge portion (132), the leading edge portion and the trailing edge portion fabricated from a first material, a filler portion (150) coupled in a gap formed by the leading edge portion and the trailing edge portion, the filler portion including a first side (203) and a second side (204), the filler portion fabricated from a second material, the second material different from the first material, a first skin coupled to the first side, the first skin fabricated from a composite material, and a second skin coupled to the second side, the second skin fabricated from a composite material.
Description
METHOD AND APPARATUS FOR A STRUCTURAL
OUTLET GUIDE VANE
BACKGROUND OF THE INVENTION
This invention relates generally to gas turbine engines, and more particularly, to a gas turbine engine vane and a method of fabricating the same.
At least one known gas turbine engine assembly includes a fan assembly that is mounted upstream from a core gas turbine engine. During operation, airflow discharged from the fan assembly is channeled downstream to the core gas turbine engine where the airflow is further compressed. The compressed airflow is then channeled into a combustor, mixed with fuel, and ignited to generate hot combustion gases. The combustion gases are then channeled to a turbine which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight.
To facilitate channeling the airflow from the fan assembly to the core gas turbine engine, at least one known gas turbine engine assembly includes an outlet guide vane assembly. The outlet guide vane assembly is configured to redirect the circumferentially flowing air discharged from the fan assembly into a substantially axial direction prior to the fan flow being channeled into the core gas turbine engine.
In addition to redirecting the fan airflow, the outlet guide vane assembly also provides structural stiffness to the fan frame. More specifically, outlet guide vane assemblies generally include a plurality of outlet guide vanes that are coupled to the fan frame.
To provide the necessary structural stiffness the fan frame, the known outlet guide vanes are forged as substantially solid vanes using a metallic material.
However, because some known outlet guide vanes are substantially solid, they increase the overall weight of the gas turbine engine assembly, and may also cause a reduction in fuel efficiency. Other known guide vanes have attempted to use a lighter filler material surrounded by a metallic portion but while lighter, these vanes are still too heavy.
BRIEF SUMMARY OF THE INVENTION
In one aspect, a method for assembling an outlet guide vane includes providing an outlet guide vane frame having a first side and a second side. The outlet guide vane includes a radially inward flange coupled to a radially outward flange by a leading edge. A trailing edge is coupled to the radially inward flange and the radially outward flange aft of the leading edge. A cavity is defined between the radially inward flange, the radially outward flange, the leading edge, and the trailing edge. A first and a second mating surface circumscribe the cavity on the first and second sides respectively. The method further includes coupling a filler portion within the cavity wherein the filler portion includes a third side and a fourth side, coupling a first skin to the third side and the first mating surface, and coupling a second skin to the fourth side and the second mating surface, wherein at least one of the first skin and the second skin are fabricated from a composite material.
In another aspect, an outlet guide vane for a gas turbine engine includes a first flange, a second flange positioned radially outwardly from the first flange, an airfoil extending between the first and second flanges and coupled thereto. The airfoil comprises a leading edge portion and a trailing edge portion. The leading edge portion and the trailing edge portion are fabricated from a first material. A
filler portion is coupled in a gap formed by the leading edge portion and the trailing edge portion. The filler portion comprises a first side and a second side and is fabricated from a second material different from the first material. A first skin is coupled to the first side and a second skin is coupled to the second side. The first and second skins are fabricated from a composite material.
In a further aspect, a gas turbine engine assembly includes a core gas turbine engine, a fan assembly disposed upstream from the core gas turbine engine, and an outlet guide vane assembly including a plurality of outlet guide vanes. The outlet guide vane assembly is positioned downstream from the fan assembly. At least one of the plurality of outlet guide vanes includes, an airfoil including a leading edge structure and a trailing edge structure each fabricated from a first material, a filler portion having a first side and a second side wherein the filler portion is positioned between the leading and trailing edge portions and fabricated from a second material that is
OUTLET GUIDE VANE
BACKGROUND OF THE INVENTION
This invention relates generally to gas turbine engines, and more particularly, to a gas turbine engine vane and a method of fabricating the same.
At least one known gas turbine engine assembly includes a fan assembly that is mounted upstream from a core gas turbine engine. During operation, airflow discharged from the fan assembly is channeled downstream to the core gas turbine engine where the airflow is further compressed. The compressed airflow is then channeled into a combustor, mixed with fuel, and ignited to generate hot combustion gases. The combustion gases are then channeled to a turbine which extracts energy from the combustion gases for powering the compressor, as well as producing useful work to propel an aircraft in flight.
To facilitate channeling the airflow from the fan assembly to the core gas turbine engine, at least one known gas turbine engine assembly includes an outlet guide vane assembly. The outlet guide vane assembly is configured to redirect the circumferentially flowing air discharged from the fan assembly into a substantially axial direction prior to the fan flow being channeled into the core gas turbine engine.
In addition to redirecting the fan airflow, the outlet guide vane assembly also provides structural stiffness to the fan frame. More specifically, outlet guide vane assemblies generally include a plurality of outlet guide vanes that are coupled to the fan frame.
To provide the necessary structural stiffness the fan frame, the known outlet guide vanes are forged as substantially solid vanes using a metallic material.
However, because some known outlet guide vanes are substantially solid, they increase the overall weight of the gas turbine engine assembly, and may also cause a reduction in fuel efficiency. Other known guide vanes have attempted to use a lighter filler material surrounded by a metallic portion but while lighter, these vanes are still too heavy.
BRIEF SUMMARY OF THE INVENTION
In one aspect, a method for assembling an outlet guide vane includes providing an outlet guide vane frame having a first side and a second side. The outlet guide vane includes a radially inward flange coupled to a radially outward flange by a leading edge. A trailing edge is coupled to the radially inward flange and the radially outward flange aft of the leading edge. A cavity is defined between the radially inward flange, the radially outward flange, the leading edge, and the trailing edge. A first and a second mating surface circumscribe the cavity on the first and second sides respectively. The method further includes coupling a filler portion within the cavity wherein the filler portion includes a third side and a fourth side, coupling a first skin to the third side and the first mating surface, and coupling a second skin to the fourth side and the second mating surface, wherein at least one of the first skin and the second skin are fabricated from a composite material.
In another aspect, an outlet guide vane for a gas turbine engine includes a first flange, a second flange positioned radially outwardly from the first flange, an airfoil extending between the first and second flanges and coupled thereto. The airfoil comprises a leading edge portion and a trailing edge portion. The leading edge portion and the trailing edge portion are fabricated from a first material. A
filler portion is coupled in a gap formed by the leading edge portion and the trailing edge portion. The filler portion comprises a first side and a second side and is fabricated from a second material different from the first material. A first skin is coupled to the first side and a second skin is coupled to the second side. The first and second skins are fabricated from a composite material.
In a further aspect, a gas turbine engine assembly includes a core gas turbine engine, a fan assembly disposed upstream from the core gas turbine engine, and an outlet guide vane assembly including a plurality of outlet guide vanes. The outlet guide vane assembly is positioned downstream from the fan assembly. At least one of the plurality of outlet guide vanes includes, an airfoil including a leading edge structure and a trailing edge structure each fabricated from a first material, a filler portion having a first side and a second side wherein the filler portion is positioned between the leading and trailing edge portions and fabricated from a second material that is
-2-different than the first material. The airfoil further includes a first skin coupled to the first side and a second skin coupled to the second side. The firs and second skins are fabricated from a composite material.
BRIEF DESCRIPTION OF THE DRAWING
Figures 1-5 show exemplary embodiments of the method and apparatus described herein.
Figure 1 is a schematic illustration of an exemplary gas turbine engine assembly;
Figure 2 is a perspective view of a fan frame assembly that includes a plurality of outlet guide vanes that may be utilized with the gas turbine engine assembly shown in Figure 1;
Figure 3 is a perspective view of an outlet guide vane that may be used with the fan frame assembly shown in Figure 2;
Figure 4 is an exploded view of the outlet guide vane of Figure 3; and Figure 5 is a cross-sectional view of the outlet guide vane shown in Figure 3.
DETAILED DESCRIPTION OF THE INVENTION
Figure 1 is a schematic illustration of an exemplary gas turbine engine assembly 10 having a longitudinal axis 11. Gas turbine engine assembly 10 includes a fan assembly 12 and a core gas turbine engine 13. Core gas turbine engine 13 includes a high pressure compressor 14, a combustor 16, and a high pressure turbine 18.
In the exemplary embodiment, gas turbine engine assembly 10 also includes a low pressure turbine 20, and a multi-stage booster compressor 22.
Fan assembly 12 includes an array of fan blades 24 extending radially outward from a rotor disk 26. Gas turbine engine assembly 10 has an intake side 28 and an exhaust side 30. Fan assembly 12, booster 22, and turbine 20 are coupled together by a first rotor shaft 31, and compressor 14 and turbine 18 are coupled together by a second rotor shaft 32.
BRIEF DESCRIPTION OF THE DRAWING
Figures 1-5 show exemplary embodiments of the method and apparatus described herein.
Figure 1 is a schematic illustration of an exemplary gas turbine engine assembly;
Figure 2 is a perspective view of a fan frame assembly that includes a plurality of outlet guide vanes that may be utilized with the gas turbine engine assembly shown in Figure 1;
Figure 3 is a perspective view of an outlet guide vane that may be used with the fan frame assembly shown in Figure 2;
Figure 4 is an exploded view of the outlet guide vane of Figure 3; and Figure 5 is a cross-sectional view of the outlet guide vane shown in Figure 3.
DETAILED DESCRIPTION OF THE INVENTION
Figure 1 is a schematic illustration of an exemplary gas turbine engine assembly 10 having a longitudinal axis 11. Gas turbine engine assembly 10 includes a fan assembly 12 and a core gas turbine engine 13. Core gas turbine engine 13 includes a high pressure compressor 14, a combustor 16, and a high pressure turbine 18.
In the exemplary embodiment, gas turbine engine assembly 10 also includes a low pressure turbine 20, and a multi-stage booster compressor 22.
Fan assembly 12 includes an array of fan blades 24 extending radially outward from a rotor disk 26. Gas turbine engine assembly 10 has an intake side 28 and an exhaust side 30. Fan assembly 12, booster 22, and turbine 20 are coupled together by a first rotor shaft 31, and compressor 14 and turbine 18 are coupled together by a second rotor shaft 32.
-3-In operation, air flows through fan assembly 12 and a first portion of the airflow is channeled through booster 22. The compressed air that is discharged from booster 22 is channeled through compressor 14 wherein the airflow is further compressed and delivered to combustor 16. Hot products of combustion (not shown in Figure 1) from combustor 16 are utilized to drive turbines 18 and 20, and turbine 20 is utilized to drive fan assembly 12 and booster 22 by way of shaft 31. Gas turbine engine assembly 10 is operable at a range of operating conditions between design operating conditions and off-design operating conditions.
A second portion of the airflow discharged from fan assembly 12 is channeled through a bypass duct 40 to bypass a portion of the airflow from fan assembly around core gas turbine engine 13. More specifically, bypass duct 40 extends between a fan casing or shroud 42 and splitter 44. Accordingly, a first portion 50 of the airflow from fan assembly 12 is channeled through booster 22 and then into compressor 14 as described above and a second portion 52 of the airflow from fan assembly 12 is channeled through bypass duct 40 to provide thrust for an aircraft, for example. Gas turbine engine assembly 10 also includes a fan frame assembly 60 to provide structural support for fan assembly 12 and is also utilized to coupled fan assembly 12 to core gas turbine engine 13.
Figure 2 is a front perspective view of fan frame assembly 60. Fan frame assembly 60 includes a plurality of outlet guide vanes 62 that extend substantially radially between a radially outward mounting flange 64 and a radially inward mounting flange 66 and are circumferentially-spaced around within bypass duct 40. Fan frame assembly also includes a plurality of struts 68 that are coupled between radially outward mounting flange 64 and radially inward mounting flange 66. In one embodiment, fan frame assembly 60 is fabricated in arcuate segments in which flanges 64 and 66 are coupled to outlet guide vanes 62 and struts 68. In the exemplary embodiment, outlet guide vanes 62 and struts 68 are coupled coaxially within bypass duct 40.
Specifically, outlet guide vanes 62 and struts 68 each coupled between flanges 64 and 66 in the same axial location as shown in Figure 2. Optionally, outlet guide vanes 62 may be coupled downstream from struts 68 within bypass duct 40.
A second portion of the airflow discharged from fan assembly 12 is channeled through a bypass duct 40 to bypass a portion of the airflow from fan assembly around core gas turbine engine 13. More specifically, bypass duct 40 extends between a fan casing or shroud 42 and splitter 44. Accordingly, a first portion 50 of the airflow from fan assembly 12 is channeled through booster 22 and then into compressor 14 as described above and a second portion 52 of the airflow from fan assembly 12 is channeled through bypass duct 40 to provide thrust for an aircraft, for example. Gas turbine engine assembly 10 also includes a fan frame assembly 60 to provide structural support for fan assembly 12 and is also utilized to coupled fan assembly 12 to core gas turbine engine 13.
Figure 2 is a front perspective view of fan frame assembly 60. Fan frame assembly 60 includes a plurality of outlet guide vanes 62 that extend substantially radially between a radially outward mounting flange 64 and a radially inward mounting flange 66 and are circumferentially-spaced around within bypass duct 40. Fan frame assembly also includes a plurality of struts 68 that are coupled between radially outward mounting flange 64 and radially inward mounting flange 66. In one embodiment, fan frame assembly 60 is fabricated in arcuate segments in which flanges 64 and 66 are coupled to outlet guide vanes 62 and struts 68. In the exemplary embodiment, outlet guide vanes 62 and struts 68 are coupled coaxially within bypass duct 40.
Specifically, outlet guide vanes 62 and struts 68 each coupled between flanges 64 and 66 in the same axial location as shown in Figure 2. Optionally, outlet guide vanes 62 may be coupled downstream from struts 68 within bypass duct 40.
-4-Fan frame assembly 60 is one of the various frame and support assemblies of gas turbine engine assembly 10 that are used to facilitate maintaining an orientation of various components within gas turbine engine assembly 10. More specifically, such frame and support assemblies interconnect stationary components and provide rotor bearing supports. Fan frame assembly 60 is coupled downstream from fan assembly 12 within bypass duct 40 such that outlet guide vanes 62 and struts 68 are circumferentially-spaced around the outlet of fan assembly 12 and extend across the airflow path discharged from fan assembly 12.
In the exemplary embodiment, outlet guide vane 100 includes an airfoil 102 that is coupled between a radially outward flange 104 and a radially inward flange 106. In the exemplary embodiment, airfoil 102, radially outward flange 104, and radially inward flange 106 are cast or forged as a unitary outlet guide vane 100.
Optionally, radially outward flange 104 and radially inward flange 106 may be coupled to airfoil 102 using a welding or brazing technique, for example. Although in the exemplary embodiment radially outward and radially inward flanges 104 and 106 are fabricated from metal, in other embodiments radially outward and radially inward flanges and 106 may be fabricated from a composite material. In still other embodiments, airfoil 102 may be fabricated partially or entirely from a composite material and may have a metallic strengthener attached to it.
Airfoil 102 includes a first sidewall 110 and a second sidewall 112. In one embodiment, either first and/or second sidewalls 110 and/or 112 may be contoured to improve aerodynamic performance. In the exemplary embodiment, first sidewall is convex and defines a suction side of airfoil 102, and second sidewall 112 is concave and defines a pressure side of airfoil 102. Sidewalls 110 and 112 are joined at a leading edge 114 and at an axially-spaced trailing edge 116 of airfoil 102.
More specifically, airfoil trailing edge 116 is spaced chordwise and downstream from airfoil leading edge 114. First and second sidewalls 110 and 112, respectively, extend longitudinally or radially outward in span from radially inner flange 106 to radially outer flange 104. In the exemplary embodiment, at least a portion of outlet guide vane 100 is fabricated utilizing a metallic material such as, but not limited to, titanium, aluminum, and/or a Metal Matrix Composite (MMC) material.
In the exemplary embodiment, outlet guide vane 100 includes an airfoil 102 that is coupled between a radially outward flange 104 and a radially inward flange 106. In the exemplary embodiment, airfoil 102, radially outward flange 104, and radially inward flange 106 are cast or forged as a unitary outlet guide vane 100.
Optionally, radially outward flange 104 and radially inward flange 106 may be coupled to airfoil 102 using a welding or brazing technique, for example. Although in the exemplary embodiment radially outward and radially inward flanges 104 and 106 are fabricated from metal, in other embodiments radially outward and radially inward flanges and 106 may be fabricated from a composite material. In still other embodiments, airfoil 102 may be fabricated partially or entirely from a composite material and may have a metallic strengthener attached to it.
Airfoil 102 includes a first sidewall 110 and a second sidewall 112. In one embodiment, either first and/or second sidewalls 110 and/or 112 may be contoured to improve aerodynamic performance. In the exemplary embodiment, first sidewall is convex and defines a suction side of airfoil 102, and second sidewall 112 is concave and defines a pressure side of airfoil 102. Sidewalls 110 and 112 are joined at a leading edge 114 and at an axially-spaced trailing edge 116 of airfoil 102.
More specifically, airfoil trailing edge 116 is spaced chordwise and downstream from airfoil leading edge 114. First and second sidewalls 110 and 112, respectively, extend longitudinally or radially outward in span from radially inner flange 106 to radially outer flange 104. In the exemplary embodiment, at least a portion of outlet guide vane 100 is fabricated utilizing a metallic material such as, but not limited to, titanium, aluminum, and/or a Metal Matrix Composite (MMC) material.
-5-As shown in Figure 3, airfoil 102 includes a first assembly surface 120, a second assembly surface 121, a leading edge portion or spar 130 and a trailing edge portion or spar 132. More specifically, airfoil 102 has a profile that tapers outwardly from leading edge 114 at least partially towards trailing edge 116 and also tapers outwardly from trailing edge 116 at least partially towards leading edge 114. The profile then tapers inwardly from the leading edge portion 130 and from trailing edge portion 132 to define first and second assembly surfaces 120 and 121. The profile then tapers further inwardly to form a cavity 140 circumscribed by first assembly surface 120 on one side of the airfoil and circumscribed by second assembly surface 121 on the other side of the airfoil.
In the exemplary embodiment, cavity 140 is sized to receive a filler 150 comprising a first side 203 and a second side 204 that is a honeycombed material. That is, filler 150 has a thickness 151 that is smaller than the thickness 144 of cavity 140 such that filler 150 is not flush with a leading edge portion 130 and/or trailing edge portion 132 when positioned within cavity 140. Two laminate portions 201 and 202 are sized to overlay the filler 150 between leading edge portion 130 and trailing edge portion 132.
Laminate portion 201 is coupled to first side 203 and first assembly surface 120 such that laminate portion 201 is substantially flush with both leading edge portion 130 and trailing edge portion 132 and laminate portion 202 is coupled to second side 204 and second assembly surface 121 such that laminate portion 202 is also substantially flush with both leading edge portion 130 and trailing edge portion 132. In some embodiments, a protective coating 122, such as polyurethane, may be applied to the outlet guide vane 100.
During fabrication, of outlet guide vane 100, outlet guide vane 100 is cast or forged to include leading edge portion 130, trailing edge portion 132, and inward and outward flange 104 and 106. Filler 150 is then coupled within cavity 140 as described above.
Then a laminate portion 201 is coupled to first side 203 and laminate portion 202 is coupled to second side 204.
Described herein, is a gas turbine engine wherein at least some known outlet guide vanes are replaced with an outlet guide vane having a substantially hollow interior portion filled with a relatively lightweight material sandwiched between two laminate
In the exemplary embodiment, cavity 140 is sized to receive a filler 150 comprising a first side 203 and a second side 204 that is a honeycombed material. That is, filler 150 has a thickness 151 that is smaller than the thickness 144 of cavity 140 such that filler 150 is not flush with a leading edge portion 130 and/or trailing edge portion 132 when positioned within cavity 140. Two laminate portions 201 and 202 are sized to overlay the filler 150 between leading edge portion 130 and trailing edge portion 132.
Laminate portion 201 is coupled to first side 203 and first assembly surface 120 such that laminate portion 201 is substantially flush with both leading edge portion 130 and trailing edge portion 132 and laminate portion 202 is coupled to second side 204 and second assembly surface 121 such that laminate portion 202 is also substantially flush with both leading edge portion 130 and trailing edge portion 132. In some embodiments, a protective coating 122, such as polyurethane, may be applied to the outlet guide vane 100.
During fabrication, of outlet guide vane 100, outlet guide vane 100 is cast or forged to include leading edge portion 130, trailing edge portion 132, and inward and outward flange 104 and 106. Filler 150 is then coupled within cavity 140 as described above.
Then a laminate portion 201 is coupled to first side 203 and laminate portion 202 is coupled to second side 204.
Described herein, is a gas turbine engine wherein at least some known outlet guide vanes are replaced with an outlet guide vane having a substantially hollow interior portion filled with a relatively lightweight material sandwiched between two laminate
-6-sheets. As such, the exemplary outlet guide vanes described herein reduce the overall weight of the gas turbine engine assembly while still maintaining structural integrity thus achieving the very challenging engine weight goals for new applications.
More specifically, the outlet guide vane 100 described herein includes two spars that form the airfoil portion of the outlet guide vane 100. The area between the spars is filled with a honeycombed material to add rigidity to the airfoil, and then sandwiched between two laminate sheets.
This written description uses examples to disclose the invention, including the best mode, and also to enable a person skilled in the art to make and use the invention.
The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.
More specifically, the outlet guide vane 100 described herein includes two spars that form the airfoil portion of the outlet guide vane 100. The area between the spars is filled with a honeycombed material to add rigidity to the airfoil, and then sandwiched between two laminate sheets.
This written description uses examples to disclose the invention, including the best mode, and also to enable a person skilled in the art to make and use the invention.
The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.
-7-
Claims (10)
1. An outlet guide vane (62,100) for a gas turbine engine, said outlet guide vane comprising:
a first flange (64);
a second flange (66) positioned radially outwardly from said first flange;
an airfoil (102) extending between said first and second flanges and coupled thereto, the airfoil comprising a leading edge portion (130) and a trailing edge portion (132), said leading edge portion and said trailing edge portion fabricated from a first material;
a filler portion (150) coupled in a gap formed by the leading edge portion and the trailing edge portion, the filler portion comprising a first side and a second side, the filler portion fabricated from a second material, said second material different from said first material;
a first skin coupled to the first side (203), the first skin fabricated from a composite material; and a second skin coupled to the second side (204), the second skin fabricated from a composite material.
a first flange (64);
a second flange (66) positioned radially outwardly from said first flange;
an airfoil (102) extending between said first and second flanges and coupled thereto, the airfoil comprising a leading edge portion (130) and a trailing edge portion (132), said leading edge portion and said trailing edge portion fabricated from a first material;
a filler portion (150) coupled in a gap formed by the leading edge portion and the trailing edge portion, the filler portion comprising a first side and a second side, the filler portion fabricated from a second material, said second material different from said first material;
a first skin coupled to the first side (203), the first skin fabricated from a composite material; and a second skin coupled to the second side (204), the second skin fabricated from a composite material.
2. An outlet guide vane (62,100) in accordance with Claim 1, wherein said first material comprises a metallic material and said second material comprises a honeycomb material.
3. An outlet guide vane (62,100) in accordance with Claim 1, wherein at least one of the radially inward flange (106), the radially outward flange (104), and the airfoil (102) comprises a composite material.
4. An outlet guide vane (62,100) in accordance with Claim 1, wherein the first material is aluminum.
5. An outlet guide vane (62,100) in accordance with Claim 1 further comprising a protective coating (122) applied to the outlet guide vane.
6. An outlet guide vane (62,100) in accordance with Claim 1 wherein said first flange (104), said second flange (106), and said airfoil (102) comprise at least one of a cast unitary outlet guide vane and a forged unitary outlet guide vane.
7. A gas turbine engine assembly (10) comprising:
a core gas turbine engine (13);
a fan assembly (12) disposed upstream from said core gas turbine engine, said fan assembly comprising a plurality of fan blades; and an outlet guide vane assembly comprising a plurality of outlet guide vanes (62,100), said outlet guide vane assembly positioned downstream from said fan assembly, at least one of said plurality of outlet guide vanes comprising:
an airfoil (102) comprising a leading edge structure and a trailing edge structure each fabricated from a first material;
a filler portion (150) comprising a first side and a second side, said filler portion positioned between the leading and trailing edge portions, the filler portion fabricated from a second material that is different than the first material;
a first skin coupled to the first side (203), the first skin fabricated from a composite material; and a second skin coupled to the second side (204), the second skin fabricated from a composite material.
a core gas turbine engine (13);
a fan assembly (12) disposed upstream from said core gas turbine engine, said fan assembly comprising a plurality of fan blades; and an outlet guide vane assembly comprising a plurality of outlet guide vanes (62,100), said outlet guide vane assembly positioned downstream from said fan assembly, at least one of said plurality of outlet guide vanes comprising:
an airfoil (102) comprising a leading edge structure and a trailing edge structure each fabricated from a first material;
a filler portion (150) comprising a first side and a second side, said filler portion positioned between the leading and trailing edge portions, the filler portion fabricated from a second material that is different than the first material;
a first skin coupled to the first side (203), the first skin fabricated from a composite material; and a second skin coupled to the second side (204), the second skin fabricated from a composite material.
8. A gas turbine engine assembly (10) in accordance with Claim 7, wherein said airfoil (102) further comprises a separable assembly.
9. A gas turbine engine assembly (10) in accordance with Claim 7, wherein the first material comprises a metallic material and the second material comprises a honeycomb material.
10. A gas turbine engine assembly (10) in accordance with Claim 7, wherein at least a portion of the airfoil (102) is fabricated from a composite material.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CA2697292A CA2697292C (en) | 2010-03-18 | 2010-03-18 | Method and apparatus for a structural outlet guide vane |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CA2697292A CA2697292C (en) | 2010-03-18 | 2010-03-18 | Method and apparatus for a structural outlet guide vane |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CA2697292A1 true CA2697292A1 (en) | 2011-09-18 |
| CA2697292C CA2697292C (en) | 2017-06-13 |
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| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA2697292A Expired - Fee Related CA2697292C (en) | 2010-03-18 | 2010-03-18 | Method and apparatus for a structural outlet guide vane |
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| CA (1) | CA2697292C (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2014031203A2 (en) | 2012-06-05 | 2014-02-27 | United Technologies Corporation | Airfoil cover system |
| EP3483396A1 (en) * | 2017-11-10 | 2019-05-15 | Safran Aircraft Engines | Stator blade ring of a turbojet engine comprising an acoustic isolating structure |
-
2010
- 2010-03-18 CA CA2697292A patent/CA2697292C/en not_active Expired - Fee Related
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2014031203A2 (en) | 2012-06-05 | 2014-02-27 | United Technologies Corporation | Airfoil cover system |
| EP2855849A4 (en) * | 2012-06-05 | 2016-11-02 | United Technologies Corp | Airfoil cover system |
| EP3483396A1 (en) * | 2017-11-10 | 2019-05-15 | Safran Aircraft Engines | Stator blade ring of a turbojet engine comprising an acoustic isolating structure |
| FR3073561A1 (en) * | 2017-11-10 | 2019-05-17 | Safran Aircraft Engines | FIXED AUB RING OF A TURBOJET ENGINE COMPRISING AN ACOUSTIC TREATMENT STRUCTURE |
| CN109763863A (en) * | 2017-11-10 | 2019-05-17 | 赛峰航空器发动机 | The ring of turbojet blade including Acoustic treatment structure |
| US10723476B2 (en) | 2017-11-10 | 2020-07-28 | Safran Aircraft Engines | Ring of turbojet vanes including an acoustic treatment structure |
Also Published As
| Publication number | Publication date |
|---|---|
| CA2697292C (en) | 2017-06-13 |
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Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| EEER | Examination request |
Effective date: 20150116 |
|
| MKLA | Lapsed |
Effective date: 20190318 |