GB2112869A

GB2112869A - Cooled airfoil

Info

Publication number: GB2112869A
Application number: GB08236941A
Authority: GB
Inventors: Paul Clarence Holden
Original assignee: Westinghouse Electric Corp
Current assignee: CBS Corp
Priority date: 1981-12-31
Filing date: 1982-12-30
Publication date: 1983-07-27
Also published as: AR231165A1; CA1193551A; IT8225000A0; IT1153921B; IT8225000A1; JPS58119902A; BE895473A

Abstract

A gas turbine airfoil blade comprises an airfoil-shaped spar 12 and a metallic shell 14 bonded thereto and defining coolant passageways 16 for convective cooling of the airfoil. The passageways 16 are so made that the cross-sectional area of each passageway gradually decreases in the downstream direction of coolant flow so that cooling air flow per unit area increases as the cooling air progresses through the passageways. <IMAGE>

Description

SPECIFICATION Shell-spar cooled airfoil having variable coolant passageway area The present invention relates generally to combustion turbine rotor blades and vanes and, more particularly, to an airfoil for a combustion turbine rotor blade or vane having an improved arrangement for forced 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 turbine blades and stationary vanes. Also, as turbine blades and vane temperatures increase with increasing 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 generally permits an increase in inlet operating temepratures while keeping the turbine blade and vane temperatures below the maximum specified operating temperature for the material from which the blade or vane is formed.

There presently exist many arrangements for cooling a turbine blade or vane. In a known prior art arrangement, 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 turbine 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 cooling 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 airfoils. While transpiration and film-cooled airfoils have certain advantages, convention-cooled airfoils are however preferable in many turbine application. For example, convection-cooled airfoils are preferred in turbines utilizing heavy oil fuels, where apertures in the working surface of transpiration and film-cooled airfoils may tend to become blocked by deposits rendering the airfoil cooling system ineffective. Convention-cooled airfoils typically have no wroking surface holes which may become blocked, but the airfoils do have enclosed coolant passageways 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 passageways. 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 specifications of the volume of cooling fluid flow and the cross-sectional area for the coolant passageway are typically governed by the worstcase conditions at the exit point for the coolant passageway. While such a 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 differential cooling effect produces a temperature gradient along the coolant passageway. This gradient may give rise to thermal stress within the airfoil, which could reduce the life potential of the airfoil. This, in turn, would require an increased cooling 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.

The invention in its broad form comprises a force cooled, fluid-directing airfoil member for a combustion turbine, the member being of the type including built in cooling passageways; an airfoilshaped spar having spaced concave and convex sides; means for conveying to said spar a supply of cooling fluid; venting exhaust means in said spar for venting from said spar an exhaust of said cooling-fluid air which has been utilized to convectively 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, said plurality of passageways conveying cooling fluid from said supply means to said exhaust means for convention cooling of said airfoil member, each of said passageways having a non-uniform cross-sectional area which decreases in downstream direction; whereby cooling fluid flows through said plurality of passageways, gradually increasing in flow per unit area as the cross-sectional areas of said plurality of passageways decreases.

In a preferred embodiment described herein, an airfoil for a combustion turbine rotor blade or stator vane is provided with a structure having improved cooling which enables better airfoil operation wherein the airfoil comprises an airfoilshaped 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 passageways are arranged with cross-sectional areas which decrease in the downstream direction, so that the flow per unit area of the cooling air gradually increases as the cooling air progresses through the passageways.As a result, the gradual heating of the cooling air as it passes along a coolant passageway is compensated by increasing the flow per unit area of the air, producing a balanced cooling effect along the exterior surface adjacent the passageway.

The invention can be better understood from the following description of a preferred example, given by way of example only, and to be read and understood in conjunction with the accompanying drawings in which: 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 1; Figure 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 in Figure 5, as it might appear on a turbine blade.

More particularly, there is shown 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 1 4 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 1 0.

The passageways 1 6 may be defined by channels in the spar 12, 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 1 2 (not shown).

The spar 1 2 defines a plurality of cavities 18, Figure 1 depicts the preferred embodiment of the airfoil 10 having three cavities 1 8a, b, c. The fore cavity 1 8a and the aft cavity 1 8c are utilized as supply cavities. The supply cavities are pressurized by a flow of cooling air from a compressor section of the turbine. Cooling air within the supply cavities 1 spa, c is delivered to a plurality of generally chordwise coolant passageways 1 6 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 1 8a, 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 1 8b or at the trailing edge 24 of the airfoil 1 0. Thus, the exhaust cavity 1 8b receives a flow of cooling air directed from passageways 1 6 from the supply cavities 1 spa, 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 alternative schemes for venting the exhaust cavity 1 8b, of the airfoil depicted in Figure 1 can be chosen suitably.

Figure 2 shows a simplified representation of a coolant passageway 16 for the airfoil 10 shown in Figure 1. The passageways 1 6 depicted in Figure 2 are not intended to show 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 a non-uniform crosssectional area of the coolant passageways 1 6.

Larger coolant passageway cross-sectional areas are employed 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 lower. As the coolant temperature rises, a balanced cooling effect is achieved by increasing the coolant flow per unit area. This is accomplished by a gradual reduction of the coolant passageway crosssectional area. Thus, a relatively constant airfoil surface temperature can be maintained and axial temperature and the problems caused thereby avoided.

The increased coolant passageway crosssectional 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 upstream 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 relationship among the hot gas, the blade wall, and the coolant along a single coolant passageway. The graph in Figure 7 demonstrates the qualitative relationship among the three temperatures for both a typical prior art coolant passageway of constant cross-sectional area and a coolant passageway structured according to the principles of the invention with variable crosssectional area. The temperature of the hot gas 30 is shown as a constant for both a constant area and a variable area coolant passageway. 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 passageway.

The temperature of the blade wall shown at 36 demonstrates the effect of decreasing the crosssectional 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 gradients and thereby decreasing the thermal stress on the airfoil.

Figure 3 shows a section of the coolant passageway of Figure 2 at a downstream point on the coolant passageway 16; Figure 4 shows a section of the same coolant passageway at an upstream point on the same coolant passageway at an upstream point on the passageway 1 6.

Figures 3 and 4 depict the preferred arrangement of variable depth grooves in the spar 12 to achieve the variable cross-sectional area of the passageway 1 6. 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 variable depth grooves in both the shell 14 and the spar 12.

Figures 5 and 6 depict an alternative embodiment of an airfoil 50 structured using the principles of the invention. This embodiment of the airfoil is preferably utilized in downstream portions of the turbine. The shell 14 and the spar 12 of the airfoil 50 define spanwise coolant passageways 52 in contrast to the chordwise coolant passageways 1 6 of the airfoil 10. In a typical application 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 convey cooling 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 tiD 62.

In accordance with the principles of the invention, 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 cooling air from the pressurized interior 58 of the airfoil 50 to the exterior of the airfoil and thereby provide a mechanism for cooling the trailing edge 64 of the airfoil 50.

Claims

1. A force cooled, fluid-directing airfoil member for a combustion turbine, the member being of the type including built in cooling passageways, comprising: an airfoil-shaped spar having spaced concave and convex sides; means for conveying to said spar a supply of cooling fluid; venting exhaust means in said spar for venting from said spar an exhaust of said cooling fluid air which has been utilized to convectively 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, said plurality of passageways conveying cooling fluid from said supply means to said exhaust means for convection cooling of said airfoil member, each of said passageways having a non-uniform crosssectional area which decreases in downstream direction; whereby cooling fluid 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 cooling fluid is air and wherein said spar has a substantially hollow interior into which cooling air is conveyed from said supply means, said spar defining a plurality of apertures through which cooling air is conveyed from said hollow interior to said plurality of passageways.

3. An airfoil member according to claim 2 wherein said hollow interior of said spar is divided into at least two cavities, at least one of said cavities receiving a flow of spent cooling air through apertures in said spar from the downstream end of said plurality of said passageways, which cooling air is thereafter conveyed to said exhaust means.

4. An airfoil membe;'according to claim 2 wherein said exhaust means comprises at least some of said plurality of passageways arranged to vent cooling air from the passageway downstream end to an exterior of said airfoil member.

5. An airfoil member according to claim 2 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.

6. An airfoil member according to claim 5 wherein said plurality of spanwise passageways vent cooling air through a tip portion of said airfoil.

7. An airfoil member according to claim 5 further comprising a plurality of passageways venting cooling air from the hollow interior of said spar through a trailing edge of said airfoil.

8. An airfoil member according to claim 1 wherein said spar defines an odd number of spanwise cavities.

9. An airfoil member according to claim 8 wherein said spanwise cavities of said spar are defined by at least two spanwise partitions attaching and extending generally normal to the sides of said spar.

10. An airfoil member according to claim 8 wherein said strut has at least three of said cavities.

11. An airfoil member according to claim 10 wherein said cavities of said spar nearest to the leading edge and the trailing edge of said airfoil member are supply cavities and succeeding adjacent cavities therebetween are alternately exhaust cavities and supply cavities.

12. An airfoil member according to claim 11 wherein at least some of said passageways defined by said shell and said spar connect each of said supply cavities to an adjacent exhaust cavity.

13. An airfoil member according to claim 12 wherein at least some of said plurality of passageways defined by said shell and said spar extend spanwise and chordwise between said supply cavities and said exhaust cavities.

14. An airfoil member according to claim 12 wherein said plurality of passageways defined by said shell and said spar are formed of channels in said shell.

15. An airfoil member according to claim 12 wherein said plurality of passageways defined by said shell and said spar are formed of channels in said spar.

16. An airfoil member according to claim 12 wherein said plurality of passageways defined by said shell and said spar are formed of a combination of channels in said shell and channels in said spar.

17. An airfoil member according to claim 12 wherein said shell comprises at least one layer of sheet metal.

1 8. An airfoil member according to claim 12 wherein said exhaust means comprises a plurality of apertures through said spar and said shell on the convex side of said spar.