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US 8,801,366 B2Granted
Stator blade for a gas turbine and gas turbine having same
AssigneeALSTOM Technology Ltd
InventorsRoland Dueckershoff, Shailendra Naik, Martin Schnieder
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AssigneeALSTOM Technology Ltd
InventorsRoland Dueckershoff, Shailendra Naik, Martin Schnieder
Estimating the term…
A stator blade for a gas turbine with sequential combustion, has a blade airfoil which extends in the radial direction between a blade tip and a shroud, with cooling passages extending inside the blade airfoil, through which a cooling medium can flow for cooling the blade and can then discharge from the stator blade into the hot gas flow flowing through the turbine. The blade airfoil has a sharply curved shape in space in the radial direction, and three cooling passages, which extend in the radial direction, arranged inside the blade airfoil in series in the hot gas flow direction and are interconnected by deflection regions, which are arranged at ends of the blade airfoil, so that the cooling medium flows through the cooling passages one after the other, with change of direction. The cooling passages follow the curvature of the blade airfoil in space in the radial direction.
A stator blade for a gas turbine, the stator blade having a blade airfoil extending in a radial direction between a blade tip and a shroud, with cooling passages extending inside the blade airfoil, through which a cooling medium flows for cooling the stator blade and then discharges from the stator blade into a hot gas flow which flows through the turbine, the blade airfoil having a sharply curved shape in space in the radial direction, a plurality of cooling passages, which extend in the radial direction, are arranged inside the blade airfoil, in series, in the hot gas flow direction and deflection regions, which are arranged at ends of the blade airfoil, are interconnected so that the cooling medium flows through the cooling passages one after the other, with change of direction, the cooling passages follow a curvature of the blade airfoil in space in the radial direction, wherein an entry of a first cooling passage, via a cooling air inlet which extends through a shroud, is in communication with an outer space which lies above the shroud, and the first cooling passage is connected to a start of a second cooling passage by a first deflection region, and a cross-sectional area at an entry of the second cooling passage is larger than a cross-sectional area at an outlet of the first cooling passage, and a throttling element, for throttling a mass flow of cooling medium which flows through the cooling air inlet, is arranged in a region of the cooling air inlet, the throttling element having the form of a plate which is provided with at least one opening, the throttling element closes off the entry of the cooling air inlet and is soldered into the shroud, wherein a ratio of the cross-sectional area at the entry of the second cooling passage to the cross-sectional area at the outlet of the first cooling passage is about 1.6 and wherein an end of the second cooling passage is connected to a start of a third cooling passage by a second deflection region, a throttling device, for the controlled extraction of cooling medium for the cooling of an outer platform and of a part of a trailing edge, is provided in the second deflection region, a cross-sectional area at an entry of the third cooling passage is smaller than a cross-sectional area at an outlet of the second cooling passage and cross-sectional areas of the three cooling passages halfway up the stator blade are in a ratio of 1:2:1.
A gas turbine with a stator blade having a blade airfoil extending in a radial direction between a blade tip and a shroud, with cooling passages extending inside the blade airfoil, through which a cooling medium flows for cooling the stator blade and then discharges from the stator blade into a hot gas flow which flows through the turbine, the blade airfoil having a sharply curved shape in space in the radial direction, a plurality of cooling passages, which extend in the radial direction, are arranged inside the blade airfoil, in series, in the hot gas flow direction and deflection regions, which are arranged at ends of the blade airfoil, are interconnected so that the cooling medium flows through the cooling passages one after the other, with change of direction, the cooling passages follow a curvature of the blade airfoil in space in the radial direction, an entry of a first cooling passage, via a cooling air inlet which extends through the shroud, is in communication with an outer space which lies above the shroud, and the first cooling passage is connected to a start of a second cooling passage by a first deflection region, and a cross-sectional area at an entry of the second cooling passage is larger than a cross-sectional area at an outlet of the first cooling passage, and a throttling element, for throttling a mass flow of cooling medium which flows through the cooling air inlet, is arranged in a region of the cooling air inlet, the throttling element having the form of a plate which is provided with at least one opening, the throttling element closes off the entry of the cooling air inlet and is soldered into the shroud, wherein the gas turbine functions with sequential combustion and has two combustion chambers in series in a flow direction, with a subsequent turbine in each case, the stator blade is arranged in the second turbine, wherein a ratio of the cross-sectional area at the entry of the second cooling passage to the cross-sectional area at the outlet of the first cooling passage is about 1.6 and wherein an end of the second cooling passage is connected to a start of a third cooling passage by a second deflection region, a throttling device, for the controlled extraction of cooling medium for the cooling of an outer platform and of a part of a trailing edge, is provided in the second deflection region, a cross-sectional area at an entry of the third cooling passage is smaller than a cross-sectional area at an outlet of the second cooling passage and cross-sectional areas of the three cooling passages halfway up the stator blade are in a ratio of 1:2:1.
Dependencies are inferred from the claim text.
Drawings are temporarily unavailable. Open the PDF at USPTO
From the US filing history. Foreign family members aren’t listed.
Hollow blades, i.e. blades with cooling or heating channels or cavities ; Heating, heat-insulating or cooling means on blades: convection cooling
Forming ring or sector: using blades
For a special turbine stage: the last stage of the turbine
Patterned: serpentine-like
About 2 years 5 months of term left · ends 12 Mar 2029 (estimated)
The present invention relates to the field of gas turbines. It refers to a stator blade for a gas turbine.
Large stationary gas turbines with sequential combustion have been proven in industrial application. In these gas turbines, two combustion chambers are arranged in series in the flow direction, and each with an associated turbine which are exposed to admission of the hot gas which is produced in the respective combustion chamber. Gas turbines of this type, which for example are offered by the Assignee of the present invention under the type designation GT24/26, are described for example in the printed publication by Joos F. et al., “Field experience with the sequential combustion system of the GT24/26 gas turbine family”, ABB Review 5/1998, p. 12-20 (1998). FIG. 1 of which is reproduced in the present application as FIG. 1. A further description of such a gas turbine originates furthermore from EP-B1-0 620 362.
FIG. 1 shows a gas turbine 10 with sequential combustion, in which a compressor 11, a first combustion chamber 14, a high-pressure turbine 15, a second combustion chamber 17 and a low-pressure turbine 18 are arranged in series along an axis 19. In summary, these gas turbines can be described as follows: The compressor 11 and the two turbines 15 (HD), 18 (ND) are part of a rotor which rotates around the axis 19. The compressor 11 compresses the inducted air, wherein this compressed air then flows into a plenum and from there flows into the first combustion chamber. This combustion chamber is operated with premix burners, as originate for example from EP-A1-0 321 809, and furthermore from EP-A2-0 704 657. The compressed air flows into the premix burners whereintermixing with at least one fuel takes place. This fuel/air mixture then flows into the first combustion chamber 14, in which this mixture combusts, forming a stable flame front. The hot gas which thus results is partially expanded in the adjoining high-pressure turbine 15, performing work, and after that flows into the second combustion chamber 17 where a further fuel feed 16 takes place. As a result of the high temperatures, which the hot gas which is partially expanded in the high-pressure turbine 15 always has, combustion takes place in the second combustion chamber 17, which is based on spontaneous ignition. The hot gas which is reheated in the second combustion chamber 17 is then expanded in a multistage low-pressure turbine 18, in which blade rows of rotor blades and stator blades are arranged in series in an alternating manner.
The invention shall subsequently be explained in more detail based on exemplary embodiments in conjunction with the drawing. All elements which are not necessary for the direct understanding of the invention have been omitted. Like elements are provided with the same designations in the different figures. The flow direction is indicated by arrows. In the drawing:
FIG. 1 shows the principle construction of a gas turbine with sequential combustion according to the prior art,
FIG. 2 shows in a side view a stator blade of a gas turbine with sequential combustion in accordance with FIG. 1 according to a preferred exemplary embodiment of the invention,
The invention should provide a remedy for the above mentioned limitations. It is therefore the object of the invention to create a stator blade which overcomes the limitations of known stator blades and enables an increase in the efficiency of the gas turbine.
The object is achieved by the entirety of the features of invention. It is preferable that the blade airfoil of the stator blade has a curved shape in space, even sharply curved at times, in the radial direction, that a number of cooling passages, which extend in the radial direction, are arranged in series in the direction of the hot gas flow and are interconnected by deflection regions, which are arranged at the ends of the blade airfoil, so that the cooling medium flows through the cooling passages one after the other, with change of direction, and that the cooling passages follow the curvature of the blade airfoil in space in the radial direction. As a result of the sharply curved shape of the blade airfoil, the interaction with the hot gas flow of the gas turbine is significantly improved. The configuration of the cooling passages, which follows the external shape of the blade airfoil, in this case enables improved cooling of the thermally highly stressed sectors of the stator blade.
One development of the invention is the entry of the first cooling passage, via a cooling air inlet which extends through the shroud, is in communication with the outer space which lies above the shroud, and a throttling element, for throttling the mass flow of cooling medium which flows through the cooling air inlet, is arranged in the region of the cooling air inlet. The entry-side throttling of the cooling medium flow which flows through the blade, by an independent throttling element, enables the precise adjustment of the cooling flow and so avoids unnecessary losses which have a negative effect upon the efficiency.