Fuel jetting nozzle assembly for use in gas turbine combustor

Abstract: A fuel jetting nozzle comprises a nozzle head (6) secured to a head plate, a fuel jetting nozzle (10) secured to the nozzle head and having fuel jetting holes (17) opened to an inside of a combustion chamber, swirling vanes (16) disposed on an outer peripheral portion of the fuel jetting nozzle for supplying air (23) as an annular swirling flow in the combustion chamber of the combustor liner, the fuel jetting holes being formed at base portions of the swirling vanes, and a cooling member composed of a plurality of cooling holes (30) formed to a front end portion of the fuel jetting nozzle so as to introduce a portion (31) of the air into an area in front of a central portion (18) of the nozzle from an area upstream of the swirling vanes.

Description: Background of the Invention The present invention relates to a gas turbine combustor comprising a fuel jetting nozzle assembly, particularly, in which a burn damage to the central portion of the extreme end portion of the fuel jetting nozzle is prevented as much as possible. Figs. 5 to 7, explained hereinafter, represents an example of a typical well-known gas turbine combustor of a conventional structure. Referring to Figs. 5 to 7, a plurality of gas turbine combustors are arranged on the outer peripheral portion of a discharge casing 2 of an air conditioner 1. A combustor liner 5 by which an internal combustion chamber 4 is enclosed is housed within the combustor casing 3, and a nozzle head 6, an igniter 7 and a flame detector, not shown, are provided in the internal combustion chamber 4. The nozzle head 6 is mounted on a head plate 8, and this head plate 8 and a flow sleeve 9 are mounted on the combustor casing 3. The head plate 8 is disposed so as to close one end of the casing 3. A fuel jetting nozzle 10 is mounted on the nozzle head 6 and prevented from rotating by a locking plate 11. The combustor liner 5 is mounted on the extreme, i.e. front, end portion of the fuel jetting nozzle 10, and a liner supporter 12 provided on the flow sleeve 9 supports the combustor liner 5. A transition piece 13 is connected to the extreme end portion (the downstream area) of the combustor liner 5. The combustor liner 5 is connected to a first-stage turbine stationary blade 14a of a gas turbine 14 by way of the transition piece 13. An air intake passage 15 is formed in the outer peripheral portion of the fuel jetting nozzle 10. A swirl vane 16 is disposed between the air intake passage 15 and the internal combustion chamber 4. Fuel jetting holes 17, through which the inside of the fuel jetting nozzle 10 is communicated with the swirl vane 16, are provided on the peripheral wall portion of the fuel jetting nozzle 10. The front side of a central end portion 18 of the fuel jetting nozzle 10 faces the inside of the internal combustion chamber 4 and forms a portion thereof. A fuel intake 19 is formed in the nozzle head 6, from which a gaseous fuel 20 is introduced into the fuel jetting nozzle 10. An air flow around the gas turbine combustor will be explained hereunder. An air 21 discharged from the air conditioner 1 flows around the transition piece 13 and is guided in a direction opposite to the flow of combustion gas 22 between the combustor liner 5 and the flow sleeve 9. The discharged air 21 is introduced into the internal combustion chamber 4 through air passages which are broadly divided into three portions. That is, the discharged air 21 is divided into primary air 23 introduced from the swirl vane 16 around the fuel jetting nozzle 10, secondary air 25 introduced from an air guide 24 provided on the trunk portion of the combustor liner 5, and tertiary air 26 for dilution purposes introduced from the holes provided downstream of the air guide 24 used for the secondary air. A stable annular vortex area, i.e. flame area, of the primary air 23 and the gaseous fuel 20 is formed in the inside of the annular swirl flow caused by the primary air 23. The stable annular vortex area stabilizes and maintains the combustion flame, and the combustion gas 22 flows to the exit area of the combustor liner 5. The primary air 23 is mixed with the tertiary air 26, and cools the combustor liner 5 and decreases the gas temperature so that the liner exit temperature becomes a temperature required for the turbine. In this viewpoint, the primary air 23, the secondary air 25 and the tertiary air 26 are allocated in various ways so as to control combustion performance. In some instances, the secondary air 25 and tertiary air 26 may not be provided. Furthermore, the primary air 23 and the secondary air 25 may be mixed with the gaseous fuel 20 beforehand and introduced into the internal combustion chamber 4. The discharged air 21 passes through a slot, not shown, used to cool the combustor liner 5 and is supplied to the internal combustion chamber 4. The details of the fuel jetting nozzle 10 are shown in Fig. 6. Some of the primary air 23 of the discharged air 21 discharged from the air compressor 1 enters from the air intake passage 15 into the internal combustion chamber 4. At this time, the air is mixed with the gaseous fuel 20 jetted from the fuel jetting holes 17, passes the swirl vane 16 disposed around the fuel jetting nozzle 10, is jetted into the internal combustion chamber 4 while it is being swirled and is then ignited. Ignition is performed by the igniter 7 shown in Fig. 5. The combustion gas 22 passes the transition piece 13 and is introduced to the first-stage turbine stationary blade 14a of the gas turbine 14, causing a turbine rotor, not shown, to rotate by using the energy thereof. The flow of gas near the outlet of the fuel jetting nozzle 10 inside the internal combustion chamber 4 is shown in Fig.7. The primary air 23 passes the swirl vanes 16 of the fuel jetting nozzle 10 and flows into the internal combustion chamber 4 while it is being swirled. The secondary air 25 which flows into the internal combustion chamber 4 through the air guide 24 provided in the trunk portion of the combustor liner 5 flows into a swirling flow 27 formed by air passing through the fuel jetting nozzle 10, forming a reverse flow, i.e. vortex flow, flame area 28 in the central portion and a reverse flow, i.e. vortex flow, flame area 29 in the outer periphery. The local temperature of the combustion gas inside the reverse flow flame area 28 in the central portion becomes a high temperature above approximately 2,000°C and a stable flame can thus be maintained. However, in the above-described conventional fuel jetting nozzle for use in a gas turbine combustor, problems arise. For example, the central end portion 18 of the fuel jetting nozzle 10 is burned by radiation and forced convection by high-temperature gas of the reverse flow flame area 28 in the central portion, and the service life of the fuel jetting nozzle 10 becomes short. A gas turbine combustor comprising a fuel jetting nozzle assembly according to the preamble of the appended claim has been disclosed in GB-A-2 101 732. The fuel injector of this assembly is fixed directly to the walls of the combustion chamber. The fuel injector comprises a housing with an open upstream end, into which a fuel duct leads. Fuel, leaving fuel outlets of the fuel duct, is mixed with compressed air within the open upstream end of the housing and leaves the housing through radial ducts at the downstream side of the air swirl vanes, which are provided at the outer periphery of the housing. The mixture of fuel and air leaving the duct is mixing with the swirl air flow, leaving the air swirl vanes and is burning within the combustion chamber. To avoid that any carbon particles which may be formed by combustion adhere to the housing within the combustion chamber, the front end of the housing is provided with an end cap which defines a plenum chamber. The plenum chamber is supplied with compressed air through ducts within the housing. The front wall facing to the combustion chamber of the end cap is formed with openings, through which swirling streams of air are leaving the plenum chamber and sweep over the front face of the end cap to avoid any adhering of carbon particles onto the end cap. The construction of these known fuel jetting nozzle assembly is rather complicated, for its housing must be formed with axial ducts for supplying the plenum chamber with compressed air. The end cap additionally complicates the structure. Summary of the Invention It is an object of the invention to provide a gas turbine combustor comprising a fuel jetting nozzle assembly, which fuel jetting nozzle assembly is of a simple construction and has a lengthened service life by effectively preventing a burn damage to the central end portion of the jetting nozzle. This object is achieved by a gas turbine combustor comprising a fuel jetting nozzle assembly according to the appended claim. With the inventive gas turbine combustor comprising a fuel jetting nozzle assembly the fuel jetting holes which are supplied directly with gaseous fuel are formed at the base portions of the swirling vanes, i.e. the burning mixture is generated directly between the swirling vaines which leads to a well prepared, fast burning combustion gas flowing within the combustion chamber. The inventive arrangement of the cooling holes as simply manufacturable bores, which lead from the room outside the fuel jetting nozzle means, supplied with compressed air, into the combustion chamber, allows an effective cooling at the central end portion of the fuel jetting nozzle by the forced convection of the cooling air. BRIEF DESCRIPTION OF THE DRAWINGS For a better understanding of the present invention and to show how the same is carried out, reference will be made, by way of preferred embodiments, to the accompanying drawings, in which: Fig. 1 is a sectional view, taken along the line I-I shown in Fig. 2, mentioned below, illustrating a first embodiment of a fuel jetting nozzle assembly for use in a gas turbine combustor according to the present invention; Fig. 2 is a view of the fuel jetting nozzle assembly for use in the gas turbine combustor shown in Fig. 1 taken in the direction of the arrow II in Fig. 1; Fig. 3 is a sectional view taken along the line III-III of Fig. 2; Fig. 4 is a view illustrating the relationship between the swirl flow inside the combustion chamber of the fuel jetting nozzle assembly for use in a gas turbine combustor and the flow of the secondary air; Fig. 5 is a sectional view illustrating a conventional gas turbine combustor; Fig. 6 is an expanded view of the essential portion of the fuel jetting nozzle provided in the gas turbine combustor shown in Fig. 5; and Fig. 7 is a view illustrating the relationship between the swirl flow inside the combustion chamber of the conventional fuel jetting nozzle and the flow of the secondary air. DESCRIPTION OF THE PREFERRED EMBODIMENTS The present invention will be explained hereunder with reference to Figs. 1 to 4. Figs. 1 to 4 illustrate an embodiment of a fuel jetting nozzle assembly for use in a gas turbine combustor according to the present invention. Components which are the same as those in a conventional fuel jetting nozzle assembly shown in Figs. 5 to 7 are given the same reference numerals, and thus an explanation thereof is omitted herein. A basic difference of the fuel jetting nozzle of the present invention from the conventional one is that a plurality of swirl vanes 16 by which fuel air is made to flow into the internal combustion chamber 4 are evenly arranged circumferentially on the outer peripheral portion of the fuel jetting nozzle 10, and that fuel jetting holes 17 are provided on the base portions of the swirl vanes 16. The fuel jetting nozzle 10 is fastened to the nozzle head 6 having the fuel intake 19. Each of cooling holes 30, formed at the base portions of the respective swirl vanes 16 and positioned between the adjacnet two fuel jetting holes 17, which reaches the front of the central portion 18 of the front end portion of the fuel jetting nozzle 10 from an area upstream of the swirl vanes 16, is provided on the peripheral wall of the fuel jetting nozzle 10. Each of the cooling holes 30 is provided with an inward angle γ with respect to the front side of the central end portion 18 of the fuel jetting nozzle 10. As shown in Fig.3, it has a swirl angle component β in the same direction as that of the swirl anglea of the swirl vanes 16. The operation of this embodiment will be explained hereunder with reference to Fig. 4. Some of the air 21, discharged from the air compressor 1, passes the swirl vanes 16 as the primary air 23, becomes an annular swirl flow and flows into the internal combustion chamber 4. As a result, a reverse flow, i.e vortex flow, flame area 28 is formed in the central portion. At the same time, some of the discharged air 21 flows into the cooling holes 30 as cooling air 31. This cooling air 31 flows out to the front side of the central flame area 28 while it takes away heat by forced convection cooling, which heat flows in from the front side of the fuel jetting nozzle 10. The air flown out to the front side of the central flame area 28 forms an air layer in front of the central end portion 18 of the fuel jetting nozzle 10. Thus, the front side is protected from fuel gas by film-cooling effect. More particularly, since the cooling hole 30 is provided with a swirl angle component β in the same direction as the swirl angle of the swirl vane 16 and also with an inward angle γ, the cooling air 31 is expanded on the front side of the central end portion 18 of the fuel jetting nozzle 10 while it is being swirled. As a result, a high film-cooling effect can be attained. According to this embodiment, as described above, the central end portion 18 of the fuel jetting nozzle 10 can be cooled by the forced convection using the cooling air 31 passing through the cooling holes 30 and by the film-cooling effect using an air layer. As a result, a burn damage to the central end portion 18 of the fuel jetting nozzle 10 can be prevented.

Claims: A gas turbine combustor comprising a fuel jetting nozzle assembly , said gas turbine combustor further comprising an outer casing (2), a combustor liner (5) disposed inside the outer casing and having a combustion chamber (4) and a head plate (8) closing one end of the outer casing, said fuel jetting nozzle assembly comprising: a fuel jetting nozzle means (10) having fuel jetting holes (17) opened to an inside of the combustion chamber; a swirling means disposed on an outer peripheral portion of the fuel jetting nozzle means for supplying air as an annular swirling flow in the combustion chamber of the combustor liner, said swirling means comprising a base body and a plurality of swirling vanes (16) integrally disposed on the base body and separated equally, said fuel jetting holes being formed at the base body of the swirling means; and a cooling means formed at a front jetting end portion of the fuel jetting nozzle means so as to introduce a portion of the air into a forward portion of a central portion of the front jetting end portion of the nozzle jetting means from an upstream side, with respect to an air flow direction, of the swirling means so as to flow the air into the combustor liner, said cooling means being composed of a plurality of cooling holes (30), having a swirling inclination component with respect to the front surface of the central portion (18) of the front end of the jetting nozzle means, characterized in that said fuel jetting nozzle means (10) is secured to a nozzle head (6) which is secured to the head plate (8), said base body of the swirling means having an upstream side front surface constituting a boundary wall structure between the combustion gas and the fuel, said fuel jetting holes (17), which are directly supplied with fuel, are formed at the base portions of the swirling vanes (16) and each of said cooling holes is positioned between two adjacent fuel jetting holes, each of said cooling holes having an inward inclination with respect to the front surface of the central portion (18) of the front end of the jetting nozzle means, with said swirling inclination component having the same direction as that of an inclination of each of the swirling vanes for causing swirling flow.

Inventors: KOBAYASHI TAKAHIRO, OKAMOTO HIROAKI, KOBAYASHI, TAKAHIRO, OKAMOTO, HIROAKI
Applicants: TOSHIBA KK, KABUSHIKI KAISHA TOSHIBA
Application year: 1992
Main classification: F23D14
IPC codes: F23R3, F23D14