US20140064955A1

US20140064955A1 - Guide vane assembly for a gas turbine engine

Info

Publication number: US20140064955A1
Application number: US13/607,147
Authority: US
Inventors: Thurmond Douglas Senter; Mark Joseph Mielke; Alan Glen Turner; Thomas Becker
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2011-09-14
Filing date: 2012-09-07
Publication date: 2014-03-06
Also published as: CA2847566A1; WO2013081714A3; WO2013081714A2; EP2756182A2

Abstract

A guide vane assembly for a gas turbine engine includes: a strut having: a root, a tip, a leading edge, a trailing edge wall, and opposed sidewalls extending between the leading edge and the trailing edge wall; and a pair of spaced-apart prongs extending axially aft from the trailing edge wall, each prong having an outer wall that is flush with a respective one of the sidewalls; and a flap positioned axially aft of the strut, the flap having a root, a tip, a leading edge, a trailing edge, and opposed sides extending between the leading and trailing edge, wherein a portion of the leading edge is disposed between the prongs, and wherein a disk-like flap button is disposed at the tip of the flap adjacent its leading edge, the flap button including a pair of spaced-apart notches formed therein.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of Provisional Application No. 61/534,826, Filed Sep. 14, 2011.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH AND DEVELOPMENT

The U.S. Government may have certain rights in this invention pursuant to contract number FA8650-07-C-2802 awarded by the Department of the Air Force.

BACKGROUND OF THE INVENTION

This invention relates generally to variable inlet guide vanes for gas turbine engines and more particularly to the mounting and configuration of such vanes.
Some gas turbine engines include stationary inlet guide vanes (“IGVs”) positioned in the inlet upstream of the fan or compressor. In many military gas turbine engines, the IGVs are configured as fan inlet strut/flap assemblies. The struts house various service leads such as oil tubes, and the flaps, which are positioned directly downstream of the struts, direct inlet flow to the downstream fan or compressor. The flaps are pivoted and have a variable effective angle in order to throttle mass flow through the engine as needed in different operating conditions.
One primary challenge with strut/flap design is optimizing the design to minimize aeromechanical stimulus of the downstream rotor. While the flap aerodynamic design can be created with aeromechanic considerations in mind, current state of the art design approaches result in several features necessary for assembly that exacerbate aeromechanics problems and reduce efficiency. In some applications, the assembly problems are compounded by the presence of an additional bypass stream from a fan-on-blade stage or “FLADE” stream. For example, a notch is frequently formed in the trailing edge of the strut at its tip, which promotes leakage between the strut and flap. Also, a typical prior art aerodynamic approach is to minimize the thickness of the strut and flap airfoils as much as possible. However, this results in a large discrepancy between the diameter of the flap leading edge and an adjacent flap button.
Accordingly, there is a need for a flap/strut assembly which minimizes aeromechanical effects while accommodating assembly procedures.

BRIEF SUMMARY OF THE INVENTION

This need is addressed by the present invention, which provides an inlet strut/flap assembly configured with closely fitting mechanical features in combination with a flap button having notches that permit assembly.
According to one aspect of the invention, a guide vane assembly for a gas turbine engine includes: a strut having: a root, a tip, a leading edge, a trailing edge wall, and opposed sidewalls extending between the leading edge and the trailing edge wall; and a pair of spaced-apart prongs extending axially aft from the trailing edge wall, each prong having an outer wall that is flush with a respective one of the sidewalls; and a flap positioned axially aft of the strut, the flap having a root, a tip, a leading edge, a trailing edge, and opposed sides extending between the leading and trailing edge, wherein a portion of the leading edge is disposed between the prongs, and wherein a disk-like flap button is disposed at the tip of the flap adjacent its leading edge, the flap button including a pair of spaced-apart notches formed therein.
According to another aspect of the invention, a guide vane assembly for a gas turbine engine includes: an array of struts, each strut including: a root, a tip, a leading edge, a trailing edge wall, and opposed sidewalls extending between the leading edge and the trailing edge wall; and a pair of spaced-apart prongs extending axially aft from the trailing edge wall, each prong having an outer wall that is flush with a respective one of the sidewalls; an arcuate shroud surrounding the tips of the struts; an array of strut extensions extending radially outward from the shroud; an outer band surrounding the strut extensions; and a flap positioned axially aft of each strut, each flap having a root, a tip, a leading edge, a trailing edge, and opposed sides extending between the leading and trailing edge, wherein a portion of the leading edge is disposed between the prongs of the respective strut, and wherein a disk-like flap button is disposed at the tip of the flap adjacent its leading edge, the flap button including a pair of spaced-apart notches formed therein; and a ring of casing segments surrounding the flaps, each casing segment including an array of stationary outer flaps extending between annular inner and outer bands, each outer flap being aligned with a respective one of the strut extensions.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine including a guide vane assembly constructed according to the present invention;

FIG. 2 is an enlarged, partially-sectioned side view of a portion of a guide vane assembly;

FIG. 3 is a top plan view of a flap of the guide vane assembly of FIG. 2;

FIG. 4 is a partial side view of the guide flap of FIG. 3;

FIG. 5 is a schematic cross-sectional view of a guide vane assembly; and

FIG. 6 is an enlarged view of a portion of FIG. 5.

DETAILED DESCRIPTION OF THE INVENTION

Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, FIG. 1 illustrates a portion of an exemplary gas turbine engine, generally designated 10. The engine 10 has a longitudinal center line or axis A and an outer stationary annular casing 12 disposed concentrically about and coaxially along the axis A. The engine 10 has a fan 14, compressor 16, combustor 18, high pressure turbine 20, and low pressure turbine 22 arranged in serial flow relationship. In operation, pressurized air from the compressor 16 is mixed with fuel in the combustor 18 and ignited, thereby generating pressurized combustion gases. Some work is extracted from these gases by the high pressure turbine 20 which drives the compressor 16 via an outer shaft 24. The combustion gases then flow into the low pressure turbine 22, which drives the fan 14 via an inner shaft 26. The fan 14, inner shaft 26, and low pressure turbine 22 are collectively considered portions of a “low pressure spool” or “LP spool” (not labeled in the figures).
A portion of the fan discharge flows through the compressor 16, combustor 18, and high-pressure turbine 20, which are collectively referred to as the “core” 28 of the engine 10. Another portion of the fan discharge flows through an annular bypass duct 30 which surrounds the core 28. The illustrated fan 14 includes, in flow sequence, a row of non-rotating fan guide vane assemblies 32 (described in greater detail below), a first stage of rotating fan blades 34, a row of non-rotating interstage vanes 36, and a second stage of rotating fan blades 38.
As described in more detail below, the guide vane assemblies 32 may have their angle of attack with respect to the airflow and their open flow area selectively changed by using an actuator 40 of a known type.
The engine 10 also includes a supplementary fan, referred to as a “FLADE” stage 44 (FLADE being an acronym for “fan on blade”), in the form of a ring of airfoils extending radially outwardly from an annular shroud 46 and driven by the fan 14 (in this case the second stage 36). The FLADE stage 44 is positioned in a fan outer duct 48 which surrounds the bypass duct 30. The FLADE stage 44 provides an additional flow stream at a different flow and pressure ratio that than of the fan 14. The FLADE stage flow is sized to provide sufficient bleed air pressure and flow for a selected aircraft bleed-air powered system of a known type (not shown). A row of variable-angle FLADE inlet guide vanes 50, operated by an actuator 52, are moveable between open and closed positions to vary the flow through the FLADE stage 44.
The fan outer duct 48 includes one or more bleed air outlets 54 which direct flow to the aircraft bleed air system. Bleed air valves 56 may also be provided to selectively close off the bleed air outlets 54 and direct the FLADE stage flow downstream through the fan outer duct 48.
An exhaust duct 58 is disposed downstream of the core 28, and receives the mixed air flow from both the core 28 and the bypass duct 30. A mixer 60 (for example a lobed or chute-type mixer) is disposed at the juncture of the core 28 and bypass duct 30 flow streams to promote efficient mixing of the two streams.
In operation, the engine 10 generates thrust for aircraft propulsion in a known manner, while the FLADE stage discharges bleed air flow through the bleed air outlets 54.
Each guide vane assembly 32 includes both a streamlined strut 62 and an airfoil-shaped flap 64. The struts 62 are spaced in an array about the circumference of a fan hub 66 (shown in FIG. 1) and the casing 12. The struts 62 are stationary and their interior may house various service leads such as oil tubes (not shown). The struts 62 are shaped to direct air entering the engine inlet around strut 62 towards the flap 64. Each strut 62 has a root 68, a tip 70, and (see FIGS. 5 and 6) a pair of sidewalls 72 and 74 that are connected at a leading edge 76 and an trailing edge wall 78. A pair of spaced apart walls extend axially aft from the trailing edge wall 78. These walls are referred to herein as “prongs” 80. Each prong 80 includes an outer wall 82 that is flush with a respective one of the sidewalls 72 and 74, an inner wall 84 parallel to the outer wall 82, and an aft wall 86 interconnecting the inner and outer walls 82 and 84. The aft walls 86 are angled and face partially inwards toward each other, so as to approximate a tangent to the surface of the adjacent flap 64.
Referring to FIG. 2, an annular shroud segment 87 is disposed outboard of the tip 70 of each strut 62. Collectively the shroud segments 87 of the adjacent struts 62 form a continuous ring. A strut extension 89 having a cross-sectional shape identical or similar to that of the strut 62 extends radially outward from the radially outer surface of each shroud segment 87. The strut extensions 89 are circumscribed by an outer band 91 which forms part of an annular FLADE duct 93 (see FIG. 1) that channels inlet air to the FLADE stage 44.
As seen in FIGS. 3 and 4, each flap 64 includes a pair of sidewalls 88 and 90 connected at a leading edge 92 and trailing edge 94. Each sidewall 88 and 90 extends in radial span between a root 95 (FIG. 1) and a tip 96. In cross-section the shape defined between the sidewalls 88 and 90 is generally an airfoil shape.
A disk-like flap button 102 is positioned at the tip 96, near the leading edge 92. The majority of the flap button 102 is generally circular in plan view, and the outer diameter of the flap button 102 is only slightly greater than the thickness of the airfoil portion of the flap 64. A pair of notches 104 are formed in the perimeter of the flap button 102.
It is noted that a typical aerodynamic approach in the prior art is to minimize the thickness of the flap as much as possible. However, this results in a large discrepancy between the diameter of the flap leading edge and an adjacent flap button The flap 64 described herein uses a relatively much thicker flap leading edge diameter in order to bring the button and flap diameters more in-line. The strut thickness is also increased accordingly.
A circular cross-section trunnion 106 extends radially outward from the flap button 102. A parallel-sided lug 108 is disposed near the outer end of the trunnion 106. A threaded stud 109 extends radially outward from the lug 108.
Referring to FIG. 2, a ring of casing segments 110 surrounds the flaps 64. Each casing segment 110 includes a stationary outer flap 112 having a cross-sectional shape identical or similar to that of the flap 64 and extending between an arcuate inner band segment 114 and an arcuate outer band segment 116. Each outer flap 112 is aligned with one of the upstream strut extensions 89. Collectively the inner and outer band segments 114 and 116 of the adjacent casing segments 110 form a continuous ring. The casing segments 110 form part of the FLADE duct 93. The casing segments 110 may be partially or wholly integral with the struts 62 and or the strut extensions 89.
Each outer flap 112 receives a bushing 118. The bushing 118 is a generally cylindrical structure with an annular flange 120 disposed at its radially outer end that abuts the exterior surface of the outer band segment 116. The bushing 118 may be press fit, threaded, or otherwise rigidly secured to the casing segment 116. The bushing 118 may be constructed of any material that will bear the operating loads on the trunnion 106 and allow it to rotate freely, and may be metallic or nonmetallic. Collectively each bushing 118 and the casing segment 110 it is installed in define a recess 122 shaped and sized to receive one of the flap buttons 102 described above. The bushing 118 is effective to react gasloads from the flap 64 transmitted through the trunnion 106, and the recess 122 restrains the flap 64 in the radial direction. This allows the flaps 64 to be selectively positionable during engine operation.
The flap 64 is assembled by inserting it radially outwardly so the trunnion 12 enters the bore of the respective bushing 118. During assembly, the flap 64 is pivoted to a position so that the notches 104 in the flap button 106 clear the prongs 80 of the upstream strut 62. Once the flap button 106 is in position in the recess 122, it will lie radially outboard of the prongs 80 and is free to pivot with the trunnion 102. An actuator arm 124 is then placed over the lug 108, and a nut 126 installed to the stud 109 to complete assembly. The actuator arm 124 is coupled to the actuator 40 depicted schematically in FIG. 1.
FIGS. 5 and 6 illustrate the assembled relationship of the strut 62 and the flap 64 in more detail. The prongs 80 extend in close proximity to a nose portion 128 of the flap 64, minimizing the gap between the two. Furthermore, the nose portion 128 of the flap 64 is shaped to minimize leakage between the strut 62 and the flap 64 in all positions of the flap 64. In particular, the nose portion 128 of the flap has a leading edge profile with two distinct profiles. One portion 130 has a circular cross-sectional shape with a radius “R”. The adjacent portion 132 is a noncircular cross-sectional airfoil shape.
The guide vane assembly described herein has several advantages over prior art configurations. Most of the mechanical loads are reacted out through the trunnion 102 itself, which spans through the FLADE duct 93 and eliminates the need for a control tube. Furthermore, eliminating or significantly reducing the notch found in prior art struts results in significantly less leakage flow and therefore better aerodynamics (lower loss) and less aeromechanical stimulus. This is especially important when the flap must turn down at high speed. The improved aerodynamics and reduced aeromechanics stimulus results in a more efficient engine with less inclination to rotor vibratory modes. Since the strut/flap is the very first component in the engine, improving its loss incrementally cascades through the entire engine and results in an even larger benefit to overall performance. Commercial advantages would manifest themselves via improved SFC and longer time on wing.
The foregoing has described a guide vane assembly for a gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation, the invention being defined by the claims.

Claims

What is claimed is:

1. A guide vane assembly for a gas turbine engine, comprising:

a strut having:

a root, a tip, a leading edge, a trailing edge wall, and opposed sidewalls extending between the leading edge and the trailing edge wall; and

a pair of spaced-apart prongs extending axially aft from the trailing edge wall, each prong having an outer wall that is flush with a respective one of the sidewalls;

a flap positioned axially aft of the strut, the flap having a root, a tip, a leading edge, a trailing edge, and opposed sides extending between the leading and trailing edge, wherein a portion of the leading edge is disposed between the prongs, and wherein a disk-like flap button is disposed at the tip of the flap adjacent its leading edge, the flap button including a pair of spaced-apart notches formed therein.

2. The guide vane assembly of claim 1 wherein the flap includes a trunnion extending radially outward from the flap button.

3. The guide vane assembly of claim 2 wherein a lug with parallel sides is disposed at an outer end of the trunnion.

4. The guide vane assembly of claim 2 wherein a threaded stud extends radially outward from an outer end of the trunnion.

5. The guide vane assembly of claim 1 wherein each prong includes: an outer wall that is flush with a respective one of the sidewalls, an inner wall parallel to the outer wall, and an aft wall interconnecting the inner and outer walls.

6. The guide vane assembly of claim 1 wherein the flap includes a nose portion adjacent the leading edge, wherein one part of the nose portion has a circular cross-sectional shape, and another part of the nose portion has a noncircular cross-sectional shape.

7. A guide vane assembly for a gas turbine engine, comprising:

an array of struts, each strut including:

an arcuate shroud surrounding the tips of the struts;

an array of strut extensions extending radially outward from the shroud;

an outer band surrounding the strut extensions;

a flap positioned axially aft of each strut, each flap having a root, a tip, a leading edge, a trailing edge, and opposed sides extending between the leading and trailing edge, wherein a portion of the leading edge is disposed between the prongs of the respective strut, and wherein a disk-like flap button is disposed at the tip of the flap adjacent its leading edge, the flap button including a pair of spaced-apart notches formed therein; and

a ring of casing segments surrounding the flaps, each casing segment including an array of stationary outer flaps extending between annular inner and outer bands, each outer flap being aligned with a respective one of the strut extensions.

8. The guide vane assembly of claim 7 wherein each flap includes a trunnion extending radially outward from the flap button.

9. The guide vane assembly of claim 8 wherein each trunnion is received in the bore of a generally cylindrical trunnion carried by one of the outer flaps.

10. The guide vane assembly of claim 9 wherein each bushing cooperatively with an outer flap defines a recess which receives a respective one of the flap buttons.

11. The guide vane assembly of claim 8 wherein a lug with parallel sides is disposed at an outer end of the trunnion.

12. The guide vane assembly of claim 8 wherein a threaded stud extends radially outward from an outer end of each trunnion.

13. The guide vane assembly of claim 7 wherein each prong includes: an outer wall that is flush with a respective one of the sidewalls, an inner wall parallel to the outer wall, and an aft wall interconnecting the inner and outer walls.

14. The guide vane assembly of claim 13 wherein the aft walls are angled and face partially inwards toward each other.

15. The guide vane assembly of claim 7 wherein each flap includes a nose portion adjacent the leading edge, wherein one part of the nose portion has a circular cross-sectional shape, and another part of the nose portion has a noncircular cross-sectional shape.