|Publication number||US3867068 A|
|Publication date||Feb 18, 1975|
|Filing date||Mar 30, 1973|
|Priority date||Mar 30, 1973|
|Also published as||CA994672A, CA994672A1, DE2413292A1|
|Publication number||US 3867068 A, US 3867068A, US-A-3867068, US3867068 A, US3867068A|
|Inventors||Robert J Corsmeier, Charles E Corrigan, Ronald E Dennis|
|Original Assignee||Gen Electric|
|Export Citation||BiBTeX, EndNote, RefMan|
|Patent Citations (2), Referenced by (36), Classifications (11)|
|External Links: USPTO, USPTO Assignment, Espacenet|
United States Patent [191 Corsmeier et al.
[ TURBOMACHINERY BLADE COOLING INSERT RETAINERS  Inventors: Robert J. Corsmeier; Charles E.
Corrigan; Ronald E. Dennis, all of Cincinnati, Ohio  Assignee: General Electric Company,
Cincinnati, Ohio  Filed? Mar. 30, 1973  Appl. No.: 346,422
FOREIGN PATENTS OR APPLICATIONS 4/1960 Great Britain 416/97 Feb. 18, 1975 Primary ExaminerEverette A. Powell, Jr. Attorney, Agent, or Firm-Derek P. Lawrence; Lee H. Sachs  ABSTRACT An improved turbomachinery blade includes impingement inserts adapted to direct a coolant toward surfaces defining an internal cavity within the blade. The inserts are provided with means which mechanically secure the inserts in place within the cavity. The securing means include a bushing having flanges of approximately the same thickness as that of the side walls of the insert, which flanges are secured within openings provided in such side walls near the bottom end of the inserts. A pin is inserted within holes provided in the root portion of the blade and a hole extending through the bushing, which is aligned with the holes in the root portion of the blade, and a collar of braze alloy is positioned around each end of the pin, thereby fluidically sealing each end of the pin and the insert and securing the pin to the blade and the insert in place within the cavity.
15 Claims, 10 Drawing Figures TURBOMACHINERY BLADE COOLING INSERT RETAINERS BACKGROUND OF THE INVENTION This invention relates generally to gas turbine engines and, more particularly, to an improved fluid cooled turbomachinery blade structure for use in high temperature gas turbines.
The invention herein described was made in the course of or under a contract, or a subcontract thereunder, with the United States Department of the Air Force.
It is well known that significant increases in gas turbine engine performance, in terms of thrust or work output per unit of fluid input, can be obtained by increasing the turbine inlet temperature of the motive fluid or hot gas stream. It is also recognized that one major limitation on turbine inlet temperature is that which is imposed by the turbine blade temperature capability. In an effort to extend turbine blade capabilities, numerous complex turbomachinery blade structures have been proposed which employone or more modes of cooling using fluid extracted from the compressor.
One such mode of cooling which is becoming increasingly popular is the provision of impingement inserts within an internal cavity defined by a hollow body portion of the turbomachinery blade. Coolant is delivered to the interior of such an insert and is expelled through a multiplicity of small holes against an internal wall of the turbomachinery blade, thereby cooling the portion of the turbomachinery blade which is exposed to the hot gas stream.
As is further known to those skilled in the art, the overall weight of a gas turbine engine, and in particular the thrust-to-weight ratio of the engine, is one of the engines most critical characteristics. The weight of the turbine rotor assembly, like that of any other turbomachinery component, must thereforebe maintained at the minimum practical levels obtainable. Not only must the wall thickness of the hollow body portion of the blade itself be maintained at a minimum thickness, but the wall thickness of the impingement insert also must be maintained at minimum practical levels. Wall thicknesses for such inserts on the order of 0.010 inch are becoming more and more prevalent. When one attempts to secure such a thin insert to the turbomachinery blade, however, problems arise due to the high centrifugal force asserted on such an insert.
Prior attempts at securing such a blade insert in place have included the provision of an enlarged flange formed integrally with the insert at the bottom thereof and the further provision of a load-bearing surface upon which such flange would rest. Such a design is shown in US. Pat. No. 3,715,170 Savage et al., which patent is assigned to the same assignee as the present application. Inserts provided with such a load-bearing flange have proven extremely successful, but in many applications such a flange is not practical. For example, as briefly mentioned above, due to the increase in tip speed of newer turbine rotors, the blade must be made as lightweight as possible to keep the disc size and weight at a minimum. One method of keeping the blade weight low is to make the airfoil walls as thin as possible. In designing such a blade, a minimum practical wall thickness is chosen for the tip portion of the blade, but the airfoil wall thickness must increase as it nears the blade platform in order to carry the increasing load of the airfoil. This centrifugal loading and the radial temperature gradient of the airfoil combine to produce blades having walls that are constant in thickness from the tip to a point above the pitch line, thereafter increase in thickness to a short distance below the pitch line, and then become essentially constant in thickness to the airfoil root portion. With these changes in wall thickness, the internal cavity formed by the hollow body portion of the blade also changes in size, with a blade as described above providing a bottle-shaped cavity with the largest opening being near the tip of the blade.
As is well known to those skilled in the art, it is highly desirable to provide a constant clearance between the impingement inserts and the inner wall of the blade. To install a baffle from the root end of the blade would be. virtually impossible if such a specified clearance must be maintained between the baffle and this bottleshaped cavity. Support embossments located on the exterior of the inserts further hinder the assembly. To further complicate matters, many airfoil portions are twisted. Since the impingement inserts cannot be installed from the root end of the airfoil, they must necessarily be installed from the blade tip end; and, this makes it virtually impossible to utilize an insert having an enlarged load-bearing flange at its inner end.
Attempts have also been made in the past to pin the bottle-shaped inserts in place, but the base wall thickness of 0.010-0.020 inch provides insufficient strength to withstand the loads imposed by the pins due to the centrifugal force exerted on the blade and insert during normal operation of the gas turbine. Another alternative is to braze the insert directly to the blade member, but the location of the braze joint in an area of only limited access makes itdifflcult to properly place the braze and to inspect the joint.
SUMMARY OF THE INVENTION It is an object of this invention, therefore, to provide a means for mechanically securing a cooling insert in a turbine blade which means overcomes the abovedescribed problems associated with prior art designs. It is a further object to provide such a positive mechanical securing means which does not substantially interfere with the coolant flow, is easy to assemble, and sub stantially eliminates leakage of coolant.
Briefly stated, the above objects are attained in the present instance by providing an insert which includes an opening which is drilled or punched through both sides thereof near the bottom end thereof. A grommettype bushing is placed in the openings and secured to both sides of the insert. A hole is then drilled through the bushing and the insert is pushed into the blade internal cavity until the hole in the bushing lines up with suitable holes in the shank portion of the blade. A pin is then inserted into the aligned holes and braze alloy is placed around each end of the pin and the insert base in order to seal the ends of the pin and the insert base and retain the pin in place. The bushing is provided with flanges of approximately the same thickness as that of the walls of the insert so as to lend itself to an optimum weldment.
DESCRIPTION OF THE DRAWINGS While the specification concludes with a series of claims which distinctly claim and particularly point out the subject matter which Applicants consider to be their invention, a complete understanding of such invention will be obtained from the following detailed description, which is given in connection with the accompanying drawings, in which:
FIG. 1 is a partial, cross-sectional view diagrammatically showing one installation of the inventive insert retaining means of the present invention;
FIG. 2 is a cross-sectional view taken along line 22 of FIG. 1; 7
FIG. 3 is a cross-sectional view taken along line 33 of FIG. 1;
FIG. 4 is an enlarged, partial, cross-sectional view showing the details of a portion of FIG. 1;
FIG. 5 is a partial, sectional view of an alternative retaining means;
FIG. 6 is a cross-sectional view taken generally along line 66 of FIG. 5;
FIG. 7 is an axial cross-sectional view of a second alternative retaining means;
FIG. 8 is a cross-sectional view, with portion deleted, taken generally along line 8-8 of FIG. 7;
FIG. 9 is a partial sectional view, similar to FIG. 5, of another alternative embodiment; and
FIG. 10 is a cross-sectional view, with portions deleted, taken generally along line 10-10 of FIG. 9.
DESCRIPTION OF A PREFERRED EMBODIMENT Referring to the drawings wherein like numerals correspond to like elements throughout, attention is directed initially to FIGS. 1 through 4 wherein a turbomachinery blade incorporating the present inventive insert retaining means is generally designated by the numeral 10. The blade 10 includes a root portion 12 which provides a pair of tangs 14 adapted to mount within a dovetail slot (not shown) associated with a gas turbine rotor disc (not shown). The blade 10 further includes an airfoil-shaped, hollow body portion 16 which is separated from the root portion 12 by means of a blade platform 18.
The airfoil-shaped, hollow body portion 16 includes a leading edge 20 and a trailing edge 22 and a pair of airfoil-shaped side walls 24 and 26 extending therebetween which are suitably formed and adapted to extract energy from a motive fluid flowing thereacross. The side walls 24 and 26 cooperate to define an internal cavity 28 which, in the present instance, is divided into two separate sections, leading edge cavity 30 and trailing edge cavity 32 by means of a rib member 34, which is formed integrally with hollow body portion 16 and extends between the side walls 24 and 26.
As most clearly shown in FIG. 2, the side walls 24 and 26 vary in thickness from a relatively thin cross section near the tip of the hollow body portion 16 to a thicker cross section near the blade platform 18. As a result of this varying thickness, the leading edge cavity 30 and the trailing edge cavity 32 are bottle-shaped cavities which are wider near the tip end of the hollow body portion 16 than near the blade platform 18.
As further shown in FIGS. 1 and 2, a thin sheet metal impingement insert 36 is positioned within the leading edge cavity 30, while a similar impingement insert 38 is located within the trailing edge cavity 32. Each of the inserts 36 and 38is shown as being a thin-walled shell having side walls 40 and 42 generally conforming to the blade side walls 24 and 26. The side walls 40 and 42 are formed with a plurality of nozzles or apertures 44 which are disposed in spaced relationship to the inner surfaces of the cavities 30 and 32 by suitable spacing means 46. As best shown in FIG. 2, the inserts 36 and 38 are closed at one end 48 and define a chamber 50 therein which is open at end 52 to receive coolant fluid.
In order to direct the coolant fluid into the chambers 50, suitable passage means 54 are provided in the root portion 12 of the blade 10. Furthermore, the open or base end 52 of the inserts 36 and 38 are sized so as to sealingly engage side walls 56 located near the top of the passage means 54. As described in greater detail hereafter, the joint b etween the base end 52 and the side walls 56 can be filled with a braze alloy to further seal this joint. In this manner. coolant delivered through the passage means 54 is directed to the chamber 50 and thus to the plurality of apertures or nozzles 44 and is not permitted to pass around the open ends 52 of the inserts 36 and 38 directly into the cavities 30 and 32. The coolant fluid, in passing through the apertures 44, impinges directly on the inner sides of the leading edge 20 and the side walls 24 and 26 of the blade 10, thereby cooling such walls in a known manner. The coolant fluid in the leading edge cavity 30 then exits through openings 60 (one of which is shown in FIG. 1) provided in a tip cap 62, which defines the outer end of the cavities 30 and 32, and through film openings (not shown) in the hollow body portion 16, while the coolant fluid in the trailing edge cavity 32 exits through openings in the tip cap and through trailing edge slots 64 located in the trailing edge 22 of the blade 10.
Referring still to FIGS. 1 through 4, means are provided for mechanically securing the inserts 36 and 38 into the position shown in FIG. 1. The means for mechanically securing the inserts in place are generally designated by the numeral 66 and include a grommettype bushing 68 which is positioned within openings 70 and 72 located in the side walls 40 and 42 of the inserts 36 and 38. Since the means for mechanically securing the inserts 36 and 38 in place are substantially identical, only the structure associated with the insert 36 will be described.
As best shown in FIGS. 3 and 4, the bushing 68 includes a pair of thin flanges 74 and 76 which are interconnected by means of an inner spool 78. The flanges 74 and 76 are sized so as to correspond in shape to that of the openings 70 and 72 associated with the side walls 40 and 42. In the embodiment shown in FIGS. 1 through 4, the openings 70 and 72 are made to an approximately circular configuration, although other configurations would also be suitable. Furthermore, as clearly shown in FIG. 3, the flanges 74 and 76 are made of approximately the same thickness as that of the side walls 40 and 42 so as to provide a configuration which lends itself to an optimum weldment. In addition. the openings 70 and 72 provide a much larger perimeter than that of the inner spool 78 so as to provide a large weld area which further enhances the strength of the weldment.
As best shown in FIGS. 3 and 4, the inner spool 78 is provided with a pair of flats 80 and 82 which extend parallel to the side walls 56 of the passage means 54 when the flanges 74 and 76 are positioned within the openings 70 and 72. In this manner, the flats 80 and 82 permit the correct volume of air to pass through the open end 52 and into the chamber 50 of the insert 36.
Prior to assembling the insert 36 into the blade 10, the bushing 68 is positioned within the insert and is welded in place. A hole 84 is then drilled through the center of the bushing 68, and the insert 36 is positioned within the hollow body portion 16 of the blade through the open end thereof. The hole 84 is aligned with a pair of holes 86 and 88 located within the root portion 12 of the blade 10. A pin 90 is then inserted through the holes 84, 86 and 88, as shown in FIG. 3.
The holes 86 and 88 in the root portion of the blade 10 are made oversized with respect to the diameter of the pin 90 so as to allow for the tolerance build-up of the various parts and also the placement of a braze collar around the ends of the pin 90. The hole 84 is sized so as to fit relatively securely around thepin 90 was to provide accurate placement of the insert 36 within the hollow body portion 16. The braze alloy collar which is placed around both ends of the pin 90 not only secures the pin 90 to the root portion 12 of the blade but also acts as a seal which precludes the flow of coolant around the pin 90 before it enters the chamber 50 associated with the insert 36. Furthermore, if desired, the base end 52 of the insert can be sized so as to provide a slight gap between the outer wall of the insert 36 and the side walls 56 of the blade. In such a case, the braze not only seals the pins 90 but also flows into the gap between the insert and the wall 56 and effectively seals this gap. In this manner, all coolant which flows through the passage means 54 must flow into the chamber 50 associated with one of the inserts 36 or 38.
In assembling the inserts into the blade, once the inserts 36 and 38 are positioned within the hollow body portion 16 and the pins 90 are located within the holes 84, 86 and 88, the braze alloy is placed around each of the pins 90. The tip cap 62 is then located in place, a suitable braze alloy is applied to the tip cap 62, and the entire blade is then heated which causes the braze alloy to melt, thereby sealing each end of the pins 90 and the base ends of the inserts in addition to securing the pins 90 and the tip cap 62 in their respective positions.
Referring now to FIGS. 5 through 10, a number of alternative embodiments of the mechanical securing means for cooling inserts are shown. As shown in FIGS. 5 and 6, a bushing 94 having a pair of relatively elongated flanges 96 and 98 could be utilized in place of the bushing 68. In such a case, the flanges could extend down to the open end of the insert. An inner spool 100 interconnects the flanges 96 and 98, and, if desired, flats 102 and 104 can be provided on either side of the spool 100. As was the case of the bushing 68, once the bushing 94 is welded to the insert 36, a hole 106 is drilled therethrough and a pin 108 is positioned therein in such a manner as to secure insert 36' to the blade 10.
As shown in FIGS. 7 and 8, in certain applications a bushing 110 is designed to fit within the open end 52 of the insert 36. In such a case, the bushing 110 would have an outer contour shaped in the form of an airfoil and sized so as to fit within an insert 36". The bushing 110 would be hollow and would provide a passageway 112 for the flow of coolant into a chamber 50" of the insert 36''.
In assembling the insert 36 into the blade, the bushing 110 could be initially brazed in place within the insert 36", and the insert 36" thereafter positioned within the blade 10. Holes 114 and 116 could then be drilled through the root portion 12 and each side of the insert 36. Pins 118 and 120 would then be positioned within the holes 114 and 116 and would act to secure the bushing and, thus, the insert 36" in place.
While the bushing 68 described in connection with FIGS. 1 through 4 provides an extremely secure assembly, in certain applications a simple cylindrical collar or a tear-shaped bushing such as that shown at in FIGS. 9 and 10 may be all that is required to secure the insert in place. In such a case, the bushing 130 would be welded or brazed to the inner sides of the insert 36" and a hole 132 would be drilled through the insert 36" and bushing 130. Once the insert 36" is positioned within the blade 10 a pin 134 would be inserted within the hole 132 and brazed in place as described above.
While a number of suitable embodiments of Applicants inventive means for securing cooling inserts to the interior of gas turbine blades have been described, it should be readily apparent to those skilled in the art that slight modifications could be made in these em bodiments without departing from the broader inventive concepts described herein. For example, while each of the above embodiments is described in connection with an open-ended blade and a removable tip cap, the inventive securing means could be readily adapted for use in applications wherein the inserts are inserted through the dovetail portion of the blade. Similarly, the shape of the flanges 74 and 76 can be readily changed without departing from the broad inventive concepts. For this reason, the appended claims are intended to cover these and all similar modifications which fall within the scope of the invention.
What is claimed is:
1. In a turbomachinery blade of the type including a root portion, a blade platform, a hollow body portion defining an internal cavity, at least one thin-walled, hollow insert adapted to be positioned within said hollow body portion, and means for delivering a coolant to the interior of said insert, the improvement comprising:
means for mechanically securing said insert to said hollow body portion, said securing means including a bushing adapted to be secured to said insert, and pin means adapted to be secured to both said bushing and to said blade.
2. The improved blade of claim ll wherein said insert comprises an impingement insert having an inlet at one end thereof and a multiplicity of relatively small holes therethrough which are directed toward surfaces defining said cavity and are adapted to impinge cooling air thereagainst.
3. The improved blade of claim 2 wherein said bushing is adapted to be positioned within said inlet, and said bushing includes at least one flange of approximately the same thickness as that of said insert wall, and said wall includes an opening for receiving said flange.
4. The improved blade of claim 3 wherein said bushing includes a second flange of approximately the same thickness as that of said insert wall, said wall includes a second opening for receiving said second flange and said first and second flanges are adapted to be secured to said insert when said flanges are positioned within said openings.
5. The improved blade of claim 3 wherein said insert includes a base end which defines said inlet, and said base end is sized so as to provide a slight gap between side walls of said internal cavity and said base end, and said gap is filled with a sealant.
6. The improved blade of claim 2 wherein said bushing comprises a sleeve adapted to be positioned within said inlet, and said sleeve is adapted to be secured to the inner walls of said insert.
7. A turbomachinery blade comprising a root portion, an airfoil-shaped, hollow body portion defining an internal cavity therein, a blade platform separating said root portion and said hollow body portion and defining the inner bounds of said airfoil, at least one thin-walled, hollow insert positioned within said hollow body portion, passage means adapted to deliver a coolant to the interior of said insert, and means for mechanically securing said insert in a position within said internal cavity, said securing means including a bushing adapted to be secured to said insert and pin means adapted to be secured to both said bushing and to said blade.
8. The turbomachinery blade recited in claim 7 wherein said insert comprises an impingement insert having an inlet at one end thereof and a multiplicity of relatively small holes therethrough which are directed toward surfaces defining said cavity and are adapted to impinge cooling air thereagainst.
9. The turbomachinery blade recited in claim 8 wherein said bushing includes at least one flange of approximately the same thickness as that of said insert wall, said wall includes an opening for receiving said flange, and said blade further includes means for securing said flange in said opening.
10. The turbomachinery blade recited in claim 9 wherein said bushing includes a second flange of approximately the same thickness as that of said insert wall, said wall includes a second opening for receiving said second flange, and means for securing said second flange in said second opening.
11. The turbomachinery blade recited in claim 9 wherein said bushing includes an inner spool, and said spool is adapted to extend across said inlet of said insert.
12. The turbomachinery blade recited in claim 11 wherein said spool includes at least one flat side.
13. The turbomachinery blade recited in claim 8 further including means for dividing said internal cavity into at least two separate cavities, each of said cavities is provided with one of said impingement inserts, and each of said cavities includes means for mechanically securing said insert to said blade.
14. The turbomachinery blade recited in claim 9 wherein said root portion of said blade includes at least one hole, said bushing includes at least one holeadapted to align with said hole in said root portion, and said pin means comprise a pin adapted to be positioned within each of said holes.
15. The improved turbomachinery blade recited in claim 14 wherein said hole in said root portion is slightly larger than the outer diameter of said pin and a braze collar is positioned within the gap between said pin and said hole.
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|U.S. Classification||416/97.00R, 416/96.00A|
|International Classification||F04D29/32, F01D5/14, F04D29/54, F01D5/18|
|Cooperative Classification||F01D5/147, F05D2260/201, F01D5/189|
|European Classification||F01D5/14C, F01D5/18G2C|