Patent Application
20100263382
1. Field of the Invention
The present invention relates to staged gas turbine engine combustion systems in which the production of undesirable combustion product components is minimized over the engine operating regime and, more particularly, to a method and apparatus for actively controlling fuel flow to a mixer assembly having a pilot mixer with primary and secondary fuel injection ports.
2. Description of Related Art
Aircraft gas turbine engine staged combustion systems have been developed to limit the production of undesirable combustion product components such as oxides of nitrogen (NOx), unburned hydrocarbons (HC), and carbon monoxide (CO) particularly in the vicinity of airports, where they contribute to urban photochemical smog problems. Gas turbine engines also are designed to be fuel efficient and have a low cost of operation.
Modern day emphasis on minimizing the production and discharge of gases that contribute to smog and to other undesirable environmental conditions, particularly those gases that are emitted from internal combustion engines, have led to different gas turbine engine combustor designs that have been developed in an effort to reduce the production and discharge of such undesirable combustion product components. Other factors that influence combustor design are the desires of users of gas turbine engines for efficient, low cost operation, which translates into a need for reduced fuel consumption while at the same time maintaining or even increasing engine output. As a consequence, important design criteria for aircraft gas turbine engine combustion systems include provisions for high combustion temperatures, in order to provide high thermal efficiency under a variety of engine operating conditions, as well as minimizing undesirable combustion conditions that contribute to the emission of particulates, and to the emission of undesirable gases, and to the emission of combustion products that are precursors to the formation of photochemical smog.
Various governmental regulatory bodies have established emission limits for acceptable levels of unburned hydrocarbons (HC), carbon monoxide (CO), and oxides of nitrogen (NOx), which have been identified as the primary contributors to the generation of undesirable atmospheric conditions. Therefore, different combustor designs have been developed to meet those criteria. For example, one way in which the problem of minimizing the emission of undesirable gas turbine engine combustion products has been attacked is through staged combustion. Staged combustors include a first stage burner for low speed and low power conditions to more closely control the character of the combustion products. A combination of first stage and second stage burners is provided for higher power outlet conditions while attempting to maintain the combustion products within the emissions limits. It will be appreciated that balancing the operation of the first and second stage burners to allow efficient thermal operation of the engine, while simultaneously minimizing the production of undesirable combustion products, is difficult to achieve. In that regard, operating at low combustion temperatures to lower the emissions of NOx, can also result in incomplete or partially incomplete combustion, which can lead to the production of excessive amounts of HC and CO, in addition to producing lower power output and lower thermal efficiency. High combustion temperature, on the other hand, although improving thermal efficiency and lowering the amount of HC and CO, often results in a higher output of NOx.
Another way that has been proposed to minimize the production of those undesirable combustion product components is to provide for more effective intermixing of the injected fuel and the combustion air. In that regard, numerous mixer designs have been proposed over the years to improve the mixing of the fuel and air. In this way, burning occurs uniformly over the entire mixture and reduces the level of HC and CO that result from incomplete combustion. Even with improved mixing, however, higher levels of undesirable NOx are formed under high power conditions when the flame temperatures are high.
One mixer design that has been utilized is known as a twin annular premixing swirler (TAPS), which is disclosed in the following U.S. Pat. Nos. 6,354,072; 6,363,726; 6,367,262; 6,381,964; 6,389,815; 6,418,726; 6,453,660; 6,484,489; and, 6,865,889. It will be understood that the TAPS mixer assembly includes a pilot mixer which is supplied with fuel during the entire engine operating cycle and a main mixer which is supplied with fuel only during increased power conditions of the engine operating cycle. While improvements in the main mixer of the assembly during high power conditions (i.e., take-off and climb) are disclosed in patent applications having Ser. Nos. 11/188,596, 11/188,598, and 11/188,470, modification of the pilot mixer is desired to improve operability across other portions of the engine's operating envelope (i.e., idle, approach and cruise) while maintaining combustion efficiency. To this end and in order to provide increased functionality and flexibility, the pilot mixer in a TAPS type mixer assembly has been developed and is disclosed in a patent application entitled “Pilot Mixer For Mixer Assembly Of A Gas Turbine Engine Combustor Having A Primary Fuel Injector And A Plurality Of Secondary Fuel Injection Ports.” This patent application, having Ser. No. 11/365,428, is owned by the assignee of the present application and hereby incorporated by reference.
Thus, there is a need to increase combustor efficiency and reduce combustion acoustic resonance for TAPS combustors at various engine operating modes and conditions. There is need to provide a TAPS mixer assembly for a gas turbine engine where the fuel injectors of the pilot mixer have an increased fuel flow range by improving the fuel flow number while not sacrificing spray stability and atomization quality of the injected fuel at low flow conditions. Sub-idle and low power conditions require a low total pilot fuel spray tip flow number and a second pilot fuel nozzle injector and circuit is needed for pilot operation to higher engine thrust conditions. Thus, it is also desirable to improve sub-idle efficiency and reduce combustion acoustic resonance. All of these concerns must be addressed while maintaining a low susceptibility to coking of the fuel injectors.
A gas turbine engine fuel nozzle assembly includes substantially concentric primary and secondary pilot fuel nozzles (referred to as a dual orifice pilot fuel injector), a main fuel nozzle spaced radially outwardly of the primary and secondary pilot fuel nozzles, and circular primary and annular secondary exits of the primary and secondary pilot fuel nozzles respectively. An exemplary embodiment of the main fuel nozzle includes a circular or annular array of radially outwardly open fuel injection orifices.
An exemplary embodiment of the fuel nozzle assembly includes a dual orifice pilot fuel injector tip including the primary and secondary pilot fuel nozzles having conical primary and secondary exit holes respectively.
The secondary pilot fuel nozzle may be radially located directly adjacent to and surrounding the primary pilot fuel nozzle or alternatively the secondary pilot fuel nozzle may be radially spaced apart from the primary pilot fuel nozzle.
A gas turbine engine fuel injector having a hollow stem may be used to support at least one fuel nozzle assembly with the substantially concentric primary and secondary pilot fuel nozzles with the primary and secondary pilot fuel nozzles having circular primary and annular secondary exits respectively.
An exemplary embodiment of the fuel injector includes a first pilot swirler located radially outwardly of and adjacent to the dual orifice pilot fuel injector tip, a second pilot swirler located radially outwardly of the first swirler, and a splitter radially positioned between the first and second pilot swirlers. A venturi is formed in a downstream portion of the splitter and includes a converging section, a diverging section, and a throat therebetween and located downstream of the primary exit of the primary pilot fuel nozzle.
A shortened version of the splitter and the venturi has a diverging section length of the diverging section in a range of 1% to 25% of a converging section length of the converging section and a blunt splitter end of the splitter wall. Cooling for the blunt splitter end may be provided by cooling holes extending from cooling hole inlets on a radially inner surface of the splitter wall and upstream of the throat to cooling hole outlets on a downstream facing surface on the blunt splitter end of the splitter wall. Alternative circumferentially skewed cooling holes extending from cooling hole inlets on a radially inner surface of the splitter wall and upstream of the throat to cooling hole outlets on an aft or downstream facing surface on the blunt splitter end of the splitter wall may be used.
A gas turbine engine fuel supply circuit incorporating the dual orifice pilot fuel injector further includes a combined pilot primary and main fuel manifold in fuel supply connection with primary and main fuel circuits in fuel supply connection with the primary pilot and main fuel nozzles respectively. A pilot secondary fuel manifold is in fuel supply connection with a secondary fuel circuit in fuel supply connection with the secondary pilot fuel nozzle. A continuously variable pressure-actuated fuel splitter valve is operably disposed between the combined pilot primary and main fuel manifold and the primary and main fuel circuits for varying a split of fuel between the primary pilot and main fuel nozzles in the primary and main fuel circuits respectively. A continuously variable pressure-actuated fuel flow valve may be operably disposed between the pilot secondary fuel manifold and the secondary pilot fuel nozzle for controlling fuel flow from the pilot secondary fuel manifold to the secondary pilot fuel nozzle in the secondary fuel circuit.
The foregoing aspects and other features of the invention are explained in the following description, taken in connection with the accompanying drawings where:
Illustrated in
Combustor 16 receives an annular stream of pressurized compressor discharge air 14 from a high pressure compressor discharge outlet 69 at what is referred to as CDP air (compressor discharge pressure air). A first portion 23 of the compressor discharge air 14 flows into the mixer assembly 40, where fuel is also injected to mix with the air and form a fuel-air mixture that is provided to the combustion zone 18 for combustion. Ignition of the fuel-air mixture 65 is accomplished by a suitable igniter 70, and the resulting combustion gases 60 flow in an axial direction toward and into an annular, first stage turbine nozzle 72. Nozzle 72 is defined by an annular flow channel that includes a plurality of radially-extending, circularly-spaced nozzle vanes 74 that turn the gases so that they flow angularly and impinge upon the first stage turbine blades (not shown) of a first turbine (not shown).
The arrows in
Referring to
A centerbody 103 is radially disposed between and supports the primary and secondary pilot fuel nozzles 58, 59 and the main fuel nozzle 61. The centerbody 103 surrounds the pilot mixer 102 and defines a chamber 105 that is in flow communication with, and downstream from, the pilot mixer 102. The pilot mixer 102 radially supports the dual orifice pilot fuel injector tip 57 at a radially inner diameter ID and the centerbody 103 radially supports the main fuel nozzle 61 at a radially outer diameter OD. The main fuel nozzle 61 is disposed within the main mixer 104 of the mixer assembly 40 and the dual orifice pilot fuel injector tip 57 is disposed within the pilot mixer 102.
The pilot mixer 102 includes a first pilot swirler 112 located radially outwardly of and adjacent to the dual orifice pilot fuel injector tip 57, a second pilot swirler 114 located radially outwardly of the first swirler 112, and a splitter 116 positioned therebetween. The splitter 116 extends downstream of the dual orifice pilot fuel injector tip 57 and a venturi 118 is formed in a downstream portion 115 of the splitter 116. The venturi 118 includes a converging section 117, a diverging section 119, and a throat 121 therebetween. The throat 121 is located downstream of a primary exit 98 of the primary pilot fuel nozzle 58. The splitter 116 has a wall thickness 125 that tapers down aft or downstream of the throat 121 through the converging section 117. The first and second pilot swirlers 112, 114 are generally oriented parallel to a centerline axis 120 of the dual orifice pilot fuel injector tip 57 and the mixing assembly 40. The first and second pilot swirlers 112, 114 include a plurality of swirling vanes 44 (illustrated schematically in
The primary and secondary pilot fuel nozzles 58, 59 have circular primary and annular secondary exits 98, 100 respectively, are operable to inject fuel in a generally downstream direction, and are often referred to as a dual orifice nozzle. The main fuel nozzle 61 is operable to inject fuel in a generally radially outwardly direction through the circular array of radially outwardly open fuel injection orifices 63. The primary pilot fuel nozzle 58 includes a primary fuel supply passage 158 which feeds fuel to a primary annular manifold 138 located adjacent a downstream end 142 of the primary pilot fuel nozzle 58. The secondary pilot fuel nozzle 59 includes a secondary fuel supply passage 159 which flows fuel to a secondary annular manifold 139 located adjacent a downstream end 143 of the secondary pilot fuel nozzle 59.
Fuel is fed from the manifold 138 into a primary fuel swirler 136 at the downstream end 142. The exemplary primary fuel swirler 136 illustrated herein is a cylindrical plug having downstream and circumferentially angled fuel injection holes 164 to pre-film a conical primary exit hole 166 of the primary pilot fuel nozzle 58 with fuel which improves atomization of the fuel. The conical primary exit hole 166 culminates at the circular primary exit 98. The primary fuel swirler 136 swirls the fuel and centrifugal force of the swirling fuel forces the fuel against a primary conical surface 168 of the conical primary exit hole 166 thus pre-filming the fuel along the primary conical surface 168.
Fuel flows from the secondary annular manifold 139 through a secondary fuel swirler 137 in the secondary fuel supply passage 159 at the downstream end 143 of the secondary pilot fuel nozzle 59. The exemplary secondary fuel swirler 137, as illustrated herein, is a circular array 180 of fuel swirling vanes 182 operable to pre-film a conical secondary exit hole 167 of the secondary pilot fuel nozzle 59 with fuel which improves atomization of the fuel. The conical secondary exit hole 167 culminates at the annular secondary exit 100. The secondary fuel swirler 137 swirls the fuel and centrifugal force of the swirling fuel forces it against a secondary conical surface 169 of the conical secondary exit hole 167 thus pre-filming the fuel along the secondary conical surface 169.
The dual orifice nozzle provides improved atomization, particularly for starting and relight after a high power fuel cut, relative to radial fuel injection of disclosed in the prior art in the U.S. patent application entitled “Pilot Mixer For Mixer Assembly Of A Gas Turbine Engine Combustor Having A Primary Fuel Injector And A Plurality Of Secondary Fuel Injection Ports.” having Ser. No. 11/365,428 referenced above. Concentric annular fuel films from the concentric primary and secondary pilot fuel nozzles 58, 59 merge together and the combined fuel is atomized by an air stream from the pilot mixer 102 which is at its maximum velocity in a plane in the vicinity of the annular secondary exit 100 and has significantly higher air velocity than at the axial plane used in the prior art.
Illustrated in
The primary pilot fuel nozzle 58 of the alternative dual orifice pilot fuel injector tip 57 includes a primary fuel supply passage 158 which feeds fuel to a primary annular manifold 138 located adjacent a downstream end 142 of the primary pilot fuel nozzle 58. The secondary pilot fuel nozzle 59 is annular and generally concentric to the primary pilot fuel nozzle 58. The primary and secondary pilot fuel nozzles 58, 59 are spaced radially apart. The first pilot swirler 112 is located radially outwardly of, adjacent to, and surrounds the primary pilot fuel nozzle 58. The second pilot swirler 114 is located radially outwardly of the first swirler 112 and the splitter 116 is radially positioned between the first and second pilot swirlers 112, 114. The secondary pilot fuel nozzle 59 is located radially outwardly of, adjacent to, and surrounds the second pilot swirler 114. A third pilot swirler 130 is located radially outwardly of, adjacent to, and surrounds the secondary pilot fuel nozzle 59.
The splitter 116 extends downstream of the dual orifice pilot fuel injector tip 57 to form a venturi 118 in a downstream portion 115 of the splitter 116. The venturi 118 includes a converging section 117, a diverging section 119, and a throat 121 therebetween. The secondary pilot fuel nozzle 59 includes a secondary fuel supply passage 159 which flows fuel to a secondary annular manifold 139 located in the secondary fuel supply passage 159 near the downstream end 143 of the secondary pilot fuel nozzle 59. The secondary fuel supply passage 159 culminates at an annular secondary pilot orifice 80 at a downstream end 82 of an annular third pilot passage 84 within which the third pilot swirler 130 is disposed.
Illustrated in
Illustrated in
Illustrated in
Illustrated in
First and second sets 190, 192 of the fuel nozzle injectors 10 are used for staging as graphically illustrated in
The first set 190 of the four fuel injectors 10 function as nozzles and located in the vicinity and close to first and second igniters 210, 212 in the combustor. The primary pilot fuel nozzles 58 in the first set 190 are operated at a relatively high fuel flow number of about 9.25 in one exemplary embodiment of the combustor for start and low power fuel enrichment. The primary pilot fuel nozzles 58 in the second set 192 are operated at a relatively low fuel flow number of about 3.7 in the exemplary embodiment of the combustor for sub-idle power levels. This provides improved sub-idle efficiency, altitude relight capability, and low power operability.
The present invention has been described in an illustrative manner. It is to be understood that the terminology which has been used is intended to be in the nature of words of description rather than of limitation. While there have been described herein, what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.
Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims:
