Patent Application
20160230584
The present disclosure is generally related to rotating assemblies for turbomachinery and, more specifically, to a variable area turbine vane row assembly.
In gas turbine engines, energy is added to the air through the processes of compression and combustion, while energy is extracted by means of a turbine. In a turbofan engine, compression is accomplished sequentially through a fan and thereafter through a low-pressure compressor and high-pressure compressor, with the fan and low-pressure compressor being driven by a low-pressure turbine and the high-pressure compressor being driven by a high-pressure turbine through concentric shaft connections. Combustion occurs between the high-pressure compressor and the high-pressure turbine. Since the energy available to the turbines far exceeds that required to maintain the compression process, the excess energy is exhausted as high velocity gases through one or more nozzles at the rear of the engine to produce thrust by the reaction principle.
Typically, turbines within an engine are comprised of fixed geometries which are designed to provide balanced performance across a wide engine operating range. By introducing to a high-pressure and/or low-pressure turbine the ability to vary the turbine inlet flow area, the turbine can be adjusted to achieve optimal engine performance at multiple engine operating points. One metric of performance for the engine may be maximum achievable thrust. The level of engine thrust may be increased by allowing the high-pressure turbine to accept more flow for a given combustor exit temperature by increasing the high-pressure turbine inlet area, as governed by the flow area of the first row of high-pressure turbine vanes. Another metric of performance may be minimized fuel consumption. It is characteristic of some of these variable cycle engines that both the high and low pressure turbines contain mechanisms to allow for variable turbine inlet flow areas.
Various designs for providing a variable turbine inlet flow area using rotatable vanes have been proposed, but improvements are still needed in the art.
In one embodiment, a variable area turbine vane row assembly is disclosed, comprising: a first fixed vane; a second fixed vane proximate the first fixed vane; and a first rotatable vane asymmetrically positioned between the first and second fixed vanes.
In a further embodiment of the above, the first rotatable vane is circumferentially biased toward the first fixed vane.
In a further embodiment of any of the above, the first rotatable vane is circumferentially biased toward a suction side of the first fixed vane.
In a further embodiment of any of the above, the first rotatable vane is circumferentially biased toward a pressure side of the first fixed vane.
In a further embodiment of any of the above, the first rotatable vane is axially biased in an aft design direction with respect to the first and second fixed vanes.
In a further embodiment of any of the above, the first rotatable vane is axially biased in a forward design direction with respect to the first and second fixed vanes.
In a further embodiment of any of the above, the first rotatable vane is both circumferentially biased toward the first fixed vane and axially biased in a forward design direction with respect to the first and second fixed vanes.
In a further embodiment of any of the above, the first rotatable vane is both circumferentially biased toward the first fixed vane and axially biased in an aft design direction with respect to the first and second fixed vanes.
In another embodiment, a variable area turbine vane row assembly is disclosed, comprising: a first fixed vane; a second fixed vane proximate the first fixed vane; and a plurality of rotatable vanes positioned between the first and second fixed vanes; wherein no other fixed vanes are positioned between the first and second fixed vanes.
In a further embodiment of the above, the plurality of rotatable vanes comprises three rotatable vanes.
In a further embodiment of any of the above, the plurality of rotatable vanes comprises a first rotatable vane and a second rotatable vane.
In a further embodiment of any of the above, the first rotatable vane is asymmetrically positioned between the first fixed vane and the second rotatable vane.
In a further embodiment of any of the above, the first rotatable vane is circumferentially biased toward the first fixed vane.
In a further embodiment of any of the above, the first rotatable vane is circumferentially biased toward a suction side of the first fixed vane.
In a further embodiment of any of the above, the first rotatable vane is circumferentially biased toward a pressure side of the first fixed vane.
In a further embodiment of any of the above, the first rotatable vane is axially biased in an aft design direction with respect to the first and second fixed vanes.
In a further embodiment of any of the above, the first rotatable vane is axially biased in a forward design direction with respect to the first and second fixed vanes.
In a further embodiment of any of the above, the first rotatable vane is both circumferentially biased toward the first fixed vane and axially biased in a forward design direction with respect to the first and second fixed vanes.
In a further embodiment of any of the above, the first rotatable vane is both circumferentially biased toward the first fixed vane and axially biased in an aft design direction with respect to the first and second fixed vanes.
Other embodiments are also disclosed.
The embodiments and other features, advantages and disclosures contained herein, and the manner of attaining them, will become apparent and the present disclosure will be better understood by reference to the following description of various exemplary embodiments of the present disclosure taken in conjunction with the accompanying drawings, wherein:
For the purposes of promoting an understanding of the principles of the invention, reference will now be made to certain embodiments and specific language will be used to describe the same. It will nevertheless be understood that no limitation of the scope of the invention is thereby intended, and alterations and modifications in the illustrated device, and further applications of the principles of the invention as illustrated therein are herein contemplated as would normally occur to one skilled in the art to which the invention relates.
The exemplary engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided, and the location of bearing systems 38 may be varied as appropriate to the application.
The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged in exemplary gas turbine 20 between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 further supports bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A which is collinear with their longitudinal axes.
The core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path C. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion. It will be appreciated that each of the positions of the fan section 22, compressor section 24, combustor section 26, turbine section 28, and fan drive gear system 48 may be varied. For example, gear system 48 may be located aft of combustor section 26 or even aft of turbine section 28, and fan section 22 may be positioned forward or aft of the location of gear system 48.
The engine 20 in one example is a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about five 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.
A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.
In a gas turbine engine, such as the gas turbine engine 20, the gases from the combustor are directed toward a pair of turbines each including one or more rows of turbine vanes, wherein the vanes in each row are spaced apart circumferentially to direct the flow of combustion gases through the turbine.
The present disclosure provides in an embodiment for asymmetrical positioning of at least one vane from the vanes located on either side of it in a variable area turbine vane row assembly. For example,
The present disclosure further encompasses in an embodiment providing a plurality of rotating vanes between a pair of fixed vanes in a vane row, regardless of whether or not the vanes are asymmetrically spaced (i.e., circumferentially biased and/or axially biased with respect to each other as well as to the adjacent fixed vanes) or symmetrically spaced. For example,
Unlike other solutions where every vane in the vane row rotates or every other vane in the vane row rotates, the presently disclosed system of multiple rotatable vanes between successive fixed vanes balances the structural benefits of having fixed vanes with the aerodynamic benefit of having every vane rotate. The number of fixed vanes and rotating vanes in any vane row assembly will depend upon the particular design constraints of the engine, such as structural needs and vane aerodynamics such as endwall gap loss, flow passage non-uniformity reduction, etc.
While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only certain embodiments have been shown and described and that all changes and modifications that come within the spirit of the invention are desired to be protected.
The present application claims the benefit of and incorporates by reference herein the disclosure of U.S. Ser. No. 61/878,458 filed Sep. 16, 2013.
This invention was made with government support under Contract No. N00014-09-D-0821-0006 awarded by the United States Navy. The government has certain rights in the invention.
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/US2014/055743 | 9/16/2014 | WO | 00 |
| Number | Date | Country | |
|---|---|---|---|
| 61878458 | Sep 2013 | US |
