Patent Publication Number: US-9885480-B2

Title: Combustion chamber of a combustor for a gas turbine

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the US National Stage of International Application No. PCT/EP2012/076604 filed Dec. 21, 2012, and claims the benefit thereof. The International Application claims the benefit of European Application No. EP12150314 filed Jan. 5, 2012. All of the applications are incorporated by reference herein in their entirety. 
     FIELD OF INVENTION 
     The present invention relates to a combustor and more particularly to combustion chamber of a gas turbine. 
     BACKGROUND OF INVENTION 
     In gas turbines, fuel is delivered from a source of fuel to a combustor where the fuel is mixed with air and ignited to produce hot combustion products which are generally known as working gases. As will be appreciated, the amount of working gas produced depends on a proper and effective mixing of the fuel and air in the combustor. 
     DE 10 2011 000879 A1 discloses a combustor for a gas turbine. The combustor comprises a combustion chamber in which a working medium consisting of fuel and air is mixed and subsequently burned. The air intake of cooling air into an annular channel is allowed by an outer shell in which airfoils allow to guide incoming air to have a swirl when entering that annular channel. 
     Currently, swirlers are used in the combustor to generate swirls in the air so that the air is properly mixed with fuel. Proper mixing of the fuel and air results in increasing the efficiency of gas turbine since the generation of the working gas by subsequent burning of the fuel and air mixture is more efficient. This also reduces the amount of NOx gases produced from the burning of the fuel and air mixture. 
     Burners with swirlers are widely known. Nevertheless several problems may occur in known combustion chambers, like the combustion chamber of DE 10 2011 000879 A1. For example pulsation and vibrations may occur within the combustion chamber. Furthermore it may be a disadvantage that the combusted fluid may be turbulent or may just be guided by a combustion liner such that the angle of attack on subsequent turbine vanes or blades is not optimal. 
     SUMMARY OF INVENTION 
     It is therefore an object of the present invention to provide an improved arrangement in a combustor to overcome the mentioned problems. 
     The object is achieved by providing a combustion chamber for a combustion chamber, a combustor, and a gas turbine according to the claims. 
     The present invention provides the combustion chamber for the combustor for a gas turbine which is an annular combustion chamber including a plurality of segments arranged annularly about an axis of the combustion chamber, each segment comprising a radial inner wall portion and a radial outer wall portion, a first section comprising an opening for the installation of a burner, and a second section at which at least one airfoil extends between the radial inner wall portion and radial outer wall portion of the segment. The first section and the second section are located at opposing first end and second end of the combustion chamber. By having the burner and the airfoil at respective first section and second section, which correspond to the opposing first end and second end of the combustion chamber space for mixing of fuel and air is increased. In addition the airfoil increases the swirling in the air passing through it which increases the mixing of fuel and air. The airfoil present at the second end guides the working medium through an exit located at the second end of the combustion chamber. 
     Each segment comprises an inner surface and an outer surface with a channel for air defined between the inner and outer surface, wherein air in the channel is conducted from the airfoil. Such an arrangement ensures that air and fuel are properly mixed inside the combustor. 
     Herein, compressed air from a compressor of the gas turbine is directed into the airfoil. 
     In one embodiment, the segment includes at least one air inlet at the second section wherein the airfoil is located such that air entering the segment through the air inlet is swirled. This arrangement increases the mixing between the fuel and the air due to increase in swirl of the air. 
     In one embodiment, the first section and the second section are located at the first end and the second end of the combustion chamber, this increase space for effective mixing of the fuel with air. 
     In one embodiment, the airfoil and the wall portion are formed of one piece of a material which increases the dimensional stability of the segment. 
     In one embodiment, the airfoil and the wall portion are cast which obviates the need for machining and welding. In addition, the airfoil and the wall portion would be a single piece and would exhibit uniform properties with increased strength. 
     In another embodiment, two adjacent segments are assigned to one burner, which enables greater mixing of air with the fuel which then is then ignited by the burner. 
     In another embodiment, each segment comprises two airfoils to increase the swirling of air in the combustion chamber. 
     In one embodiment, the outer surface of the segment is brazed which ensures that the air from the compressor is kept within the combustor. 
     In one embodiment, the airfoil and the wall portions are formed from an alloy, which increases strength of the segment and are capable of withstanding high temperatures. 
     In one embodiment, the alloy is Nickel based gamma prime strengthened alloy. The creep strength of this type of casting alloy is significantly higher than those in traditional combustor alloys which results in improved dimensional stability. In addition, gamma prime alloy is ductile and thus imparts strength to the matrix without lowering the fracture toughness of the alloy. 
     In another embodiment, the alloy is IN738LC. IN738LC is a nickel based superalloy which exhibits compatibility with currently used thermal barrier coating systems. 
     In another embodiment, the alloy is CM247CC. CM247CC is also a nickel based superalloy which is also compatible with currently existing thermal barrier coating systems, as well as the ability to form a layer of protective alumina which provides a significant improvement in oxidation resistance as compared to other alloys. 
     The above-mentioned and other features of the invention will now be addressed with reference to the accompanying drawings of the present invention. The illustrated embodiments are intended to illustrate, but not limit the invention. The drawings contain the following figures, in which like numbers refer to like parts, throughout the description and drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic diagram of a gas turbine; and 
         FIG. 2  is a schematic diagram of a combustor and its combustion chamber, in accordance with aspects of the present technique. 
         FIG. 3  is a schematic end view of the annular combustor looking at a second section. 
         FIG. 4  is a schematic end view of the annular combustor looking at a first section. 
     
    
    
     DETAILED DESCRIPTION OF INVENTION 
       FIG. 1  is a schematic diagram of a gas turbine  10  depicting internal components. The gas turbine  10  includes a rotor  13  which is mounted such that it can rotate along an axis of rotation  12 , has a shaft  11  and is also referred to as a turbine rotor. 
     The gas turbine  10  includes an intake housing  14 , a compressor  15 , a combustor  16  having a combustion chamber  20 , a turbine  18 , and an exhaust-gas housing  19  following one another along the rotor  13 . The combustion chamber  20  is an annular combustion chamber with a plurality of coaxially arranged burners  17 . 
     The annular combustion chamber  20  is in communication with an annular hot-gas passage  21 , where, by way of example, four successive turbine stages  22  form the turbine  18 . 
     It may be noted that each turbine stage  22  is formed, for example, from two blade or vane rings. As seen in the direction of flow of a working medium  23  from the combustion chamber  20  to the turbine  18 , in the hot gas passage  21  a row  25  of guide vanes  40  is followed by a row  35  formed from rotor blades  30 . The guide vanes  40  are secured to an inner housing  48  of a stator  53 , whereas the rotor blades  30  of the row  35  are fitted to the rotor  13  for example by means of a turbine disk  43 . 
     A generator not shown in  FIG. 1  is coupled to the rotor  13 . During the operation of the gas turbine  10 , the compressor  15  sucks in air  45  through the intake housing  14  and compresses it. The compressed air provided at the turbine-side end of the compressor  15  is passed to the burners  17 , where it is mixed with a fuel. The mix is then burnt in the combustion chamber  20 , forming the working medium  23 . From there, the working medium  23  flows along the hot-gas passage  21  past the guide vanes  40  and the rotor blades  30 . The working medium  23  is expanded at the rotor blades  30 , transferring its momentum, so that the rotor blades  30  drive the rotor  13  and the latter in turn drives the generator coupled to it. 
     In addition, while the gas turbine  10  is in operation, the components which are exposed to the hot working medium  23  are subjected to thermal stresses. The guide vanes  40  and the rotor blades  30  of the first turbine stage  22 , as seen in the direction of flow of the working medium  23 , together with the heat shield bricks which line the annular combustion chamber  20 , are subject to the highest thermal stresses. These components are typically cooled by a coolant, such as oil. 
     As will be appreciated, the components of the gas turbine  10  are made from a material such as superalloys which are iron-based, nickel-based or cobalt-based. More particularly, the turbine vanes  40  and/or blades  30  and components of the combustion chamber  20  are made from the superalloys mentioned hereinabove. 
     The combustion chamber  20  which is an annular combustion chamber  20  in the presently contemplated configuration includes a multiplicity of burners  17  arranged circumferentially around the axis of rotation  12  and open out into a common combustion chamber space and generates flames. To achieve a high efficiency, the combustion chamber  20  is designed for a temperature of the working medium  23  of approximately 1000 degree Celsius to 1600 degree Celsius. To allow a long service life even with these operating parameters, which are unfavorable for the materials, the combustion chamber wall is provided, on its side which faces the working medium  23 , with an inner lining formed from heat shield elements. 
     Referring now to  FIG. 2 , a schematic diagram of the combustor  16  and its combustion chamber  20 , respectively, is depicted in accordance with aspects of the present technique. The combustor  16  includes the combustion chamber  20  which in the presently contemplated configuration is an annular combustion chamber which includes a plurality of segments arranged circumferentially around the axis  12 .  FIG. 2  shows a cross section through one of those segments. As an example, a total of twenty segments would form the combustion chamber  20 . Each segment includes an inner wall portion  54  and an outer wall portion  56 . 
     It may be noted that the inner wall portion  54  and the outer wall portion  56  are positioned radially outwards from the axis  12 . 
     In accordance with aspects of the present technique, the segment has a first section  62  and a second section  64 , with the burner installed at an opening  63  at the first section  62  and an airfoil  52  such as a guide vane at the second section  64 . 
     It may be noted however, that the first section may be at the first end and the second section may be at the second end, wherein the first end and the second end are opposing each other. For the purpose of explanation the terms “first section” and “first end” and the “second section” and “second end” are used interchangeably. 
     As previously noted, the combustion chamber  20  includes the opening  63  at the first end  62  as depicted in  FIG. 2 . A burner  17  is installed at the opening  63  at the first end  62 . Air from the compressor  15  is directed via a panel  72  and through the airfoil  52  in to the combustion chamber  20  and mixed with fuel. Fuel is directed into the combustion chamber via a fuel pipe  69 . The air and fuel mixture is ignited by the burner  17  to produce the working medium  23 . 
     In accordance with aspects of the present technique, the airfoil  52  is present at the second end  64 . The airfoil  52  extends between the inner wall portion  54  and an outer wall portion  56 . The compressed air from the compressor  15  is directed into the airfoil  52  as indicated by reference numeral  51 . Air  51  in the airfoil  52  could also be swirled to create turbulence. 
     The combustor segment includes an inner surface  60  and an outer surface  58  forming a channel  70  there between to conduct air from the airfoil  52  to the channel  70 . Air is mixed with a fuel supplied through the fuel pipe  69  and is ignited by the burner  17  to generate flames  68  and hence produce the working medium  23  for the turbine. This working medium  23  is guided through an exit by the airfoil  52  present at the second end  64  out of the combustion chamber  20 . 
     Additionally the combustor  16  may include cooling holes, or cooling pipes at the end walls to supply cooling air to cool the walls of the combustion chamber  20 . 
     As previously noted, the panel  72  is located at the first section or the first end  62  inside the combustion chamber  20  which acts as a Helmholtz panel to draw air into the combustion chamber  20 . The panel  72  alongwith the airfoil  52  acts as a Helmholtz resonator and will keep the air inside the chamber  20  to ensure effective mixing of the air with the fuel and hence better combustion is achieved. 
       FIG. 3  is an end view of the annular combustor  16  looking upstream at the second section  64 .  FIG. 4  is an end view of the annular chamber  16  looking downstream at the first section  62 . As previously noted and as can be seen in  FIGS. 3 and 4 , the combustion chamber  20  includes a plurality of segments  60  separated at respective interfaces  62 . In  FIG. 3  an example embodiment having sixteen segments  60  are shown. The segments  60  are arranged adjacent to each other in a manner such that two segments  60  are assigned to one burner  17 . In addition, each segment  60  includes two airfoils  52  located adjacent to each other. The inner wall portion  54 , the outer wall portion  56  and the airfoil  52  in a segment  60  are formed of one piece of a material. More particularly, the airfoil  52 , the inner wall portion  54  and the outer wall portion  56  are cast to produce a single piece material. 
     In accordance with the aspects of the present technique, the airfoil  52  and the wall portions  54 ,  56  are made of material such as alloys, for example nickel-based superalloy. These alloys are capable of withstanding high temperatures which may exceed 650 degree centigrade. The airfoil  52  and the wall portions  54 ,  56  are cast from the same type of alloy such as, Nickel-based gamma prime strengthened alloy. 
     It may be noted that the inner wall  54  and the outer wall  56  may be coated with a thermal barrier coating to protect against the high temperatures of the hot gas. Hence it may be noted that the alloys in the present technique are chosen which are compatible with the thermal barrier coatings. Furthermore, it may be noted that alloys such as Nickel-based gamma prime strengthened alloys include a higher quantity of aluminum than the traditional alloys used in the combustors. The presence of aluminum increases the life time of the thermal barrier coatings that are applied to the wall. 
     Additionally, the alloys for casting the segments of the combustion chamber are chosen which have a better castability and are capable of casting large components such as the segments of combustion chamber  20 , such as IN738LC, which is a nickel-based super alloy and has a chemical composition in wt % as Cobalt 8.59, Chromium 16.08, Aluminum 3.43, Silicon 0.18, Carbon 0.11, Phosphorus 0.01, Iron 0.50, Boron 0.05, Sulfur 0.01, Tungsten 2.67, Tantalum 1.75, Nobelium 0.90, Titanium 3.38, Manganese 0.03, Copper 0.03 and Nickel as remaining. 
     Alternatively, alloy such as CM247CC, which is also a nickel based superalloy may be used for casting the segment. This alloy has a composition in wt % as Cobalt 10, Chromium 8, Molybdenum 0.5, Tungsten 9.5, Aluminum 5.65, Tantalum 3, Hafnium 1.5, Zirconium 0.1, Carbon 0.1 and Nickel as remaining. 
     Although the invention has been described with reference to specific embodiments, this description is not meant to be construed in a limiting sense. Various modifications of the disclosed embodiments, as well as alternate embodiments of the invention, will become apparent to persons skilled in the art upon reference to the description of the invention. It is therefore contemplated that such modifications can be made without departing from the embodiments of the present invention as defined.