GB2152150A

GB2152150A - Anti-icing inlet guide vane

Info

Publication number: GB2152150A
Application number: GB08428338A
Authority: GB
Inventors: John Paul Petrowicz; Verna E Irons Executrix Irons
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 1983-12-27
Filing date: 1984-11-09
Publication date: 1985-07-31
Also published as: FR2557201A1; IT8424117A0; DE3446206A1; JPS60175706A; GB8428338D0; CA1237587A

Abstract

A thin inlet guide vane is investment cast of a high thermal conductivity material. Internal flow chambers include pins 76 bridging the flow chambers. The pins 76 provide strength and support to sheathing members, aid in the transfer of heat from air in the flow chambers to the sheathing members and provide air flow control. The vane may be formed with reverse camber (Fig. 6) and be formed with a twist in the range 2 to 25 degrees. <IMAGE>

Description

SPECIFICATION Inlet guide vane The present invention relates to gas turbine engines and, more particularly, to inlet guide vanes for gas turbine engines.

BACKGROUND OF THE INVENTION Inlet guide vanes are conventionally employed at the inlet to a gas turbine engine to control the amount and rotation of air fed to a compressor. The angle of the inlet guide vanes with respect to the inflowing air is variable in dependence on engine speed. At low speed, the inlet guide vanes are rotated to their most closed positions and, at full speed, are rotated to their most open positions.

Inlet guide vanes, being at the engine inlet, receive air at ambient temperature and humidity. Thus, inlet guide vanes are subject to icing if measures are not taken to prevent such a phenomenon. If ice were permitted to build up on the inlet guide vanes, it could separate from the inlet guide vanes, enter the rapidly rotating engine components thereby causing damage. It has, therefore, become customary to provide anti-icing in inlet guide vanes.

Anti-icing for inlet guide vanes may be accomplished by tapping off a flow of air from the compressor of the gas turbine engine at a point where the work done on the air has raised its temperature to, for example, from about 1 50 to about 500 degrees F and valving this heated air to a plenum surrounding the inlet guide vanes. Each inlet guide vane includes one or more flow chambers leading from its spindle, which is located in the plenum, to its vane wherein the air is distributed for heating the vane and is finally exhausted to the engine inlet.

Forming flow chambers in the vane presents several problems necessitating design compromises which have prevented taking fullest advantage of available engine efficiency.

The cross section of a vane portion of an inlet guide vane is kept relatively thin to permit as free a flow of air past it as possible. For example, the maximum cross section of a vane in certain small engines may be oneeighth inch and may be as thin as onesixteenth inch. Such small dimensions present a problem in forming flow chambers within the vane. It has heretofore been customary to forge*a vane from a beryllium copper material and then to machine flow chambers into one surface of the forged piece. A cover plate is then brazed atop the flow chambers to form one of the vane surfaces. Since the vane tapers to a very thin cross section toward its trailing edge, it is not possible by conventional methods to carry the flow chambers all the way to the trailing edge.Instead, the cover plate terminates well short of the trailing edge and the air in the flow chambers exit to the surface at that point.

Economic constraints limit the complexity of machining that can be employed to form the flow chambers in the vane. Furthermore, the requirements for machining and brazing have constrained the shape of the cover plate to a plane thus limiting substantially one complete side of the vane to be planar rather than permitting a more efficient curved aerodynamic shape. In fact, for greatest aerodynamic efficiency, a complex curvature should be employed. Such complex curvature must be foregone in the structure of the prior art due to the requirement for a planar cover plate.

Even when a planar cover plate is employed, brazing is not always perfect and consequently voids may be formed between the cover plate and the surface to which it should be brazed. Such voids reduce heat transmission and may permit cool spots to develop.

In another area of a gas turbine engine, namely in the turbine stage, blades are formed of a ferrous material by investment casting. The investment casting produces a hollow core in the turbine blades for the passage of cooling fluid such as air therethrough. The process of investment casting involves the preparation of a ceramic core made of, for example, silica (SiO2) about which the ferrous material is cast. After solidification of the ferrous material, the silica core is leached out using a caustic such as potassium or sodium hydroxide.

Inlet guide vanes are formed of a high thermal conductivity material rather than a ferrous metal. Copper. with about one percent beryllium alloyed therein, produces a relatively high strength, high conductivity material for use in inlet guide vanes. It was not previously recognized that investment casting could be successfully employed with such material for use in inlet guide vanes.

OBJECTS OF THE INVENTION Accordingly, it is an object of the present invention to provide an inlet guide vane which overcomes the drawbacks of the prior art.

It is another object of the invention to provide an inlet guide vane having its vane portion formed in a single piece by investment casting.

It is a further object of the invention to provide an inlet guide vane having internal flow chambers of improved shape for providing heat transfer to the surface of the inlet guide vane.

It is still a further object of the invention to provide an inlet guide vane with stiffer support for the surface of flow chambers therein.

SUMMARY OF THE INVENTION According to an embodiment of the invention, there is provided an inlet guide vane comprising a one-piece cast vane, a spindle attached to the vane, an air channel in the spindle, an air flow chamber in the vane between first and second opposed sheathing members of the vane communicating with the air channel, a plurality of pins bridging the air flow chamber integral with the first and second opposed sheathing members and, an air exit from the air flow chamber between the first and second sheathing members at a trailing edge of the vane. The first sheathing member has a first curved aerodynamic shape, and the second sheathing member having a second curved aerodynamic shape.

According to a feature of the invention, there is provided an inlet guide vane as described above, with its one-piece cast vane being cast of a copper and beryllium alloy having from about 0.5 to about 2 percent beryllium. In addition, the first and second curved aerodynamic shapes define a twist about an axis of the vane. The twist is from about 2 to about 25 degrees. The vane has a first cross section concave in a first direction at a first end of the vane and a second cross section concave in a second direction opposite to the first direction at a second end of the vane.

Briefly stated, according to an embodiment of the invention, there is provided an inlet guide vane which is investment cast in one piece of a high thermal conductivity material with an array of pins connecting opposed sheathing members of flow chambers in the vane. The pins support and stabilize the sheathing members, aid in providing heat conduction from the air in the flow chambers to the surface of each sheathing member and provide flow control of the air. Use of investment casting permits the use of a shape which is optimized for aerodynamic performance without the shape being constrained by fabrication requirements.

The above, and other objects, features and advantages of the present invention will become apparent from the following description read in conjunction with the accompanying drawings, in which like reference numerals designate the same elements.

BRIEF DESCRIPTION OF THE DRAWINGS FIGURE 1 is a simplified schematic diagram of a gas turbine engine to which the present invention may be applied.

FIGURE 2 is a cross section of an inlet guide vane assembly showing a side view with cut-away portion of an inlet guide vane according to the prior art.

FIGURE 3 is a cross section of the inlet guide vane taken along line 3-3 of Figure 2.

FIGURE 4 is a cross section taken along line 4-4 of Figure 2.

FIGURE 5 is a cross section of an inlet guide vane assembly showing a side view with cut-away portion of an inlet guide vane according to an embodiment of the present invention.

FIGURE 6 is a bottom view of the inlet guide vane of Figure 5.

FIGURE 7 is a cross section taken along line 7-7 of Figure 5.

DETAILED DESCRIPTION OF THE PRE FERRED EMBODIMENT Referring now to Figure 1, there is shown, generally at 10, a gas turbine engine in which the present invention may be applied. Gas turbine engine 10 includes an axial compressor 1 2 which receives a supply of air from an inlet guide vane assembly 1 4. Axial compressor 1 2 compresses the incoming air and feeds it to a combustor 1 6 where its energy is increased by the burning of fuel. The products of combustion and heated air from combustor 1 6 are fed to a turbine 1 8 which drives axial compressor 1 2 through a mechanical connection 20 as well as providing output power either on a shaft or a jet exhaust (not shown).

Compressed air, which has been heated by the work done on it during compression is tapped from axial compressor 1 2 and fed in a conduit 22 to an anti-icing valve 24 which meters a supply of heated air in a conduit 26 to a plenum 28 surrounding inlet guide vane assembly 1 4.

An inlet guide vane control 30 provides mechanical control of the angie of inlet guide vanes in inlet guide vane assembly 1 4 in response to an engine speed signal from an engine speed signal generator 32. Anti-icing valve 24 may also be responsive to an engine speed signal produced by engine speed signal generator 32 to control the metering of air fed to conduit 26 as well as anti-icing air fed to other components in the engine (not shown) as indicated by a conduit 34.

Referring now to Figure 2, there is shown an inlet guide vane 36 according to the prior art. Inlet guide vane 36 includes a vane 38 to which is attached a spindle 40 passing upward through plenum 28. An actuating lever 42 of a conventional type is affixed to the end of spindle 40 by conventional means such as, for example, by a threaded nut 43. Vane 38 is rotatable by actuating lever 42 about an axis defined by spindle 40 and an inner pivot 44.

A cross channel 46 in spindle 40 admits heated air from plenum 28 into an axial channel 48 which conveys the heated air along paths shown by dashed arrows into a flow chamber 50 in vane 38. Flow chamber 50 is formed with ribs 52 appropriately shaped and positioned both to channel the airflow and to provide a mounting base for brazing a cover plate 54.

Referring now also to Figures 3 and 4, a recess shelf 56 permits insetting cover plate 54 flush with a surface 58 of vane 38.

Recess shelf 56 forms a step 60 surrounding cover plate 54 to which cover plate 54 is brazed in the assembly operation. Rib 52 is recessed to the same depth as step 60 so that cover plate 54 is also supported by, and brazed to, rib 52.

Referring now to Figure 3, an exit channel 62 vents the hot air from flow chamber 50 to a trailing surface 64 of vane 38.

As previously noted, successful brazing limits the shape of cover plate 54, and consequently the side of vane 38 containing cover plate 54 to a plane surface. In addition, flow chamber 50 must stop well short of trailing edge 66 of vane 38 due to the thinness of the material approaching trailing edge 66. Also, due to the limited cross sectional dimension of vane 38, cover plate 54 is necessarily a thin plate. In order to retain reasonable control over machining complexity for the formation of flow chambers 50, relatively large areas of cover plate 54 remain unsupported and subject to vibration and subsequent cracking failures. Furthermore, air flow through flow chambers 50 may be relatively smooth and thus fail to provide maximum heat transfer to the surfaces of vane 38.

Referring now to Figure 5, an inlet guide vane 68, according to an embodiment of the present invention, is shown. A vane 70 of inlet guide vane 68 is formed in a single piece by investment casting with the internal structure shown formed during the casing process.

Spindle 40 may be cast at the same time of the same material as vane 70, or preferably of a higher strength material such as steel and attached to vane 70 by brazing for greater durability especially in the attachment area of actuating lever 42. A plurality of ribs 72 define air flow chambers 74. A plurality of pins 76 bridge flow chambers 74 both to support the opposed sheathing members 86 and 88 of vane 70 and also to transfer heat from the air inside flow chambers 74 to sheathing members 86 and 88. It will be noted that, where necessary or desirable, pins 76 are placed in staggered rows so that air passing between one pair of pins 76 impacts directly on a pin 76 in the succeeding row.

This contributes to air turbulence in flow chamber 74 and thereby enhances heat transfer to the surfaces, Referring now to Figure 7, a typical chord near the midspan of vane 70 is shown. The maximum cross section thickness 90 of vane 70 should be as small as possible to permit the free flow of air past vanes 70. In the present invention, this maximum thickness will be between 65 and 1 25 mils, with a minimum thickness near trailing edge 78 of about 45 mils. In a preferred embodiment, the thickness 92 of each sheathing member 86 and 88 is between 10 and 25 mils with the smaller values near trailing edge 78. The width 94 of air flow chamber 74 between sheathing members 86 and 88 is 20 to 30 mils with the smaller values near trailing edge 78.

Flow chamber 74 extends fully to an air exit 77 at trailing edge 78 of vane 70 so that the heated air is retained in vane 70 to trailing edge 78 without being discharged before it reaches that location. This improves de-icing because the heated air is not forced to exit the vane before reaching trailing edge 78. It will also be noted in Figure 5 that the perimeter of flow chamber 74 can be formed in a more complex manner than is economically or practically feasible when fabricating a two-piece vane. For example, a notch 80 is formed in the bottom of flow chamber 74 to permit air to flow about an adjacent pin 76. Other advantageous shapes including special means for channeling de-icing air to locations where it is most needed can be envisioned by one skilled in the art in the light of the present disclosure.Pins 76 can be positioned as needed or desired for three purposes: 1) support of sheathing members of vane 70 2) heat transfer from the air to the outer surface of the sheathing members 3) flow control With regard to flow control, the closeness of spacing of pins 76 establishes the flow resistance in a particular portion of flow chamber 74. In this manner, the flow rate to selected portions of vane 70 can be controlled.

Vane 70 is cast of a copper and beryllium alloy in which the amount of beryllium is from about 0.1 to about 10 percent. As the beryllium content is reduced, the strength of vane 70 is also reduced. As the beryllium content is increased, the thermal conductivity is reduced. In the preferred embodiment, a beryllium content of from about 0.5 to about 2 percent is employed.

Referring now to Figure 6, it will be noted that vane 70, instead of being planar as was necessary in the prior art vane of Figure 2, is cambered and twisted for improved aerodynamic efficiency. Camber is a measure of the curvature or arch of an axial cross section of an airfoil. Twist about the axis of an airfoil is a measure of the change in angle, from the airfoil root to its tip, of the airfoil chord relative to the engine center line. Total twist is in the range of 2 to 25 degrees. In the preferred embodiment, the twist is from about 10 to about 1 8 degrees and, in the most preferred embodiment, the twist is from about 1 2 to about 1 6 degrees.

In the perspective of Figure 6, the extreme bottom 82 (radially inward portion) includes a cross section which is concave facing downward in the figure whereas the extreme top 84 of vane 70 has a cross section which is concave facing upward. This change in curvature from bottom 82 to top 84 is known as reverse camber. The camber at the extreme top 84 (radially outward) is in the direction of rotation of the compressor blades in the succeeding stage. This tends to reduce the relative velocity between the air at the outer tips of the first stage compressor rotor blades where they are closest to sonic velocity.The reverse camber at the extreme bottom 82 (radially inward) provides counterswirl to the air entering near the roots of the first compressor rotor stage so that this air, after passing the blades of the first compressor rotor stage, impacts the first compressor stator stage with a lower Mach number.

One typical cross section is shown in Figure 7 at a location just beyond the transition from concave downward to concave upward. A first sheathing member 86 of vane 70 is shown concave upward. A second sheathing member 88 is shown convex downward. Each sheathing member has a different curved aerodynamic shape. These shapes are determined wholly by the aerodynamic requirements of vane 70 and are unconstrained by any need for planarity as was the case in the prior art.

Referring again to Figure 5, the use of both ribs 72 and pins 76 is considered a transitional embodiment which may be wholly replaced by an embodiment using pins alone appropriately spaced and positioned for the purposes enumerated in the preceding. Although such a pin-only embodiment is not shown or described herein, the disclosure herein is sufficient to fully enable one skilled in the art to understand and to make and use an embodiment employing pins alone.

Having described specific preferred embodiments of the invention with reference to the accompanying drawings, it is to be understood that the invention is not limited to those precise embodiments, and that various changes and modifications may be effected therein by one skilled in the art without departing from the scope or spirit of the invention as defined in the appended claims.

Claims

1. An inlet guide vane comprising: a one-piece cast vane, said vane being cast of a copper and beryllium alloy having from about 0.5 to about 2 percent beryllium; a spindle attached to said vane; an air channel in said spindle; an air flow chamber in said vane between first and second opposed sheathing members of said vane communicating with said air channel; a plurality of pins bridging said air flow chamber integral with said first and second opposed sheathing members; and an air exit from said air flow chamber between said first and second sheathing members at a trailing edge of said vane; wherein said first sheathing member has a first curved aerodynamic shape and said second sheathing member has a second curved aerodynamic shape, said first and second curved aerodynamic shapes defining a twist about an axis of said vane, said twist being from about 2 to about 25 degrees; and wherein said vane has a first cross section concave in a first direction at a first end of said vane and a second cross section concave in a second direction opposite to said first direction at a second end of said vane.

2. An inlet guide vane according to claim 1 wherein the maximum thickness of said onepiece cast vane is 1 25 mils.

3. An inlet guide vane according to claim 2 wherein the width of said air flow chamber between said first and second opposed sheathing members is between 20 and 30 mils.

4. An inlet guide vane comprising.

a one-piece cast vane: a spindle attached to said vane; an air channel in said spindle; an air flow chamber in said vane between first and second opposed sheathing members of said vane communicating with said air channel; a plurality of pins bridging said air flow chamber integral with said first and second opposed sheathing members; and an air exit from said air flow chamber between said first and second sheathing members at a trailing edge of said vane: wherein said first sheathing member has a first curved aerodynamic shape, and said second sheathing member has a second curved aerodynamic shape.

5. An inlet guide vane according to claim 4 wherein said first and second curved aerodynamic shapes define a twist about an axis of said vane.

6. An inlet guide vane according to claim 5 wherein said vane includes a first cross section concave in a first direction at a first end of said vane and a second cross section concave in a second direction generally opposite to said first direction at a second end of said vane.

7. An inlet guide vane according to claim 5 wherein said twist is from about 2 to about 25 degrees.

8. An inlet guide vane according to claim 5 wherein said twist is from about 10 to about 18 degrees.

9. An inlet guide vane according to claim 6 wherein said twist is from about 1 2 to about 1 6 degrees.

10. An inlet guide vane according to claim 4 wherein said one-piece cast vane is cast of copper and beryllium alloy.

11. An inlet guide vane according to claim 10 wherein said alloy contains from about 0.5 to about 2 percent beryllium.

1 2. An inlet guide vane according to claim 4 wherein the maximum thickness of said one-piece cast vane is 1 25 mils.

1 3. An inlet guide vane according to claim 1 2 wherein the width of said air flow chamber between said first and second sheathing members is between 20 and 30 mils.

14. A guide vane substantially as hereinbefore described with reference to and as illustrated in Figures 5 to 7 of the drawings.