Cooling system for gas turbine combustor

Abstract: A combustor of a gas turbine includes a combustion tube acting as a combustion chamber. In the combustion tube, a cylindrical heat insulating member is disposed and forms a wall of the combustion chamber. Numerous cooling air passages are formed in the heat insulating member and extending in the axial direction between cooling air inlet disposed at one end of the heat insulating member and cooling air outlet disposed at the other end of the heat insulating member. The heat insulating member is attached to the inner surface of the combustion tube at its cooling air inlet side end and a seal ring is disposed around the outer circumference of the heat insulating member at a portion between the cooling air inlet side end and the cooling air outlet side end. The seal ring is made of heat resistant material and prevents cooling air flowing through the annular space between the inner surface of the combustion tube and the outer surface of the heat insulating member. Since cooling air flows through the annular space around the outer surface of the heat insulating member, all of cooling air flows through the cooling passages in the heat insulating member and is used for cooling the heat insulating member. Therefore, the cooling capacity of the heat insulating member increases.

Description: BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine and, specifically, the present invention relates to a cooling system for a gas turbine combustor. 2. Description of the Related Art In general, there are two types of cooling systems for the wall of a gas turbine combustor. One is a compound air cooling system employing both convective cooling and film air cooling using air as cooling fluid and another is steam cooling system using steam as cooling fluid. In an actual gas turbine, one of the compound air cooling system and the steam cooling system is selected according to the temperature of combustion gas at the inlet of turbine. Fig. 1 is a sectional view showing a general construction of a gas turbine combustor employing a compound air cooling system. In Fig. 1, reference numeral 1 designates a combustor of a gas turbine as a whole. The combustor 1 consists of a combustion tube 5b which acts as a combustion chamber for burning fuel injected from fuel nozzles 33 and a tail pipe 5a which directs combustion gas generated in the combustion tube to the first stage stator of the turbine. The combustor tube 5b and the tail pipe 5a are made as separate parts and joined together to form a combustor 1. Fuel is injected into the combustion tube 5b from the main nozzles 33 as a premixed air-fuel mixture. The air-fuel mixture is ignited by a pilot flame formed by a pilot nozzle 31 and generates a premixed flame in the combustion tube. Fig. 2 is a enlarged section of the wall of the combustor tube 5b employing a conventional compound air cooling system. As can be seen from Fig. 2, in an actual gas turbine, the combustor tube 5b is formed by joining a plurality of cylindrical shells 55 having different diameters. The respective shells 55 are aligned in the axial direction and are joined to each other through stepped diameter portions thereof. Each of the shells 55 acts as a structural member forming the combustor tube 5b. A heat insulating member 155 is disposed at inside of each cylindrical shell 55 in order to protect the shell from the flame in the combustor tube and, thereby, preventing a strength degradation of the shell as a structural member. In the conventional cooling system, fin-rings are used for the heat insulating members 155. The fin-ring consists of a cylindrical member having numerous grooves on the outer surface thereof extending in the axial direction. Each of the fin-rings 155 is held inside of the shell by attaching one end thereof to the smaller diameter portion of the corresponding shell 55 (i.e., a fuel nozzle side end of the shell 55), for example, by brazing. In this system, pressurized air in the casing 7 (Fig. 1) is introduced from inlet openings 57 distributed around the smaller diameter portion of the shell 55 into the space between the shell 55 and the fin-ring 155. Air introduced into the space passes through the axial grooves outside of the fin-ring 155 and cools fin-ring 155 by convective cooling. After passing through the axial grooves, air is injected from the outlet 159 at the end of the fin-ring 155 in the direction along the inner surface of the heat insulating member (in Fig. 2, indicated by reference numeral 155b) adjacent thereto. Thus, the wall surface of the combustion chamber, i.e., the inner surface of the adjacent fin-ring 155b is cooled by the film of the injected air. On the other hand, Fig. 4 is a sectional view similar to Fig. 1 showing a gas turbine combustor employing a conventional steam convective cooling system. Since the heat-transfer coefficient of air is relatively low, sufficient cooling can not be obtained by convective cooling and, usually, a compound air cooling system using both convective cooling and film air cooling is employed in the air cooling system. However, compound air cooling system has its inherent problem. In the compound air system, air used for film air cooling is injected into the combustion tube and mixes with combustion gas. This causes dilution of combustion gas and lowers its turbine inlet temperature and, thereby, causes deterioration in the gas turbine output and efficiency. In order to prevent this problem, the combustor in Fig. 4 employs steam cooling system using steam convective cooling instead of compound air cooling. Since the heat-transfer coefficient of steam is larger than that of air, the combustor is sufficiently cooled solely by convective cooling in the steam cooling system. In Fig. 4, reference numerals the same as those in Fig. 1 denote elements similar to those in Fig. 1. The combustor in Fig. 4 is a one-piece construction in which the combustion tube 5b and the tail pipe 5a are formed as an integral part. Therefore, the combustion tube 5 in the combustor 1 in Fig. 4 has outlet 52 at one end thereof in order to supply combustion gas to the first stage stator of the turbine. The combustion tube 5 in Fig. 4 has a double-wall construction including an outer shell (outer wall) and an inner shell (inner wall). The space between the outer shell and inner shell acts as a passage for cooling steam. Cooling steam is supplied to the cooling steam passage between the outer and the inner shells from a steam inlet connection 507 disposed near the center of the length of the combustion tube 5. The steam introduced into the cooling passage is divided into two streams flowing in the directions opposite to each other. Namely, a portion of the cooling steam flows through an upstream cooling passage in the wall of the combustion tube 5 from the inlet 507 in the upstream direction (i.e., towards the main nozzle 33 side) and other portion of the cooling steam flows through a downstream cooling passage in the wall of the combustion tube 5 from the inlet 507 in the downstream direction (i.e., towards the outlet 52 of the combustion tube). Cooling steam outlet pipes 509a and 509b are connected to the cooling steam passage at the upstream (main nozzle 33 side) end and the downstream (outlet 52 side) end of the combustion tube 5, respectively, in order to collect cooling steam after it cooled the combustor walls. Since the heat-transfer coefficient of steam is relatively large, the walls of the combustor are sufficiently cooled by convective cooling using cooling steam. The conventional compound air cooling system and the steam cooling system as explained above include respective drawbacks. In the first place, in the compound air cooling system using the fin-rings, consumption of cooling air is large. Fig. 3 is a cross sectional view taken along the line IX-IX in Fig. 2. As explained before, the fin-ring 155 is provided with grooves extending along the axial direction on the outer surface thereof. When the fin-ring 155 is attached to the shell 55, an annular clearance 155c must be disposed between the shell 55 and fin-ring 155 in order to avoid contact between the shell 55 and fin-ring due to thermal expansion of the fin-ring. When the manufacturing tolerance and the tolerance in the assembling of the combustor are taken into account, the required width of the clearance 155c becomes almost the same as the depth of the grooves 155b in some cases. Therefore, in the conventional compound air cooling system, since a relatively large clearance 155c between the outer surface of the fin-ring 155b and the inner surface of the shell 55, a large amount of cooling air passes through the clearance 155c in the axial direction and flows into the combustion chamber without passing through the grooves 155b. In other words, a large portion of the cooling air introduced from the inlet 57 flows into the combustion chamber without being used for cooling the fin-ring 155. Consequently, in order to obtain sufficient convective cooling of the fin-ring 155, the amount of cooling air supplied from the inlet 57 must be increased so that a sufficient amount of air passes through the grooves 155b. Further, a large amount of cooling air which passes through the annular clearance 155c and does not contribute to convective cooling of the fin-ring 155 also flows into the combustion chamber and dilutes the combustion gas. Therefore, the drop of the combustion gas temperature due to introduction of cooling air becomes large in the conventional compound air cooling system. The width of the annular clearance 155c may be reduced if the tolerances of machining and assembly of the shell 55 and fin-ring 155 are smaller. However, smaller tolerance in machining and assembly of these parts causes an increase in the cost and time required for manufacturing the combustion tube 5. Further, in the conventional compound air cooling system using the fin-ring 155, since the fin-ring 155 is attached to the shell 55 at only one end thereof, it is difficult to increase the structural strength of the combustion tube assembly 5. On the other hand, the problems such as those in the conventional compound air cooling system as explained above do not occur in the steam cooling system in Fig. 4. However the steam cooling system also has an inherent problem of a large consumption of cooling steam. When the steam cooling system is used, as explained in Fig. 4, cooling steam is introduced into the cooling passage of the combustion tube 5 from the steam inlet 507 disposed near the center of the length of the combustion tube 5 and passes through the upstream cooling passage and the downstream cooling passage in directions opposite to each other. Usually, cooling steam is supplied to both an upstream and a downstream cooling passage at a same supply conditions. However, although cooling steam is supplied at the same supply conditions, the heat loads on the upstream cooling passage and the downstream cooling passage are not the same and, usually, the heat load on the downstream cooling passage is larger than that on the upstream cooling passage. Air-fuel mixture is injected into the combustion tube 5 from the main nozzles 33 at the upstream end thereof and it burns while it flows towards the outlet 52 of the combustion tube 5. Therefore, since the combustion of air-fuel mixture is not completed in the upstream half of the combustion tube 5, the temperature of the combustion gas is relatively low in the upstream half of the combustion tube 5. On the other hand, since combustion of air-fuel mixture is completed at the downstream half of the combustion tube 5, the temperature of combustion gas is higher at the downstream half of the combustion tube 5 than at the upstream half thereof. Consequently, heat load on the downstream cooling passage becomes higher than that on the upstream cooling passage. However, since the supply conditions of the cooling steam to both cooling passage are the same, the supply conditions of cooling steam must be adjusted to meet the requirement of the downstream cooling passage where the heat load becomes the maximum in the conventional steam cooling system. This means that the wall of the upper half of the combustion tube 5 is cooled more than necessary (i.e. excessive cooling occurs at the upper half of the combustion tube). Therefore, in the conventional steam cooling system, an excess amount of cooling steam is required to cool the upper half of the combustion tube 5 excessively. EP-A-0 624 757 discloses a cooling system for a combustor of a gas turbine which utilizes air impingement cooling of the outer liner of the combustor that is subjected to hot combustion gases. EP-A-0 199 534 discloses a liner structure for a gas turbine combustion chamber which is similar to the structure disclosed in Figure 2. US-A-5 950 417 discloses another cooling system for a combustor of a gas turbine in which the combustor unit casing is cooled by an augmented air and oxygen stream discharged into the burning zone. SUMMARY OF THE INVENTION In view of the problems in the related art as set forth above the object of the present invention is to provide a cooling system for a gas turbine combustor which is capable of reducing consumption of cooling steam without lowering the cooling capacity. According to the present invention there is provided a cooling system for a combustor of a gas turbine as defined in claim 1. According to the present invention, there is provided a cooling system for a combustor of a gas turbine which generates combustion gas by burning fuel for driving a turbine comprising a cylindrical combustion tube having an inlet end and an outlet end and forming a combustion chamber therein for burning fuel supplied from the inlet end thereof and supplying combustion gas to a turbine from the outlet end thereof, a plurality of cooling steam passages formed in the wall of the combustion tube and extending along the length of the combustion tube between the inlet end and the outlet end, the cooling steam passages including first cooling steam passages in which cooling steam flows in a first direction and second cooling steam passages in which cooling steam flows in a second direction opposite to the first direction. According to the invention, two groups of cooling steam passages, i.e., the first cooling steam passages and the second cooling steam passages are provided in the wall of the combustion tube. In the first and the second cooling steam passages, cooling steam flows in the direction opposite to each other. Therefore, the average of the temperatures of cooling steam flowing through both first and second cooling steam passages become uniform along the length of the combustion tube and excessive cooling of the inlet side end of the combustion tube does not occur. Thus, the combustion tube is suitably cooled with smaller amount of cooling steam and the amount of cooling steam required for cooling the combustion tube can be reduced. BRIEF DESCRIPTION OF THE DRAWINGS The present invention will be better understood from the description, as set forth hereinafter, with reference to the accompanying drawings in which: Fig. 1 is a longitudinal section view showing a general construction of a gas turbine combustor employing a compound air cooling system; Fig. 2 is an enlarged section view of the wall of a gas turbine combustor employing a conventional compound air cooling system; Fig. 3 is a cross sectional view taken along the line IX-IX in Fig. 2; and Fig. 4 is a longitudinal section view of a gas turbine combustor employing a conventional steam cooling system; Fig. 5 is a longitudinal section view showing a general construction of a gas turbine combustor according to an embodiment of the present invention; Figs. 6A and 6B schematically show the arrangement of cooling steam passages in the combustor in Fig. 5. DESCRIPTION OF THE PREFERRED EMBODIMENT Hereinafter, embodiments of the cooling system of the gas turbine combustor according to the present invention will be explained with reference to Figs. 1 through 6. Fig. 1 is a longitudinal section view showing a general construction of a gas turbine combustor according to the first embodiment of the present invention. A compound air cooling system is used for the combustor in this embodiment. In Fig. 1, reference numeral 1 designates a combustor of a gas turbine as a whole. The combustor 1 consists of a combustor portion 5 and a nozzle portion 3. The nozzle portion 3 contains fuel nozzles for injecting fuel into the combustion portion 5. In this embodiment, the combustion portion 5 includes a combustion tube 5b which acts as a combustion chamber for burning fuel injected from fuel nozzles and a tail pipe 5a which directs the combustion gas generated in the combustion tube 5b to the turbine inlet. The combustion portion 5 in this embodiment is formed by assembling the combustion tube 5b and the tail pipe 5a. The combustor 1 is held in a turbine casing 7 by supports not shown in the drawing. In an actual gas turbine, plural combustors as shown in Fig. 1 are arranged around the turbine at a regular interval and each supplies combustion gas to the turbine. Combustion air (the arrow 11 in Fig. 1) is compressed from the atmosphere by a gas turbine compressor driven by the turbine and supplied to the casing 7. Combustion air 11 enters to the nozzle portion 3 of the combustor 1 through combustion air inlets 35 disposed around the nozzle portion 3 and is injected into the combustion portion 5 of the combustor 1 through a pilot nozzle 31 and main nozzles 33. The nozzle portion 3 includes a cylindrical shell, a pilot nozzle 31 disposed at the center of the shell and plural main nozzles 33 (in this embodiment, eight main nozzles are used) disposed at a regular interval around the pilot nozzle 33. The pilot nozzle 31 is provided with a pilot fuel nozzle 31a for injecting fuel into combustion tube 5b. Fuel injected from the pilot fuel nozzle 31a is ignited by an igniter (not shown in the drawing) and, while it mixes with combustion air flowing through the pilot nozzle 31, burns in a diffusion combustion mode and forms a pilot flame. In this case, the main nozzles 33 are formed as premix nozzles which form a premixed air-fuel mixture by mixing fuel injected from the fuel nozzles 33 into combustion air in the nozzle portion 3 before injecting premixed air-fuel mixture into combustion tube 5b. The main nozzles 33 are provided with swirlers 37. The swirlers 37 are formed as vanes having aerofoil shape cross sections and each is disposed at a predetermined angle to the combustion air flow in the nozzle 33 in order to impart tangential velocity to the combustion air flowing through the nozzle 33. Therefore, a swirl of combustion air is formed in the main nozzle 33 by the swirlers 37. In this embodiment, a gaseous fuel such as town gas and natural gas and liquid fuel such as heavy oil or gas oil are used for fuel. Fuel is injected into the swirl of the combustion air in the nozzle 33 from injection holes (not shown in the drawing) disposed on the-fuel nozzle 33 downstream of the swirlers 37. Therefore, the mixing of the injected fuel and combustion air is promoted by the swirl of combustion air and a uniform pre-mixture of air and fuel is formed in the main nozzle 33. A funnel shaped pilot cone 36 is provided at the outlet of the pilot nozzle 31. The pilot cone 36 prevents the air-fuel mixture injected from the main nozzles 33 from contacting the pilot flame at the portion close to the outlet of the main nozzles 33 in order to avoid back firing in the main nozzles 33. A nozzle extension 34 is provided at the outlet of each main nozzle in order to guide the air-fuel mixture along the side surface of the pilot cone 36. In Fig. 1, numeral 9 designates an air bypass valve. The air bypass valve 9 is a gate valve directly connecting the tail pipe 5b to the interior of the casing 7. The air bypass valve 9 controls the amount of combustion air supplied to the combustion portion 5 from the nozzle portion 3 in accordance with a load of the gas turbine. For example, when the amount of combustion air required for the combustion in the combustion portion 5 (such as during the start up and light load operation of the gas turbine), the air bypass valve 9 is opened by an actuator (not shown) in order to supply combustion air in the casing 7 directly to the interior of the tail pipe 5a. Since this causes a decrease in the combustion air flowing through the main nozzle 33, a suitable pre-mixture of air and fuel is formed in the main nozzle 33. The air-fuel mixture injected from the main nozzles 33 around the pilot cone 36 contacts the pilot flame formed by the pilot nozzle 31 at the periphery of the outlet of pilot cone 36 and is ignited by the pilot flame. Therefore, the air-fuel mixture injected from the main nozzles 33 forms a premixed combustion flame at a portion sufficiently distant from the outlet of the main nozzle 33. The combustion gas generated by the premixed combustion flame flows through the tail pipe 5a and inlet stator blades (not shown) and flows into the rotor blades of the turbine in order to rotate the same. The construction of the combustor as explained above is basically the same as the construction of the combustor employing a conventional compound air cooling system. Fig. 5 is a longitudinal section view similar to Fig. 1 which shows a gas turbine combustor according to the embodiment of the present invention. In this embodiment, cooling of the combustor is performed using cooling steam. In Fig. 5, reference numerals the same as those in Fig. 1 designate elements similar to those in Fig. 1. In this embodiment, the construction of the combustor is different from that of Fig. 1 in that the combustion portion 5 of the combustor 1 is a one-piece construction in which the combustion tube 5b and the tail pipe 5b are formed as an integral part. Therefore, the combustion portion 5 will be referred to as a combustion tube 5 in this embodiment. In this embodiment, similarly to the combustor in Fig. 10, the wall of the combustion tube 5 is cooled by cooling steam along whole length of the combustion tube 5. However, although cooling steam is supplied at the middle portion of the combustor in Fig. 1 (i.e., the portion where the combustion tube 5b and the tail pipe 5a in Fig. 1 is joined) and divided into two flows of cooling steam (i.e., a flow towards upstream and flow towards downstream), a pair of cooling steam inlet pipes 511, 513 and a pair of cooling steam outlet pipes 515, 517 are connected to the combustion tube 5 in this embodiment. Namely, the cooling steam inlet pipe 511 and outlet pipe 515 are connected to the wall of the combustion tube 5 at the outlet end (the turbine end) thereof, and the cooling steam inlet pipe 513 and the outlet pipe 517 are connected to the wall of the combustion tube 5 at the inlet end (the main fuel nozzle side end) thereof. In this embodiment, two groups of cooling steam passages, i.e., first cooling steam passage 523 which is connected to the inlet pipe 511 and outlet pipe 517 and second cooling passages 521 connected to the inlet pipe 513 and outlet pipe 515 are formed within the wall of the combustion tube 5. Cooling steam flows through the first cooling steam passages 523 in the upstream direction (i.e., from the outlet side to the inlet side of the combustion tube 5) and in the downstream direction (i.e., from the inlet side to the outlet side of the combustion tube 5). Figs. 6A and 6B schematically show the arrangement of the first cooling steam passages 523 and the second cooling steam passages 521 in the wall of the combustion tube 5. Fig. 6A shows the arrangement of the passages along the circumference of the combustion tube 5 at the outlet portion 52 thereof. As shown in Fig. 6A, the first cooling steam passages 523 and the second cooling steam passages 521 are disposed alternately around the circumference of combustion tube 5. Fig. 6B is a section view taken along the line "B"-"B" in Fig. 6A. As shown in Fig. 6B, an annular steam inlet manifold 513a which connects the second cooling steam passages 521 to the steam inlet pipe 513 and an annular steam outlet manifold 517a which connects the first cooling steam passages 523 to the steam outlet pipe 517 are formed around the outlet portion 52 of the combustion tube 5. Although not shown in the drawing, an annular steam inlet manifold 511a and an annular steam outlet manifold 515a are formed around the inlet portion of the combustion tube 5. The arrangement of the steam manifolds at the inlet of the combustion tube 5 is similar to the steam manifolds in Figs. 6A and 6B. The annular inlet manifold 511a connects the first cooling steam passages 523 to the steam inlet pipe 511 and an annular steam outlet manifold 515a connects the first cooling steam passages 523 to the steam outlet pipe 515. In this embodiment, flow rates of cooling steam in the first and second cooling steam passages 523, 521 are adjusted in such a manner that the inlet steam temperatures in the inlet manifolds 511a and 513a are the same, and the outlet steam temperatures in the outlet manifolds 515a and 517a are the same. For example, the steam temperature in the inlet manifolds 511a and 513a (i.e., the steam temperature at the inlet of the cooling steam passages 521 and 523) is adjusted to about 300°C and the steam temperature in the outlet manifolds 515a and 517a (i.e., the steam temperature at the outlet of the cooling steam passages 521 and 523) is adjusted to about 600°C. Therefore, the average temperatures of the steam flowing through the cooling steam passages 521 and 523 becomes nearly uniform along the length of the cooling steam passages 521 and 523. As explained before, fuel injected in the combustion tube 5 burns while it flows through the combustion tube 5. Therefore, the temperature of the combustion gas is not uniform along the axis of the combustion tube 5. Namely, the temperature of the combustion gas is relatively low at the inlet side of the combustion tube 5 and relatively high at the outlet side of the combustion tube 5. On the other hand, the wall temperature of the combustion tube 5 (the metal temperature) changes in accordance with the steam temperature in the cooling steam passages. More specifically, difference between the metal temperature of the combustion tube 5 and the steam temperature in the cooling steam passages is nearly constant regardless of the combustion gas temperature in the combustion tube 5. In this embodiment, as explained above, since the cooling steam in the cooling steam passages 521 and 523 flows in the direction opposite to each other, the average steam temperature in the cooling steam passages 521 and 523 becomes nearly constant along the length of the combustion tube 5. Therefore, the metal temperature of the combustion tube 5 also becomes nearly constant along the length of the combustion tube 5 and, since the excessive cooling of the wall of the combustion tube does not occur, the consumption of the cooling steam can be reduced in the present embodiment.

Claims: A cooling system for a combustor of a gas turbine comprising: a cylindrical combustion tube (5) having an inlet end and an outlet end and forming a combustion chamber therein for burning fuel supplied from the inlet end thereof and supplying combustion gas to a turbine from the outlet end thereof; the cooling system further comprising a plurality of cooling steam passages (523,521) formed in the wall of the combustion tube (5) and extending along the length of the combustion tube (5) between the inlet end and the outlet end, said cooling steam passages including first cooling steam passages (523) in which cooling steam is adapted to flow in a first direction; the cooling system being characterized in that said cooling passages further include second cooling steam passages (521) in which cooling steam is adapted to flow in a second direction opposite to the first direction. The cooling system for a gas turbine combustor as set forth in claim 1, wherein the first cooling steam passages (523) and the second cooling steam passages (521) are disposed alternately in a circumferential direction of the combustion tube (5).

Inventors: MANDAI SHIGEMI, SUENAGA KIYOSHI, TANAKA KATSUNORI, MANDAI, SHIGEMI, SUENAGA, KIYOSHI, TANAKA, KATSUNORI, Mandai, Shigemi, Takasago Res. &amp; Develop. Center, Suenaga, Kiyoshi, Takasago Res. &amp; Develop. Center, Tanaka, Katsunori, Takasago Res. &amp; Develop. Center
Applicants: MITSUBISHI HEAVY IND LTD, MITSUBISHI HEAVY INDUSTRIES, LTD.
Application year: 2001
Main classification: F23R3
IPC codes: F23L7, F23R3, F02C7