This invention relates in general to a gas turbine engine and structure for directing compressed air directly on a blade ring.
Controlling gas turbine engine blade tip clearance is desirable so as to maintain engine structural integrity and efficient performance. Turbine efficiency improves as the clearance or gap between turbine blade tips and a surrounding static structure is reduced. The static structure comprises a blade ring coupled to an engine casing and a ring segment coupled to the blade ring via isolation rings. The ring segment is exposed to hot working gases passing through the gas turbine. During engine startup, the turbine blades radially expand quickly due to a rapid increase in the temperature of the hot working gases impinging and centrifugal forces acting on the blades. Also during start-up, the blade ring expands radially outward away from the blade tips as the temperature of the blade ring increases. However, the temperature of the blade ring increases to its steady state temperature at a slower rate than that of the blades during engine start-up. The diameter of the blade ring and the length of the blades are designed so that during engine startup, the tips of the blades do not contact an inner surface of the static structure ring segment. However, during steady-state operation, the gap between the blade tips and the static structure ring segment increases due to the blade ring temperature increasing.
In accordance with a first aspect of the present invention, a gas turbine engine is provided comprising a compressor for generating compressed air. The compressed air may increase in temperature from ambient when the gas turbine engine begins operation to an elevated temperature. The gas turbine engine may further comprise a turbine comprising a plurality of rows of vanes; a plurality of rows of rotatable blades; at least one static structure comprising a blade ring surrounding a corresponding row of vanes and a corresponding row of blades; and fluid structure for receiving compressed air from the compressor and extending toward the one stationary blade ring for discharging the compressed air directly against a surface of the blade ring at least during an initial startup period of the gas turbine engine such that the compressed air impinges on the blade ring surface.
The temperature of the compressed air may quickly increase to the elevated temperature after the gas turbine engine begins operation such that it transfers energy in the form of heat to the stationary blade ring during ramp-up of the gas turbine engine from about 0% load to about 100% load, thereby causing the stationary blade ring to move radially away from the corresponding row of blades.
The fluid structure may comprise at least one impingement pipe located adjacent the blade ring surface. The at least one impingement pipe may comprise a plurality of openings positioned so as to discharge the compressed air toward the blade ring surface. The at least one impingement pipe may extend circumferentially. The at least one static structure may further comprise a ring segment coupled to the blade ring and positioned between the blade ring and the corresponding row of blades.
The vanes of the corresponding row of vanes may comprise cooling passages which communicate with at least one corresponding opening in the one blade ring such that the compressed air passes through the vane passages after impinging upon the blade ring surface. The gas turbine engine may still further comprise a plurality of static structures comprising blade rings, each static structure surrounding a corresponding row of vanes and a corresponding row of blades.
In accordance with a second aspect of the present invention, a gas turbine engine is provided comprising a compressor for generating compressed air, a turbine and fluid structure. The turbine may comprise a plurality of rows of vanes; a plurality of rows of rotatable blades; and at least one static structure comprising a blade ring surrounding a corresponding row of vanes and a corresponding row of blades. Each of the vanes of the corresponding row of vanes may comprise a cooling passage. The blade ring may include at least one opening for communicating with the cooling passages of the corresponding row of vanes.
The fluid structure may receive compressed air from the compressor and extend toward the stationary blade ring for discharging the compressed air directly against a surface of the blade ring such that the compressed air impinges on the blade ring surface and then passes through the at least one opening in the stationary blade ring and into the cooling passages of the corresponding row of vanes. The temperature of the compressed air may quickly increase to the elevated temperature after the gas turbine engine begins operation such that it transfers energy in the form of heat to the stationary ring during ramp up of the gas turbine engine, thereby causing the stationary ring to move radially away from the corresponding row of blades. The compressed air may further function to cool the stationary ring during steady state operation of the gas turbine engine.
The fluid structure may comprise at least one impingement pipe located adjacent the blade ring surface. The at least one impingement pipe may comprise a plurality of openings positioned so as to direct the compressed air toward the blade ring surface.
The gas turbine engine may still further comprise a plurality of static structures comprising blade rings, each static structure surrounding a corresponding row of vanes and a corresponding row of blades. The fluid structure may discharge the compressed air in a direction away from the at least one opening in the blade ring.
In accordance with a third aspect of the present invention, a process for operating a gas turbine engine is provided. The gas turbine engine may comprise a compressor for generating compressed air and a turbine. The turbine may comprise a plurality of rows of vanes; a plurality of rows of rotatable blades; and at least one static structure comprising a blade ring surrounding a corresponding row of vanes and a corresponding row of blades. The process comprises discharging compressed air directly against a surface of the blade ring at least during an initial startup period of the gas turbine engine such that the compressed air impinges on the blade ring surface so as to increase the temperature of the blade ring surface. The discharging step may comprise discharging the compressed air continuously during substantially the entire operation of the gas turbine engine.
While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:
In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.
Reference is now made to
In the illustrated embodiment, the turbine 14 comprises a plurality of rows of vanes 20 and a plurality of rows of rotatable blades 22, see
In the illustrated embodiment, first, second, third and fourth static structures 26A-26D comprising first, second, third and fourth blade rings 28A-28D are provided. The first blade ring 28A generally surrounds the first row 20A of vanes 20 and the first row 22A of blades 22, the second blade ring 28B generally surrounds the second row 20B of vanes 20 and the second row 22B of blades, the third blade ring 28C generally surrounds the third row 20C of vanes 20 and the third row 22C of blades 22, and the fourth blade ring 28D generally surrounds the fourth row 20D of vanes 20 and the fourth row 22D of blades 22. Each of the blade rings 28A-28D comprises first and second generally semi-circular halves which are bolted together at their horizontal joints at assembly to form a complete cohesive blade ring (only the first halves of the blade rings 28A-28D are illustrated in
The first static structure 26A further comprises a first ring segment 30A, the second static structure 26B further comprises a second ring segment 30B, the third static structure 26C further comprises a third ring segment 30C and the fourth static structure 26D further comprises a fourth ring segment 30D. The first, second, third and fourth ring segments 30A-30D are generally axially aligned with and radially spaced a small distance from the first, second, third and fourth rows 22A-22D of blades 22.
Each vane 20 of the first, second, third and fourth rows 20A-20D of vanes comprises a vane platform 32A-32D.
The first, second, third and fourth ring segments 30A-30D and the first, second, third and fourth vane platforms 32A-32D cooperate to form an axially and circumferentially-extending wall that prevents hot gases from reaching the blade rings 28A-28D. Isolation rings 34 are coupled to the blade rings 28A-28D, the ring segments 30A-30D and the vane platforms 32A-32D so as to couple the ring segments 30A-30D and vane platforms 32A-32D to the blade rings 28A-28D. The ring segments 30A-30D and vane platforms 32A-32D are radially spaced from the blade rings 28A-28D to reduce heat transfer from the ring segments 30A-30D and vane platforms 32A-32D to the blade rings 28A-28D.
An impingement plate 36A-36D may be coupled to corresponding isolation rings 34 and located between each of the first, second, third and fourth rows 20A-20D of vanes 20 and a corresponding blade ring 28B-28D.
The turbine casing 38 of the illustrated embodiment fully surrounds the blade rings 28A-28D, see
As schematically shown in
The supply conduits 62, 64 extend through the turbine casing 38 so as to allow compressed air to enter the semi-cylindrical halves 40, 42 of the turbine casing 38, see
The fluid structure 16 further comprises, in the illustrated embodiment, circumferentially extending first and second impingement pipes 68 and 70 coupled to the impingement manifold 66. In the illustrated embodiment, the first and second impingement pipes 68, 70 are axially spaced from one another and located in the annular cavity 66A defined between the turbine casing 38 and the second blade ring 28B.
The annular cavity 66A may not extend 360 degrees, i.e., it may be restricted at 0 and 180 degree positions so as to define separate upper and lower cavity sections. In such an embodiment, each impingement pipe 68, 70 may comprise upper and lower halves received in the upper and lower cavity sections. Further, the manifold 66 may comprise upper and lower separate halves received in the upper and lower cavity sections.
Each impingement pipe 68, 70 comprises a plurality of openings 68A, 70A. As illustrated in
In the illustrated embodiment, the compressed air is discharged directly against the facing surfaces 128E and 128F and travels along those surfaces 128E and 128F so as to increase the heat transfer coefficient between the compressed air and the blade ring outer surface 78. The compressed air then flows into the openings 76 in the stationary blade ring 28B, which are generally located at a central axial location of the blade ring 28B in the illustrated embodiment. After flowing through the openings 76 and the impingement plate 36, the compressed air flows into cooling passages 80A provided in each vane 20 of the second row 20B of vanes 20. The cooling passage 80A extends from the vane platform 32B facing the blade ring 28B, into the vane 20 in a radial direction. The cooling passages 80A terminate at a radially-spaced row of discharge bores 80B extending to a trailing edge of the vane 20, see
Each impingement pipe 68, 70 may be insulated in order to reduce undesired heating or cooling of the compressed air before impingement onto the blade ring 28B.
The main conduit 56 may include a first electronically controlled proportional valve 60 (shown only in
The fluid structure 16 of the present invention preferably increases the heat transfer coefficient between the compressed air and the blade ring 28B in order to avoid the thermal expansion lag of the blade ring 28B during engine start-up, as found in the prior art.
In contrast,
Referring again to
At about 2500 seconds, the gas turbine engine reaches 100% load and begins steady-state operation at about 3000 seconds, see
It is further contemplated that the fluid structure 16 may also comprise third and fourth impingement pipes similar to the first and second impingement pipes 68 and 70, which may be positioned within an annular cavity defined between the turbine casing and the third blade ring 28C so as to increase the heat transfer coefficient between the compressed air and the third blade ring 28C during engine start-up. It is still further contemplated that the fluid structure 16 may additionally comprise fifth and sixth impingement pipes similar to the first and second impingement pipes 68 and 70, which may be positioned within an annular cavity defined between the turbine casing and the fourth blade ring 28D so as to increase the heat transfer coefficient between the compressed air and the fourth blade ring 28D during engine start-up.
While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.