Compressor air provided to combustion chamber plenum and turbine guide vane

Abstract

A gas turbine having a combustion chamber with exhaust section through which combustion gas is exhaustable, plenum chamber and compressor are provided. The plenum chamber is coupled to the compressor wherein a first quantity of compressed fluid is injectable therein at a radially inner wall of the plenum chamber. A guide vane section with at least one airfoil is coupled to the exhaust section so combustion gas is flowable against the airfoil. The exhaust section and guide vane section are housed inside the plenum chamber. The airfoil has a first flow chamber where a second quantity of compressed fluid is flowable through the guide vane section from the compressor in the direction from the inner wall to a outer wall of the plenum chamber before being discharged. The second quantity of compressed fluid streamable through the guide vane section is larger than the first quantity of the compressed fluid.

Claims

1. A gas turbine comprising a combustion chamber with an exhaust section through which a combustion gas exhausts, a plenum chamber, a compressor for generating a compressed fluid, wherein the plenum chamber is coupled to the compressor such that a first quantity of the compressed fluid is injectable into the plenum chamber at a radially inner wall of the plenum chamber, a guide vane section comprising an airfoil, a vane inner platform and a vane outer platform, wherein the guide vane section is coupled to the exhaust section such that the combustion gas flows against the airfoil, wherein the exhaust section and the guide vane section are housed inside the plenum chamber, wherein the airfoil comprises a first flow chamber which is configured such that a second quantity of the compressed fluid flows through the guide vane section from the compressor in a direction from the radially inner wall to a radially outer wall of the plenum chamber before being discharged into the plenum chamber in a radially outward direction, and wherein the second quantity of the compressed fluid is larger than the first quantity of the compressed fluid, and a diffuser, wherein the diffuser is directly attached to the vane outer platform such that the second quantity of the compressed fluid is guidable from the radially outer end of the first flow chamber through the diffuser and into the plenum chamber in a radially outward direction.

2. The gas turbine according to claim 1, wherein the diffuser comprises a first flow cross-section and a second flow cross-section, wherein the first flow cross-section corresponds to a flow cross-section at the radially outer end of the first flow chamber, wherein the first flow cross-section differs to the second flow cross-section.

3. The gas turbine according to claim 2 wherein the second flow cross-section has a larger flow cross-section than the first flow cross-section.

4. The gas turbine according to claim 1, wherein the airfoil comprises an inner wall section which forms the first flow chamber, wherein the inner wall section comprises turbulence enhancing elements such that turbulences in the compressed fluid which is flowable through the first flow chamber are generated.

5. The gas turbine according to claim 1, wherein the first flow chamber of the airfoil comprises a first volume (V1), wherein the airfoil further comprises a second flow chamber with a second volume (V2), and wherein the second flow chamber is formed within the airfoil such that a part of the second quantity of the compressed fluid is flowable along a direction from the radially outer wall to the radially inner wall of the plenum chamber and/or along the direction from the radially inner wall to the radially outer wall of the plenum chamber.

6. The gas turbine according to claim 5, wherein the first flow chamber and the second flow chamber are formed inside the airfoil such that a ratio of the first volume (V1) of the first flow chamber to the second volume (V2) of the second flow chamber is more than 70 to 30.

7. The gas turbine according to claim 6 wherein the ratio of the first volume (V1) of the first flow chamber to the second volume (V2) of the second flow chamber is more than 80 to 20.

8. The gas turbine according to claim 6, wherein the second quantity of the compressed fluid comprises a majority of all compressed fluid generated by the compressor.

9. The gas turbine according to claim 6, further comprising an inlet conduit comprising: a conduit inlet configured to receive the compressed fluid directly from the compressor; a conduit outlet configured to deliver the second quantity of the compressed fluid directly to the airfoil.

10. The gas turbine according to claim 5, wherein the second flow chamber is fed by compressed fluid which first cooled a vane inner platform and/or a vane outer platform of the guide vane section.

11. The gas turbine according to claim 5, wherein the airfoil comprises a trailing edge, wherein the airfoil is formed such that a further predefined quantity of compressed fluid flowing through the second flow chamber exhausts into the exhaust section.

12. The gas turbine according to claim 5, wherein the airfoil comprises a dividing wall which separates the first flow chamber and the second flow chamber, wherein the dividing wall comprises at least one through-hole configured to guide a predefined quantity of the second quantity of the compressed fluid flowing through the first flow chamber to the second flow chamber.

13. The gas turbine according to claim 1, further comprising a rotor blade section comprising a blade inner platform and a blade airfoil, wherein the rotor blade section is located downstream and adjacent to the guide vane section, and wherein the rotor blade section is coupled to the guide vane section such that compressed fluid is collected from compressed fluid used to cool a vane inner platform of the guide vane section and is routed to the blade inner platform and the blade airfoil of the rotor blade section.

14. The gas turbine according to claim 1, wherein the combustion chamber is a can-type combustion chamber or an annular-type combustion chamber.

15. A method for guiding compressed fluid in a gas turbine, the method comprising generating a compressed fluid by a compressor, injecting a first quantity of the compressed fluid into a plenum chamber at a radially inner wall of the plenum chamber which is coupled to the compressor, guiding a combustion gas against an airfoil of a guide vane section, wherein the guide vane section is coupled to an exhaust section of a combustion chamber through which exhaust section the combustion gas exhausts, wherein the exhaust section and the guide vane section are housed inside the plenum chamber, and guiding a second quantity of the compressed fluid through a first flow chamber of the airfoil of the guide vane section from the compressor in a direction from the radially inner wall to a radially outer wall of the plenum chamber and then the through a diffuser that is directly attached to a radially outer end of the first flow chamber before discharging the second quantity of the compressed fluid from the diffuser into the plenum chamber in a radially outward direction, wherein the second quantity of the compressed fluid is larger than the first quantity of the compressed fluid.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) The aspects defined above and further aspects of the present invention are apparent from the examples of embodiment to be described hereinafter and are explained with reference to the examples of embodiment. The invention will be described in more detail hereinafter with reference to examples of embodiment but to which the invention is not limited.

(2) FIG. 1 shows a sectional view of a gas turbine according to an exemplary embodiment of the present invention;

(3) FIG. 2 shows a schematical view of a cross-section of an airfoil according to an exemplary embodiment of the present invention; and

(4) FIG. 3 shows a schematical view of the guide vane section and the rotating blade section according to an exemplary embodiment of the present invention.

DETAILED DESCRIPTION

(5) The illustrations in the drawings are schematically. It is noted that in different figures, similar or identical elements are provided with the same reference signs.

(6) FIG. 1 shows a gas turbine 100 according to an exemplary embodiment of the present invention. The gas turbine 100 comprises a can-annular combustor arrangement 122 comprising an array of can-combustors 124, each can-combustor 124 comprising a combustion chamber 101 with an exhaust section 102 through which a combustion gas 103 is exhaustible. Moreover, the gas turbine 100 comprises a plenum chamber 104 and a compressor 105 for generating a compressed fluid 106. The plenum chamber 104 is coupled to the compressor 105 such that a first quantity 107 of the compressed fluid 106 is injectable into the plenum chamber 104 at a radially inner wall 111 of the plenum chamber 104. The gas turbine 100 further comprises a guide vane section 113 with at least one airfoil 114, wherein the guide vane section 113 is coupled to the exhaust section 102 such that the combustion gas 103 is flowable towards the airfoil 114. The exhaust section 102 and the guide vane section 113 are housed inside the plenum chamber 104.

(7) The airfoil 114 comprises a first flow chamber 201 (see FIG. 2) which is formed such that a second quantity 108 of the compressed fluid 106 is flowable through the guide vane section 113 from the compressor 105 in the direction from the radially inner wall 111 to a radially outer wall 112 of the plenum chamber 104 before being discharged into the plenum chamber 104. In particular, the compressed fluid 106 is flowable through the guide vane section 113 from the compressor 105 in the direction from a vane inner platform 119 to a vane outer platform 120 of the guide vane section 113. The second quantity 108 of the compressed fluid 106 which is streamable through the guide vane section 113 is larger than the first quantity 107 of the compressed fluid 106.

(8) In FIG. 1 the rotary axis 118 of a turbine shaft is shown, wherein the rotary axis 118 defines the axial direction which is parallel to the rotary axis and defines the radial direction which is perpendicular to the axial direction and which runs through the center point of the turbine shaft.

(9) The compressor 105 generates the compressed fluid 106. The compressed fluid 106 is guided by a channel which may be formed by a conduit, e.g. a tube, or by a volume between radially adjacent inner and outer parts of the gas turbine 111 to the plenum chamber 104.

(10) The plenum chamber 104 surrounds at least partially the combustion chamber 101 and its exhaust section 102 as well as the guide vane section 113, a rotary blade section and for example further downstream located elements of the gas turbine 100. Between the walls 111,112 of the plenum chamber 104 and the components of the gas turbine 100 which are located inside the plenum chamber 104 a volume for the compressed fluid 106 is generated. Inside the volume, the compressed fluid 106 may flow along the components of the gas turbine 100 which are located inside the volume of the plenum chamber 104.

(11) A first quantity 107 of the compressed fluid 106 is injected or bled off according to the present invention at the radially inner located wall 111 of the plenum chamber 104. The second quantity 108 of the compressed fluid 106 is guided from the compressor 105 to an inlet conduit 121 which connects a supply conduit of the compressed fluid 106 with the first flow chamber 201. For example, the inlet conduit 121 is fixed to the vane inner platform 119. The flow cross section of the inlet conduit 121 at the interface to the first flow chamber 201 is equal or smaller than the flow cross section of the inlet conduit 121 at its upstream end.

(12) If the flow cross section of the inlet conduit 121 at the interface to the first flow chamber 201 is smaller than the flow cross section of the upstream end a beneficial change in fluid characteristics (pressure and velocity changes) of the compressed fluid 106 streaming into the first flow chamber 201 is achieved i.e. the inlet losses are minimized and the flow resistance in the interface between the inlet conduit 121 and the first flow chamber 201 is reduced, because no flanges or edges protrude in the flow cross sections at the interface region.

(13) Furthermore, the inlet conduit 121 may in particular have a gradual change in flow cross section from an oval or polygonal shape at its inlet to match the profile of the cross section of the first flow chamber 201 at the interface, for example. The second quantity 108 flows via the first flow chamber 201, the interior part of the airfoil 114, from a radially inner section to a radially outer section of the airfoil 114 before being injected into the volume of the plenum chamber 104.

(14) At the radially outer end 130 of the first flow chamber 201 which is located at the radially outer section, i.e. at the radially vane outer platform 120, of the airfoil 114 a diffuser 115 is installed. The diffuser 115 may comprise a first flow diameter 132 and a second flow diameter 134, wherein the first flow diameter 132 corresponds to a flow diameter 136 at the radially outer end of the first flow chamber 201, wherein the first flow diameter 132 differs to the second flow diameter 134. In particular, the second flow diameter 134 comprises a larger diameter in comparison to the first flow diameter 132. Preferably, the first flow diameter 132 matches the shape of the profile of the cross section 138 of the first flow chamber 201. Hence, kinetic and pressure losses of the second quantity 108 caused by flowing through the first flow chamber 201 may be reduced. The first quantity 107 and the second quantity 108 are mixed together after flowing through the first flow chamber 201 inside the plenum chamber 104, wherein a part of the mixed compressed fluid 109 may flow along the exhaust section 102 and the outer wall of the combustion chamber 101 for cooling purposes. After approaching an axial end at an upstream location of the combustion chamber 101, the compressed fluid 109 is injected into a burner or a swirler, mixed with fuel and is thus injected into the inside of the combustion chamber 101. In the combustion chamber 101 the mixture of compressed fluid 109 and fuel is burnt such that a combustion gas 103 is generated. A portion of the compressed fluid 109 is used to cool the components of the combustion chamber 101 and the exhaust section 102.

(15) The combustion gas 103 flows with at least one component along a downstream direction inside the combustion chamber 101. At a downstream end section of the exhaust section 102, the guide vane section 113 is located, such that the combustion gas 103 is guided by the airfoil 114 of the guide vane section 113.

(16) A further part of the mixed compressed fluid 110 may be used for cooling further downstream located components of the gas turbine 100, e.g. additional guide vane sections, casing components and/or additional rotor blade sections.

(17) A blade airfoil 116 of a rotor blade section is located adjacent and downstream of the airfoil 114 inside the end section of the exhaust section 102. The rotor blade section may be a first stage of a high pressure turbine of the gas turbine 100.

(18) The rotary blade section comprises a blade inner platform 117 from which the blade airfoil 116 extends inside the exhaust section 102. The blade inner platform 117 may comprise a circumferential run or a segmented design of several blade inner platforms 117, wherein to the blade inner platform 117 also an individual or a plurality of blade airfoils 116 may be attached.

(19) FIG. 2 shows a cross-section of the airfoil 114 having an inner wall section 150 and turbulence enhancing features 152. The airfoil 114 comprises the first flow chamber 201 and a second flow chamber 202. Along a cord of the airfoil 114, the first flow chamber 201 is located closer to a leading edge of the airfoil 114, whereas the second flow chamber 202 is located closer to the trailing edge 204 of the airfoil 114 in comparison to the first flow chamber 201.

(20) Moreover, FIG. 2 shows that the first volume VI of the first flow chamber 201 is much larger than a second volume V2 of the second flow chamber 202. Hence, a high amount of compressed fluid, such as the second quantity 108 of the compressed fluid 106, is flowable through the first flow chamber 201 without having high fluid flow losses which may be caused due to a restrictor effect, for example. In the second flow chamber 202 a compressed fluid 206 may flow from an outer or an inner volume of the plenum chamber 104 through the second flow chamber 202.

(21) More particularly, before the second quantity 108 enters the second volume V2 of the second flow chamber 202, the second quantity 108 has already cooled the platform of e.g. the guide vane, in particular the vane inner platform 119. Hence, the flow of the cooling fluid cools along a predefined flow direction one component after the other in a serial manner, such that a serial cooling flow is generated. Such a serial cooling flow, which passes along its flow a plurality of adjacent turbine components, has the benefit that it is not necessary to release the cooling air into the hot working fluid flow after each component required to be cooled and hence the used cooling air flow can be minimized.

(22) The preferred arrangement is for a portion the second quantity 108 to be used for cooling the vane outer platform 120 of the guide vane section 113 before the second quantity 108 of the compressed fluid 106 enters the second volume V2. As indicated in FIG. 2, the flow direction of the second quantity 108 in the first flow chamber 201 and the flow direction of a compressed fluid in the second flow chamber 202 are then opposed with respect to each other.

(23) In particular, the second quantity 108 of compressed fluid flows along a radial direction away from the rotary axis 118, whereas the compressed fluid 206 inside the second flow chamber 202 flows along the radial direction to the rotary axis 118.

(24) In some circumstances it may be preferable to cool the vane inner platform 119 and the vane outer platform 120 using the second quantity 108, wherein a first amount of the second quantity 108 cools the inner platform 119 and a further second amount of the second quantity 108 cools the outer platform 120. After cooling the inner platform 119 and the outer platform 120, the first amount and the second amount respectively of the cooling air is fed to the second volume V2 at a corresponding end (radial inner and radial outer end). At the corresponding end, the first amount and the second amount flows e.g. in opposite directions towards each other.

(25) Furthermore, a dividing wall 205 may separate the first volume VI of the first flow chamber 201 from the second volume V2 of the second flow chamber 202. In the dividing wall 205 a through-hole 203 may be formed such that additional mass flow from the second quantity 108 of compressed fluid 106 in the first flow chamber 201 may flow to the second flow chamber 202 for increasing the mass flow or adjusting the flow distribution inside the second flow chamber 202.

(26) At the trailing edge 204, a further opening may be formed such that the compressed fluid 206 inside the second flow chamber 202 may stream outside into the volume of the exhaust section so that the exhausted compressed fluid 206 is mixed with the combustion gas 103.

(27) FIG. 3 shows an enlarged view of the guide vane section 113 and the rotary blade section. The combustion gas 103 streams against the airfoil 114 and streams further downstream against the blade airfoil 116. The rotary blade section comprises a blade inner platform 117 from which the blade airfoil 116 extends inside the exhaust section 102. To the blade inner platform 117 may one or more blade airfoil (s) 116 be attached. A multiple of blade inner platforms 117 makes up a circumferential run around the turbine disc.

(28) By an air channel i.e. a nozzle 301 and a further air channel 302 the compressed fluid 306, which exits the collection point for the portion of the second quantity 108 that has been used to cool the vane inner platform 119 of the guide vane section 113, may be guided to the blade inner platform 117 and onwards to the blade airfoil 116 and its interior for cooling purposes, via the cooling entrance in a turbine disc channel 302 and the blade root.

(29) It should be noted that the term “comprising” does not exclude other elements or steps and “a” or “an” does not exclude a plurality. Also elements described in association with different embodiments may be combined. It should also be noted that reference signs in the claims should not be construed as limiting the scope of the claims.