Cooling air configuration in a gas turbine engine

The present invention relates to cooling air configurations in a gas turbine engine, wherein at least a portion of cooling air provided into a turbine section is provided into a turbine disc bore and bypasses an upstream turbine stage.

In a turbomachine, such as a gas turbine engine, air is pressurized in a compressor section then mixed with fuel and burned in a combustion section to generate hot combustion gases. The hot combustion gases are expanded within a turbine section of the engine where energy is extracted to provide output power used to produce electricity. The hot combustion gases travel through a series of stages when passing through the turbine section. A stage typically includes a row of stationary airfoils, i.e., vanes, followed by a row of rotating airfoils, i.e., blades, where the blades extract energy from the hot combustion gases for providing output power.

In accordance with a first aspect of the present invention, a method is provided for providing cooling air from a source of cooling air through a cooling air circuit in a turbine section of a gas turbine engine. A first portion of cooling air is provided from the source of cooling air along a first path of the cooling air circuit to a plurality of blades associated with a stage of the turbine section. A second portion of cooling air is provided from the source of cooling air along a second path of the cooling air circuit. The second path includes a turbine disc bore where the cooling air provides cooling to a radially innermost portion of at least one turbine disc that forms a part of a rotor of the engine. The second path is independent from the first path such that the second portion of cooling air bypasses the stage and is not mixed with the first portion of cooling air in the cooling air circuit after leaving the source of cooling air.

In accordance with a second aspect of the present invention, a method is provided for providing cooling air from a source of cooling air through a cooling air circuit in a turbine section of a gas turbine engine. A first portion of cooling air is provided from the source of cooling air along a first path of the cooling air circuit to a plurality of blades associated with a first stage of the turbine section. A second portion of cooling air is provided from the source of cooling air along a second path of the cooling air circuit. The second path includes a turbine disc bore where the cooling air provides cooling to a radially innermost portion of at least one turbine disc that forms a part of a rotor of the engine. The second path is independent from the first path such that the second portion of cooling air bypasses the first stage and is not mixed with the first portion of cooling air in the cooling air circuit after leaving the source of cooling air. A third portion of cooling air is provided from the source of cooling air along a third path of the cooling air circuit to a plurality of blades associated with a second stage of the turbine section, the second stage being located downstream from the first stage with respect to a hot gas flowpath that is defined within the turbine section and that extends generally parallel to a longitudinal axis of the engine.

While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein:

FIG. 1 is a schematic illustration, partially in cross section, of a portion of a turbine engine including a cooling air configuration according to an aspect of the present invention; and

FIG. 2 is a schematic illustration, partially in cross section, of a portion of a turbine engine including a cooling air configuration according to another aspect of the present invention.

In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, specific preferred embodiments in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention.

Referring to FIG. 1 , a portion of a gas turbine engine 10 including an upper half of a turbine section 12 is schematically shown. The exemplary turbine section 12 illustrated in FIG. 1 includes first, second, third, and fourth stages 14 A, 14 B, 14 C, 14 D, wherein each stage 14 A-D includes a row of stationary turbine vanes 16 A-D and a row of rotating turbine blades 18 A-D positioned downstream from each respective row of vanes 16 A-D with respect to a direction of hot gas flow through a hot gas flowpath 20 defined within the turbine section 12 and extending generally parallel to a longitudinal axis L A of the engine 10 . As shown in FIG. 1 , each row of blades 18 A-D is mounted to a respective blade disc structure 22 A-D, which, in turn, is mounted to a respective turbine disc 24 A-D, wherein turbine discs 24 A-D each form a part of a rotor 26 of the engine 10 . The term “blade disc structure” as used herein refers to any structure located between the blades and the turbine discs, including but not limited to, roots, platforms, disc attachments, etc.

Also shown in FIG. 1 is a cooling air circuit 30 constructed in accordance with an aspect of the present invention. Cooling air, which may comprise compressor discharge air, is provided into the cooling air circuit 30 from a source of cooling air 32 as will be described herein. The cooling air provided to the cooling air circuit 30 from the source of cooling air 32 may optionally be cooled in a conventional air cooler (not shown) before being provided to the source of cooling air 32 , which, in the embodiment shown, comprises an annular source cavity 34 located radially between the hot gas flowpath 20 and a turbine disc bore 36 that forms part of the cooling air circuit 30 . In the embodiment shown, the source cavity 34 is located directly radially inwardly from the first stage row of vanes 16 A, and the turbine disc bore 36 is defined between the turbine discs 24 A-D and a central, rotatable shaft 38 of the engine 10 .

The cooling air circuit 30 according to this embodiment further comprises a first passage 40 that extends axially and radially outwardly from the source cavity 34 through the first turbine disc 24 A to the blade disc structure 22 A associated with the first stage row of blades 18 A; a second passage 42 that extends axially and radially inwardly from the source cavity 34 through a seal disc 44 to a radially inner portion of an auxiliary cavity 46 , wherein the radially inner portion of the auxiliary cavity 46 is located in close proximity to and is in fluid communication with the turbine disc bore 36 ; a third passage 48 that extends generally axially from the source cavity 34 through the first turbine disc 24 A to a first cooling air cavity 50 A located axially between the source cavity 34 and the second stage row of blades 18 B; and a fourth passage 52 that extends generally radially inwardly from the source cavity 34 through a gap located between the seal disc 44 and the first turbine disc 24 A to a radially outer portion of the auxiliary cavity 46 . The auxiliary cavity 46 is defined between the seal disc 44 and the first turbine disc 24 and is located radially inwardly from the source cavity 34 . It is noted that the second passage 42 could extend directly to the turbine disc bore 36 without departing from the scope and spirit of the invention.

The cooling air circuit 30 further comprises a fifth passage 58 that extends generally radially outwardly from the first cooling air cavity 50 A through the second turbine disc 24 B to the blade disc structure 22 B associated with the second stage row of blades 18 B; a sixth passage 60 that extends generally axially from the first cooling air cavity 50 A through the second turbine disc 24 B to a second cooling air cavity 50 B located axially between the first cooling air cavity 50 A and the third stage row of blades 18 C; and a seventh passage 62 that extends generally radially inwardly from the first cooling air cavity 50 A through a gap located between the first turbine disc 24 A and the second turbine disc 24 B to a first rotor disc cavity 64 A that is defined between the first turbine disc 24 A and the second turbine disc 24 B and is located radially between the first cooling air cavity 50 A and the turbine disc bore 36 .

The cooling air circuit 30 still further comprises an eighth passage 66 that extends generally radially outwardly from the second cooling air cavity 50 B through the third turbine disc 24 C to the blade disc structure 22 C associated with the third stage row of blades 18 C; and a ninth passage 68 that extends generally radially inwardly from the second cooling air cavity 50 B through a gap located between the second turbine disc 24 B and the third turbine disc 24 C to a second rotor disc cavity 64 B that is defined between the second turbine disc 24 B and the third turbine disc 24 C and is located radially between the second cooling air cavity 50 B and the turbine disc bore 36 .

The cooling air circuit 30 also comprises a third rotor disc cavity 64 C that is in fluid communication with the turbine disc bore 36 and is located radially between a third cooling air cavity 50 C and the turbine disc bore 36 ; a tenth passage 70 that extends generally radially outwardly from the third rotor disc cavity 64 C through a gap between the third turbine disc 24 C and the fourth turbine disc 24 D to the third cooling air cavity 50 C of the cooling air circuit 30 , which is located axially between the second cooling air cavity 50 B and the fourth stage row of blades 18 D; and an eleventh passage 72 that extends generally radially outwardly from the third cooling air cavity 50 C through the fourth turbine disc 24 D to the blade disc structure 22 D associated with the fourth stage row of blades 18 D.

Seals 78 A, 78 B are provided between the respective first and second rotor disc cavities 64 A, 64 B and the rotor disc bore 36 for substantially preventing leakage therebetween.

A method for providing cooling air from the source of cooling air 32 , i.e., the source cavity 34 in the embodiment shown, through the cooling air circuit 30 will now be described.

A first portion CA 1 of cooling air is provided from the source cavity 34 along a first path P 1 of the cooling air circuit 30 to the first stage row of blades 18 A, wherein the first stage 14 A is also referred to herein as an upstream stage. The first path P 1 according to this embodiment comprises the first passage 40 , which delivers the first portion CA 1 of cooling air to the first stage blade disc structure 22 A, which in turn delivers the first portion CA 1 of cooling air to the first stage row of blades 18 A. The first portion CA 1 of cooling air is used to cool the first stage row of blades 18 A in any known manner and then may exit the first stage row of blades 18 A and be swept up by the hot gas flowing through the hot gas flowpath 20 . It is noted that the first stage blade disc structure 22 A is schematically illustrated in FIG. 1 and could include any suitable configuration for delivering the first portion CA 1 of cooling air to the first stage row of blades 18 A.

A second portion CA 2 of cooling air is provided from the source cavity 34 along a second path P 2 of the cooling air circuit 30 . The second path P 2 according to this embodiment comprises the second passage 42 , which delivers the second portion CA 2 of cooling air to the radially inner portion of the auxiliary cavity 46 . The second portion CA 2 of cooling air then passes into the turbine disc bore 36 from the auxiliary cavity 46 , although the second passage 42 could extend directly to the turbine disc bore 36 as noted above. The second path P 2 according to this embodiment further comprises the turbine disc bore 36 , wherein the second portion CA 2 of cooling air provides cooling to radially innermost portions of the turbine discs 24 A-D while passing through the turbine disc bore 36 .

The second path P 2 according to this embodiment still further comprises the third rotor disc cavity 64 C, the tenth passage 70 , the third cooling fluid cavity 50 C, and the eleventh passage 72 . The eleventh passage 72 delivers the second portion CA 2 of cooling air to the fourth stage blade disc structure 22 D, which in turn discharges the second portion CA 2 of cooling air to the hot gas flowpath 20 , wherein the fourth stage 14 D is also referred to herein as a downstream stage or a final stage. It is noted that the fourth stage blade disc structure 22 D could deliver the second portion CA 2 of cooling air to the fourth stage row of blades 18 D for cooling the fourth stage row of blades 18 D in any known manner, wherein the second portion CA 2 of cooling air could then exit the fourth stage row of blades 18 D and be swept up by the hot gas flowing through the hot gas flowpath 20 .

According to this embodiment of the invention, the second path P 2 is independent from the first path P 1 , such that the second portion CA 2 of cooling air bypasses the first stage 14 A and is not mixed with the first portion CA 1 of cooling air in the cooling air circuit 30 after leaving the source cavity 34 , although the first and second portions CA 1 , CA 2 of cooling air may once again convene upon being swept up by the hot gas flowing through the hot gas flowpath 20 . Hence, all of the cooling provided by the second portion CA 2 of cooling air is used to cool structure along the second path P 2 ) the fourth stage blade disc structure 22 D, and, optionally, the fourth stage row of blades 18 D.

A third portion CA 3 of cooling air is provided from the source cavity 34 along a third path P 3 of the cooling air circuit 30 . The third path P 3 according to this embodiment comprises the third, fifth, sixth, seventh, eighth, and ninth passages 48 , 58 , 60 , 62 , 66 , 68 , the first and second cooling air cavities 50 A, 50 B, and the first and second rotor disc cavities 64 A, 64 B.

More specifically, the third passage 48 delivers the third portion CA 3 of cooling air from the source cavity 34 to the first cooling air cavity 50 A. A first allotment of the third portion CA 3 of cooling air is provided to the second stage blade disc structure 22 B via the fifth passage 58 . The second stage blade disc structure 22 B in turn delivers the first allotment of the third portion CA 3 of cooling air to the second stage row of blades 18 B, wherein the second stage 14 B is also referred to herein as an intermediate stage. The first allotment of the third portion CA 3 of cooling air is used to cool the second stage row of blades 18 B in any known manner and then may exit the second stage row of blades 18 B and be swept up by the hot gas flowing through the hot gas flowpath 20 . It is noted that the second stage blade disc structure 22 B is schematically illustrated in FIG. 1 and could include any suitable configuration for delivering the first allotment of the third portion CA 3 of cooling air to the second stage row of blades 18 B.

A second allotment of the third portion CA 3 of cooling air is provided from the first cooling air cavity 50 A to the second cooling air cavity 50 B via the sixth passage 60 . Some of the second allotment of the third portion CA 3 of cooling air is provided to the third stage blade disc structure 22 C via the eighth passage 66 . The third stage blade disc structure 22 C in turn delivers this cooling air to the third stage row of blades 18 C, wherein the third stage 14 C is also referred to herein as an intermediate stage. This cooling air is used to cool the third stage row of blades 18 C in any known manner and then may exit the third stage row of blades 18 C and be swept up by the hot gas flowing through the hot gas flowpath 20 . It is noted that the third stage blade disc structure 22 C is schematically illustrated in FIG. 1 and could include any suitable configuration for delivering cooling air to the third stage row of blades 18 C.

The remainder of the second allotment of the third portion CA 3 of cooling air in the second cooling air cavity 50 B is provided into the second rotor disc cavity 64 B via the ninth passage 68 .

A third allotment of the third portion CA 3 of cooling air is provided from the first cooling air cavity 50 A to the first rotor disc cavity 64 A via the seventh passage 62 .

According to this embodiment of the invention, the third path P 3 is independent from the first and second paths P 1 , P 2 , such that the third portion CA 3 of cooling air bypasses the first stage 14 A and is not mixed with the first or second portions CA 1 , CA 2 of cooling air in the cooling air circuit 30 after leaving the source cavity 34 , although the first, second, and third portions CA 1 , CA 2 , CA 3 of cooling air may once again convene upon being swept up by the hot gas flowing through the hot gas flowpath 20 . Hence, all of the cooling provided by the third portion CA 3 of cooling air is used to cool the structure along the third path P 3 , the second and third stage blade disc structures 22 B, 22 C, and the second and third stage rows of blades 18 B, 18 C.

A fourth portion CA 4 of cooling air, also referred to herein as an auxiliary portion of cooling air, is provided from the source cavity 34 along a fourth path P 4 of the cooling air circuit 30 , also referred to herein as an auxiliary path. The fourth path P 4 according to this embodiment comprises the fourth passage 52 , which delivers the fourth portion CA 4 of cooling air to the radially outer portion of the auxiliary cavity 46 . The fourth portion CA 4 of cooling air then passes through the auxiliary cavity 46 and is mixed with the second portion CA 2 of cooling air for entry into the turbine disc bore 36 with the second portion CA 2 of cooling air. The fourth path P 4 according to this embodiment further comprises the turbine disc bore 36 , wherein the fourth portion CA 4 of cooling air, together with the second portion CA 2 of cooling air, provides cooling to the radially innermost portions of the turbine discs 24 A-D while passing through the turbine disc bore 36 .

The fourth path P 4 according to this embodiment still further comprises the third rotor disc cavity 64 C, the tenth passage 70 , the third cooling fluid cavity 50 C, and the eleventh passage 72 . The eleventh passage 72 delivers the fourth portion CA 4 of cooling air, together with the second portion CA 2 of cooling air, to the fourth stage blade disc structure 22 D, which in turn discharges the second and fourth portions CA 2 , CA 4 of cooling air to the hot gas flowpath 20 , although the fourth stage blade disc structure 22 D could deliver the second and fourth portions CA 2 , CA 4 of cooling air to the fourth stage row of blades 18 D for providing cooling thereto.

Referring now to FIG. 2 , a portion of a gas turbine engine 110 including an upper half of a turbine section 112 is schematically shown. The exemplary turbine section 112 illustrated in FIG. 2 includes first, second, third, and fourth stages 114 A, 114 B, 114 C, 114 D, wherein each stage 114 A-D includes a row of stationary turbine vanes 116 A-D and a row of rotating turbine blades 118 A-D positioned downstream from each respective row of vanes 116 A-D with respect to a direction of hot gas flow through a hot gas flowpath 120 defined within the turbine section 12 and extending generally parallel to a longitudinal axis L A of the engine 110 . As shown in FIG. 2 , each row of blades 118 A-D is mounted to a respective blade disc structure 122 A-D, which, in turn, is mounted to a respective turbine disc 124 A-D, wherein turbine discs 124 A-D each form a part of a rotor 126 of the engine 110 .

Also shown in FIG. 2 is a cooling air circuit 130 constructed in accordance with another aspect of the present invention. Cooling air, which may comprise compressor discharge air, is provided into the cooling air circuit 130 from a source of cooling air 132 as will be described herein. The cooling air provided to the cooling air circuit 130 from the source of cooling air 132 may optionally be cooled in a conventional air cooler (not shown) before being provided to the source of cooling air 132 , which, in the embodiment shown, comprises an annular source cavity 134 located radially between the hot gas flowpath 120 and a turbine disc bore 136 that forms part of the cooling air circuit 130 . In the embodiment shown, the source cavity 134 is located directly radially inwardly from the first stage row of vanes 116 A, and the turbine disc bore 136 is defined between the turbine discs 124 A-D and a central, rotatable shaft 138 of the engine 110 .

The cooling air circuit 130 according to this embodiment further comprises a first passage 140 that extends axially and radially outwardly from the source cavity 134 through the first turbine disc 124 A to the blade disc structure 122 A associated with the first stage row of blades 118 A; a second passage 142 that extends axially and radially inwardly from the source cavity 134 through a seal disc 144 to a radially inner portion of an auxiliary cavity 146 , wherein the radially inner portion of the auxiliary cavity 146 is located in close proximity to and is in fluid communication with the turbine disc bore 136 ; a third passage 148 that extends axially and radially inwardly from the source cavity 134 through the first turbine disc 124 A to a first rotor disc cavity 150 A located radially between a first cooling air cavity 154 A and the turbine disc bore 136 ; and a fourth passage 152 that extends generally radially inwardly from the source cavity 134 through a gap located between the seal disc 144 and the first turbine disc 124 A to a radially outer portion of the auxiliary cavity 146 . The auxiliary cavity 146 is defined between the seal disc 144 and the first turbine disc 124 and is located radially inwardly from the source cavity 134 . It is noted that the second passage 142 could extend directly to the turbine disc bore 136 without departing from the scope and spirit of the invention.

The cooling air circuit 130 further comprises a fifth passage 158 that extends axially and radially outwardly from the first rotor disc cavity 150 A through the second turbine disc 124 B to the blade disc structure 122 B associated with the second stage row of blades 118 B; a sixth passage 160 that extends generally axially from the first rotor disc cavity 150 A through the second turbine disc 124 B to a second rotor disc cavity 150 B located radially between a second cooling air cavity 154 B and the turbine disc bore 136 ; and a seventh passage 162 that extends generally radially outwardly from the first rotor disc cavity 150 A through a gap located between the first turbine disc 124 A and the second turbine disc 124 B to the first cooling air cavity 154 A, which is defined between the first turbine disc 124 A and the second turbine disc 124 B and is located axially between the source cavity 134 and the second stage row of blades 118 B.

The cooling air circuit 130 still further comprises an eighth passage 166 that extends axially and radially outwardly from the second rotor disc cavity 150 B through the third turbine disc 124 C to the blade disc structure 122 C associated with the third stage row of blades 118 C; and a ninth passage 168 that extends generally radially outwardly from the second rotor disc cavity 150 B through a gap located between the second turbine disc 124 B and the third turbine disc 124 C to the second cooling air cavity 154 B, which is defined between the second turbine disc 124 B and the third turbine disc 124 C and is located axially between the first cooling air cavity 154 A and the third stage row of blades 118 C.

The cooling air circuit 130 also comprises a third rotor disc cavity 150 C that is in fluid communication with the turbine disc bore 136 and is located radially between a third cooling air cavity 154 C and the turbine disc bore 136 ; a tenth passage 170 that extends generally radially outwardly from the third rotor disc cavity 150 C through a gap between the third turbine disc 124 C and the fourth turbine disc 124 D to the third cooling air cavity 154 C of the cooling air circuit 130 , which is located axially between the second cooling air cavity 154 B and the fourth stage row of blades 118 D; and an eleventh passage 172 that extends axially and radially outwardly from the third rotor disc cavity 150 C through the fourth turbine disc 124 D to the blade disc structure 122 D associated with the fourth stage row of blades 118 D.

Seals 178 A, 178 B are provided between the respective first and second rotor disc cavities 150 A, 150 B and the rotor disc bore 136 for substantially preventing leakage therebetween.

A method for providing cooling air from the source of cooling air 132 , i.e., the source cavity 134 in the embodiment shown, through the cooling air circuit 130 will now be described.

A first portion CA 1 of cooling air is provided from the source cavity 134 along a first path P 1 of the cooling air circuit 130 to the first stage row of blades 118 A, wherein the first stage 114 A is also referred to herein as an upstream stage. The first path P 1 according to this embodiment comprises the first passage 140 , which delivers the first portion CA 1 of cooling air to the first stage blade disc structure 122 A. The first stage blade disc structure 122 A in turn delivers the first portion CA 1 of cooling air to the first stage row of blades 118 A. The first portion CA 1 of cooling air is used to cool the first stage row of blades 118 A in any known manner and then may exit the first stage row of blades 118 A and be swept up by the hot gas flowing through the hot gas flowpath 120 . It is noted that the first stage blade disc structure 122 A is schematically illustrated in FIG. 2 and could include any suitable configuration for delivering the first portion CA 1 of cooling air to the first stage row of blades 118 A.

A second portion CA 2 of cooling air is provided from the source cavity 134 along a second path P 2 of the cooling air circuit 130 . The second path P 2 according to this embodiment comprises the second passage 142 , which delivers the second portion CA 2 of cooling air to the radially inner portion of the auxiliary cavity 146 . The second portion CA 2 of cooling air then passes into the turbine disc bore 136 from the auxiliary cavity 146 , although the second passage 142 could extend directly to the turbine disc bore 136 as noted above. The second path P 2 according to this embodiment further comprises the turbine disc bore 136 , wherein the second portion CA 2 of cooling air provides cooling to radially innermost portions of the turbine discs 124 A-D while passing through the turbine disc bore 136 .

The second path P 2 according to this embodiment still further comprises the third rotor disc cavity 150 C, the tenth passage 170 , the third cooling fluid cavity 154 C, and the eleventh passage 172 . The tenth passage 170 delivers some of the second portion CA 2 of cooling air from the third rotor disc cavity 150 C to the third cooling fluid cavity 154 C. The eleventh passage 172 delivers the remainder of the second portion CA 2 of cooling air from the third rotor disc cavity 150 C to the fourth stage blade disc structure 122 D, which in turn discharges the second portion CA 2 of cooling air to the hot gas flowpath 120 , wherein the fourth stage 114 D is also referred to herein as a downstream stage or a final stage. It is noted that the fourth stage blade disc structure 122 D could deliver the second portion CA 2 of cooling air to the fourth stage row of blades 118 D for cooling the fourth stage row of blades 118 D in any known manner, wherein the second portion CA 2 of cooling air could then exit the fourth stage row of blades 118 D and be swept up by the hot gas flowing through the hot gas flowpath 120 .

According to this embodiment of the invention, the second path P 2 is independent from the first path P 1 , such that the second portion CA 2 of cooling air bypasses the first stage 114 A and is not mixed with the first portion CA 1 of cooling air in the cooling air circuit 130 after leaving the source cavity 134 , although the first and second portions CA 1 , CA 2 of cooling air may once again convene upon being swept up by the hot gas flowing through the hot gas flowpath 120 . Hence, all of the cooling provided by the second portion CA 2 of cooling air is used to cool structure along the second path P 2 , the fourth stage blade disc structure 122 D, and, optionally, the fourth stage row of blades 118 D.

A third portion CA 3 of cooling air is provided from the source cavity 134 along a third path P 3 of the cooling air circuit 130 . The third path P 3 according to this embodiment comprises the third, fifth, sixth, seventh, eighth, and ninth passages 148 , 158 , 160 , 162 , 166 , 168 , the first and second rotor disc cavities 150 A, 150 B, and the first and second cooling air cavities 154 A, 154 B.

More specifically, the third passage 148 delivers the third portion CA 3 of cooling air from the source cavity 134 to the first rotor disc cavity 150 A. A first allotment of the third portion CA 3 of cooling air is provided to the second stage blade disc structure 122 B via the fifth passage 158 . The second stage blade disc structure 122 B in turn delivers the first allotment of the third portion CA 2 of cooling air to the second stage row of blades 118 B, wherein the second stage 114 B is also referred to herein as an intermediate stage. The first allotment of the third portion CA 3 of cooling air is used to cool the second stage row of blades 118 B in any known manner and then may exit the second stage row of blades 118 B and be swept up by the hot gas flowing through the hot gas flowpath 120 . It is noted that the second stage blade disc structure 122 B is schematically illustrated in FIG. 2 and could include any suitable configuration for delivering the first allotment of the third portion CA 3 of cooling air to the second stage row of blades 118 B.

A second allotment of the third portion CA 3 of cooling air is provided from the first rotor disc cavity 150 A to the second rotor disc cavity 150 B via the sixth passage 160 . Some of the second allotment of the third portion CA 3 of cooling air is provided to the third stage blade disc structure 122 C via the eighth passage 166 . The third stage blade disc structure 122 C in turn delivers this cooling air to the third stage row of blades 118 C, wherein the third stage 114 C is also referred to herein as an intermediate stage. This cooling air is used to cool the third stage row of blades 118 C in any known manner and then may exit the third stage row of blades 118 C and be swept up by the hot gas flowing through the hot gas flowpath 120 . It is noted that the third stage blade disc structure 122 C is schematically illustrated in FIG. 2 and could include any suitable configuration for delivering cooling air to the third stage row of blades 118 C.

The remainder of the second allotment of the third portion CA 3 of cooling air in the second rotor disc cavity 150 B is provided into the second cooling air cavity 154 B via the ninth passage 168 .

A third allotment of the third portion CA 3 of cooling air is provided from the first rotor disc cavity 150 A to the first cooling air cavity 154 A via the seventh passage 162 .

According to this embodiment of the invention, the third path P 3 is independent from the first and second paths P 1 , P 2 , such that the third portion CA 3 of cooling air bypasses the first stage 114 A and is not mixed with the first or second portions CA 1 , CA 2 of cooling air in the cooling air circuit 130 after leaving the source cavity 134 , although the first, second, and third portions CA 1 , CA 2 , CA 3 of cooling air may once again convene upon being swept up by the hot gas flowing through the hot gas flowpath 120 . Hence, all of the cooling provided by the third portion CA 3 of cooling air is used to cool structure along the third path P 3 , the second and third stage blade disc structures 122 B, 122 C, and the second and third stage rows of blades 118 B, 118 C.

A fourth portion CA 4 of cooling air, also referred to herein as an auxiliary portion of cooling air, is provided from the source cavity 134 along a fourth path P 4 of the cooling air circuit 130 , also referred to herein as an auxiliary path. The fourth path P 4 according to this embodiment comprises the fourth passage 152 , which delivers the fourth portion CA 4 of cooling air to the radially outer portion of the auxiliary cavity 146 , wherein the fourth portion CA 4 of cooling air then passes through the auxiliary cavity 146 and is mixed with the second portion CA 2 of cooling air for entry into the turbine disc bore 136 with the second portion CA 2 of cooling air. The fourth path P 4 according to this embodiment further comprises the turbine disc bore 136 , wherein the fourth portion CA 4 of cooling air, together with the second portion CA 2 of cooling air, provides cooling to the radially innermost portions of the turbine discs 124 A-D while passing through the turbine disc bore 136 .

The fourth path P 4 according to this embodiment still further comprises the third rotor disc cavity 150 C, the tenth passage 170 , the third cooling fluid cavity 154 C, and the eleventh passage 172 . The eleventh passage 172 delivers some of the fourth portion CA 4 of cooling air, together with some of the second portion CA 2 of cooling air, to the fourth stage blade disc structure 122 D, which in turn discharges this cooling air to the hot gas flowpath 120 , although the fourth stage blade disc structure 122 D could deliver this cooling air to the fourth stage row of blades 118 D for providing cooling thereto.

According to the present invention, it is believed that adequate cooling is provided to the radially innermost portions of at least the first, second, and third turbine discs 24 A-C (FIG. 1 ) and 124 A-C ( FIG. 2 ) so as to reduce thermal stresses experienced by these components and other components in and around the turbine disc bore 36 ( FIG. 1) and 136 ( FIG. 2 ). Such reduction of thermal stresses is believed to effect an increase of the useful lifespan of these components.

Additionally, in the configuration disclosed in FIG. 2 , belly band seals 80 A, 80 B, 80 C, which are provided for sealing the cooling air cavities 50 A-C in the embodiment of FIG. 1 , can be removed. Specifically, these seals 80 A-C are not required in the configuration illustrated in FIG. 2 , as these seals 80 A-C are provided in FIG. 1 to ensure that adequate cooling air is provided to the respective rows of blades 18 A-C. Since the cooling air provided to the rows of blades 118 A-C illustrated in FIG. 2 is provided directly from the rotor disc cavities 150 A-C, the amount of cooling air provided to the rows of blades 118 A-C can be controlled by changing the diameters of the passages that extend between the rotor disc cavities 150 A-C and the cooling air cavities 154 A-C.

Further, it is noted that the dimensions and directions of the passages and cavities illustrated in FIGS. 1 and 2 and described herein are exemplary, and the present invention is not intended to be limited to the dimensions and directions illustrated and described.

While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.