BACKGROUND OF THE INVENTION
This invention relates generally to fuel injectors used in gas turbine engine combustors and, in particular, to a fuel injector design primarily for aerospace applications, which produces a stable flame at low power, generating low CO and UHC pollutants, and also provides enhanced fuel-air mixing to reduce NOx emissions at high-power.
Atmospheric pollution concerns and effects worldwide have led to enactment of increasingly stricter emission controls and standards requiring most industries to significantly reduce the emissions of pollutants. The strict emission controls have required implementation of numerous design changes in gas turbines used for industrial, power generation and propulsion applications. Thus typical gas turbine engines are now required to operate efficiently over a wide range of conditions while at the same time producing minimal quantities of noxious emissions. The common precursors to gas turbine engines cause atmospheric pollution include Carbon Monoxide (CO), and Unburned Hydrocarbons (UHC) at low engine power conditions, and Nitrous Oxide (NOx) at intermediate and high engine power conditions.
Reductions in gas turbine emissions of NOx have been obtained primarily by the reduction of flame temperatures in the combustor. Some of the techniques employed include lean burn pre-mix combustors in which the fuel to air ratio is reduced as far as possible in the higher operating range; staged combustors, whereby fuel is admitted to the combustion chamber at different axial locations thereby staging the combustion process; lean-direct injection involving methods of injecting fuel and air into a combustion chamber where the mixture is fuel-lean, or below the stoichiometric fuel/air ratio; and other related techniques known to those skilled in the art as “rich burn” and “quick quench”.
Gas turbine propulsion engines employ annular and can-type combustors to burn fuel. The fuel is metered and sprayed into the combustor through a single or a plurality of fuel nozzles along with combustion air having a designated amount of swirl. In the typical gas turbine engine, flame stability, variable cycle operation, and emission control dominate combustor design requirements. The characteristics of a given fuel injector under light-up and low speed conditions are different to those under full power conditions. Consequently a fuel injector is often a compromise between two designs to enable it to operate under both of these conditions.
Typically, fuel is supplied through one supply duct under starting or low power conditions and through another or through both fuel supply ducts under high power conditions. During light up and low speed conditions, only the pilot fuel injectors are used whereas both the pilot and the main fuel injectors are used under higher speed conditions. The pilot combustion stage is usually long in comparison with the main combustion stage. Consequently, the residence time of the fuel in the pilot stage is comparatively long, limiting emissions of hydrocarbons and carbon monoxide. Conversely, the residence time of the fuel in the main stage is comparatively short, limiting emissions of the oxides of nitrogen.
Many techniques have been utilized in the prior art to enable efficient gas turbine engine operation over a wide range of conditions while at the same time producing minimal quantities of noxious emissions. Of particular interest in this regard are the following references and examples:
U.S. Pat. No. 4,701,124 discloses a can-type combustor intended for industrial applications and designed to use oil or natural gas as the fuel. The fuel is injected at a compound angle in the main fuel chamber at the air swirler vane exit plane.
U.S. Pat. No. 5,062,792 also targets industrial burners and uses oil or natural gas as the fuel. In this design, the oil is used to fuel the pilot circuit, while the natural gas is used to fuel the main circuit.
U.S. Pat. No. 5,069,029 involves a staged combustor, wherein fuel is admitted to the combustion chamber at different axial locations, thereby staging the combustion process. Axial-type air swirlers are utilized for both pilot and main fuel circuits.
U.S. Pat. No. 5,816,050 discloses a can-type combustor with pilot fuel system that injects the fuel axially into the swirling air stream, which is kept separate from the main air stream by a containment wall. Separate fuel injectors are utilized for the pilot and main systems, the latter injecting the fuel near the outer diameter of the air swirler.
U.S. Pat. No. 5,862,668 teaches a radial-staged combustion system, whereby the pilot system is operated at low power (inward and radial location) and the main and pilot are operated at high power conditions. The combustor is considered a premix type system due to the longer residence time of fuel and air in the premixing tubes.
U.S. Pat. No. 6,151,899 involves a can-type combustion chamber whereby main fuel is injected axially into the swirling air passage, and pilot fuel is injected in two discreet axial and radial locations.
U.S. Pat. No. 6,158,223 also teaches a can-type combustion chamber providing for a pilot fuel injection point in the same plane as the main fuel/air injection point. The design utilizes a plurality of main air swirlers for each pilot air swirler.
U.S. patent application Ser. No. 09/492,678, filed Jan. 27, 2000, for Combustor with Fuel Preparation Chambers, by the same inventor as this application, and also assigned to the assignee of this application, discloses an annular combustor having fuel preparation chambers mounted in the dome of the combustor with the objective of uniformly mixing fuel and air so as to reduce NOx formed by the ignition of the fuel/air mixture.
With the exception of the last reference, none of the above-cited prior art is specifically intended for lightweight, aerospace type applications, and some suffer from one or more of the following disadvantages:
a) aerospace type fuels cannot be utilized.
b) fuel-air residence times prior to entering the combustion chamber are relatively long, thus increasing the danger of autoignition or flashback situations, which are a safety concern for aerospace applications.
c) complexity in fabrication and operation.
d) inability to retrofit the system on existing engines.
e) the fuel injection system is usable for can type or annular combustors, but not for both.
For the foregoing reasons, as can be seen, there is a need for a simple, retrofittable, gas turbine engine fuel injector system primarily for aerospace applications, which produces a clean stable flame at low power, provides enhanced fuel-air mixing to reduce noxious emissions at high-power, and can be employed on both can-type and annular type combustion systems.
SUMMARY OF THE INVENTION
In one aspect of the present invention, a fuel injector system is provided to produces a stable flame at low power operations, generating low CO and UHC pollutants.
In another aspect of the present invention, a fuel injector system is provided to produce enhanced fuel-air mixing to reduce NOx emissions at intermediate and high power operations.
In a further aspect of the present invention, a gas turbine engine fuel injection system, comprised of a single fuel injection body incorporating both pilot circuit and main circuit fuel injectors, is provided that produces minimal quantities of CO, UHC and NOx emissions.
In yet another aspect of the present invention, a simple, low cost gas turbine engine fuel injection system is provided that is capable of producing minimal quantities of noxious emissions, and which can be retrofitted into existing products, including engines for aerospace applications.
In a still further aspect of the present invention, a gas turbine engine fuel injection system is provided that is capable of producing minimal quantities of noxious emissions, and which can be utilized on annular as well as can-type combustors.
The present invention is comprised of one fuel injection body to supply both pilot and main fuel systems. The fuel injection body is further comprised of a pilot fuel circuit and a main fuel circuit, both of which inject fuel at essentially the same axial and radial location. The recessed pilot fuel injection site is along the combustor centerline into a swirling air passage produced by axial air swirlers. The main fuel is injected radially through a plurality of injection sites, at a compound angle, into the inner diameter of a swirling air passage produced by radial air swirlers. One main air swirler is utilized for each pilot air swirler. The fuel/air residence time prior to entering the combustion chamber is relatively short, minimizing the likelihood of auto ignition.
A method for producing a low noxious emission gas in a combustor is also disclosed. The method includes injecting pilot circuit fuel spray into a swirling passage along the combustor centerline for low power operations. For intermediate and high power operations, main circuit fuel spray, together with pilot circuit fuel spray, is discharged into a radial swirler air passage, and thereafter fluidly communicates with a single combustion chamber. The swirling mixture of pilot circuit and main fuel and air enters the combustion chamber, is expanded, and ignited to form low emission gas.
These and other objects, features and advantages of the present invention, are specifically set forth in, or will become apparent from, the following detailed description of a preferred embodiment of the invention when read in conjunction with the accompanying drawings.