This application is entitled to the benefit of British Patent Application No. GB 0724612.7, filed on Dec. 19, 2007.
The present invention relates to rotor blades.
Rotor blades are used in gas turbine engines to interact with combustion gases to convert kinetic energy of the combustion gases into rotation of the rotor. The efficiency of the engine is affected by the manner in which the combustion gases flow around the rotor blades.
Examples of the present invention provide a rotor blade having a trailing edge and a tip, the tip having an outer face, which includes at least two channels, each of the channels extending to an outlet in the vicinity of the trailing edge.
The blade may have a leading edge, at least one of the channels extending to the outlet from the vicinity of the leading edge. At least one of the channels may have an inlet in the vicinity of the leading edge. At least two channels may have an inlet in common. The outer face may include a common upstream channel, which bifurcates to provide the said two channels.
There may be at least two channels, which extend generally alongside each other over the uncovered turning region of the blade. There may be at least two channels, which extend to the outlet from a position approximately midway between the leading and trailing edges of the blade.
The two channels may be defined by at least three walls upstanding at the tip. One or more of the walls may be canted toward the pressure face of the blade.
The blade may further comprise at least one recess in the outer face, the recess being closed to the trailing edge.
In another aspect, examples of the invention provide a gas turbine engine comprising at least one rotor blade as aforesaid.
a to 8d are details corresponding with
Referring to
The gas turbine engine 10 operates in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 which produce two air flows: a first air flow into the intermediate pressure compressor 13 and a second air flow which provides propulsive thrust. The intermediate pressure compressor compresses the air flow directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high pressure compressor 14 is directed into the combustor 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive, the high, intermediate and low pressure turbines 16, 17 and 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16, 17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12 by suitable interconnecting shafts 26, 28, 30.
A ring of static nozzle guide vanes 50 are provided upstream of the blades 42, to further improve the flow characteristics of combustion gases through the blades 42, thereby increasing the efficiency of the stage.
The tips 44 rotate in close proximity with a fixed component 52, which may be a lining or shroud segment. The pressure difference at the two surfaces of the blades 42 creates a tendency for combustion gas leakage from the pressure side to the suction side, around the tip 44. Examples can now be described for arrangements to control leakage flow around the tips 44.
The two channels 58 are defined by walls 66 (
In the event that any gas leaks past the middle wall 66 to the second channel 58, a similar vortex effect will tend to occur, again resulting in drainage of the leakage gas along the channel 58 to the outlet 60, and creating crossflow to inhibit further leakage to the second channel 58.
The channels 58 therefore provide a form of gutter effect, tending to direct leakage gas to the outlet 60.
Leakage can be further inhibited in an optional manner indicated in
The two channels 58a are again defined by walls 66a generally as illustrated in
Leakage can be further inhibited by the use of canted walls, in the manner indicated in
The two channels 58b are again defined by walls 66b generally as illustrated in
Leakage can be further inhibited by the use of canted walls, in the manner indicated in
The views of
In
In
In any of the examples described above, the tip 44 could be formed to define at least one recess 76 in the outer face, as illustrated in
The walls 66 may carry abrasive material on the upper edges. This allows the engine to be set up with closer tolerances between the tip 44 and the fixed component 52, allowing the abrasive material to abrade the fixed component 52, in the event of contact. This closer tolerance further inhibits leakage flow. The wall 66 may be locally widened to accommodate abrasive material, if required.
It is envisaged that the examples illustrated above, and alternative examples, will exhibit improved blade tip sealing properties and thus reduce blade tip leakage flow. A reduction in blade tip leakage flow is expected to result in fewer losses arising from aerodynamic mixing of high and low pressure gases around the aerofoil and this, in turn, is expected to result in improved aerodynamic efficiency of each blade and thus of the stage, resulting in improved specific fuel consumption for the engine.
Improved blade tip sealing and reduced blade tip leakage flow is also expected to reduce the temperature of components of the blade, particularly at the tip, resulting in improved life expectancy and allowing the weight of components to be reduced to allow higher blade speeds or a reduction in stress within the components.
Passages can be incorporated within any of the examples described, to convey cooling air to outlets at appropriate positions.
Many variations and alternatives can be envisaged to the examples described above. Different wall shapes could be envisaged, and different wall sections could be used, such as canted, sloping or stepped sections. The depth and width of the various channels may be substantially constant along the length of the channel, or may be varied at different positions along the length. The exposed surfaces of the tip may be protected with anti-abrasion coatings.
In the examples described above, each gutter channel leads to a respective outlet, but a common outlet could be provided for one or more channels.
The examples described above can be used in turbine blades for aero engines, marine engines or industrial engines. The arrangements can be incorporated within cooled or uncooled turbine blades. The arrangements can be incorporated within turbine blades designed for contact or noncontact at their tips.
Number | Date | Country | Kind |
---|---|---|---|
0724612.7 | Dec 2007 | GB | national |
Number | Name | Date | Kind |
---|---|---|---|
2010094 | Leinweber | Aug 1935 | A |
3635585 | Metzler, Jr. | Jan 1972 | A |
3854842 | Caudill | Dec 1974 | A |
4390320 | Eiswerth | Jun 1983 | A |
4424001 | North et al. | Jan 1984 | A |
4606701 | McClay et al. | Aug 1986 | A |
5217349 | Succi | Jun 1993 | A |
5356265 | Kercher | Oct 1994 | A |
5503527 | Lee et al. | Apr 1996 | A |
5564902 | Tomita | Oct 1996 | A |
5733102 | Lee et al. | Mar 1998 | A |
5997251 | Lee | Dec 1999 | A |
6027306 | Bunker | Feb 2000 | A |
6059530 | Lee | May 2000 | A |
6190129 | Mayer et al. | Feb 2001 | B1 |
6616406 | Liang | Sep 2003 | B2 |
7513743 | Liang | Apr 2009 | B2 |
20010048878 | Willett et al. | Dec 2001 | A1 |
20050232771 | Harvey et al. | Oct 2005 | A1 |
20080170946 | Brittingham et al. | Jul 2008 | A1 |
Number | Date | Country |
---|---|---|
1591624 | Apr 2004 | EP |
1693552 | Jan 2006 | EP |
1865149 | Dec 2007 | EP |
1903183 | Mar 2008 | EP |
2075129 | Nov 1981 | GB |
Number | Date | Country | |
---|---|---|---|
20090162200 A1 | Jun 2009 | US |