Note: Descriptions are shown in the official language in which they were submitted.
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SHELL~SPAR COOLED AIRFOIL HAVING VA~IABL~
COOLANT PASSAGEWAY ARRA
CROSS REFERENCE TO RELATED APPLICATIONS
i: C~n~di~ A~ enti~tled "Shell-Spar Cooled Airfoil
Using Multiple Spar Cavities," filed by P. C. ~olden and
copendiny herewith.
BACKGROUND OF THE INVENTION
The present invention relatas generally to com-
bustion turbine rotor blades and vanes and, more particu-
larly, to an airfoil for a combustion turbine rotor blade
or vane having an improved arrangement for fluid cooling.
It is well established that greater operating
efficiency and improved power output of a combustion
turbine may be achieved through higher inlet operating
temperatures. Inlet operating temperatures are limited,
however, by the maximum temperature tolerable to the
rotating turb.ine blades and stationary vanes. Also, as
turbine blade and vane temperatures increase with increas-
ing inlet gas temperature, the vulnerability of the blades
and vanes to damage from -the tension and stresses which
normally accompany turbine operation also increases.
Cooling the blades and vanes permits an increase in inlet
operating temperatures while keeping the turbine hlade and
vane temperatures below the maximum specified opsratincJ
temperature for the material from which the blade or vane
is formed.
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There presently exist many arrangements for
`~ cooling a turbine blade or vane. In a typical arrange-
ment, cooling air is drawn from a compressor section of
; the turbine and passed through channels within the turbine
to reach the blades or vanes. In the case of turbine
blades, cooling air drawn from the compressor section may
typically pass through channels along the t~rbine rotor to
reach each of several turbine rotor discs. Each rotor
disc may define a plurality of channels communicating
cooling air to a plurality of blade roots secured within
the periphery of each rotor disc. Cooling channels within
each of the turbine blades communicate coolin~ air from
the blade root throughout an airfoil portion of the blade.
Similar arrangements typically communicate cooling air to
the turbine vane airfoils.
Typical prior art airfoil cooling arrangements
include transpiration, film, and convection-cooled air-
foils. While transpiration and film-cooled airfoils have
certain advantages, convection-cooled airfoils are pre-
ferable in many turbine applications. For example,convection-cooled airfoils are preferred in turbines
utilizing heavy oil fuels, where apertures in the workiny
surface of transpiration and film-cooled airfoils may tend
to become blocked by deposits rendering the airfoil cool-
ing system ineffective. Convection-cooled airfoils typi-
cally have no working surface holes which may become
blocked, but the airfoils do have enclosed coolant passage-
ways which can give rise to other types of problems.
Convection-cooled airfoils typically comprise a
plurality of coolant passageways arranged to promote
convective cooling of the exterior surface of the airfoil
by means of a coolant fluid flowing through the passage-
ways. Because the cooling fluid gradually heats up as it
passes along a coolant passageway, the cooling fluid is
warmest and thus least effective at the exit point for the
coolant passageway. As a result, the minimum specifica-
tions for the volume of cooling fluid flow and the cross-
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sectional area for the coolant passageway are typicallygoverned by the worst-case conditions at the exit point
for the coolant passageway.
While this procedure assures adequate cooling at
the exit point for a coolant passageway, it generally
results in over-cooling upstream portions of the coolant
passageway. The resultant disproportionate cooling effect
produces a temperature gradient along the coolant passag0-
way. This gradient may give rise to thermal stress within
the airfoil, which could reduce the life potential of the
airfoil. This, in turn, would re~uire an increased cool-
ing fluid flow to compensate.
Hence, prior art convection-cooled airfoils do
not appear to be equipped to deal effectively with the
disproportionate cooling effect described above. The
inadequacy of the prior art is compounded by the present
trend toward increasing the inlet operating temperatures
of a combustion turbine so as to improve turbine power and
efficiency.
SUM~ARY OF THE INVENTION
Accordingly, an airfoil for a combustion turbine
rotor hlade or stator vane is provided with a structure
having improved cooling which enables better airfoil
operation. The airfoil comprises an airfoil-shaped spar
and a metallic shell of one or more layers of sheet metal
bonded to and enclosing the spar. The shell and the spar
define therebetween a plurality of coolant passageways
which conduct cooling air for convective cooling of the
airfoil. The passayeways are arranged with cross-
sectional areas which decrease in the downstream direc-
tion, so that the flow per unit area of the cooling air
gradually increases as the cooling air progresses throuyh
the passageways. As a result, the gradual heating of the
cooling air as it passes along a coolant passageway is
compensated by increasing th~;flow per unit area of the
`` air, producing a balanced ~4~ effect along the exterior
surface adjacent the passageway.
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BRIEF DESCRI~TION O~ THE DRAWINGS
Figure 1 depicts in cross-section an airfoil for
a combustion turbine rotor blade or stator vane;
Figure 2 shows in cross-section a simplified
representation of a coolant passageway within a wall of
the airfoil depicted in Figure l;
Eigure 3 shows a sectional view of the airfoil
wall depicted in Figure 2;
Figure 4 shows a second sectional view of the
airfoil wall depicted in Figure 2;
Figure 5 depicts in cross-section an alternative
embodiment of an airfoil for combustion turbine rotor
blade or stator vane; and
Figure 6 shows in elevation the airfoil depicted5 in Figure 5, as it might appear on a turbine blade.
DESCRIPTION OF THE PREFERRED EMBODIMENT
More particularly, there is ~ own in Figure 1 a
sectional view of an airfoil 10 for~combustion turbine
rotor blade or stator vane. The airfoil 10 comprises a
frame-like, airfoil-shaped strut, or spar, 12 to which is
bonded one or more layers of sheet metal to form a shell
14 which encloses the spar 12. Coolant passageways 16,
arranged as further described below, are formed by the
conjunction of the spar 12 and the shell 14 so as to
promote convection-cooling of the airfoil 10. The pas-
sageways 16 may be defined by channels in the spar ~2, as
shown in Figures 2, 3 and 4, or by channels in the shell
14 (not shown), or by a combination of channels in both
the shell 14 and the spar 12 (not shown).
The spar 12 defines a plurality of cavities 18.
Figure 1 depicts -the preferred embodiment of the airfoil
10 having three cavities 18a, b, c. The fore cavity 18a
and the aft cavity 18c are utilized as supply cavities.
The supply cavities are pressurized by a flow of cooling
air from a compressor section of the turbine. Cooling alr
within the supply cavities 13a, c is delivered to a plux-
ality of generally chordwise coolant passageways 16
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through a plurality of apertures 20 in the spar 12. The
apertures 20 are arranged in one or more spanwise columns
extending the length of the airfoil 10.
Each aperture 20 in the spar 12 of the supply
cavities 18a, c delivers a flow of cooling air to one or
more passageways 16, which terminate at either an aperture
22 in the spar 12 within an exhaust cavity 18b or at the
trailing edge 24 of the airfoil 10. Thus, the exhaust
cavity 18b receives a ~low of cooling air directed from
passageways 16 from the supply cavities l~a, c and vents
this cooling air, for example, in the case of a rotor
blade, through an opening at a radially outer tip (not
shown) of the airfoil 10. The structure and airflow
characteristics, including alternate schemes for venting
the exhaust cavity 18b, of the airfoil depicted in Figure
1 is described in further detail in Canadian application
~17,215, assigned to the assignee of the present
application.
Figure 2 shows a simplified representation of a
coolant passageway 16 for the airfoil 10 shown in Figure
1. The features of the passageway 16 depicted in Figure 2
are not intended as a scaled drawing, but are distorted to
more readily demonstrate the preferred structure. In
accordance with the principles of the invention, a more
balanced airfoil cooling effect is obtained by variation
of the cross-sectional area of the coolant passageways 16.
Larger coolant passageway cross-sectional areas are em-
ployed near the supply cavity inlet apertures 20. The
larger cross-sectional areas result in lower coolant flow
per unit area at a point where the coolant temperature is
lo~er. As the coolant temperature rises, a balanced
cooling effect is achieved by increasing the coolant ~low
per lmit area. This is accomplished by a gradual reduc-
tlon of the coolant passageway cross-sectional area.
Thus, a relatively constant airfoil surface temperature
can be maintained and axial temperature gradients and the
problems incurred thereby avoided.
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The increased coolant passageway cross-sectional
area in the upstream portions of a passageway results in a
decreased pressure drop in these areas, which in turn
reduces the coolant flow requirement and improves the
operating efficiency of the combustion turbine. Coolant
supply pressure is determined by the turbine aerodynamic
design. By utilizing a lower pressure drop in the up-
stream portions of the coolant passageway, higher coolant
flow per unit area and resultant higher coolant heat
transfer coefficients may be utilized in the downstream
portions of the coolant passageway without exceeding the
available supply pressure. Higher coolant heat transfer
coefficients permit use of higher coolant temperature
rises and thereby result in still further reduction in
coolant flow.
Figure 7 demonstrates the temperature rela-
tionship among the hot gas, the blade wall, and the cool-
ant along a single coolant passageway. The graph in
Figure 7 demonstrates the qualitative relationship among
the three temperatures for both a typical prior art cool-
ant passageway of constant cross-sectional area and a
coolant passageway structured according to the principles
of the invention with variable cross-sectional area. The
temperature of the hot gas 30 is shown as a constant for
both a constant area and a variable area coolant passage-
way. The blade wall temperature shown at 32 evidences the
imbalanced cooling effect of a typical constant cross-
sectional area coolant passageway. The temperature of the
coolant shown at 34 gradually increases as it progresses
through the constant cross-sectional area coolant passage-
way.
The temperature of the blade wall shown at 36
demonstrates the effect of decreasing the cross-sectional
area of the coolant passageway as the coolant temperature,
shown at 38, increases. The result is a balanced cooling
effect on the blade wall, decreasing or eliminating axial
temperature qradients and thereby decreasing the thermal
stress on the airfoil.
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Figure 3 shows a section of the coolant passage-
way of Figure 2 at a downstream point on the coolant
passageway 16; Figure 4 shows a section of the same cool-
ant passageway at an upstream polnt on the passageway 16.
Figures 3 and 4 depict the preferred arrangement of vari-
able depth grooves in the spar 12 to achieve the variable
cross-sectional area of the passageway 16. Although not
shown in the drawings, it is envisioned that the same
effect may be achieved by use of variable depth grooves in
the shell 14 or by use of varlable depth grooves in both
the shell 14 and the spar 12.
Figures 5 and 6 depict an alternative embodiment
of an airfoil 50 structured according to the principles of
the invention. This embodiment of the airfoil is prefer-
ably utili2ed in downstream portions of the turbine. Theshell 14 and the spar 12 of the airfoil 50 define spanwise
coolant passageways 52 in contrast to the chordwise cool-
ant passageways 16 of the airfoil 10~ In a typical appli-
cation of the airfoil 50, cooling air may be forced
through one or more coolant channels 54 in a blade root 56
to reach a pressurized hollow interior 58 of the airfoil
50. Apertures 60 through the spar 12 along the base of
the airfoil 50 near the blade root 56 communicate coolin~
air to the plurality of spanwise coolant passageways 52.
The spanwise coolant passageways 52 carry the cooling air
radially outward from the entrance apertures 60 to exit at
a blade tip 62.
In accordance with the principles of the inven-
tion, the cross-sectional areas of the passageways 52
gradually decrease in the radially outward, downstream
direction. Cooling air flowing through the passageways 52
thereby gradually increases in flow per unit area as its
temperature increases, resulting in a substantially
balanced cooling effect.
A trailing edge 64 of the spar 12 defines a
plurality of chordwise passageways 66 arranged in a single
spanwise column. The chordwise passageways 66 deliver
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cooling air from the pressurized interior 58 of the air-
foil 50 to the exterior of the airfoil and thereby provide
a mechanism for cooling the trailing edge 64 of the air-
foil 50.
