WO2012092277A1 - Gas turbine engine and variable camber vane system - Google Patents
Gas turbine engine and variable camber vane system Download PDFInfo
- Publication number
- WO2012092277A1 WO2012092277A1 PCT/US2011/067393 US2011067393W WO2012092277A1 WO 2012092277 A1 WO2012092277 A1 WO 2012092277A1 US 2011067393 W US2011067393 W US 2011067393W WO 2012092277 A1 WO2012092277 A1 WO 2012092277A1
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- WO
- WIPO (PCT)
- Prior art keywords
- airfoil
- crown
- gas turbine
- turbine engine
- groove
- Prior art date
Links
- 239000012530 fluid Substances 0.000 claims description 8
- 238000007789 sealing Methods 0.000 claims description 5
- 229920000642 polymer Polymers 0.000 claims description 4
- 239000004963 Torlon Substances 0.000 claims description 3
- 229920003997 Torlon® Polymers 0.000 claims description 3
- 229920003223 poly(pyromellitimide-1,4-diphenyl ether) Polymers 0.000 claims description 3
- 239000002861 polymer material Substances 0.000 claims description 2
- 238000000034 method Methods 0.000 abstract description 2
- 239000007789 gas Substances 0.000 description 23
- 238000002485 combustion reaction Methods 0.000 description 12
- 238000012986 modification Methods 0.000 description 3
- 230000004048 modification Effects 0.000 description 3
- 238000004891 communication Methods 0.000 description 2
- 238000009434 installation Methods 0.000 description 2
- 239000000463 material Substances 0.000 description 2
- 239000004962 Polyamide-imide Substances 0.000 description 1
- 230000004075 alteration Effects 0.000 description 1
- 238000004200 deflagration Methods 0.000 description 1
- 238000005474 detonation Methods 0.000 description 1
- 238000005516 engineering process Methods 0.000 description 1
- 239000000284 extract Substances 0.000 description 1
- RLQJEEJISHYWON-UHFFFAOYSA-N flonicamid Chemical compound FC(F)(F)C1=CC=NC=C1C(=O)NCC#N RLQJEEJISHYWON-UHFFFAOYSA-N 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 229920002312 polyamide-imide Polymers 0.000 description 1
- 230000001737 promoting effect Effects 0.000 description 1
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
- F01D17/00—Regulating or controlling by varying flow
- F01D17/10—Final actuators
- F01D17/12—Final actuators arranged in stator parts
- F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
- F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
- F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/08—Sealings
- F04D29/083—Sealings especially adapted for elastic fluid pumps
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
- F04D29/544—Blade shapes
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
- F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
Definitions
- the present invention relates to gas turbine engines, and more particularly, to gas turbine engines with variable camber vane systems.
- One embodiment of the present invention is a unique variable camber vane system for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and variable camber vane systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.
- FIG. 1 schematically depicts some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention.
- FIG. 2 schematically depicts some aspects of a non-limiting example of a fan system for a gas turbine engine in accordance with an embodiment of the present invention.
- FIG. 3 depicts some aspects of a non-limiting example of a variable camber guide vane system in accordance with an embodiment of the present invention.
- FIG. 4 depicts some aspects of the variable camber guide vane system of FIG. 3.
- FIG. 5 depicts some aspects of a non-limiting example of a seal strip in accordance with an embodiment of the present invention.
- gas turbine engine 10 is an aircraft propulsion power plant.
- gas turbine engine 10 may be a land-based or marine engine.
- gas turbine engine 10 is a multi-spool turbofan engine. In other words,
- gas turbine engine 10 may be a single or multi-spool turbofan, turboshaft, turbojet, turboprop gas turbine or combined cycle engine.
- Gas turbine engine 10 includes a fan system 12, a compressor system 14, a diffuser 16, a combustion system 18 and a turbine system 20.
- Compressor system 14 is in fluid communication with fan system 12.
- Diffuser 16 is in fluid communication with compressor system 14.
- Combustion system 18 is flu idly disposed between compressor system 14 and turbine system 20.
- Fan system 2 includes a fan rotor system 22.
- fan rotor system 22 includes one or more rotors (not shown) that are powered by turbine system 20.
- Compressor system 14 includes a compressor rotor system 24.
- compressor rotor system 24 includes one or more rotors (not shown) that are powered by turbine system 20.
- Turbine system 20 includes a turbine rotor system 26.
- turbine rotor system 26 includes one or more rotors (not shown) operative to drive fan rotor system 22 and compressor rotor system 24.
- Turbine rotor system 26 is driving coupled to compressor rotor system 24 and fan rotor system 22 via a shafting system 28.
- shafting system 28 includes a plurality of shafts that may rotate at the same or different speeds and directions. In some embodiments, only a single shaft may be employed.
- combustion system 1 8 includes a combustion liner (not shown) that contains a continuous combustion process.
- combustion system 18 may take other forms, and may be, for example, a wave rotor combustion system, a rotary valve combustion system, or a siinger combustion system, and may employ deflagration and/or detonation combustion processes.
- the hot gases exiting combustor 1 8 are directed into turbine system 20, which extracts energy in the form of mechanical shaft power to drive fan system 12 and compressor system 14 via shafting system 28.
- the hot gases exiting turbine system 20 are directed into a nozzle (not shown), and provide a component of the thrust output by gas turbine engine 10.
- Fan system 12 includes a variable guide vane system 40 having a variable inlet guide vane stage 42 and a variable outlet guide vane stage 44 disposed on either side of a rotating fan stage 46.
- Variable inlet guide vane stage 42 is operative to guide air into rotating fan stage 46, and to selectively vary the incidence angle of the air flow entering rotating fan stage 46.
- Variable outlet guide vane stage 44 is operative to guide air exiting rotating fan stage 46, and to selectively vary the incidence angle of the air flow exiting rotating fan stage 46.
- Variable inlet guide vane stage 42 and variable outlet guide vane stage 44 are actuated by an actuation system (not shown).
- variable guide vane system 40 may also or alternatively be employed as part of compressor system 14.
- variable guide vane system 40 includes both variable inlet and outlet guide vane stages, other embodiments may include only a variable inlet guide vane stage or a variable outlet guide vane stage.
- variable inlet guide vane stage 42 in accordance with an embodiment of the present invention is illustrated. It will be understood that some embodiments of variable outlet guide vane stage 44 may be similar to variable inlet guide vane stage 42, and hence, the following description of variable inlet guide vane stage 42 is also applicable to aspects of some embodiments of variable outlet guide vane stage 44.
- Variable inlet guide vane stage 42 includes an outer band 50, an inner band 52 and plurality of vanes 54. Outer band 50 defines an outer flowpath wall of variable inlet guide vane stage 42.
- Inner band 52 defines an inner flowpath wall of variable inlet guide vane stage 42.
- Vanes 54 are airfoils that extend between outer band 50 and inner band 52, and are spaced apart circumferentially. in one form, vanes 54 extend in the radial direction between outer band 50 and inner band 52. In other embodiments, vanes 54 may extend between outer band 50 and inner band 52 at other angles.
- Each vane 54 includes an airfoil portion 56 and an airfoil portion 58.
- Airfoil portion 56 extends between a tip portion 60 and a root portion 62.
- airfoil portion 56 includes the trailing edge 64 of vane 54.
- airfoil portion 56 may be formed with a leading edge of vane 54 instead of trailing edge 64, e.g., for use in variable outlet guide vane 44.
- Airfoil portion 58 extends between a tip portion 66 and a root portion 68.
- airfoil portion 58 includes the leading edge 70 of vane 54.
- airfoil portion 58 may be formed with a trailing edge instead of leading edge 70, e.g., for use in variable outlet guide vane 44.
- airfoil portion 56 is fixed, i.e., stationary.
- airfoil portion 56 may be movable, e.g., pivotable about an axis so as to be able to vary the angle of the trailing edge of vane 54.
- airfoil portion 58 is variable, being configured to pivot about a pivot axis 72 with respect to airfoil portion 56, to provide a variable camber for vane 54.
- airfoil portion 58 may be fixed.
- airfoil portion 58 is coupled to an actuation system (not shown) that is operative to selectively position airfoil portion 58 at a desired incidence angle.
- airfoii portion 56 may also or alternatively be coupled to an actuation system (not shown) that is operative to selectively position airfoil portion 56 at a desired incidence angle.
- Outer band 50 includes a plurality of spaced apart openings 78.
- Inner band 52 includes a plurality of spaced apart openings 80. Openings 78 and 80 are operative to receive pivot shafts 74 and 76, respectively, and retain airfoil portions 58 in the engine axial, circumferential and radial direction.
- pivot shafts 74 and 76 retain airfoil portion 58 in outer band 50 and inner band 52 via anti-friction bushings 82 and 84.
- Anti-friction bushings 82 and 84 are operative to provide bearing surfaces for pivot shafts 74 and 76. Other embodiments may not include anti-friction bushings 82 and 84.
- Airfoil portion 58 is operative to rotate in rotation directions 86 about pivot axis 72.
- Vane 54 has a pressure side 90 and a suction side 92, wherein the pressure on pressure side 90 exceeds that of suction side 92.
- the pressure differential between pressure side 90 and suction side 92 may vary, e.g., depending upon vane 54 camber and engine operating conditions.
- the pressure differential between pressure side 90 and suction side 92 provides an impetus to flow from pressure side 90 to suction side 92, e.g., between airfoil portion 56 and airfoil portion 58.
- vanes 54 include a sealing arrangement 94 operative to seal between airfoil portion 56 and airfoil portion 58.
- Sealing arrangement 94 includes a seal strip 96 arranged to seal against fluid flow between airfoil portion 56 and airfoil portion 58 during the operation of engine 10, and to accommodate movement of one or both of airfoil portions 56 and 58, e.g., rotation of airfoil portion 58 about pivot axis 72, while sealing against fluid flow.
- seal strip 96 is a rigid structure that does not substantially deform in use or installation. In other embodiments, seal strip 96 may be a flexible structure. In one form, seal strip 96 is formed of a polymeric material, such as Vespel (commercially available from DuPont Engineering Polymers, located in Newark, Delaware, U.S.A.) and/or Torlon polyamide-imide (commercially available from Solvay Advanced
- seal strip 96 may be formed of other materials.
- seal strip 96 is disposed in a groove 98.
- groove 98 is disposed in a face 1 00 of airfoil portion 56 that faces airfoil portion 58.
- seal strip 96, groove 98 and face 100 extend between tip portion 60 and root portion 62 of airfoil portion 56.
- seal strip 96, groove 98 and/or face 100 may extend only partially between tip portion 60 and root portion 62.
- Face 100 is formed with a radius 02 centered on pivot axis 72. In one form, face 100 is formed integrally with airfoil portion 56.
- face 100 may be formed separately and affixed to airfoil portion 56.
- seal strip 96 is partially installed in groove 98, that is, leaving a portion 108 of seal strip 96 extending beyond face 100 of airfoil portion 56.
- Seal strip 96 has a width 104 greater than a width 106 of groove 98, and is installed into groove 98 with an interference fit, e.g., 0.001 -0.002 inch. The amount of interference may vary with the needs of the application.
- Airfoil portion 58 includes a crown 1 10 facing face 00 of airfoil portion 56. In one form, crown 1 10 is formed integrally with airfoil portion 58.
- crown 1 10 may be formed separately and affixed to airfoil portion 58. Crown 1 10 is formed with a radius 1 12 centered on pivot axis 72. In one form, crown 1 10 extends between tip portion 66 and root portion 68 of airfoil portion 58, and is positioned opposite groove 98. In other embodiments, crown 110 may extend only partially between tip portion 66 and root portion 68. In one form, face 100 of airfoil portion 56 is concave, and is operative to receive therein crown 1 10 opposite groove 98 in a nested arrangement. In other embodiments, face 100 may be convex.
- crown 1 10 of airfoil portion 58 is convex, and is operative to be received into face 100 in a nested arrangement.
- crown 1 10 may be convex, e.g., an inverted crown.
- the depicted embodiment includes groove 98 and seal strip 96 being located in face 100, in other embodiments, groove 98 and seal strip 96 may be located in crown 1 10.
- Seal strip 96 includes a rubbing surface 1 14.
- rubbing surface 1 14 is disposed opposite radius 1 12 of crown 1 10, and is operative to contact and seal against radius 1 12 of crown 110 of airfoil portion 58.
- rubbing surface 1 14 may rub against crown 1 10, e.g., until wear of seal strip 96 resulting from rotation of airfoil portion 58 reduces or eliminates contact between seal strip 96 and crown 1 10.
- rubbing surface 1 14 may be configured to be in close proximity to crown 1 10, but without any rubbing contact.
- seal strip 96 may be installed in crown 1 10, and rubbing surface 1 14 may be configured to seal against face 100.
- Rubbing surface 1 14 is preformed prior to installation into airfoil portion 56 , e.g., machined.
- rubbing surface 1 14 is configured as a radius 1 16 centered about pivot axis 72, e.g., the same radius as radius 1 12 of crown 1 0.
- radius 1 16 may be the same radius as radius 102 of face 100 or any other radius suitable for the application, in still other embodiments, other shapes for rubbing surface 1 14 may be employed.
- rubbing surface 1 14 is concave. In other embodiments, rubbing surface 1 14 may take other forms, and may be, for example, convex.
- Embodiments of the present invention include a variable camber vane system for a gas turbine engine, comprising: a first airfoil portion having a first tip portion, a first root portion, a face extending at least partially between the first tip portion and the first root portion, and a groove in the face extending at least partially between the first tip portion and the first root portion, wherein the groove has a groove width; a second airfoil portion arranged to rotate with respect to the first airfoil portion about a pivot axis, wherein the second airfoil portion includes a second tip portion; a second root portion; and a crown extending at least partially between the second tip portion and the second root portion, wherein the crown includes a crown radius centered about the pivot axis and positioned opposite the groove; and a seal strip having a seal width greater than the groove width and a rubbing surface opposite the crown radius, wherein the seal strip is at least partially disposed in the groove with an interference fit; and wherein the seal strip is arranged to seal against fluid
- the seal strip is a rigid structure.
- the rubbing surface has a rubbing surface radius the same as the crown radius.
- the crown is formed integrally with the second airfoil portion.
- the face is formed integrally with the first airfoil portion.
- the face is concave and operative to receive the crown therein.
- the first airfoil portion is stationary.
- first airfoil portion and the second airfoil portion form at least part of an inlet guide vane having a fixed leading edge and a variable trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
- first airfoil portion and the second airfoil portion form at least part of an outlet guide vane having a variable leading edge and a fixed trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
- Embodiments of the present invention include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a camber of the variable camber vane system , wherein a first of the airfoil portions includes a groove and a second of the airfoil portions includes a crown having a crown radius; and a seal strip at least partially disposed in the groove with an interference fit, wherein the seal strip includes a rubbing surface opposite the crown radius and operative to seal against fluid flow between the first of the airfoil portions and the second of the airfoil portions.
- the rubbing surface contacts the crown at the crown radius.
- the seal strip is formed of a polymer material.
- the seal strip is formed of at least one of Vespel and
- the at least two airfoil portions form an inlet guide vane.
- the at least two airfoil portions form an outlet guide vane.
- Embodiments include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a camber of the variable camber vane system, wherein a first of the airfoil portions includes a groove; and wherein a second of the airfoil portions includes a crown having a crown radius; and a seal strip disposed in the groove; wherein the seal strip has a rubbing surface radius preformed thereon and configured for sealing engagement with the crown.
- the seal strip is a rigid structure formed of a polymer.
- the crown radius is convex, and the rubbing surface radius is concave.
- the seal strip is fitted in the groove with an interference fit.
- the crown is nested within the first of the airfoil portions opposite the groove.
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- Engineering & Computer Science (AREA)
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- Structures Of Non-Positive Displacement Pumps (AREA)
Abstract
One embodiment of the present invention is a unique variable camber vane system for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and variable camber vane systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.
Description
GAS TURBINE ENGINE AND VARIABLE CAMBER VANE SYSTEM
Government Rights
The present application was made with the United States government support under Contract No. FA8650-07-C-2803, awarded by the United States Air Force. The United States government may have certain rights in the present application.
Field of the Invention
The present invention relates to gas turbine engines, and more particularly, to gas turbine engines with variable camber vane systems.
Background
Gas turbine engines with variable camber vane systems remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology.
Summary
One embodiment of the present invention is a unique variable camber vane system for a gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and variable camber vane systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.
Brief Description of the Drawings
The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein:
FIG. 1 schematically depicts some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention.
FIG. 2 schematically depicts some aspects of a non-limiting example of a fan system for a gas turbine engine in accordance with an embodiment of the present invention.
FIG. 3 depicts some aspects of a non-limiting example of a variable camber guide vane system in accordance with an embodiment of the present invention.
FIG. 4 depicts some aspects of the variable camber guide vane system of FIG. 3.
FIG. 5 depicts some aspects of a non-limiting example of a seal strip in accordance with an embodiment of the present invention.
Detailed Description
For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or
modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention.
Referring to the drawings, and in particular FIG. 1 , a non-limiting example of a gas turbine engine 10 in accordance with an embodiment of the present invention is depicted. In one form, gas turbine engine 10 is an aircraft propulsion power plant. In other embodiments, gas turbine engine 10 may be a land-based or marine engine. In one form, gas turbine engine 10 is a multi-spool turbofan engine. In other
embodiments, gas turbine engine 10 may be a single or multi-spool turbofan, turboshaft, turbojet, turboprop gas turbine or combined cycle engine.
Gas turbine engine 10 includes a fan system 12, a compressor system 14, a diffuser 16, a combustion system 18 and a turbine system 20. Compressor system 14 is in fluid communication with fan system 12. Diffuser 16 is in fluid communication with compressor system 14. Combustion system 18 is flu idly disposed between compressor system 14 and turbine system 20. Fan system 2 includes a fan rotor system 22. In
various embodiments, fan rotor system 22 includes one or more rotors (not shown) that are powered by turbine system 20. Compressor system 14 includes a compressor rotor system 24. In various embodiments, compressor rotor system 24 includes one or more rotors (not shown) that are powered by turbine system 20. Turbine system 20 includes a turbine rotor system 26. In various embodiments, turbine rotor system 26 includes one or more rotors (not shown) operative to drive fan rotor system 22 and compressor rotor system 24. Turbine rotor system 26 is driving coupled to compressor rotor system 24 and fan rotor system 22 via a shafting system 28. In various embodiments, shafting system 28 includes a plurality of shafts that may rotate at the same or different speeds and directions. In some embodiments, only a single shaft may be employed.
During the operation of gas turbine engine 10, air is drawn into the inlet of fan 12 and pressurized by fan 12. Some of the air pressurized by fan 12 is directed into compressor system 14, and the balance is directed into a bypass duct (not shown). Compressor system 14 further pressurizes the air received from fan 12, which is then discharged into diffuser 16. Diffuser 16 reduces the velocity of the pressurized air, and directs the diffused airflow into combustion system 18. Fuel is mixed with the pressurized air in combustion system 18, which is then combusted. In one form, combustion system 1 8 includes a combustion liner (not shown) that contains a continuous combustion process. In other embodiments, combustion system 18 may take other forms, and may be, for example, a wave rotor combustion system, a rotary valve combustion system, or a siinger combustion system, and may employ deflagration and/or detonation combustion processes. The hot gases exiting combustor 1 8 are directed into turbine system 20, which extracts energy in the form of mechanical shaft
power to drive fan system 12 and compressor system 14 via shafting system 28. The hot gases exiting turbine system 20 are directed into a nozzle (not shown), and provide a component of the thrust output by gas turbine engine 10.
Referring to FIG. 2, a non-limiting example of some aspects of fan system 12 in accordance with an embodiment of the present invention is schematically depicted. Fan system 12 includes a variable guide vane system 40 having a variable inlet guide vane stage 42 and a variable outlet guide vane stage 44 disposed on either side of a rotating fan stage 46. Variable inlet guide vane stage 42 is operative to guide air into rotating fan stage 46, and to selectively vary the incidence angle of the air flow entering rotating fan stage 46. Variable outlet guide vane stage 44 is operative to guide air exiting rotating fan stage 46, and to selectively vary the incidence angle of the air flow exiting rotating fan stage 46. Variable inlet guide vane stage 42 and variable outlet guide vane stage 44 are actuated by an actuation system (not shown). Although described herein as with respect to fan system 12, it will be understood that variable guide vane system 40 may also or alternatively be employed as part of compressor system 14. In addition, although variable guide vane system 40 includes both variable inlet and outlet guide vane stages, other embodiments may include only a variable inlet guide vane stage or a variable outlet guide vane stage.
Referring to FIGS. 3-5, a non-limiting example of some aspects of variable inlet guide vane stage 42 in accordance with an embodiment of the present invention is illustrated. It will be understood that some embodiments of variable outlet guide vane stage 44 may be similar to variable inlet guide vane stage 42, and hence, the following description of variable inlet guide vane stage 42 is also applicable to aspects of some
embodiments of variable outlet guide vane stage 44. Variable inlet guide vane stage 42 includes an outer band 50, an inner band 52 and plurality of vanes 54. Outer band 50 defines an outer flowpath wall of variable inlet guide vane stage 42. Inner band 52 defines an inner flowpath wall of variable inlet guide vane stage 42. Vanes 54 are airfoils that extend between outer band 50 and inner band 52, and are spaced apart circumferentially. in one form, vanes 54 extend in the radial direction between outer band 50 and inner band 52. In other embodiments, vanes 54 may extend between outer band 50 and inner band 52 at other angles.
Each vane 54 includes an airfoil portion 56 and an airfoil portion 58. Airfoil portion 56 extends between a tip portion 60 and a root portion 62. In one form, airfoil portion 56 includes the trailing edge 64 of vane 54. In other embodiments, airfoil portion 56 may be formed with a leading edge of vane 54 instead of trailing edge 64, e.g., for use in variable outlet guide vane 44. Airfoil portion 58 extends between a tip portion 66 and a root portion 68. In one form, airfoil portion 58 includes the leading edge 70 of vane 54. In other embodiments, airfoil portion 58 may be formed with a trailing edge instead of leading edge 70, e.g., for use in variable outlet guide vane 44. In one form, airfoil portion 56 is fixed, i.e., stationary. In other embodiments, airfoil portion 56 may be movable, e.g., pivotable about an axis so as to be able to vary the angle of the trailing edge of vane 54. In one form, airfoil portion 58 is variable, being configured to pivot about a pivot axis 72 with respect to airfoil portion 56, to provide a variable camber for vane 54. In other embodiments, airfoil portion 58 may be fixed. In one form, airfoil portion 58 is coupled to an actuation system (not shown) that is operative to selectively position airfoil portion 58 at a desired incidence angle. In other embodiments, airfoii
portion 56 may also or alternatively be coupled to an actuation system (not shown) that is operative to selectively position airfoil portion 56 at a desired incidence angle.
Extending from airfoil portion 58 are pivot shafts 74 and 76, which establish pivot axis 72. Outer band 50 includes a plurality of spaced apart openings 78. Inner band 52 includes a plurality of spaced apart openings 80. Openings 78 and 80 are operative to receive pivot shafts 74 and 76, respectively, and retain airfoil portions 58 in the engine axial, circumferential and radial direction. In one form, pivot shafts 74 and 76 retain airfoil portion 58 in outer band 50 and inner band 52 via anti-friction bushings 82 and 84. Anti-friction bushings 82 and 84 are operative to provide bearing surfaces for pivot shafts 74 and 76. Other embodiments may not include anti-friction bushings 82 and 84. Airfoil portion 58 is operative to rotate in rotation directions 86 about pivot axis 72.
During the operation of engine 10, air flows past vanes 54 in the general direction illustrated as direction 88. Vane 54 has a pressure side 90 and a suction side 92, wherein the pressure on pressure side 90 exceeds that of suction side 92. The pressure differential between pressure side 90 and suction side 92 may vary, e.g., depending upon vane 54 camber and engine operating conditions. The pressure differential between pressure side 90 and suction side 92 provides an impetus to flow from pressure side 90 to suction side 92, e.g., between airfoil portion 56 and airfoil portion 58. It is desirable to reduce or prevent leakage between airfoil portion 56 and airfoil portion 58, e.g., leakage flow from pressure side 90 to suction side 92, e.g., in order to improve fan 12 and engine 10 efficiency. Accordingly, vanes 54 include a sealing arrangement 94 operative to seal between airfoil portion 56 and airfoil portion 58. Sealing arrangement 94 includes a seal strip 96 arranged to seal against fluid flow
between airfoil portion 56 and airfoil portion 58 during the operation of engine 10, and to accommodate movement of one or both of airfoil portions 56 and 58, e.g., rotation of airfoil portion 58 about pivot axis 72, while sealing against fluid flow.
In one form, seal strip 96 is a rigid structure that does not substantially deform in use or installation. In other embodiments, seal strip 96 may be a flexible structure. In one form, seal strip 96 is formed of a polymeric material, such as Vespel (commercially available from DuPont Engineering Polymers, located in Newark, Delaware, U.S.A.) and/or Torlon polyamide-imide (commercially available from Solvay Advanced
Polymers, located in Alpharetta, Georgia, U.S.A.). In other embodiments, seal strip 96 may be formed of other materials. In one form, seal strip 96 is disposed in a groove 98. In one form, groove 98 is disposed in a face 1 00 of airfoil portion 56 that faces airfoil portion 58. In one form, seal strip 96, groove 98 and face 100 extend between tip portion 60 and root portion 62 of airfoil portion 56. In other embodiments, seal strip 96, groove 98 and/or face 100 may extend only partially between tip portion 60 and root portion 62. Face 100 is formed with a radius 02 centered on pivot axis 72. In one form, face 100 is formed integrally with airfoil portion 56. In other embodiments, face 100 may be formed separately and affixed to airfoil portion 56. In one form, seal strip 96 is partially installed in groove 98, that is, leaving a portion 108 of seal strip 96 extending beyond face 100 of airfoil portion 56. Seal strip 96 has a width 104 greater than a width 106 of groove 98, and is installed into groove 98 with an interference fit, e.g., 0.001 -0.002 inch. The amount of interference may vary with the needs of the application.
Airfoil portion 58 includes a crown 1 10 facing face 00 of airfoil portion 56. In one form, crown 1 10 is formed integrally with airfoil portion 58. In other embodiments, crown 1 10 may be formed separately and affixed to airfoil portion 58. Crown 1 10 is formed with a radius 1 12 centered on pivot axis 72. In one form, crown 1 10 extends between tip portion 66 and root portion 68 of airfoil portion 58, and is positioned opposite groove 98. In other embodiments, crown 110 may extend only partially between tip portion 66 and root portion 68. In one form, face 100 of airfoil portion 56 is concave, and is operative to receive therein crown 1 10 opposite groove 98 in a nested arrangement. In other embodiments, face 100 may be convex. In one form, crown 1 10 of airfoil portion 58 is convex, and is operative to be received into face 100 in a nested arrangement. In other embodiments, crown 1 10 may be convex, e.g., an inverted crown. Although the depicted embodiment includes groove 98 and seal strip 96 being located in face 100, in other embodiments, groove 98 and seal strip 96 may be located in crown 1 10.
Seal strip 96 includes a rubbing surface 1 14. In one form, rubbing surface 1 14 is disposed opposite radius 1 12 of crown 1 10, and is operative to contact and seal against radius 1 12 of crown 110 of airfoil portion 58. During movement of airfoil portion 58, e.g., when changing the camber of vane 54 by rotating airfoil portion 58 about pivot axis 72, rubbing surface 1 14 may rub against crown 1 10, e.g., until wear of seal strip 96 resulting from rotation of airfoil portion 58 reduces or eliminates contact between seal strip 96 and crown 1 10. In other embodiments, rubbing surface 1 14 may be configured to be in close proximity to crown 1 10, but without any rubbing contact. In still other
embodiments, seal strip 96 may be installed in crown 1 10, and rubbing surface 1 14 may be configured to seal against face 100.
Rubbing surface 1 14 is preformed prior to installation into airfoil portion 56 , e.g., machined. In one form, rubbing surface 1 14 is configured as a radius 1 16 centered about pivot axis 72, e.g., the same radius as radius 1 12 of crown 1 0. In other embodiments, radius 1 16 may be the same radius as radius 102 of face 100 or any other radius suitable for the application, in still other embodiments, other shapes for rubbing surface 1 14 may be employed. In one form, rubbing surface 1 14 is concave. In other embodiments, rubbing surface 1 14 may take other forms, and may be, for example, convex.
Embodiments of the present invention include a variable camber vane system for a gas turbine engine, comprising: a first airfoil portion having a first tip portion, a first root portion, a face extending at least partially between the first tip portion and the first root portion, and a groove in the face extending at least partially between the first tip portion and the first root portion, wherein the groove has a groove width; a second airfoil portion arranged to rotate with respect to the first airfoil portion about a pivot axis, wherein the second airfoil portion includes a second tip portion; a second root portion; and a crown extending at least partially between the second tip portion and the second root portion, wherein the crown includes a crown radius centered about the pivot axis and positioned opposite the groove; and a seal strip having a seal width greater than the groove width and a rubbing surface opposite the crown radius, wherein the seal strip is at least partially disposed in the groove with an interference fit; and wherein the seal
strip is arranged to seal against fluid flow between the first airfoil portion and the second airfoil portion.
In a refinement, the seal strip is a rigid structure.
In another refinement, the rubbing surface has a rubbing surface radius the same as the crown radius.
In yet another refinement, the crown is formed integrally with the second airfoil portion.
In still another refinement, the face is formed integrally with the first airfoil portion.
In yet still another refinement, the face is concave and operative to receive the crown therein.
In a further refinement, the first airfoil portion is stationary.
In a yet further refinement, the first airfoil portion and the second airfoil portion form at least part of an inlet guide vane having a fixed leading edge and a variable trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
In a still further refinement, the first airfoil portion and the second airfoil portion form at least part of an outlet guide vane having a variable leading edge and a fixed trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
Embodiments of the present invention include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a
camber of the variable camber vane system , wherein a first of the airfoil portions includes a groove and a second of the airfoil portions includes a crown having a crown radius; and a seal strip at least partially disposed in the groove with an interference fit, wherein the seal strip includes a rubbing surface opposite the crown radius and operative to seal against fluid flow between the first of the airfoil portions and the second of the airfoil portions.
In a refinement, the rubbing surface contacts the crown at the crown radius.
In another refinement, the seal strip is formed of a polymer material.
In yet another refinement, the seal strip is formed of at least one of Vespel and
Torlon.
In still another refinement, the at least two airfoil portions form an inlet guide vane.
In a further refinement, the at least two airfoil portions form an outlet guide vane.
Embodiments include a gas turbine engine, comprising: at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including: at least two airfoil portions adapted to vary a camber of the variable camber vane system, wherein a first of the airfoil portions includes a groove; and wherein a second of the airfoil portions includes a crown having a crown radius; and a seal strip disposed in the groove; wherein the seal strip has a rubbing surface radius preformed thereon and configured for sealing engagement with the crown.
In a refinement, the seal strip is a rigid structure formed of a polymer.
In another refinement, the crown radius is convex, and the rubbing surface radius is concave.
In yet another refinement, the seal strip is fitted in the groove with an interference fit.
In still another refinement, the crown is nested within the first of the airfoil portions opposite the groove.
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as "a," "an," "at least one" and "at least a portion" are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language "at least a portion" and/or "a portion" is used the item may include a portion and/or the entire item unless specifically stated to the contrary.
Claims
1 . A variable camber vane system for a gas turbine engine, comprising:
a first airfoil portion having a first tip portion, a first root portion, a face extending at least partially between the first tip portion and the first root portion, and a groove in the face extending at least partially between the first tip portion and the first root portion, wherein the groove has a groove width;
a second airfoil portion arranged to rotate with respect to the first airfoil portion about a pivot axis, wherein the second airfoil portion includes a second tip portion; a second root portion; and a crown extending at least partially between the second tip portion and the second root portion, wherein the crown includes a crown radius centered about the pivot axis and positioned opposite the groove; and
a seal strip having a seal width greater than the groove width and a rubbing surface opposite the crown radius, wherein the seal strip is at least partially disposed in the groove with an interference fit; and wherein the seal strip is arranged to seal against fluid flow between the first airfoil portion and the second airfoil portion.
2. The variable camber vane system of claim 1 , wherein the seal strip is a rigid structure.
3. The variable camber vane system of claim 1 , wherein the rubbing surface has a rubbing surface radius the same as the crown radius.
4. The variable camber vane system of claim 1 , wherein the crown is formed integrally with the second airfoil portion.
5. The variable camber vane system of claim 1 , wherein the face is formed integrally with the first airfoil portion.
6. The variable camber vane system of claim 1 , wherein the face is concave and operative to receive the crown therein.
7. The variable camber vane system of claim 1 , wherein the first airfoil portion is stationary.
8. The variable camber vane system of claim 7, wherein the first airfoil portion and the second airfoil portion form at least part of an inlet guide vane having a fixed leading edge and a variable trailing edge; wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
9. The variable camber vane system of claim 7, wherein the first airfoil portion and the second airfoil portion form at least part of an outlet guide vane having a variable leading edge and a fixed trailing edge: wherein the first airfoil portion includes the leading edge; and wherein the second airfoil portion includes the trailing edge.
A gas turbine engine, comprising at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including:
at least two airfoil portions adapted to vary a camber of the variable camber vane system, wherein a first of the airfoil portions includes a groove and a second of the airfoil portions includes a crown having a crown radius; and a seal strip at least partially disposed in the groove with an interference fit, wherein the seal strip includes a rubbing surface opposite the crown radius and operative to seal against fluid flow between the first of the airfoil portions and the second of the airfoil portions.
1 1 . The gas turbine engine of claim 1 0, wherein the rubbing surface contacts the crown at the crown radius.
12. The gas turbine engine of claim 10, wherein the seal strip is formed of a polymer material .
13. The gas turbine engine of claim 12, wherein the seal strip is formed of at least one of Vespel and Torlon.
14. The gas turbine engine of claim 10, wherein the at least two airfoil portions form an inlet guide vane.
15. The gas turbine engine of claim 1 0, wherein the at least two airfoil portions form an outlet guide vane.
16. A gas turbine engine, comprising :
at least one of a fan and a compressor having a variable camber vane system, the variable camber vane system including :
at least two airfoil portions adapted to vary a camber of the variable camber vane system, wherein a first of the airfoil portions includes a groove; and wherein a second of the airfoil portions includes a crown having a crown radius; and
a seal strip disposed in the groove; wherein the seal strip has a rubbing surface radius preformed thereon and configured for sealing engagement with the crown.
17. The gas turbine engine of claim 16, wherein the seal strip is a rigid structure formed of a polymer.
18. The gas turbine engine of claim 16, wherein the crown radius is convex, and wherein the rubbing surface radius is concave.
19. The gas turbine engine of claim 16, wherein the seal strip is fitted in the groove with an interference fit.
20. The gas turbine engine of claim 16, wherein the crown is nested within the first of the airfoil portions opposite the groove.
Priority Applications (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
EP11853198.7A EP2659112B1 (en) | 2010-12-27 | 2011-12-27 | Gas turbine engine and variable camber vane system |
CA2822965A CA2822965C (en) | 2010-12-27 | 2011-12-27 | Gas turbine engine and variable camber vane system |
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US12/978,843 | 2010-12-27 | ||
US12/978,843 US20120163960A1 (en) | 2010-12-27 | 2010-12-27 | Gas turbine engine and variable camber vane system |
Publications (1)
Publication Number | Publication Date |
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WO2012092277A1 true WO2012092277A1 (en) | 2012-07-05 |
Family
ID=46317012
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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PCT/US2011/067393 WO2012092277A1 (en) | 2010-12-27 | 2011-12-27 | Gas turbine engine and variable camber vane system |
Country Status (4)
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US (1) | US20120163960A1 (en) |
EP (1) | EP2659112B1 (en) |
CA (1) | CA2822965C (en) |
WO (1) | WO2012092277A1 (en) |
Families Citing this family (19)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20140064955A1 (en) * | 2011-09-14 | 2014-03-06 | General Electric Company | Guide vane assembly for a gas turbine engine |
EP2905425B1 (en) * | 2014-02-06 | 2019-07-10 | United Technologies Corporation | Variable vane and seal arrangement |
US9617864B2 (en) * | 2014-07-21 | 2017-04-11 | United Technologies Corporation | Seal assembly for a guide vane assembly |
CA2959993A1 (en) * | 2014-09-12 | 2016-03-17 | General Electric Company | Axi-centrifugal compressor with variable outlet guide vanes |
US9879560B2 (en) * | 2015-05-15 | 2018-01-30 | United Technologies Corporation | Vane strut positioning and securing systems |
US10704418B2 (en) | 2016-08-11 | 2020-07-07 | General Electric Company | Inlet assembly for an aircraft aft fan |
US10252790B2 (en) | 2016-08-11 | 2019-04-09 | General Electric Company | Inlet assembly for an aircraft aft fan |
US10259565B2 (en) | 2016-08-11 | 2019-04-16 | General Electric Company | Inlet assembly for an aircraft aft fan |
US10253779B2 (en) | 2016-08-11 | 2019-04-09 | General Electric Company | Inlet guide vane assembly for reducing airflow swirl distortion of an aircraft aft fan |
US10662782B2 (en) * | 2016-11-17 | 2020-05-26 | Raytheon Technologies Corporation | Airfoil with airfoil piece having axial seal |
US10794200B2 (en) | 2018-09-14 | 2020-10-06 | United Technologies Corporation | Integral half vane, ringcase, and id shroud |
US10781707B2 (en) | 2018-09-14 | 2020-09-22 | United Technologies Corporation | Integral half vane, ringcase, and id shroud |
CN111102012B (en) * | 2018-10-25 | 2024-07-09 | 中国科学院工程热物理研究所 | Blade adopting self-adaptive coanda jet and manufacturing method |
US11174916B2 (en) | 2019-03-21 | 2021-11-16 | Pratt & Whitney Canada Corp. | Aircraft engine reduction gearbox |
US11268453B1 (en) | 2021-03-17 | 2022-03-08 | Pratt & Whitney Canada Corp. | Lubrication system for aircraft engine reduction gearbox |
US11802490B2 (en) | 2021-08-25 | 2023-10-31 | Rolls-Royce Corporation | Controllable variable fan outlet guide vanes |
US11788429B2 (en) | 2021-08-25 | 2023-10-17 | Rolls-Royce Corporation | Variable tandem fan outlet guide vanes |
US11686211B2 (en) | 2021-08-25 | 2023-06-27 | Rolls-Royce Corporation | Variable outlet guide vanes |
US11879343B2 (en) | 2021-08-25 | 2024-01-23 | Rolls-Royce Corporation | Systems for controlling variable outlet guide vanes |
Citations (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3990810A (en) | 1975-12-23 | 1976-11-09 | Westinghouse Electric Corporation | Vane assembly for close coupling the compressor turbine and a single stage power turbine of a two-shaped gas turbine |
GB1504463A (en) | 1975-10-14 | 1978-03-22 | Westinghouse Canada Ltd | Adjustable vane assembly for a gas turbine engine |
EP0924389A2 (en) | 1997-12-18 | 1999-06-23 | United Technologies Corporation | Variable inlet guide vane for a gas turbine engine |
DE10352789A1 (en) | 2003-11-12 | 2005-06-09 | Mtu Aero Engines Gmbh | Gas turbine for use in aircraft, has sealing device to seal gap between pre-guiding units and pre-guiding vanes, where vanes and units are integrated in turbine and/or compressor of turbine |
US20070119150A1 (en) * | 2005-11-29 | 2007-05-31 | Wood Peter J | Turbofan gas turbine engine with variable fan outlet guide vanes |
Family Cites Families (2)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US3887297A (en) * | 1974-06-25 | 1975-06-03 | United Aircraft Corp | Variable leading edge stator vane assembly |
US8632302B2 (en) * | 2009-12-07 | 2014-01-21 | Dresser-Rand Company | Compressor performance adjustment system |
-
2010
- 2010-12-27 US US12/978,843 patent/US20120163960A1/en not_active Abandoned
-
2011
- 2011-12-27 EP EP11853198.7A patent/EP2659112B1/en active Active
- 2011-12-27 CA CA2822965A patent/CA2822965C/en active Active
- 2011-12-27 WO PCT/US2011/067393 patent/WO2012092277A1/en active Application Filing
Patent Citations (7)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
GB1504463A (en) | 1975-10-14 | 1978-03-22 | Westinghouse Canada Ltd | Adjustable vane assembly for a gas turbine engine |
US4097187A (en) * | 1975-10-14 | 1978-06-27 | Westinghouse Canada Limited | Adjustable vane assembly for a gas turbine |
US3990810A (en) | 1975-12-23 | 1976-11-09 | Westinghouse Electric Corporation | Vane assembly for close coupling the compressor turbine and a single stage power turbine of a two-shaped gas turbine |
EP0924389A2 (en) | 1997-12-18 | 1999-06-23 | United Technologies Corporation | Variable inlet guide vane for a gas turbine engine |
US6045325A (en) * | 1997-12-18 | 2000-04-04 | United Technologies Corporation | Apparatus for minimizing inlet airflow turbulence in a gas turbine engine |
DE10352789A1 (en) | 2003-11-12 | 2005-06-09 | Mtu Aero Engines Gmbh | Gas turbine for use in aircraft, has sealing device to seal gap between pre-guiding units and pre-guiding vanes, where vanes and units are integrated in turbine and/or compressor of turbine |
US20070119150A1 (en) * | 2005-11-29 | 2007-05-31 | Wood Peter J | Turbofan gas turbine engine with variable fan outlet guide vanes |
Non-Patent Citations (1)
Title |
---|
See also references of EP2659112A4 |
Also Published As
Publication number | Publication date |
---|---|
EP2659112A4 (en) | 2018-03-07 |
CA2822965A1 (en) | 2012-07-05 |
CA2822965C (en) | 2020-02-11 |
EP2659112A1 (en) | 2013-11-06 |
US20120163960A1 (en) | 2012-06-28 |
EP2659112B1 (en) | 2020-10-07 |
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