Note: Descriptions are shown in the official language in which they were submitted.
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INLET GUIDE VANE
The present invention relates to gas turbine
engines and, more particularly, to inlet guide vanes for
gas turbine er.gines.
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
isvariable 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 150 to about 500 degrees F and
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valving this heated air to a plenum surrounding this
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 one-eighth inch and may be as thin as one-sixteenth
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
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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 cure 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 (Si02)
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 INYENTION
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 proyide
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
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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.
S UMMARY OF THE I NVENT I ON
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 brid~ing 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, accordin~ 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.
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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.
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'DETA'ILED DESCR'IPTION O~'THE'PREFERR _ 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 12 which receives a
supply of air from an inlet guide vane assembly 14.
Axial compressor 12 compresses the incoming air and
feeds it to a combustor 16 where its energy is increased
by the burning of fuel. The products of combustion and
heated air from combustor 16 are fed to a turbine 18
which drives axial compressor 12 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 12 and fed in a conduit 22 to an anti-icing
valve 24 which meters a suppl~ of heated air in a
conduit 26 to a plenum 28 surrounding inlet guide vane
assembly 14.
2~ An inlet guide vane control 30 provides
mechanical control of the angle of inlet guide vanes in
inlet guide vane assembly 14 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
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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,
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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 125 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
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g
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 ~low 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 E'igure 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,
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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
18 degrees and, in the most preferred embodiment, the
twist is from about 12 to about 16 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 o1ltward) 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 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.
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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 aloneA
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.
