Patent Publication Number: US-2013232987-A1

Title: Gas turbine fuel injector with insulating air shroud

Description:
TECHNICAL FIELD 
     The present disclosure relates generally to a fuel injector for a gas turbine engine, and more particularly, to a gas turbine fuel injector with an insulating air shroud. 
     BACKGROUND 
     Gas turbine engines produce power by extracting energy from a flow of hot gas produced by combustion of fuel in a stream of compressed air. In general, turbine engines have an upstream air compressor coupled to a downstream turbine with a combustion chamber (“combustor”) in between. Energy is released when a mixture of compressed air and fuel is ignited in the combustor. The resulting hot gases are directed over blades of the turbine, spinning the turbine, thereby, producing mechanical power. In typical turbine engines, one or more fuel injectors direct some type of liquid or gaseous hydrocarbon fuel (such, diesel fuel or natural gas) into the combustor for combustion. Some embodiments of fuel injectors are designed to direct both a liquid and a gaseous fuel into the combustor. In these embodiments, the turbine engine may operate on one fuel as the primary fuel with the other fuel used during periods of unavailability of the primary fuel. For example, some gas turbine engines may normally operate on natural gas fuel. In these turbine engines, diesel fuel may be used during periods of natural gas unavailability. The fuel is mixed with compressed air (from the air compressor), in the fuel injector, and delivered to the combustor for combustion. This compressed air, which may exceed 800° F. (426.7° C.) in temperature, may surround sections of the fuel injector, and may create a hot ambient environment for the fuel injector. Combustion of the fuel in the combustor creates hot gases exceeding 2000° F. (1093.3° C.), which may heat surrounding surfaces. The heat released due to combustion may also heat fuel injectors, which may be coupled to the combustor. 
     Fuel injectors include fuel lines and fuel galleries that are used to direct the fuel to the fuel injector and deliver the fuel to the combustor. In a fuel injector that is configured to deliver both liquid and gaseous fuel to combustor, separate fuel lines may deliver the liquid and gaseous fuel to the fuel injector. When the turbine engine operates on gaseous fuel, the liquid fuel may remain in the fuel lines and galleries. In some embodiments, the liquid fuel may be purged from the liquid fuel lines and galleries. However, even in these embodiments, the liquid fuel may exist as a coating on these purged lines and galleries. Due to operating conditions of the fuel injector, the liquid fuel in the liquid fuel lines and galleries may be exposed to ambient temperatures of about 500° F.-800° F. (260° C.-426.7° C.) and injector surface temperatures of 1000° F.-2000° F. (537.8° C.-1093.3° C.). This high temperature may lead to coking of the liquid fuel in the lines and galleries. Over time, the coke may deposit on the lines and galleries and lead to flow restrictions and inoperable conditions. 
     U.S. Pat. No. 7,117,675 (&#39;675 patent), a patent issued to Kaplan et al. on Oct. 10, 2006, describes a cooling system for gas turbine liquid fuel components to prevent coking. In the system of the &#39;675 patent, a sleeve surrounds a liquid fuel component and a device is used to provide a current of cool air through a space between the liquid fuel component and the sleeve. In the cooling system of the &#39;675 patent the sleeve surrounding the liquid fuel component includes a plurality of spacers for centering the sleeve around the liquid fuel component to create an annulus between the sleeve and the liquid fuel component, through which the current of cool air flows. The current of cool air that is used to cool the liquid fuel component is directed to the annular space using a conduit connected between the cool air device and the sleeve. Although the cooling system of the &#39;675 patent may prevent coking of the liquid fuel within the liquid fuel component, it may have some drawbacks. For instance, using a cool air device to blow cool air around the liquid fuel component may increase the complexity and cost of operating the turbine engine. In addition using individual sleeves to provide an annular space around each liquid fuel component may introduce design complexities when space is limited. 
     SUMMARY 
     In one aspect, a fuel injector for a gas turbine engine is disclosed. The fuel injector includes an injector housing extending from a first end to a second end along a longitudinal axis. The second end of the housing is fluidly coupled to a combustor of the turbine engine and the housing includes a liquid fuel gallery annularly disposed about the longitudinal axis. The fuel injector also includes a stem extending longitudinally from the first end of the housing to a third end. The stem includes a liquid tube configured to deliver liquid fuel to the fuel injector. The fuel injector also includes an annular shell extending along the longitudinal axis from the first end to the third end and circumferentially disposed about the stem. The fuel injector further includes an insulating air shroud formed inside the shell. The air shroud includes a layer of air between the shell and the stem. 
     In another aspect, a method of operating a gas turbine engine is disclosed. The method includes delivering liquid fuel to a combustor of the turbine engine through one or more liquid fuel carrying components of a fuel injector coupled to the combustor, and combusting the liquid fuel in the combustor. The method also includes providing an insulating air shroud around one or more of the liquid fuel carrying components, and generating eddy air currents in the insulating air shroud in response to the combustion. The eddy air currents expel heated air from the insulating air shroud and draw cooler air into the insulating air shroud. The method further includes maintaining a temperature of the one or more liquid fuel carrying components below a threshold temperature as a result of the generation of the eddy air currents. 
     In yet another aspect, a method of assembling a fuel injector to a gas turbine engine is disclosed. The method includes fluidly coupling a second end of an injector housing to a combustor of the turbine engine. The housing extends from a first end to the second end along a longitudinal axis and the housing includes a stem that extends longitudinally from the first end to a third end. The stem includes a liquid tube configured to deliver liquid fuel to the fuel injector. The method also includes coupling an annular shell to the housing at the first end. The shell extends along the longitudinal axis from the first end to the third end and is circumferentially disposed about the stem to form an insulating air shroud inside the shell. The air shroud includes a layer of air between the shell and the stem. The method further includes coupling the annular shell to an outer casing of the turbine engine at the third end to form a compressed air space in an area outside the shell. The shell prevents flow of air between the compressed air space and the air shroud. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is an illustration of an exemplary disclosed gas turbine engine system; 
         FIG. 2  is a cross-sectional view of a fuel injector of the turbine engine of  FIG. 1 ; 
         FIGS. 3A and 3B  illustrate cross-sectional views of the first end and second end respectively of the fuel injector of  FIG. 2 ; and 
         FIG. 4  is an a cross-sectional view of a embodiment of a shell of the fuel injector of  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  illustrates an cut away view of an exemplary gas turbine engine (turbine engine)  100 . Turbine engine  100  may have, among other systems, a compressor system  10 , a combustor system  20 , a turbine system  70 , and an exhaust system  90 . In general, compressor system  10  may compress incoming air to a high pressure, combustor system  20  may mix the compressed air with a fuel and burn the mixture to produce high-pressure, high-velocity gas, and turbine system  70  may extract energy from the high-pressure, high-velocity gas flowing from the combustor system  20 . 
     Compressor system  10  may include any device capable of compressing air. In some embodiments this may include an axial flow compressor that produces a continuous flow of compressed air. The axial flow compressor may include rotating and stationary components that cooperate to compress air to the required pressure. A central shaft  12 , disposed concentrically about a longitudinal axis  88 , may drive a central drum  14  of compressor system  10 . The central drum  14  may have a number of annular aerofoils  16  attached thereon in rows along longitudinal axis  88 . These aerofoils  16  may rotate between similar rows of stationary aerofoils  16  attached to a stationary tubular casing of compressor system  10 . Typically, the rotating aerofoils  16  are called “rotors” and the stationary aerofoils  16  are called “stators.” Atmospheric air may enter compressor system  10 , and pass through these aerofoils  16 . As the air flows through aerofoils  16 , the air may get compressed and air pressure may increase. Along with increased pressure, the compressed air exiting aerofoils  16  may have a high temperature. The high pressure and high temperature air may exit compressor system  10  through an outlet port  18 . A pair of rotating and stationary aerofoils is called a stage. In general, the pressure and temperature of air exiting outlet port  18  may depend, among others, on the number of stages of compressor system  10 . In some embodiments, the pressure and temperature of air exiting compressor system  10  may exceed 200 psi and 800° F. (426.7° C.) respectively. 
     Combustor system  20  may be connected to outlet port  18  of compressor system  10 . Combustor system  20  may include an annular combustor  50  disposed about longitudinal axis  88 . In some embodiments, combustor system  20  may include multiple substantially cylindrical combustors (called can-type combustors) arranged in a circular array pattern about longitudinal axis  88 . In some embodiments, combustor system  20  may include combustors that are a hybrid of annular and can-type combustors (combination type combustor). Although an annular combustor  50  is depicted in  FIG. 1 , the disclosed fuel injector with an insulating shroud may be applicable with any type of combustor. Outlet port  18  of compressor system  10  may deliver compressed air into an enclosure  22  formed by an outer casing  24  around central shaft  12 . Compressed air from enclosure  22  may be directed into one or more fuel injectors  30  coupled to combustor  50  and annularly positioned about longitudinal axis  88 . 
       FIG. 2  illustrates a cross-sectional view of a fuel injector  30  coupled to combustor  50 . Fuel injector  30  may be positioned in enclosure  22  with a first end  45  coupled to combustor  50  and a second end  35  coupled to outer casing  24 . High pressure and high temperature compressed air from compressor system  10  may surround fuel injector  30  in enclosure  22 . In some cases, the temperature of compressed air in enclosure  22  may exceed 800° F. (426.7° C.). This high temperature compressed air may heat external surfaces of fuel injector  30 . 
     The compressed air in enclosure  22  may be directed into fuel injector  30  through an air swirler  42 . Air swirler  42  may include a plurality of straight or curved blades attached to a housing  30   a  of fuel injector  30  to swirl the incoming compressed air. The number of blades in air swirler  42  may vary with application. Although air swirler  42  of  FIG. 2  is illustrated as a radial swirler, air swirler  42  in general, may have a radial or an axial configuration. A radial swirler is an air swirler in which compressed air from compressor system  10  may be directed to the curved blades radially, while an axial swirler is an air swirler in which the compressed air may be directed to the curved blades axially. 
     A plurality of liquid fuel nozzles  58  attached to housing  30   a  may inject liquid fuel into the swirled air stream from air swirler  42 . Although liquid fuel nozzles  58  positioned upstream of air swirler  42  are depicted in  FIG. 2 , in some embodiments, these liquid fuel nozzles  58  may take the form of small tubes attached to air swirler  42 . Fuel injector  30  may also include gas ports (not shown) to deliver the gaseous fuel to combustor  50 . In some embodiments, these gas ports may include a plurality of small holes located on air swirler  42 . When turbine engine  100  operates using gaseous fuel, fuel gas may be injected into the swirled air stream through these gas ports. Swirling the incoming air into fuel injector  30 , using air swirler  42 , may help mix the fuel with the compressed air and deliver a premixed mixture of fuel and air to combustor  50 . This premixed fuel-air mixture may be delivered to combustor  50  through a premix barrel  32  of fuel injector  30  that may be coupled to combustor  50 . 
     Fuel injector  30  may also include a pilot assembly  40  disposed radially inwards of premix barrel  32 . In some embodiments, pilot assembly  40  and premix barrel  32  may be aligned along a second longitudinal axis  98  of fuel injector  30 . Pilot assembly  40  may include components configured to inject a stream of pressurized fuel into combustor  50 . In embodiments of fuel injector  30  configured to deliver both liquid and gaseous fuel to combustor  50 , pilot assembly  40  may be configured to inject a stream of pressurized liquid and gaseous fuel into combustor  50 . Pilot assembly  40  may also include components configured to deliver a stream of compressed air along with the pressurized fuel into combustor  50 . In addition, swirl features (not shown) may also be located within pilot assembly  40  to swirl the compressed air delivered to pilot assembly  40 . 
     Combustor  50  may include an ignition device (not shown), such as a torch igniter, to ignite the fuel delivered to combustor  50 . The premixed fuel-air mixture delivered through premix barrel  32 , and the pressurized stream of fuel and air delivered through pilot assembly  40 , may ignite in combustor  50  to create combustion flames. Once ignited, a continuous stream of fuel delivered through fuel injector  30  may sustain the combustion flame. An average temperature of the combustion flame may, in some cases, exceed 2000° F. (1093.3° C.). The flame may heat surfaces of combustor  50  and first end  45  of fuel injector  30  proximate the flame. This heat may be transferred to relatively cooler regions of the fuel injector  30  by standard modes of heat transfer (such as, conduction, convection, and radiation). A cooling air flow may be maintained through a space between multiple walls (not shown) of combustor  50  to keep the combustor surfaces at a safe operating temperature. 
     Fuel injector  30  may include fuel supply conduits that deliver fuel to fuel injector  30 . These conduits may form a stem  34  extending longitudinally from second end  35  along second longitudinal axis  98 . The stem  34  may include a main gas tube  48 , a pilot gas tube, main liquid fuel tube  54 , and pilot liquid tube  44 . It is contemplated that, in some embodiments, stem  34  may include less than, or more than, these afore mentioned conduits. In some embodiments, stem  34  may extend along second longitudinal axis  98  from second end  35  towards housing  30   a.  Main gas tube  48  may supply gaseous fuel from a gaseous fuel manifold (not shown) to a main gas gallery  52  included in fuel injector housing  30   a.  Main gas gallery  52 , annularly positioned around second longitudinal axis  98 , may deliver gaseous fuel to the swirled air stream in premix barrel  32 . Main gas gallery  52  may also supply gaseous fuel to pilot assembly  40 . In some embodiments, a separate pilot gas tube included in stem  34  may supply gaseous fuel to pilot assembly  40 . 
     Liquid fuel tube  54  may supply liquid fuel from a liquid fuel supply (not shown) to a main liquid gallery  56  included in housing  30   a.  Main liquid gallery  56  may include an annular channel around second longitudinal axis  98 . Main liquid gallery  56  may be fluidly coupled to liquid fuel nozzle  58  and may deliver liquid fuel to the swirled air stream in premix barrel  32  to create the premixed fuel-air mixture. 
     Pilot liquid tube  44  may direct liquid fuel from outside fuel injector  30  to pilot assembly  40 . Pilot liquid tube  44  may be an elongate assembly extending from second end  35  to first end  45  along second longitudinal axis  98 . The liquid fuel delivered to pilot assembly  40  through pilot liquid tube  44  may be sprayed into combustor  50  through a nozzle coupled to first end  45  of pilot liquid tube  44 . Compressed air may also be injected into combustor  50  alongside the fuel spray through openings around pilot liquid tube  44 . This liquid fuel and compressed air spray may form the pressurized stream of fuel and air delivered to combustor  50  through pilot assembly  40 . 
     Heat transferred from the combustion flame (in combustor  50 ) and the compressed air (in enclosure  22 ) to the relatively cooler regions of fuel injector  30  may heat the liquid fuel carrying components of fuel injector  30 . The term “liquid fuel carrying components” are generally used to include any component of fuel injector  30  that is configured to deliver liquid fuel to combustor  50 . In some embodiments, these liquid fuel carrying components may include liquid fuel tube  54 , main liquid gallery  56 , liquid fuel nozzle  58 , and pilot liquid tube  44 . It is contemplated that, in some embodiments, liquid fuel carrying components may include additional liquid fuel carrying components, or less than all the afore mentioned liquid fuel carrying components. It may be desirable to keep the temperature of some (or all) of these liquid fuel carrying components below a threshold temperature during operation of the turbine engine  100 . In general, this threshold temperature may be any value of temperature. In some embodiments, the threshold temperature may be about 400° F. (204.4° C.). Maintaining a temperature of the liquid fuel carrying components below about 400° F. (204.4° C.) may prevent coking of the liquid fuel in the liquid fuel carrying components. 
     A shell  72  may be coupled to fuel injector  30  to form an insulating air shroud  74  around the liquid fuel carrying components to keep their temperature below about 400° F. (204.4° C.). Shell  72  may extend longitudinally from second end  35  of fuel injector  30  to a third end  65 , proximate air swirler  42 . Shell  72  may be coupled to housing  30   a  at third end  65  and to a circular disk  62  at second end  35 . In some embodiments, shell  72  may be brazed to housing  30   a  at third end  35 . However, other methods of coupling shell  72  to housing  30   a  are also contemplated.  FIGS. 3A and 3B  illustrate sections of fuel injector  30  at third end  65  and second end  35 , respectively. In the description that follows, reference will be made to both  FIGS. 3A and 3B . Circular disk  62  may be coupled to stem  34  and may include passageways to pass stem  34  there-through. Air gaps  76  (shown in  FIG. 3B ) may be formed between stem  34  and circular disk  62 . These air gaps  76  may vent insulating air shroud  74  to atmosphere outside outer casing  24 . 
     Insulating air shroud  74  may include a space formed between shell  72  and stem  34  of fuel injector  30 . Insulating air shroud  74  may include a layer of air that shields the liquid fuel carrying components from the temperature of the combustor  50  and the temperature of the compressed air in enclosure  22 . The air in insulating air shroud  74  may get heated by the heat transferred from combustor  50  and enclosure  22 . The heated air proximate third end  65  may interact with cooler air towards second end  35 . The interaction of heated air with the cooler air may create natural eddy currents within the space. These eddy currents may allow the heated air in the space to escape through air gap  76 . These eddy currents may also draw in cooler air atmospheric air (from the atmosphere outside outer casing  24 ) into insulating air shroud  74  through air gap  76 . The eddy currents may keep air in insulating air shroud  74  relatively cool, and maintain the temperature of the liquid fuel carrying components below about 400° F. (204.4° C.). 
       FIG. 4  illustrates a cross-sectional view of an exemplary shell  72  used in an application. Shell  72  may be made of any material that will survive the temperatures and stresses induced during operation of turbine engine  100 . In some embodiments, shell  72  may be made of a stainless steel alloy, such as, for example  316 L stainless steel alloy. Shell  72  may enclose substantially all the liquid fuel carrying components within insulating air shroud  74 . Although the size and shape of shell  72  may depend upon the application, in some embodiments, shell  72  may have a length  82  between about 9 to 10 inches (22.9 to 25.4 centimeters). Shell  72  may have a generally tubular shape with a first diameter  84  at the second end  35 , and a second diameter  86  at third end  65 , respectively. At a location between the second end  35  and third end  65 , shell  72  may have a third diameter  92  less than first diameter  84  and second diameter  86 . Although in general, these diameters may depend upon the application, in some embodiments shell  72  may have first diameter  84  between about 3.5 to 4.5 inches (8.9 to 11.4 centimeters), second diameter between about 4 to 5 inches (10.2 to 12.7 centimeters), and third diameter between about 1.5 to about 2.5 inches (3.8 to 6.4 centimeters). The resulting shape of shell  72  may provide an insulating air shroud  74  where eddy currents may be established to keep a temperature of the liquid fuel carrying components below about 400° F. (204.4° C.) while reducing the overall size of shell  72 . 
     Shell  72  may include a flange section  78  at second end  35  of fuel injector  30 . The flange section  78  may extend substantially perpendicularly away from second longitudinal axis  98 . In some embodiments, flange section  78  may include fastener holes  78 A annularly in a circular array about second longitudinal axis  98 . The flange section  78  may be used to couple fuel injector  30  to outer casing  24  of turbine engine  100  (shown in  FIG. 2 ). In some embodiments, fasteners (not shown) passing through fastener holes  78 A in flange section  78  may be used to attach fuel injector  30  to outer casing  24 . Structural loads from fuel injector  30  may be transferred to outer casing  24  primarily through shell  72 . Although, in the exemplary embodiments described herein, insulating air shroud  74  is configured to maintain a temperature of the liquid fuel carrying components below 400° F. (204.4° C.), in general, an insulating air shroud of the current disclosure may be configured to maintain a temperature of any component of a turbine engine fuel injector below any threshold temperature. 
     INDUSTRIAL APPLICABILITY 
     The disclosed gas turbine fuel injector with an insulating air shroud may be applicable to any turbine engine where it is desirable to maintain a temperature of selected regions of the fuel injector below a desired temperature. In an embodiment of a fuel injector that is configured to deliver liquid fuel to the turbine engine, an insulating air shroud may be used to maintain the temperature of all or selected liquid fuel carrying components below about 400° F. (204.4° C.), and thereby prevent coking of the liquid fuel. The operation of a gas turbine engine with a fuel injector having liquid fuel carrying components maintained below about 400° F. (204.4° C.) will now be described. 
     During operation of turbine engine  100 , air may be drawn into turbine engine  100  and compressed in compressor system  10  (see  FIG. 1 ). Compression of the air may increase a temperature of the compressed air to about 800° F. (426.7° C.). The compressed air may be directed to an enclosure  22  of the turbine engine  100 . The hot compressed air in enclosure  22  may heat a fuel injector  30  located in enclosure  22 . The compressed air from enclosure  22  may be directed to a combustor  50  of combustor system  20  through fuel injector  30 . Fuel may be mixed with the compressed air as it flows through fuel injector  30  into combustor  50 . The fuel-air mixture may burn in combustor  50  producing a temperature of about 2250° F. (1232.2° C.). 
     A shell  72  may be coupled with fuel injector  30  to shield the liquid fuel carrying components (liquid fuel tube  54 , main liquid gallery  56 , liquid fuel nozzle  58 , and pilot liquid tube  44  of  FIG. 2 ) of fuel injector  30  from the heat of combustion and the hot compressed air in enclosure  22 . Shell  72  may couple with housing  30   a  of fuel injector  30  to form an insulating air shroud  74  around the liquid fuel carrying components. The air in the insulating air shroud  74 , proximate third end  65 , may get heated by the combustion of the fuel-air mixture in combustor. This heated air in the insulating air shroud may interact with cooler air near the second end  35  and set up eddy currents within the insulating air shroud  74 . These eddy air currents may expel hot air from the insulating air shroud  74  and draw cooler air into the insulating shroud  74  to maintain the temperature of the liquid fuel components below about 400° F. (204.4° C.). 
     Creating an insulating air shroud around liquid fuel carrying components of the fuel injector enables the temperature of these components to be maintained below 400° F. (204.4° C.), and thereby prevent coking. Although temperatures of regions in close proximity to the liquid fuel carrying components may be at a significantly higher temperature, the insulating air shroud keeps the liquid fuel components relatively cool. Since cooling of the liquid fuel carrying components occurs due to a natural phenomenon of air within the insulating air shroud (that is, without the aid of external air moving means), the cost associated with preventing coke formation in liquid fuel components of the turbine engine may be low. Additionally, the shell that creates the insulating air shroud may be designed to meet the space requirements of fuel injectors  30 . 
     It will be apparent to those skilled in the art that various modifications and variations can be made to the disclosed fuel injector with insulating air shroud. Other embodiments will be apparent to those skilled in the art from consideration of the specification and practice of the disclosed fuel injector with insulating air shroud. It is intended that the specification and examples be considered as exemplary only, with a true scope being indicated by the following claims and their equivalents.