Abstract:
A gas fuel nozzle for mounting in a combustor wall of a gas turbine engine, with an at least partially radially-directed array of gas fuel outlets extending beyond an air flow head having an array of compressed air jet apertures around the gas fuel outlets. The air flow head also has a deflector for creating an axial flow of air for deflecting in an axial direction the radially-injected gas fuel.

Description:
TECHNICAL FIELD 
   The invention relates to a natural gas fuel nozzle for a gas turbine engine and, more particularly, a nozzle adapted to permit conversion from a liquid fuel nozzle configuration to a natural gas configuration on an aero derivative gas turbine engine with minimal change to the design of other components. 
   BACKGROUND OF THE ART 
   The technical field to which the invention relates is a gas turbine engine combustor with the flexibility of using liquid fuel or natural gas fuel nozzles interchangeably. Many combustors for industrial engines in the prior art include dual fuel nozzles that are mounted in combustor walls and can receive both liquid fuel and natural gas fuel at the same time for mixing with compressed air. 
   The invention however is most advantageously applied to an aero derivative industrial gas turbine engine. Such engines are used for stationary industrial applications but incorporate the standard components from aircraft gas turbine engine designs for efficiency and economy in manufacturing and maintenance. An important feature of an aero derivative industrial gas turbine engine is the flexibility of utilizing liquid fuel or natural gas fuel as desired. An important advantage as well is that fuel nozzles operating on natural gas can be fitted into the same combustor interface as a liquid fuel nozzle. However, liquid fuel nozzles, and especially swirl-type nozzles, are not suitable for use with natural gas because natural gas molecules are much smaller than particles of sprayed liquid fuel aerosol. As a result, gas molecules would be trapped in the swirl envelope created by a conventional liquid fuel nozzle. Also, the heat distribution around the nozzle may be different than for a liquid fuel, resulting in hot spots on the combustor. The difficulty is, however, that if a typical natural gas fuel nozzle is used, the combustor would require redesign of the combustor and other components relative to the engine&#39;s aero-engine equivalent. 
   In the prior art, a conventional dual fuel nozzle is often used so that operators can select between natural gas fuel and liquid fuel without changing nozzles. However, such nozzles are relative complex requiring multiple bores, multiple manifolds and complex fuel and air mixing jets. Since nozzles are often replaced and coke build-up requires frequent maintenance, there are advantages to use of simple interchangeable nozzles that can be manufactured at minimal cost. There is a need, however, for a simple gas fuel nozzle which may be used in an aero derivative engine with minimal change to other components of the engine. There is also a need for a gas fuel nozzle which is capable of delivering a fuel/air mixture which is as similar as possible to the liquid fuel nozzle for which the aero version of the gas turbine engine was originally designed. 
   It is an object of the invention to permit simple conversion of existing liquid fuel nozzles to natural gas fuel nozzles with minimal nozzle, combustor or other design changes. 
   Further objects of the invention will be apparent from review of the disclosure, drawings, and description of the invention below. 
   DISCLOSURE OF THE INVENTION 
   The invention, in one aspect, provides a gas fuel nozzle for mounting in a combustor end wall of a gas turbine engine, with a gas fuel delivery member having a gas fuel supply duct with a laterally directed array of gas fuel outlets extending beyond an air flow head about the gas fuel delivery member with a circumferential array of compressed air jet apertures. The fuel outlets are oriented to eject gas fuel with a radial component from the gas fuel delivery member and the air flow head has a deflector opening for creating an axial flow of air for deflecting the gas fuel ejected from the gas fuel delivery member in an axial direction. 
   In another aspect, the invention provides a method of delivering gas fuel to a gas turbine engine combustor, which involves injecting the gas fuel into the combustor with a radial component relative to an axial axis of the combustor and directing pressurized air at the injected gas fuel to deflect the injected gas fuel along the axis of the combustor. 

   
     DESCRIPTION OF THE DRAWINGS 
     In order that the invention may be readily understood, one embodiment of the invention is illustrated by way of example in the accompanying drawings. 
       FIG. 1  is an axial cross-sectional view through a typical industrial gas turbine engine, showing the general arrangement of its component parts. 
       FIG. 2  is an enlarged axial sectional view through the combustor section of  FIG. 1  incorporating fuel nozzles according to the present invention. 
       FIG. 3  is a radial detail view along line  3 — 3  of FIG.  2 . 
       FIG. 4  is a perspective view of a single natural gas fuel nozzle with a central upstanding gas fuel delivery member having rectangular gas outlet ports ejecting gas radially and two planar surfaces on the outer edges of the airflow head creating a deflecting curtain of compressed air to protect the adjacent combustor walls. 
   

   Further details of the invention and its advantages will be apparent from the detailed description included below. 
   DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS 
     FIG. 1  shows an axial cross-section through a typical industrial gas turbine (IGT) engine. It will be understood that the invention may be applicable to almost any type of IGT engine with a combustor and fuel nozzles. Air intake into the engine  1  passes in an inlet  3  into a compressor portion  5 , through a diffuser  6  and then into a plenum  7  that surrounds a combustor  8 . Fuel is supplied to the combustor  8  through fuel nozzles  9 , which also mixes fuel with air from the plenum  7  as it is injected into the combustor  8  as a fuel air mixture that is then ignited. A portion of the compressed air within the plenum  7  is admitted into the combustor  8  through orifices in the side walls to create a cooling air curtain along the combustor walls or is used for cooling to eventually mix with the hot gases from the combustor and pass over the nozzle guide vane  10  and turbines  11  before exiting the engine  1  as exhaust. 
     FIG. 2  shows an enlarged view of the reverse flow combustor  8  (though, of course the invention is not limited to this configuration). Of particular advantage, the invention provides a gas fuel nozzle  12  that is configured to be mounted in the combustor end wall  13  without requiring modification of the other components (i.e. combustor, engine casing etc.) which are used with the aero version of the engine. For example, where an aero derivative industrial gas turbine engine is to be fueled with natural gas, improved efficiency in manufacture and maintenance results. Using the natural gas fuel nozzle  12  of the invention, it is not necessary to manufacture or design a combustor  8  that is specifically adapted for natural gas fuel or for stationary operation in an industrial setting, in light of the differences between the behaviour of liquid and natural gas fuels in swirl nozzles, described above. The combustor end wall  13  including nozzle mounting mechanism such as floating collars and combustor liners need not be modified from conventional aircraft combustor design. As a result, the aero derivative industrial engine requires only the minor modification of changing fuel nozzles from an aircraft engine when the present invention is employed. The combustor  8  with end wall  13  and inner wall  14  and outer wall  15  can remain identical and capitalize on existing manufacturing and maintenance facilities using aero derivative components. 
   In  FIGS. 2 and 3 , the gas fuel nozzle  12  according to the invention is mounted in the end wall  13  of the combustor  8  in a configuration similar to aircraft gas turbine engines having a combustor  8  with a series of nozzles  12  mounted in the end wall  13  in a spaced apart circumferential array. The nozzle  12  is fed with natural gas fuel in this embodiment via the fuel tube  9  (though any suitable method of providing fuel may be employed) and then the fuel is ejected from a central gas fuel delivery member  16  radially as indicated with arrows in  FIGS. 2 and 3 . Air from the plenum  7  passes through inner and outer walls  14  and  15  of the combustor as well as through the nozzle  12  to axially deflect and mix with the gas fuel as indicated with arrows in  FIG. 2  for example. 
     FIG. 4  shows details of the nozzle  12 . The gas fuel delivery member  16  has a plurality of rectangular gas fuel outlets  17 . In the embodiment shown the gas fuel delivery member  16  has a sidewall with radially open ports  17  and the downstream end of the gas fuel delivery member  16  is capped. However in order to direct gas fuel having a radial component as illustrated, those skilled in the art will appreciate that there are other means by which this function can be accomplished. For example, with a plurality of circular holes in the side wall and auxiliary holes in the end cap of the gas fuel delivery members  16 , or with a conical deflector. In all cases however, the fuel outlets  17  are oriented to eject gas fuel with a radial component from the gas fuel delivery member  16 . To overcome the difficulty of gas molecules becoming trapped in the swirl envelope due to their small size (described above), the natural gas fuel is ejected radially from the gas fuel delivery member  16  with a sufficiently high velocity to create the required circulation within the combustor. 
   As seen in  FIGS. 2 and 3 , the combustor walls  14  and  15  are relatively close to the nozzle  12 . Therefore, by ejecting gas fuel radially as described, the combustor walls  14  and  15  and any liners or other structures on the walls  14  and  15  would be excessively subjected to hot gases moving radially, detrimentally causing hot spots. Therefore, in order to counteract these radial flows towards the combustor wall  14  and  15 , a deflecting curtain of air is provided through deflector apertures  28  created by the planar surface  19  disposed on the radially outer edge of the air flow head  20 . The nozzle  12  is oriented so that a deflector aperture  18  is positioned adjacent the outer combustor wall  15  and a second deflector aperture  18  is positioned adjacent the inner combustor wall  14 . The air that progresses from the plenum  7  into the combustor  8  through the apertures  18  provides a deflecting air curtain for protecting the combustor walls  14  and  15  and helps in recirculation of the fuel and air as well as keeping the flame on. 
   As indicated in  FIGS. 4 and 3 , in order to provide sufficient air fuel mixture and circulation of air within the combustor  8 , the airflow head  20  which surrounds the gas fuel delivery members  16  includes axially directed bores  21  and radially directed bores  22  which open to impinge on an upstream portion of the gas fuel delivery member  16 . The radially directed bores  22  are positioned within a countersunk conical recess, which surrounds the gas fuel delivery member  16  and creates airflow axially along the side wall of the gas fuel delivery member  16 . 
   The outer ring of axially directed bores  21  further serves to create an axial flow deflecting the radial flow of gas through gas fuel outlets  17  (preferably rectangular, though other shapes may be used) to create appropriate air fuel mixture, aid in circulation within the combustor  8  and direct the gas fuel mixture into the central portion of the combustor for ignition. 
   However, to further deflect the radial flow of gas fuel exiting through gas fuel outlets  17 , the airflow head  20  also includes planar surfaces  19 , that create deflector apertures  18  for deflecting the gas fuel ejected from the gas fuel delivery member  16  towards an axial direction with resulting flow of compressed air entering the combustor  8  from the plenum  7 . One skilled in the art will recognize that the deflector apparatus of the present invention can be modified as required to correct the hot spots and other design problems that may occur in a particular combustor configuration. 
   As a result, the nozzle  12  has a relatively simple construction compared to conventional liquid fuel nozzles, or dual fuel nozzles. In addition, the combustor walls  14 ,  15 , and  13  and other components from an aero derivative gas turbine engine may be utilized without modification. Converting an engine model from liquid fuel nozzles to gas fuel nozzles  12  for use as an aero derivative IGT does not therefore represent a substantial additional burden on the manufacturer, and thus permits efficiencies in manufacture and maintenance of such products. The present invention also advantageously capable of delivering a fuel/air mixture which is similar to that delivered by the liquid fuel nozzle for which the aero version of the gas turbine engine was originally designed. 
   Although the above description relates to a specific preferred embodiment as presently contemplated by the inventor, it will be understood that the invention in its broad aspect includes mechanical and functional equivalents of the elements described herein.