Patent Application
20060275107
This application discloses subject matter related to copending US patent applications “HAMMERHEAD FLUID SEAL” (APPLICANT REFERENCE NUMBER EH-11279) and “BLADE NECK FLUID SEAL” (APPLICANT REFERENCE NUMBER EH-11507) filed concurrently herewith.
(1) Field of the Invention
The invention relates to gas turbine engines, and more specifically to a seal for providing a fluid leakage restriction between components within such engines.
(2) Description of the Related Art
Gas turbine engines operate by burning a combustible fuel-air mixture in a combustor and converting the energy of combustion into a propulsive force. Combustion gases are directed axially rearward from the combustor through an annular duct, interacting with a plurality of turbine blade stages disposed within the duct. The blades transfer the combustion gas energy to one or more blades mounted on disks, rotationally disposed about a central, longitudinal axis of the engine. In a typical turbine rotor assembly, there are multiple, alternating stages of stationary vanes and rotating blades disposed in the annular duct.
Since the combustion gas temperature may reach 2000 degrees Fahrenheit or more, some blade and vane stages are cooled with a lower temperature cooling air for improved durability. Air for cooling the first-stage blades bypasses the combustor and is directed to an inner diameter cavity located between a first-stage vane support and a first-stage rotor assembly. The rotational force of the rotor assembly pumps the cooling air radially outward and into a series of conduits within each blade, thus providing the required cooling.
Since the outboard radius of the inner cavity is adjacent to the annular duct carrying the combustion gasses, it must be sealed to prevent leakage of the pressurized cooling air into the combustion gas stream. This area of the inner cavity is particularly challenging to seal due to the differences in thermal and centrifugal growth between the stationary, first-stage vane support and the rotating, first stage rotor assembly. In the past, designers have attempted to seal the outboard radius of inner cavities with varying degrees of success.
An example of such an outboard radius seal is a labyrinth seal. In a typical configuration, a multi-step labyrinth seal separates the inner cavity into two regions of approximately equal size, an inner region and an outer region. Cooling air in the inner region is pumped between the rotating disk and labyrinth seal into the hollow conduits of the blades while the outer region communicates with the annular duct carrying the combustion gases. A labyrinth seal's lands must be pre-grooved to prevent interference between the knife-edge teeth and the lands during a maximum radial excursion of the rotor. By designing the labyrinth seal for the maximum radial excursion of the rotor assembly, the leakage restriction capability is reduced during low to intermediate radial excursions of the rotor assembly. Any cooling air that leaks by the labyrinth seal is pumped through the outer region and into the annular duct by the rotating disk. This centrifugal pumping action increases the temperature of the disk in the area of the blades and creates parasitic drag, which reduces overall turbine efficiency. The rotating knife-edges also add additional rotational mass to the gas turbine engine, which further reduces engine efficiency.
Another example of such an outboard radius seal is a brush seal. As this example illustrates, a brush seal separates the inner cavity into two regions, an inner region and a smaller, outer region. A freestanding sideplate assembly defines a disk cavity, which is in fluid communication with the inner region. Cooling air in the inner region enters the disk cavity and is pumped between the rotating sideplate and disk to the hollow conduits of the blades. The seal's bristle to land contact pressure increases during the maximum radial excursions of the rotor and may cause the bristles to deflect and ‘set’ over time, reducing the leakage restriction capability during low to intermediate rotor excursions. Any cooling air that leaks by the brush seal is pumped into the outer region by the rotating disk. This centrifugal pumping action increases the temperature of the disk in the area of the blades and creates parasitic drag, which reduces overall turbine efficiency. The freestanding sideplate and minidisk also adds rotational mass to the gas turbine engine, which further reduces engine efficiency.
Although each of the above mentioned seal configurations restrict leakage of cooling air under certain engine operating conditions, a consistent leakage restriction is not maintained throughout all the radial excursions of the rotor. The seals may also increase the temperature of the disk and cooling air due to centrifugal pumping, reduce engine efficiency due to parasitic drag and add additional engine weight. What is needed is a seal that maintains a more consistent leakage restriction throughout all the radial excursions of the rotor, without negatively affecting disk and cooling air temperature, engine efficiency or engine weight.
In accordance with an embodiment of the present invention, there is provided a seal for restricting leakage of pressurized cooling air from an inner cavity flanked by a vane support and a bladed rotor assembly. The seal comprises a segmented ring defined by the bladed rotor assembly and a land defined by the vane support. The bladed rotor assembly includes a disk rotationally disposed about a central axis of the engine. The disk includes a radially outermost rim and a plurality of slots circumferentially spaced about the rim for accepting an equal plurality of blades. An interrupted rim region extends radially outward from a radius circumscribing a radially innermost floor of each slot to the outermost rim. The segmented ring extends from the interrupted rim region to define a segregated inner and outer cavity. The circumferential land is located radially above the inner cavity, proximate to the segmented ring. The segmented ring spans across the inner cavity, interacting with the land to define the seal.
By locating the seal radially outboard and in the interrupted rim region of the disk, temperature rise and parasitic drag due to duct placement and centrifugal pumping are minimized. Also, engine rotating mass is reduced with the elimination of freestanding sideplates and complex, multi-step labyrinth seal hardware as well.
Other features and advantages will be apparent from the following more detailed descriptions, taken in conjunction with the accompanying drawings, which illustrate by way of an example a seal in accordance with specific embodiments of the invention.
a-6h illustrate a series of enlarged schematics illustrating various seals of
The major sections of a typical gas turbine engine 10 of
Referring now to
During the operation of the engine 10, pressurized cooling air 40 is pumped into the inner cavity 50 by a duct 70, where a major portion of the cooling air 40 is dedicated to internally cooling the blades 58. The cooling air 40 enters the blades 58 via a series of radially extending conduits 72 communicating with a plenum 74 flanked by the blade attachment 66 and the disk 56. The cooling air 40 exits the blade 58 via a series of film holes 76. To ensure a continuous flow of cooling air 40 through the blades 58, the pressure of the cooling air 40 must remain greater than the pressure of the combustion gases 24 or the combustion gases 24 may backflow into the film holes 76, potentially affecting the durability of the blades 58.
An exemplary seal 80 in accordance with an embodiment of the invention separates the inner cavity 50 from the annular duct 30, thus ensuring adequate cooling air 40 pressure throughout all engine-operating conditions. The seal 80 is located radially inward of the annular duct 30, defining an outer cavity 82 therebetween. Since the outer cavity 82 is relatively small, any leakage of cooling air 40 through the seal 80 is subject to relatively minimal centrifugal pumping by the rotor assembly 54, prior to mixing with the combustion gases 24. This level of centrifugal pumping has limited negative impact on disk 56 temperature and aerodynamic drag, thus improving engine efficiency.
The exemplary seal 80 of
The segmented ring 86 is radially located in an interrupted rim region 110 of the disk 58. The interrupted rim region 110 extends radially outward from a radius 112 circumscribing a floor 114 of each slot 62 to the outer rim 60. As best shown in
A runner 170, also known as a knife-edge, extends outward from a segmented ring 86 as shown in
Referring now to
With the rotor assembly 54 installed in the high pressure turbine 18 as shown in
Although an exemplary seal 80 is shown positioned between a stationary member and a rotating member, it is to be understood that an exemplary seal 80 may also be located between two rotating members or two stationary members as well.
While the present invention has been described in the context of specific embodiments thereof, other alternatives, modifications and variations will become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications and variations as fall within the broad scope of the appended claims.
