Patent Application
20170298739
The present application relates to gas turbines, and more particularly to a device to route cooling air to a turbine blade. A system to improve a flow of rotor cooling air and a method to improve a flow of rotor cooling air to a turbine blade are also provided.
During operation of the gas turbine, turbine blades are exposed to extremely high temperatures. Various methods are employed for their cooling, including routing rotor cooling air to the turbine blades. Traditionally, an air separator is used to separate the air into two paths, one leading into the row one turbine disc, also referred to as the turbine disc one, for cooling of the row one blade platform and the other path leading to the rotor for cooling of the rotor discs and turbine blades. After the air separator routes the air to the rotor, the air is then brought up to the rotational speed of the rotor. This process incurs undesirable aerodynamic losses as the work of the rotor associated with bringing the air up to rotational speed is high. A pre-swirler device may be used to impart tangential momentum in order to get the rotor cooling air up to the rotational speed of the rotor quicker than the process used with the air separator. Using the pre-swirler device to swirl the incoming rotor cooling air reduces losses and improves the overall efficiency of the gas turbine which leads to the improved cooling ability of the rotor cooling air to cool the turbine blades.
Briefly described, aspects of the present disclosure relate to a device to route cooling air to the turbine blade.
A first aspect of provides a device to route cooling air to a turbine blade. The device includes a seal ring having an L-shaped cross section configured to abut a turbine disc. The seal ring comprises a radial portion extending radially with respect to a rotor and an axial portion extending axially with respect to the rotor. The seal ring also comprises a plurality of radial cooling holes disposed within the radial portion of the seal ring and arranged circumferentially around the seal ring. The plurality of cooling holes route cooling air from a device configured to impart tangential momentum to the cooling air to a turbine blade in order to cool the turbine blade.
A second aspect provides a system to improve a flow of rotor cooling air to a turbine blade. The system includes a swirler device configured to swirl a rotor cooling air with a rotation of the gas turbine. The system also includes a turbine disc and an L-shaped seal ring abutting the turbine disc and configured to route the rotor cooling air through a plurality of radial cooling holes within the seal ring from the swirler device to a turbine blade in order to cool the turbine blade.
A third aspect of provides a method to improve a flow or rotor cooling air to a turbine blade. The method includes swirling rotor cooling air such that the cooling air is rotating at the speed of the rotor and routing the swirled cooling air to a turbine blade through a radial hole in an L-shaped ring for cooling.
To facilitate an understanding of embodiments, principles, and features of the present disclosure, they are explained hereinafter with reference to implementation in illustrative embodiments. Embodiments of the present disclosure, however, are not limited to use in the described systems or methods.
The components and materials described hereinafter as making up the various embodiments are intended to be illustrative and not restrictive. Many suitable components and materials that would perform the same or a similar function as the materials described herein are intended to be embraced within the scope of embodiments of the present disclosure.
The pre-swirler device, as described above, would physically replace the air separator in existing gas turbines. In order to integrate the pre-swirler into existing gas turbines, an additional device may be needed to replace the functionality of the air separator; that is to separate the cooling air into the two paths, one path into the row one turbine blade platform and the other path to the turbine rotor discs and turbine blades. Additionally, the additional device may be needed to provide a sealing function for the pre-swirler housing. The additional device and its functionality separating the cooling air into two paths may also be incorporated into the design of a new engine.
The seal ring device 10 is disposed between the pre-swirler device 30 and the turbine disc one 20 and above the cavity 90 into which the swirled rotor cooling air F enters at the speed of the rotor after being expelled by the pre-swirler nozzle 40. The turbine one disc 20 includes multiple radial cooling passages, of which one is illustrated in the Figures, 50 through which a portion of the rotor cooling air flows radially to the turbine blade for its cooling. Additionally, an axial cooling passage 60 exists in the turbine disc one 20 for a further portion of the rotor cooling air F to flow in order to cool the further stages of turbine discs and turbine blades.
A contour of the radially interior surface of the seal ring 10 may be optimized using computational fluid dynamics such that the pressure loss of the rotor cooling air is reduced and the performance of the rotor cooling air to cool the turbine blades is improved. In the embodiment shown in
A radially exterior surface of the axial portion 130 may be adapted to accommodate the pre-swirler sealing 70. The pre-swirler sealing 70 may be designed to minimize the leakage of cooling air through the seal 70. Additionally, the pre-swirler sealing 70 keeps the cool air at a higher pressure within the cavity 90 in order to force the cool air into the turbine blades. In the embodiment shown in
In the embodiment shown in
The seal ring 10 may comprise the same or similar material as the turbine one disc 20. Using the same or similar material for the seal ring 10 as that of the turbine one disc 20 would prevent significant differences in the rate of thermal expansion between the two components during operation of the gas turbine. A significant difference in the rate of thermal expansion may cause the misalignment of the radial cooling hole 11 and the radial cooling hole 50 of the turbine disc one 20 such that the amount of the cooling air reaching the turbine 1 blade would decrease, for example. The seal ring 10 may thus comprise a low alloy steel which is traditionally used for the turbine disc material.
The seal ring 10 is configured to abut the turbine disc 20 such that the cooling passage 50 of the turbine one disc is aligned with the radial cooling hole 110 of the seal ring 10. The seal ring 10 may also comprise attachment means to attach the seal ring 10 to the turbine one disc 20. As illustrated in
Referring to
The seal ring 10 includes a plurality of radial cooling holes 110 extending radially through the radial portion 120 of the seal ring 10. These radial cooling holes 110 may be aligned with radial cooling passages 50 within the turbine disc one 20 such that the cooling air F is efficiently routed from the pre-swirler device 30 to the turbine blade. The turbine disc 20 may also include an axial cooling passage 60 which routes cooling air to turbine blades in a flow direction downstream from the turbine disc one 20. An example of a cooling air split between the radial cooling passage 50 and the axial cooling passage 60 may be 50% through the radial cooling passage 50 and 48% through the axial cooling passage 60 with approximately 2% lost through leakage. The turbine blades themselves control the amount of cooling air flow they consume. The more cooling holes 110 each turbine blade includes, the higher amount of cooling air flow the turbine blade takes in.
Referring to
In an embodiment, the method includes attaching the seal ring 10 to a turbine disc one 20. The seal ring 10 may be attached to the turbine disc one 20 using through bolts, sheer pins, interference fits, or by welding as described above. In an embodiment, a plurality of through holes 150 may be positioned in a radial portion 120 of seal ring 10 through which a bolt or sheer pin may be inserted, for example, and fastened in order to securely attach the seal ring 10 to the turbine disc one 20. In another embodiment, the seal ring 10 is welded to the turbine disc one 20.
The attaching of the seal ring 10 may include aligning a plurality of radial cooling holes 110 in the seal ring 10 with a corresponding cooling passage 50 in the turbine one disc 20 such that the flow of cooling air cools the row one turbine blade. An interference fit may be provided between the seal ring 10 and the turbine disc 20 by heating up the turbine disc 20 to center its cooling passage 60 with the radial cooling hole 110 of the seal ring 10.
In an embodiment, especially when retrofitting an existing gas turbine with a seal ring 10, the turbine disc 20 may need to be machined in order to accommodate the geometry of the seal ring 10 such that the seal ring 10 abuts the turbine disc 20. The machining would precede the attaching of the seal ring 10.
While embodiments of the present disclosure have been disclosed in exemplary forms, it will be apparent to those skilled in the art that many modifications, additions, and deletions can be made therein without departing from the spirit and scope of the invention and its equivalents, as set forth in the following claims.
