GOVERNMENT LICENSE RIGHTS

This invention was made with United States Government support under contract number DE-FE0023975 awarded by Department of Energy. The United States Government has certain rights in the invention.

TECHNICAL FIELD

The present invention relates generally to a gas turbine engine, and more specifically to an air-cooled turbine rotor blade for a closed loop cooling circuit in an industrial gas turbine engine.

BACKGROUND

In a gas turbine engine, compressed air is burned with a fuel in a combustor to produce a very high temperature gas stream that is then passed through a turbine that uses some of the power produced to drive a compressor and, in the case of an industrial gas turbine engine, an electric generator to produce electrical power. The higher the hot gas stream temperature entering the turbine, the more power the turbine can generate and the more efficient will be the engine. The turbine inlet temperature is limited to the material properties of the parts exposed to the hot gas stream and to the effectiveness of cooling of the parts. The coolant used to cool the turbine hot parts such as rotor blades and stator vanes is typically compressed air bled off from the compressor, and thus the more compressed air used for cooling lowers the efficiency of the engine since the work done on compressing the cooling air is discharged into the hot gas stream and does not perform any work on the compressor or electric generator. When most or all of the cooling air for the turbine hot parts is discharged into the hot gas stream, this is referred to as an open cooling circuit for the turbine.

FIG. 1 shows a prior art turbine rotor blade with an internal cooling circuit (that shown in US 8,864,469 issued to Liang on 10/21/2014 and entitled TURBINE ROTOR BLADE WITH SUPER COOLING). The FIG. 1 rotor blade cooling circuit includes two distinct cooling circuits with one for cooling the leading edge region of the airfoil and a second that includes a five-pass serpentine cooling circuit for cooling the mid-chord region with discharge for cooling the trailing edge region of the airfoil. The leading edge region cooling circuit 10 includes a single pass radial extending channel that feeds multiple rows of film cooling holes on the leading edge region and discharges all of the cooling air out onto the airfoil external surface as film cooling air. The mid-chord region serpentine cooling circuit includes legs 11-15 that pass upward and then downward to produce convection cooling of the hot wall surfaces on the pressure side and suction side walls of the airfoil. One or more rows of film cooling holes are connected to the serpentine legs to discharge film cooling air on the hottest surfaces of the airfoil walls. The remaining cooling air passing into the last leg 15 is discharged through the trailing edge region through a row of exit holes 16 that discharge out the trailing edge of the airfoil. All of the cooling air supplied to the airfoil through the root is discharged into the hot gas stream to form an open loop cooling circuit.

FIG. 2 shows a prior art leading edge region cooling circuit in which cooling air is supplied to a channel, from where it is passed through metering and

impingement holes 18 into a leading edge region cooling channel 19, and then discharged through rows of film cooling holes 21 and 22 to provide film cooling to hot surfaces of the airfoil. SUMMARY

A turbine rotor blade with a closed loop cooling circuit for an industrial gas turbine engine in which the cooling air for the rotor blade is delivered to a combustor of the engine instead of discharged into the turbine hot gas stream. The rotor blade includes a serpentine-flow cooling circuit in which the supply leg and discharge leg are both located in the blade root.

In a first embodiment, the closed loop cooling circuit includes a first four-pass aft-flowing serpentine flow cooling circuit in a forward section of the blade airfoil and a second four-pass forward-flowing serpentine flow cooling circuit in an aft section of the blade airfoil.

In a second embodiment, the closed loop cooling circuit includes a six-pass serpentine-flow cooling circuit with a forward flowing direction having a first leg located adjacent to a trailing edge region of the blade airfoil and a last leg located adjacent to a leading edge region of the blade airfoil.

In both embodiments, an option is to use a row of exit holes or slots along the trailing edge and connected to the closed loop cooling circuit to provide cooling to the trailing edge in which some of the total cooling air is discharged into the turbine hot gas stream.

For example, in one embodiment, an air cooled turbine rotor blade for a gas turbine engine includes: a rotor blade airfoil extending from a rotor blade root and a platform and having a rotor blade tip; a leading edge region, a trailing edge region opposite the leading edge region, a pressure side wall, and a suction side wall opposite the pressure side wall, each of the pressure side wall and the suction side wall extending between the leading edge region and the trailing edge region; and a closed loop cooling circuit formed within the rotor blade airfoil, the closed loop cooling circuit including a plurality of legs, the plurality of legs being an even number of legs that includes a first leg to supply cooling air to the closed loop cooling circuit through the rotor blade root and a last leg to discharge cooling air from the closed loop cooling circuit through the rotor blade root.

In one aspect of the embodiment, the closed loop cooling circuit includes a first serpentine-flow cooling circuit having an even number of legs and a second serpentine-flow cooling circuit having an even number of legs.

In one aspect of the embodiment, the first serpentine-flow cooling circuit is an aft-flowing four-pass serpentine flow cooling circuit located in a forward side of the rotor blade airfoil; and the second serpentine-flow cooling circuit is a forward-flowing four-pass serpentine flow cooling circuit located in an aft side of the rotor blade airfoil.

In one aspect of the embodiment, the closed loop cooling circuit includes a six-pass serpentine-flow cooling circuit having a first leg located adjacent to the trailing edge region of the rotor blade airfoil and a sixth leg located adjacent to the leading edge region of the rotor blade airfoil.

In one aspect of the embodiment, the air cooled turbine rotor blade further includes a row of exit holes extending along the trailing edge region of the rotor blade airfoil and being connected to one of the plurality of legs of the closed loop cooling circuit.

In one aspect of the embodiment, the air cooled turbine rotor blade further includes a row of film cooling holes extending along the leading edge region of the rotor blade airfoil and connected to one of the plurality of legs of the closed loop cooling circuit.

BRIEF DESCRIPTION OF THE DRAWINGS

A more complete understanding of the present invention, and the attendant advantages and features thereof, will be more readily understood by reference to the following detailed description when considered in conjunction with the accompanying drawings wherein:

FIG. 1 shows a prior art turbine rotor blade with a cooling circuit;

FIG. 2 shows a prior art turbine rotor blade with a leading edge region cooling circuit;

FIG. 3 shows a first embodiment of a turbine rotor blade with a closed loop cooling circuit for an industrial gas turbine engine;

FIG. 4 shows a second embodiment of a turbine rotor blade with a closed loop cooling circuit for an industrial gas turbine engine; and

FIG. 5 shows a twin spool turbocharged industrial gas turbine engine with a closed loop cooling circuit for a turbine stator vane of the present invention.

DETAILED DESCRIPTION

The present invention is a gas turbine engine with a closed loop or substantially closed loop turbine rotor blade cooling circuit in which cooling air for the rotor blade is discharged into the combustor instead of into the turbine hot gas stream. By substantially closed loop, the inventors mean that most of the cooling air is passed into and then exits from the rotor blade is not discharged into the turbine hot gas stream, but is reused in the combustor. A small amount of the cooling air can be used for film cooling of the leading edge region, such as through film cooling holes at the leading edge region, or at discharge slots in the trailing edge region to cool these parts of the blade. Thus, this kind of substantially closed loop cooling circuit may still be considered to be a closed loop cooling circuit, as discussed below. The closed loop cooled turbine rotor blade is intended for use in a twin spool industrial gas turbine engine in which the engine efficiency and the engine power output is greater than any of the prior art industrial engines currently existing.

The twin spool industrial gas turbine engine with closed loop cooling is shown in FIG. 5 and includes a high spool 61 that directly drives an electric generator 62 and a low spool or a turbocharger 63 that supplies compressed air to a high pressure compressor of the high spool 61. Cooling air for a turbine stator vane or rotor blade is bled off from a compressed air line connecting the low pressure compressor of the low spool 63 to the high pressure compressor of the high spool 61. At least one intercooler 64 is used to cool the compressed air from the low spool 63. FIG. 5 shows a stator vane being cooled, but a rotor blade can also be cooled in which the cooling air passes through the rotor shaft and returns through the rotor shaft to be discharged into the combustor of the high spool 61.

FIG. 3 shows a first embodiment of the air cooled turbine rotor blade 23 for the closed loop cooling circuit of the present invention. The turbine rotor blade 23 includes eight radial or spanwise-extending cooling passages (which may also be referred to herein as legs) formed within the walls of the rotor blade airfoil 24 and extending from the platform to the rotor blade tip. FIGS. 3 and 4 show a cross- sectional view of a rotor blade airfoil 24, with the cross-sectional cut made in a chordwise direction, from the leading edge to the trailing edge. Airflow is represented by arrows in FIGS. 3 and 4, with upward-pointing arrows representing airflow from the rotor blade root to the rotor blade tip, and with downward-pointing arrows representing airflow from the rotor blade tip to the rotor blade root. The pressure side wall 25 and the suction side wall 26 each extend between the leading edge region 27 and the trailing edge region 28.

The FIG. 3 embodiment includes two separate and distinct four-pass cooling circuits, with a first four-pass cooling circuit to cool the forward section of the rotor blade airfoil 24 and the second four-pass cooling circuit to cool the aft section of the rotor blade airfoil 24. The first four-pass cooling circuit includes four cooling passages or legs 31-34, with the first leg 31 located adjacent to the leading edge region 27 of the rotor blade airfoil 24. The second four-pass cooling circuit includes four cooling passages or legs 41-44, with the first leg 41 located adjacent to the trailing edge region 28 of the rotor blade airfoil 24. The first four-pass cooling circuit 31-34 is an aft- flowing serpentine-flow cooling circuit while the second four-pass cooling circuit 41-44 is a forward-flowing serpentine-flow cooling circuit. In both four-pass serpentine-flow cooling circuits, the first leg 31, 41 and the last leg 34, 44 flow through the rotor blade root. Thus, only an even number of legs or passages can be used in the closed loop rotor blade cooling circuit, because the cooling air must be supplied to and discharge from the rotor blade root and not at the rotor blade tip, as in the prior art FIG. 1 rotor blade, which must use an odd number of legs or channels.

In the FIG. 3 rotor blade cooling circuit, an option is to use a row of exit holes or slots 45 in or along the trailing edge of the rotor blade airfoil and connected to the first leg 41 of the forward-flowing serpentine-flow cooling circuit to provide cooling to the trailing edge region 28 of the rotor blade airfoil 24. Just a small amount of the total cooling air flows out through these exit holes 45 and into the turbine hot gas stream, and thus this circuit is still considered to be closed loop.

FIG. 4 shows a second embodiment of the air cooled turbine rotor blade 23 for the closed loop cooling circuit of the present invention. The rotor blade airfoil 24 includes a single six-pass serpentine-flow cooling circuit with cooling passages or legs 51-56 that form a forward- flowing serpentine-flow cooling circuit. The first leg 51 supplies cooling air through the rotor blade root and flows along the trailing edge region 28 of the rotor blade airfoil 24. The remaining legs 52-56 flow in a serpentine path to the last leg 56, which is positioned adjacent to the leading edge region 27 of the rotor blade airfoil. As in the FIG. 3 embodiment, the legs 51-56 extend from the rotor blade platform to the rotor blade tip to provide for cooling of the walls of the entire rotor blade airfoil 24. As in the FIG. 3 embodiment, the cooling air is supplied to the rotor blade airfoil 24 and discharged from the rotor blade airfoil 24 through the rotor blade root.

As an option, the trailing edge region 28 can be cooled using a row of exit holes or slots 45 in the trailing edge region 28 and connected to the first leg 51. A second option is the use of at least one row of film cooling holes 30 in the leading edge region 27 by which cooling air can be supplied to a leading edge region cooling supply channel 57 from the rotor blade root or bled off from the last leg 56. The FIG. 4 embodiment also has an even number of legs or channels in the serpentine circuit, because the supply and discharge of the cooling air to and from the turbine rotor blade 23 must pass through the rotor blade root.

In one embodiment, an air cooled turbine rotor blade for a gas turbine engine includes: a rotor blade airfoil (24) extending from a rotor blade root and a platform and having a rotor blade tip; a leading edge region (27), a trailing edge region (28) opposite the leading edge region (27), a pressure side wall (25), and a suction side wall (26) opposite the pressure side wall (25), each of the pressure side wall (25) and the suction side wall (26) extending between the leading edge region (27) and the trailing edge region (28); and a closed loop cooling circuit formed within the rotor blade airfoil (24), the closed loop cooling circuit including a plurality of legs, the plurality of legs being an even number of legs that includes a first leg (31 , 41 , 51) to supply cooling air to the closed loop cooling circuit through the rotor blade root and a last leg (34, 44, 56) to discharge cooling air from the closed loop cooling circuit through the rotor blade root.

In one aspect of the embodiment, the closed loop cooling circuit includes a first serpentine-flow cooling circuit having an even number of legs (31-34) and a second serpentine-flow cooling circuit having an even number of legs (41-44).

In one aspect of the embodiment, the first serpentine-flow cooling circuit is an aft-flowing four-pass serpentine flow cooling circuit located in a forward side of the rotor blade airfoil (24); and the second serpentine-flow cooling circuit is a forward- flowing four-pass serpentine flow cooling circuit located in an aft side of the rotor blade airfoil (24).

In one aspect of the embodiment, the closed loop cooling circuit includes a six-pass serpentine-flow cooling circuit having a first leg (51) located adjacent to the trailing edge region (28) of the rotor blade airfoil (24) and a sixth leg (56) located adjacent to the leading edge region (27) of the rotor blade airfoil (24).

In one aspect of the embodiment, the air cooled turbine rotor blade further includes a row of exit holes (45) extending along the trailing edge region (28) of the rotor blade airfoil (24) and being connected to one of the plurality of legs of the closed loop cooling circuit. In one aspect of the embodiment, the air cooled turbine rotor blade further includes a row of film cooling holes (30) extending along the leading edge region (27) of the rotor blade airfoil (24) and connected to one of the plurality of legs of the closed loop cooling circuit.

It will be appreciated by persons skilled in the art that the present invention is not limited to what has been particularly shown and described herein above. In addition, unless mention was made above to the contrary, it should be noted that all of the accompanying drawings are not to scale. A variety of modifications and variations are possible in light of the above teachings without departing from the scope and spirit of the invention, which is limited only by the following claims.