Spar and shell constructed turbine blade

Abstract

A blade for a rotor of a gas turbine engine is constructed with a spar and shell configuration. The spar is constructed in an integral unit or multi-portions and includes a first wall adjacent to the pressure side and a second wall adjacent to the suction side, a tip portion extending in the spanwise direction and extending beyond the first wall and the second wall and a root portion extending longitudinally, an attachment portion having a central opening for receiving the root portion and a platform portion. The root portion fits into the central opening and is secured therein by a pin extending transversely through the attachment and the root portion. The shell fits over the spar and is supported thereto by a plurality of complementary hooks extending from the spar and shell. The ends of the shell fit into grooves formed on the tip portion and the platform. The shell is made from a high temperature resistant material, such as Molybdenum or Niobium, and is formed from a wire EDM process.

Description

FEDERAL RESEARCH STATEMENT

None.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to internally cooled turbine blades for a gas turbine engine, and more specifically to a turbine blade made from a spar and shell construction.

2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98

As one skilled in the gas turbine technology recognizes, the efficiency of the engine is enhanced by operating the turbine at a higher temperature and by increasing the turbine's pressure ratio. Another feature that contributes to the efficacy of the engine is the ability to cool the turbine with a lesser amount of cooling air. The problem that prevents the turbine from being operated at higher temperatures is the limitation of the structural integrity of the turbine component parts that are jeopardized in its high temperature, hostile environment. Scientist and engineers have attempted to combat the structural integrity problem by utilizing internal cooling and selecting high temperature resistance materials. The problem associated with internal cooling is twofold. One, the cooling air that is utilized for the cooling comes from the compressor that has already expended energy to pressurize this air and the spent air in the turbine cooling process in essence is a deficit in engine efficiency. The second problem is that the cooling is through cooling passages and holes that are in the turbine blade which, obviously, adversely affects the blade's structural prowess. Because of the tortuous path that is presented to the cooling air, the pressure drop that is a consequence thereof requires higher pressure and more air to perform the cooling that would otherwise take a lesser amount of air given the path becomes less tortuous to the cooling air. While there are materials that are available and can operate at a higher temperature that is heretofore been used, the problem is how to harness these materials so that they can be used efficaciously in the turbine environment.

To better appreciate these problems it would be worthy of note to recognize that traditional blade cooling approaches include the use of cast nickel based alloys with load-bearing walls that are cooled with radial flow channels and re-supply holes in conjunction with film discharge cooling holes. Examples of these types of blades are exemplified by the following patents that are incorporated herein by reference. U.S. Pat. No. 4,257,737 granted to D. E. Andress et al on Mar. 24, 1981 entitled “Cooled Rotor Blade”; U.S. Pat. No. 4,753,575 granted to J. L. Levengood et al on Jun. 28, 1988 entitled “Airfoil with Nested Cooling Channels”; U.S. Pat. No. 5,476,364 granted to R. J. Kildea on Dec. 19, 1995 entitled “Tip Seal and Anti-Contamination for Turbine Blades”; and U.S. Pat. No. 5,700,131 granted to Hall et al on Dec. 23, 1997 entitled “Cooled Turbine Blades for a Gas Turbine Engine”.

Also well known by those skilled in this technology is that the engine's efficiency increases as the pressure ratio of the turbine increases and the weight of the turbine decreases. Needless to say these parameters have limitations. Increasing the speed of the turbine also increases the airfoil loadings and, of course, satisfactory operation of the turbine is to stay within given airfoil loadings. The airfoil loadings are governed by cross sectional area of the airfoil of the turbine multiplied by the velocity of the tip of the turbine squared. Obviously, the rotational speed of the turbine has a significant impact on the loadings.

The spar/shell construction contemplated by this invention affords the turbine engine designer the option of reducing the amount of cooling air that is required in any given engine design and in addition, allowing the designer to fabricate the shell from exotic high temperature materials that heretofore could not be cast or forged to define the surface profile of the airfoil section. In other words, by virtue of this invention, the skin can be made from Niobium or Molybdenum or their alloys, where the shape is formed by a well known electric discharge process (EDM) or a wire EDM process. In addition, because of the efficacious cooling scheme of this invention, the shell portion could be made from ceramics, or more conventional materials and still present an advantage to the designer because a lesser amount of cooling air would be required.

BRIEF SUMMARY OF THE INVENTION

An object of this invention is to provide a turbine rotor for a gas turbine engine that is constructed with in a spar/shell configuration.

Another object of the present invention to provide for a turbine blade that can make use of higher temperature materials than are presently used in turbine blades.

Another object of the present invention is to form a shell from a high temperature resistant material that cannot be cast or machined into a thin wall airfoil.

A feature of this invention is an inner spar that extends from the root of the blade to the tip and is joined to the attachment at the root by a pin or rod or the like.

Another feature of this invention is that the shell and/or spar can be constructed from a high temperature material such as ceramics, Molybdenum or Niobium (columbium) or a lesser temperature resistive material such as Inco 718, Waspaloy or the well known single crystal material currently being used in gas turbine engines. For existing types of engine designs where it is desirable of providing efficacious turbine blade cooling with the use of compressor air at lower amounts and obtaining the same degree of cooling. For advanced engine designs where it is desirable to utilize more exotic materials such as Niobium or Molybdenum the shell and spar can be made out of these materials or the spar can be made from a lesser exotic material that is more readily cast or forged.

The material of the shell may be taken from a group consisting of stainless steel, molybdenum, niobium, ceramics, molybdenum alloys, or niobium alloys. The material of the spar may be taken from a group consisting of stainless steel, molybdenum, niobium, ceramics, molybdenum alloys, or niobium alloys.

Another feature of this invention for engine designs that require higher turbine rotational speeds, the spar can be made form a dual spar system where the outer spar extends a shorted distance radially relative to the inner spar and defines at the junction a mid span shroud and the shell is formed in an upper section and a lower section where each section is joined at the mid span shroud. The pin in this arrangement couples the inner spar and outer spar at the attachment formed at the root of the blade. This design can utilized the same materials that are called out in the other design.

A feature of this invention is an improved turbine blade that is characterized as being easy to fabricate, provide efficacious cooling with lesser amounts of cooling air than heretofore known designs, provides a shell or shells that can be replaced and hence affords the user the option of repair or replace. The materials selected can be conventional or more esoteric depending on the specification of the engine.

The foregoing and other features of the present invention will become more apparent from the following description and accompanying drawings.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

FIG. 1 is an exploded view in perspective showing the details of one embodiment of this invention.

FIG. 2 is a perspective view illustrating the assembled turbine blade of the embodiment depicted in FIG. 1 of this invention.

FIG. 3 is a section taken from sectional lines 3-3 of FIG. 2.

FIG. 4 is a section taken along the sectional lines 4-4 of FIG. 3 illustrating the attachment of the shell to the strut of this invention.

FIG. 5 is a perspective view illustrating a second embodiment of this invention.

FIG. 6 is a section view in elevation taken along the sectional lines of 6-6 of FIG. 5.

DETAILED DESCRIPTION OF THE INVENTION

While this invention is described in its preferred embodiment in two different, but similar configurations so as to take advantage of engine's that are designed at higher speeds than are heretofore encountered, this invention has the potential of utilizing conventional materials and improving the turbine rotor by enhancing its efficiency by providing the desired cooling with a lesser amount of compressor air, and affords the designer to utilize a more exotic material that has higher resistance temperatures while also maintaining the improved cooling aspects. Hence, it will be understood to one skilled in this technology, the material selected for the particular engine design is a option left open to the designer while still employing the concepts of this invention. For the sake of simplicity and convenience only a single blade in each of the embodiments is described although one skilled in this art that the turbine rotor consists of a plurality of circumferentially spaced blades mounted in a rotor disk that makes up the rotor assembly.

This disclosure is divided into two embodiments employing the same concept of a spar and shell configuration of a turbine blade, where one of the embodiments includes a single spar and the other embodiment includes a double spar to accommodate higher turbine rotational speeds. FIGS. 1 through 4 are directed to one of the embodiments of a turbine blade generally illustrated as reference numeral 10 as comprising a spar generally elliptical shaped spar 12 extending longitudinally or in the radial direction from the root portion 14 to the tip 16 with a downwardly extending portion 18 that fairs into a rectangular shaped projection 26 that is adapted to fit into the attachment 20. The spar 12 spans the camber stations extending along the airfoil section defined by the shell 28. The attachment 20 may include a fir tree attachment portion 22 that fits into a complementary fir tree slot formed in the turbine disk (not shown). The attachment 20 may be formed with the platform 24 or the platform may be formed separately and joined thereto and projects in the circumferential direction to abut against the platform in the adjacent blade in the turbine disk. A seal, such as a feather seal (not shown), may be mounted between platforms of adjacent blades to minimize or eliminate leakage around the individual blades.

The spar may be formed as a single unit or may be made up in complementary parts and as for example it may be formed in two separate portions that are joined at the parting plane along the leading edge facing portion 30 and trailing edge facing portion 32 and extending the longitudinal axis 31. Spar 12 is attached to the attachment 20 by the pin 34 which fits through the hole 29 in the attachment 20 and the aligned hole 31 formed in the extending portion 18. Pin 34 carries the head 36 that abuts against the face 38 of the attachment 20 and includes the flared out portion 40 at the opposing end of head 36. This arrangement secures the spar 12 and assures that the load on the blade 10 is transmitted from the airfoil section though the attachment 20 to the disk (not shown). The tip of blade may be sealed by a cap 44 that may be formed integrally with the spar 12 or may be a separate piece that is suitably joined to the top end of the spar 12. It should be appreciated that this design can accommodate a squealer cap, if such is desired. The material of the spar will be predicated on the usage of the blade and in a high temperature environment the material can be a molybdenum or niobium and in a lesser temperature environment the material can be a stainless steel like Inco 718 or Waspaloy or the like.

Shell 48 extends over the surface of the spar 12 and is hollow in the central portion 50 and spaced from the outer surface of spar 12. The shell defines the pressure side 52, the suction side 54, the leading edge 56 and the trailing edge 58. As mentioned in the above paragraph the shell 48 may be made from different materials depending on the specification of the gas turbine engine. In the higher temperature requirements, the shell preferably will be made from Molybdenum or Niobium and in a lesser temperature environment the shell 48 may be made from conventional materials. If the material selected cannot be cast or forged, then the shell will be made from a blank and the contour will be machined by a wire EDM process. The shell can be made in a single unit or can be made into two halves divided along the longitudinal axis, similar to the spar 12. As best seen in FIG. 1, the attachment 20 is made to include a stud portion 88 that complements the contoured surface of spar 12 and the contoured surface of shell 48. Additionally the shell 48 and spar 12 carry complementary male and female hooks 60 and 62. The top edge 80 of shell 48 is supported by the cap 44 and fits into an annular groove 82 so that the upper edge 84 of shell 48 bears against the shoulder 86. The lowered edge 88 fits into an annular complementary groove 90 formed on the upper edge of platform 24 and bears against the opposing surfaces of the groove 90 and the outer surface of the attachment 20.

As mentioned in the above paragraphs, one of the important features of this invention is that it affords efficacious cooling, i.e. cooling that requires a lesser amount of air. This can be readily seen by referring to FIG. 3. As shown the cooling air is admitted through the inlet 66, the central opening formed in the spar 12 at the bottom face 68 of the attachment 20, and flows in a straight passage or cavity 70 without having to flow through tortuous paths. The air that is admitted into cavity 70 flows out of the feed holes 72 into the space or cavity 74 defined between the spar 12 and the shell 48. Again, there are virtually no tortuous passages that are typically found in heretofore known designs and hence the pressure drop is decreased requiring lesser amount of air at a lower pressure, all of which enhances the cooling efficiency of the blade. The air from the feed holes 72, which may be formed integrally in the spar or drilled therein, can serve to impinge on the inner wall of the shell 48 but primarily feeds the space 74. It should be understood that this design can include film cooling holes (as for example holes 71 and 73) formed in the shell 48 on both the pressure surface 52 and the suction surface 54 and may also include a shower head (depicted as holes 75) on the leading edge and cooling holes (depicted as 77) on the trailing edge 58. The design and number of all of these cooling holes i.e. shower head, film cooling, feed holes and the like are predicated on the particular specification of the engine.

The other embodiment depicted in FIGS. 5 and 6 is similarly constructed and is adapted to handle a higher rotational speed of the turbine. In this embodiment the shell 104 that is equivalent to shell 48 depicted in FIGS. 14 is formed into two halves, the upper halve 106 and the lower halve 108 and the attachment 110 that is equivalent to the attachment 20 is extended in the longitudinal and upwardly direction to extend almost midway along the airfoil portion of the blade to form another spar 112. This spar 112 surrounds the lower portion 114 of spar 12 (like numerals in all the Figs. depict like or similar elements) and is contiguous thereto along its inner surface. A ledge or platen 116 is formed integrally therewith at the top end and extends in the spanwise direction. Shell 106 and shell 108 are formed in an elliptical-like shape to define the airfoil for defining the pressure surface 52, suction surface 54, leading edge 56 and trailing edge 58. A groove 115 formed at the upper edge 117 of shell 106 bears against the outer edge 118 of cap 120 which is the equivalent to cap 16 of FIGS. 13 except it is a squealer cap. Obviously, when the blade is rotating the shell 106 is loaded against the cap 120 and this force is transmitted to the disk via the spar 12 and lower portion 114. The lower edge 122 bears against the platen 116 and can be suitably attached thereto by a suitable braze or weld. The lower shell 108 is similarly formed like shell 106 and defines the lower portion of the airfoil. Lower shell 108 includes the groove 130 formed in the increased diameter portion 132 of shell 108 and serves to receive the outer edge 134 of platen 116. The lower edge 136 of shell 108 fits into an annular groove 138 formed in the platform 24. While not shown in these Figs. the male and female hooks associated with the spar and shell is also utilized in this embodiment and this portion of the drawings are incorporated herein by reference. The stud is like the embodiment depicted in FIGS. 13 is affixed to the attachment via pin 34.

The cooling arrangement of the embodiment depicted in FIGS. 5 and 6 is almost identical to the cooling configuration of the embodiment depicted in FIGS. 14. The only difference is that since the platen 116 forms a barrier between the upper shell 106 and lower shell 108, the cooling air to the lower portion of the airfoil is directed from the inlet 66 and passage 70 via the radial spaced holes 150 consisting of the aligned holes in the spars 12 and lower portion 114 that feeds space 156, and the holes 152 formed in the upper portion of the spar 12 that feed the space 158. As is the case with the embodiment of FIGs. 14, the shell may include a shower head at the leading edge, cooling passages at the trailing edge, holes at the tip for cooling and discharging dirt and foreign particles in the coolant and film cooling holes at the surface of the pressure side and suction side.

Although this invention has been shown and described with respect to detailed embodiments thereof, it will be appreciated and understood by those skilled in the art that various changes in form and detail thereof may be made without departing from the spirit and scope of the claimed invention.

Claims

1. A shell for use in a turbine airfoil constructed from a spar and shell; the shell comprising: the shell having an airfoil shape with a leading edge and a trailing edge, and a pressure side and a suction side extending between the two edges;the shell being formed from an electric discharge machining process; and,the shell being a thin shell wall to allow for near wall cooling of the inner surface of the shell.

2. The shell of claim 1, and further comprising: the shell being formed form a relatively high temperature resistant material that cannot be cast or machined as a thin wall shell.

3. The shell of claim 2, and further comprising: the shell being formed from Molybdenum or Niobium.

4. The shell of claim 1, and further comprising: the electric discharge machining process is a wire electric discharge machining process.

5. The shell of claim 1, and further comprising: the shell includes at least one hook extending from the inner surface of the shell to secure the shell to a spar of the turbine airfoil.

6. The shell of claim 5, and further comprising: the hook extends along the spanwise length of the shell.

7. The shell of claim 5, and further comprising: the hook is also formed from the electric discharge machining process and is formed as a single piece along with the shell.

8. The shell of claim 1, and further comprising: the shell includes a plurality of exit cooling holes in the trailing edge region of the shell to discharge cooling air out from the shell.

9. The shell of claim 5, and further comprising: the shell includes a second hook in the same side of the shell as the first hook; and,the two hooks face in opposite directions such that a chordwise movement of the shell with respect to a spar is limited.

10. The shell of claim 1, and further comprising: the shell has substantially the same thickness from the platform end to the tip end of the shell.

11. A turbine blade for use in a gas turbine engine, the turbine blade comprising: an attachment portion forming a platform and having a central opening;a spar having a tip on one end;a shell secured in place between the spar tip and the attachment portion; and,the shell being formed form a relatively high temperature resistant material that cannot be cast or machined as a thin wall shell.

12. The turbine blade of claim 11, and further comprising: the shell being a thin wall shell formed from an electric discharge machining process.

13. The turbine blade of claim 12, and further comprising: the electric discharge machining process is a wire electric discharge machining process.

14. The turbine blade of claim 11, and further comprising: the shell being formed from Molybdenum or Niobium.

15. The turbine blade of claim 11, and further comprising: the shell and the spar both include hook means extending from the shell and spar to secure the shell to a spar of the turbine airfoil.

16. The turbine blade of claim 11, and further comprising: the shell has substantially the same thickness from the platform end to the tip end of the shell.

17. The turbine blade of claim 11, and further comprising: the spar having an internal cooling supply passage, a plurality of exit cooling holes in the tip, and a plurality of near wall cooling holes to discharge cooling air from the internal cooling supply passage and onto the backside wall of the shell to provide near wall cooling for the blade.

18. The turbine blade of claim 17, and further comprising: the plurality of near wall cooling holes are located on the pressure side and the suction side of the blade.

19. The turbine blade of claim 17, and further comprising: the shell includes a row of trailing edge region exit cooling holes to discharge cooling air from a space formed between the spar and the shell.