CA1193551A - Shell-spar cooled airfoil having variable coolant passageway area - Google Patents
Shell-spar cooled airfoil having variable coolant passageway areaInfo
- Publication number
- CA1193551A CA1193551A CA000417841A CA417841A CA1193551A CA 1193551 A CA1193551 A CA 1193551A CA 000417841 A CA000417841 A CA 000417841A CA 417841 A CA417841 A CA 417841A CA 1193551 A CA1193551 A CA 1193551A
- Authority
- CA
- Canada
- Prior art keywords
- airfoil
- cooling air
- spar
- passageways
- airfoil member
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
Classifications
-
- 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
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
- Motor Or Generator Cooling System (AREA)
Abstract
ABSTRACT OF THE DISCLOSURE
The invention comprises a convection-cooled airfoil for a combustion turbine rotor blade or stator vane. The airfoil comprises an airfoil-shaped spar and a metallic shell bonded thereto, the spar and the shell defining therebetween a plurality of coolant passageways for convective cooling of the airfoil. The passageways are arranged with cross-sectional areas which gradually decrease in the downstream direction so that cooling air flow per unit area increases as the cooling air progresses through the passageways.
The invention comprises a convection-cooled airfoil for a combustion turbine rotor blade or stator vane. The airfoil comprises an airfoil-shaped spar and a metallic shell bonded thereto, the spar and the shell defining therebetween a plurality of coolant passageways for convective cooling of the airfoil. The passageways are arranged with cross-sectional areas which gradually decrease in the downstream direction so that cooling air flow per unit area increases as the cooling air progresses through the passageways.
Description
355~
1 50,089 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.
""~
~L~ 55?.~
1 50,089 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.
""~
~L~ 55?.~
2 50,089 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-:~9'.~S~
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-:~9'.~S~
3 50,089 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.
s~
~ 50,089 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 depicted 5 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 3~5~
50,08~
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.
S5.~
6 50,089 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.
s~
7 50,089 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 8 50,089 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.
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.
s~
~ 50,089 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 depicted 5 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 3~5~
50,08~
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.
S5.~
6 50,089 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.
s~
7 50,089 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 8 50,089 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.
Claims (6)
1. A cooled, fluid-directing airfoil member for a combustion turbine, comprising:
an airfoil-shaped spar defining spaced concave and convex sides;
means for communicating to said spar a supply of cooling air;
means for venting from said spar exhaust cooling air which has been utilized to connectively cool said airfoil member;
a metallic shell enclosing and bonded to said spar;
and a plurality of passageways defined between said shell and said spar on both said convex and said concave sides, said plurality of passageways communicating cooling air from said supply means to said exhaust means for convection cooling of said airfoil member and respectively having cross-sectional areas which decrease in the cooling air down-stream direction;
said spar having a substantially hollow interior into which cooling air is communicated from said supply means, said spar defining a plurality of distributed apertures through which cooling air is communicated from said hollow interior to said plurality of passageways;
whereby cooling air flows through said plurality of passageways, gradually increasing in flow per unit area as the cross-sectional areas of said plurality of passageways decreases.
an airfoil-shaped spar defining spaced concave and convex sides;
means for communicating to said spar a supply of cooling air;
means for venting from said spar exhaust cooling air which has been utilized to connectively cool said airfoil member;
a metallic shell enclosing and bonded to said spar;
and a plurality of passageways defined between said shell and said spar on both said convex and said concave sides, said plurality of passageways communicating cooling air from said supply means to said exhaust means for convection cooling of said airfoil member and respectively having cross-sectional areas which decrease in the cooling air down-stream direction;
said spar having a substantially hollow interior into which cooling air is communicated from said supply means, said spar defining a plurality of distributed apertures through which cooling air is communicated from said hollow interior to said plurality of passageways;
whereby cooling air flows through said plurality of passageways, gradually increasing in flow per unit area as the cross-sectional areas of said plurality of passageways decreases.
2. An airfoil member according to claim 1 wherein said exhaust means comprises at least some of said plurality of passageways arranged to vent cooling air from the passageway downstream end directly to the exterior of said airfoil member.
3. An airfoil member according to claim 1 wherein said plurality of passageways extend substantially spanwise from the fluid-entry apertures at the base of said spar to an exit point near an outermost radial portion of said airfoil; and wherein means are provided for venting cooling air from said plurality of spanwise passageways through a tip portion of said airfoil.
4. An airfoil member according to claim 3 further comprising a plurality of passageways venting cooling air from the hollow interior of said spar directly through the trailing edge of said airfoil.
5. A cooled, fluid-directing airfoil member for a combustion turbine, said airfoil member having a leading edge and a trailing edge, and comprising:
a hollow, airfoil-shaped spar defining spaced concave and convex sides and having at least two spanwise cavities;
means for communicating to at least one of said cavities a supply of cooling air;
means for venting from at least one of said cavities exhaust cooling air which has been utilized to cool said air-foil member;
a metallic shell enclosing and bonded to said spar;
and a plurality of passageways having cross-sectional areas which decrease in the cooling air downstream direction, said plurality of passageways extending substantially chordwise between said shell and said spar on both said convex and concave sides and being arranged to communicate cooling air from said supply cavity to said exhaust cavity.
a hollow, airfoil-shaped spar defining spaced concave and convex sides and having at least two spanwise cavities;
means for communicating to at least one of said cavities a supply of cooling air;
means for venting from at least one of said cavities exhaust cooling air which has been utilized to cool said air-foil member;
a metallic shell enclosing and bonded to said spar;
and a plurality of passageways having cross-sectional areas which decrease in the cooling air downstream direction, said plurality of passageways extending substantially chordwise between said shell and said spar on both said convex and concave sides and being arranged to communicate cooling air from said supply cavity to said exhaust cavity.
6. An airfoil member according to claim 5 wherein said exhaust means comprises an opening at a tip portion of said airfoil member communicating cooling air from said exhaust cavities to the exterior of said airfoil member and wherein said passageways are arranged to communicate cooling air from said supply cavity through the airfoil trailing edge.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US33648981A | 1981-12-31 | 1981-12-31 | |
| US336,489 | 1981-12-31 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1193551A true CA1193551A (en) | 1985-09-17 |
Family
ID=23316331
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000417841A Expired CA1193551A (en) | 1981-12-31 | 1982-12-16 | Shell-spar cooled airfoil having variable coolant passageway area |
Country Status (6)
| Country | Link |
|---|---|
| JP (1) | JPS58119902A (en) |
| AR (1) | AR231165A1 (en) |
| BE (1) | BE895473A (en) |
| CA (1) | CA1193551A (en) |
| GB (1) | GB2112869A (en) |
| IT (1) | IT1153921B (en) |
Families Citing this family (8)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CA1263243A (en) * | 1985-05-14 | 1989-11-28 | Lewis Berkley Davis, Jr. | Impingement cooled transition duct |
| JP2609635B2 (en) * | 1987-10-23 | 1997-05-14 | 財団法人電力中央研究所 | Ceramic stationary blade |
| DE19860787B4 (en) * | 1998-12-30 | 2007-02-22 | Alstom | Turbine blade with cooling channels |
| DE19939179B4 (en) * | 1999-08-20 | 2007-08-02 | Alstom | Coolable blade for a gas turbine |
| US7452189B2 (en) * | 2006-05-03 | 2008-11-18 | United Technologies Corporation | Ceramic matrix composite turbine engine vane |
| US7871246B2 (en) * | 2007-02-15 | 2011-01-18 | Siemens Energy, Inc. | Airfoil for a gas turbine |
| US9200534B2 (en) * | 2012-11-13 | 2015-12-01 | General Electric Company | Turbine nozzle having non-linear cooling conduit |
| US9297267B2 (en) * | 2012-12-10 | 2016-03-29 | General Electric Company | System and method for removing heat from a turbine |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS5310206A (en) * | 1976-07-16 | 1978-01-30 | Mitsubishi Electric Corp | Information collector |
| GB1584259A (en) * | 1976-08-16 | 1981-02-11 | Iro Ab | Methods and apparatus for knitting machine control systems |
-
1982
- 1982-12-16 CA CA000417841A patent/CA1193551A/en not_active Expired
- 1982-12-23 BE BE0/209809A patent/BE895473A/en not_active IP Right Cessation
- 1982-12-28 AR AR291721A patent/AR231165A1/en active
- 1982-12-28 IT IT25000/82A patent/IT1153921B/en active
- 1982-12-29 JP JP57235107A patent/JPS58119902A/en active Pending
- 1982-12-30 GB GB08236941A patent/GB2112869A/en not_active Withdrawn
Also Published As
| Publication number | Publication date |
|---|---|
| IT8225000A1 (en) | 1984-06-28 |
| IT1153921B (en) | 1987-01-21 |
| GB2112869A (en) | 1983-07-27 |
| AR231165A1 (en) | 1984-09-28 |
| IT8225000A0 (en) | 1982-12-28 |
| BE895473A (en) | 1983-06-23 |
| JPS58119902A (en) | 1983-07-16 |
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