Patent Description:
Gas turbine engines, such as those which power modern commercial and military aircrafts, include a compressor for pressurizing a supply of air, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine for extracting energy from the resultant combustion gases. The combustor generally includes radially spaced apart inner and outer liners that define an annular combustion chamber therebetween. Arrays of circumferentially distributed combustion air holes penetrate multiple axial locations along each liner to radially admit the pressurized air into the combustion chamber. A plurality of circumferentially distributed fuel injectors axially project into a forward section of the combustion chamber to supply the fuel for mixing with the pressurized air.

Combustion of hydrocarbon fuel in the presence of pressurized air may produce nitrogen oxide (NOX) emissions that are subject to excessively stringent controls by regulatory authorities, and thus may be sought to be minimized. Lean-staged liquid-fueled aeroengine combustors can provide low NOx and particulate matter emissions, but are also prone to combustion instabilities. There are several mechanism that may cause combustion instabilities in radial-staged lean combustors including heat release concentrated in the front of the combustor, and weak flame holding at certain operating conditions where main stage air dilutes the pilot stage fuel-air ratio.

Prior art includes <CIT> which covers the features specified in the preamble of claim <NUM>, <CIT>, <CIT> and <CIT>.

A method of controlling a pilot/main fuel schedule to a combustor of a gas turbine engine according to one aspect of the present invention is claimed in claim <NUM>.

Features of embodiments are recited in the dependent claims
The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise.

The gas turbine engine <NUM> is disclosed herein as a two-spool turbo fan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. The fan section <NUM> drives air along a bypass flowpath while the compressor section <NUM> drives air along a core flowpath for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. Although depicted as a turbofan in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an intermediate pressure compressor ("IPC") between a Low Pressure Compressor ("LPC") and a High Pressure Compressor ("HPC"), and an intermediate pressure turbine ("IPT") between the high pressure turbine ("HPT") and the Low pressure Turbine ("LPT").

The engine <NUM> generally includes a low spool <NUM> and a high spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine static structure <NUM> via several bearing structures <NUM>. The low spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure compressor ("LPC") <NUM> and a low pressure turbine ("LPT") <NUM>. The inner shaft <NUM> drives the fan <NUM> directly or through a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low spool <NUM>. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

The high spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor ("HPC") <NUM> and high pressure turbine ("HPT") <NUM>. A combustor <NUM> is arranged between the high pressure compressor <NUM> and the high pressure turbine <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.

Core airflow is compressed by the LPC <NUM> then the HPC <NUM>, mixed with the fuel and burned in the combustor <NUM>, then expanded over the HPT <NUM> and the LPT <NUM>. The turbines <NUM>, <NUM> rotationally drive the respective low spool <NUM> and high spool <NUM> in response to the expansion. The main engine shafts <NUM>, <NUM> are supported at a plurality of points by bearing structures <NUM> within the static structure <NUM>. It should be understood that various bearing structures <NUM> at various locations may alternatively or additionally be provided.

In one non-limiting example, the gas turbine engine <NUM> is a high-bypass geared aircraft engine. In a further example, the gas turbine engine <NUM> bypass ratio is greater than about six (<NUM>:<NUM>). The geared architecture <NUM> can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about <NUM>, and in another example is greater than about <NUM>:<NUM>. The geared turbofan enables operation of the low spool <NUM> at higher speeds which can increase the operational efficiency of the LPC <NUM> and LPT <NUM> and render increased pressure in a fewer number of stages.

A pressure ratio associated with the LPT <NUM> is pressure measured prior to the inlet of the LPT <NUM> as related to the pressure at the outlet of the LPT <NUM> prior to an exhaust nozzle of the gas turbine engine <NUM>. In one non-limiting embodiment, the bypass ratio of the gas turbine engine <NUM> is greater than about ten (<NUM>:<NUM>), the fan diameter is significantly larger than that of the LPC <NUM>, and the LPT <NUM> has a pressure ratio that is greater than about five (<NUM>:<NUM>). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

In one embodiment, a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio. The fan section <NUM> of the gas turbine engine <NUM> is designed for a particular flight condition - typically cruise at about <NUM> Mach and about <NUM>,<NUM> feet (<NUM>). This flight condition, with the gas turbine engine <NUM> at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fan section <NUM> without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine <NUM> is less than <NUM>. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of ("Tram" / <NUM>)<NUM>. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine <NUM> is less than about <NUM> fps (<NUM>/s). With reference to <FIG>, the combustor section <NUM> generally includes a combustor <NUM> with an outer combustor liner assembly <NUM>, an inner combustor liner assembly <NUM> and a diffuser case module <NUM>. The outer combustor liner assembly <NUM> and the inner combustor liner assembly <NUM> are spaced apart such that a combustion chamber <NUM> is defined therebetween. The combustion chamber <NUM> is generally annular in shape.

The outer combustor liner assembly <NUM> is spaced radially inward from an outer diffuser case <NUM>-O of the diffuser case module <NUM> to define an outer annular plenum <NUM>. The inner combustor liner assembly <NUM> is spaced radially outward from an inner diffuser case <NUM>-I of the diffuser case module <NUM> to define an inner annular plenum <NUM>. It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.

The combustor liner assemblies <NUM>, <NUM> contain the combustion products for direction toward the turbine section <NUM>. Each combustor liner assembly <NUM>, <NUM> generally includes a respective support shell <NUM>, <NUM> which supports one or more liner panels <NUM>, <NUM> mounted to a hot side of the respective support shell <NUM>, <NUM>. Each of the liner panels <NUM>, <NUM> may be generally rectilinear and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material and are arranged to form a liner array. In one disclosed non-limiting embodiment, the liner array includes a multiple of forward liner panels 72A and a multiple of aft liner panels 72B that are circumferentially staggered to line the hot side of the outer shell <NUM> (also shown in <FIG>). A multiple of forward liner panels 74A and a multiple of aft liner panels 74B are circumferentially staggered to line the hot side of the inner shell <NUM> (also shown in <FIG>).

The combustor <NUM> further includes a forward assembly <NUM> immediately downstream of the compressor section <NUM> to receive compressed airflow therefrom. The forward assembly <NUM> generally includes an annular hood <NUM>, a bulkhead assembly <NUM>, a multiple of forward fuel nozzles <NUM> (one shown) and a multiple of swirlers <NUM> (one shown). The multiple of fuel nozzles <NUM> (one shown) and the multiple of swirlers <NUM> (one shown) define an axial pilot fuel injection system <NUM> that directs the fuel-air mixture into the combustor chamber generally along an axis F.

The bulkhead assembly <NUM> includes a bulkhead support shell <NUM> secured to the combustor liner assemblies <NUM>, <NUM>, and a multiple of circumferentially distributed bulkhead liner panels <NUM> secured to the bulkhead support shell <NUM>. The annular hood <NUM> extends radially between, and is secured to, the forwardmost ends of the combustor liner assemblies <NUM>, <NUM>. The annular hood <NUM> includes a multiple of circumferentially distributed hood ports <NUM> that accommodate the respective forward fuel nozzles <NUM> and direct air into the forward end of the combustion chamber <NUM> through a respective swirler <NUM>. Each forward fuel nozzle <NUM> may be secured to the diffuser case module <NUM> and project through one of the hood ports <NUM> and through the respective swirler <NUM>. Each of the fuel nozzles <NUM> is directed through the respective swirler <NUM> and the bulkhead assembly <NUM> along a respective axis F.

The forward assembly <NUM> introduces core combustion air into the forward section of the combustion chamber <NUM> while the remainder enters the outer annular plenum <NUM> and the inner annular plenum <NUM>. The multiple of fuel nozzles <NUM> and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber <NUM>.

Opposite the forward assembly <NUM>, the outer and inner support shells <NUM>, <NUM> are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in the HPT <NUM> to define a combustor exit <NUM>. The NGVs 54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section <NUM> to facilitate the conversion of pressure energy into kinetic energy. The combustion gases are also accelerated by the NGVs 54A because of their convergent shape and are typically given a "spin" or a "swirl" in the direction of turbine rotor rotation. The turbine rotor blades absorb this energy to drive the turbine rotor at high speed.

With reference to <FIG>, a multiple of cooling impingement holes <NUM> penetrate through the support shells <NUM>, <NUM> to allow air from the respective annular plenums <NUM>, <NUM> to enter cavities 106A, 106B formed in the combustor liner assemblies <NUM>, <NUM> between the respective support shells <NUM>, <NUM> and liner panels <NUM>, <NUM>. The cooling impingement holes <NUM> are generally normal to the surface of the liner panels <NUM>, <NUM>. The air in the cavities 106A, 106B provides cold side impingement cooling of the liner panels <NUM>, <NUM> that is generally defined herein as heat removal via internal convection.

A multiple of cooling film holes <NUM> penetrate through each of the liner panels <NUM>, <NUM>. The geometry of the film holes, e. g, diameter, shape, density, surface angle, incidence angle, etc., as well as the location of the holes with respect to the high temperature main flow also contributes to effusion film cooling. The liner panels <NUM>, <NUM> with a combination of impingement holes <NUM> and film holes <NUM> may sometimes be referred to as an Impingement Film Floatliner assembly. It should be appreciated that other liner panel assemblies inclusive of a single panel.

The cooling film holes <NUM> allow the air to pass from the cavities 106A, 106B defined in part by a cold side <NUM> of the liner panels <NUM>, <NUM> to a hot side <NUM> of the liner panels <NUM>, <NUM> and thereby facilitate the formation of a film of cooling air along the hot side <NUM>. The cooling film holes <NUM> are generally more numerous than the impingement holes <NUM> to promote the development of a film cooling along the hot side <NUM> to sheath the liner panels <NUM>, <NUM>. Film cooling as defined herein is the introduction of a relatively cooler airflow at one or more discrete locations along a surface exposed to a high temperature environment to protect that surface in the immediate region of the airflow injection as well as downstream thereof.

A multiple of dilution holes <NUM> may penetrate through both the respective support shells <NUM>, <NUM> and liner panels <NUM>, <NUM> along a common axis downstream of the forward assembly <NUM> to quench the hot gases by supplying cooling air radially into the combustor. That is, the multiple of dilution holes <NUM> provide a direct path for airflow from the annular plenums <NUM>, <NUM> into the combustion chamber <NUM>.

With reference to <FIG>, a radial main fuel injection system <NUM> communicates with the combustion chamber <NUM> downstream of the axial pilot fuel injection system <NUM> generally transverse to axis F of an Axially Controlled Stoichiometry (ACS) Combustor. The radial main fuel injection system <NUM> introduces a portion of the fuel required for desired combustion performance, e.g., emissions, operability, durability, as well as to lean-out the fuel contribution provided by the axial pilot fuel injection system <NUM>. In one disclosed non-limiting embodiment, the radial main fuel injection system <NUM> is positioned downstream of the axial pilot fuel injection system <NUM> and upstream of the multiple of dilution holes <NUM>.

The radial main fuel injection system <NUM> generally includes a radially outer fuel injection manifold <NUM> (illustrated schematically) and/or a radially inner fuel injection manifold <NUM> (illustrated schematically) with a respective multiple of outer fuel nozzles <NUM> and a multiple of inner fuel nozzles <NUM>. The radially outer fuel injection manifold <NUM> and/or a radially inner fuel injection manifold <NUM> may be mounted to the diffuser case module <NUM> and/or to the shell <NUM>, <NUM>, however, various mount arrangements may alternatively or additionally provided.

Each of the multiple of outer fuel nozzles <NUM> inner fuel nozzles <NUM> are located within a respective mixer <NUM>, <NUM> to mix the supply of fuel with the pressurized air within the diffuser case module <NUM>. As defined herein, a "mixer" as compared to a "swirler" may generate, for example, zero swirl, a counter-rotating swirl, a specific swirl which provides a resultant swirl or a residual net swirl which may be further directed at an angle. It should be appreciated that various combinations thereof may alternatively be utilized.

The radial main fuel injection system <NUM> may include only the radially outer fuel injection manifold <NUM> with the multiple of outer fuel nozzles <NUM>; only the radially inner fuel injection manifold <NUM> with the multiple of inner fuel nozzles <NUM>; or both (shown). It should be appreciated that the radial main fuel injection system <NUM> may include single sets of outer fuel nozzles <NUM> and inner fuel nozzles <NUM> (shown) or multiple axially distributed sets of, for example, relatively smaller fuel nozzles.

The radial main fuel injection system <NUM> may be circumferentially arranged in a multiple of configurations. In one disclosed non-limiting embodiment, the multiple of outer fuel nozzles <NUM> and the multiple of inner fuel nozzles <NUM> are circumferentially arranged so that the nozzles <NUM>, <NUM> are directly opposed (<FIG>). In another disclosed non-limiting embodiment, the multiple of outer fuel nozzles <NUM> and the multiple of inner fuel nozzles <NUM> are circumferentially staggered so that the nozzles <NUM>, <NUM> are not directly opposed (<FIG>). Furthermore, the nozzles <NUM>, <NUM> may be angled perpendicularly (<FIG>), tangentially (<FIG>), or at an angle such as downstream (<FIG>) relative to the cross flow from the fuel nozzles <NUM> of the axial pilot fuel injection system <NUM> that are directed along axis F.

Alternatively still, the multiple of outer fuel nozzles <NUM> may be positioned through the outer liner <NUM> opposite or staggered relative to a non-fueled mixer <NUM>' on the inner liner <NUM> (<FIG>). That is, the non-fueled mixer <NUM>' provides airflow but not fuel.

The lean-staging is accomplished by axially distributing the fuel injection with a front-end pilot injector and a downstream main injector to axially distribute the heat release similar to an RQL designs, but with lean/lean combustion to enable low NOx and PM emissions. This is different than radial staged designs where all the fuel is injected at the front-end of the combustor. Moving the heat release away from the front-end can be a pressure anti-node for longitudinal acoustic modes to decrease coupling with these modes.

With respect to <FIG>, the forward fuel nozzles <NUM> are circumferentially spaced apart between about <NUM>% - <NUM>% of a bulkhead height B. The bulkhead height B as defined herein is the radial distance between the liner panels <NUM>, <NUM> at the forward end of the combustion chamber <NUM> at the bulkhead liner panels <NUM> of bulkhead assembly <NUM>. The multiple of outer fuel nozzles <NUM> and the inner fuel nozzles <NUM> are axially spaced a distance D between <NUM>% - <NUM>% of the bulkhead height B aft of the forward fuel nozzles <NUM>.

The multiple of outer fuel nozzles <NUM> are radially spaced a distance R from the inner fuel nozzles <NUM> at between about <NUM>% - <NUM>% of the bulkhead height B. It should be understood that the distance R may be with respect to the liner panels <NUM>, <NUM> should the radial main fuel injection system <NUM> only utilize outer fuel nozzles <NUM> (<FIG>) or inner fuel nozzles <NUM> (<FIG>).

With respect to <FIG>, the multiple of outer fuel nozzles <NUM> and multiple of inner fuel nozzles <NUM> may be arranged circumferentially in-line with the forward fuel nozzles <NUM>. Alternatively, the multiple of outer fuel nozzles <NUM> and multiple of inner fuel nozzles <NUM> may be arranged circumferentially between the forward fuel nozzles <NUM> at, for example, quarter pitch (<FIG>). The multiple of outer fuel nozzles <NUM> and/or the multiple of inner fuel nozzles <NUM> may be spaced apart a distance C of between <NUM>% - <NUM>% of the bulkhead height B circumferentially, which alternatively, may be defined as about <NUM> - <NUM> fuel jet diameters. It should be appreciated that various circumferential and other relationships may be utilized and that fuel jet diameter and bulkhead sizing are but examples thereof.

Alternatively still, with respect to <FIG>, the multiple of outer fuel nozzles <NUM> may be more numerous than the forward fuel nozzles <NUM>. In this disclosed non-limiting embodiment, twice the number of outer fuel nozzles <NUM> as compared to the forward fuel nozzles <NUM>. The multiple of outer fuel nozzles <NUM> include both in-line and circumferentially distributed forward fuel nozzles <NUM>.

With reference to <FIG>, the axial pilot fuel injection system <NUM>, the radial main fuel injection system <NUM> and the multiple of dilution holes <NUM> define a forward combustion zone <NUM> axially between the bulkhead assembly <NUM> and the forward section of the radial main fuel injection system <NUM>, as well as a downstream combustion zone <NUM> between the forward section of the radial main fuel injection system <NUM> and the combustor exit <NUM>. The downstream combustion zone <NUM> is axially proximate the multiple of dilution holes <NUM>.

In one disclosed non-limiting embodiment, the axial pilot fuel injection system <NUM> provides about <NUM>% - <NUM>% of the combustor airflow, the radial main fuel injection system <NUM> provides about <NUM>% - <NUM>% of combustor airflow while the multiple of dilution holes <NUM> provide about <NUM>% - <NUM>% of the combustor airflow. It should be appreciated that these ranges of combustor airflow may define a total combustor airflow less than <NUM>% with the remainder being cooling air flow. It should be further appreciated that generally as the combustor airflow from the axial pilot fuel injection system <NUM> increases, the radial main fuel injection system <NUM> decreases and vice-versa with the balance being from the multiple dilution holes <NUM>. In one specific example , the axial pilot fuel injection system <NUM> provides about <NUM>% of the combustor airflow, the radial main fuel injection system <NUM> provides about <NUM>% of combustor airflow while the multiple of dilution holes <NUM> provide about <NUM>% of the combustor airflow with the remainder being cooling airflow.

In one disclosed non-limiting embodiment, the forward combustion zone <NUM> defines about <NUM>% to <NUM>% of the total combustor chamber <NUM> volume and the downstream combustion zone <NUM> defines about <NUM>% to <NUM>% of the total combustor chamber <NUM> volume.

In one disclosed non-limiting embodiment, the downstream combustion zone <NUM> forms an axial length L of about <NUM>% - <NUM>% a height H of the combustion chamber <NUM> between the liners <NUM>, <NUM> at the radial main fuel injection system <NUM> location. The height H as defined herein is the radial distance between the liner panels <NUM>, <NUM> within the combustion chamber <NUM> proximate the radial main fuel injection system <NUM> location. It should be appreciated that various combinations of the above-described geometries may be provided.

With reference to <FIG>, a pilot/main fuel schedule controls how the fuel flow may be shifted between the axial pilot fuel injection system <NUM> and the radial main fuel injection system <NUM> to alter the heat release distribution and convective time delays associated with each zone and enable de-tuning away from instabilities. The approach can be used to mitigate both Rayleigh gain type thermoacoustic instabilities and entropy mode type instabilities. The range of fuel shifting possible will be constrained by other combustor requirements for emissions, efficiency, LBO, etc. Movement of the heat release away from the front-end which can be a pressure anti-node for longitudinal modes decreases coupling with these modes.

For any given operating condition with required total fuel flow, the fuel percentage split between the axial pilot fuel injection system <NUM> and the radial main fuel injection system <NUM> is scheduled accordingly, based on, for example, engine power level or other operating condition to, for example, mitigate combustor tones or control other combustor performance metrics. The engine operating parameter may include, for example, at least one of engine power, throttle position, total fuel flow, and an aircraft flight condition and the combustor metric may include at least one of combustor tones, emissions, combustor efficiency, lean blow-out margin, and altitude re-light capability, etc..

A nominal fuel schedule is correlated as a percentage of the axial pilot fuel injection system <NUM>, e.g., pilot fuel flow, with respect to an operating condition, e.g., engine power include at least one of combustor tones, emissions, combustor efficiency, lean blow-out margin, and altitude re-light capability, etc..

In this example, a nominal fuel schedule is correlated as a percentage of the axial pilot fuel injection system <NUM>, e.g., pilot fuel flow, with respect to an operating condition, e.g., engine power level. A low range and a high range envelope maintains the axial pilot fuel injection system <NUM>, e.g., pilot fuel flow, and the radial main fuel injection system <NUM>, e.g., main zone fuel flow, Fuel-Air (F/A) ratio within a desired limit for combustor operability. According to the invention, the fuel split is further optimized within this high-low range to mitigate combustor tones while also meeting other combustor metrics such as emissions, combustor efficiency, lean blow-out margin, altitude re-light capability, etc..

In operation, a fuel-rich combustion environment in the forward combustion zone <NUM> is provided for low power operations. During high power operations, a fuel-lean combustion environment is maintained in both the forward combustion zone <NUM> and the downstream combustion zone <NUM>. During engine idle, the method includes the step of selectively distributing the fuel being supplied between the forward combustion zone <NUM> and the downstream combustion zone <NUM> with <NUM>% to <NUM>% as the axial pilot fuel injection system <NUM> of fuel and with <NUM>% to <NUM>% as the radial main fuel injection system <NUM> flow of fuel. During approach, <NUM>% to <NUM>% as the flow of fuel may be supplied by the axial pilot fuel injection system <NUM> and <NUM>% to <NUM>% by the radial main fuel injection system <NUM>. At higher power operation of the gas turbine engine, <NUM>% to <NUM>% of the flow of fuel is supplied by the axial pilot fuel injection system <NUM> with <NUM>% to <NUM>% from the radial main fuel injection system <NUM>. Higher power operation of the gas turbine engine may include engine operation at cruise, engine operation at climb, and engine operation at take-off.

The gas turbine combustor and the method for operating the gas turbine combustor as disclosed herein provides for lower NOx emissions at low, mid and high power operation at generally equivalent weight and operability relative to a typical conventional gas turbine combustor. At low power, such as idle and approach, the forward combustion zone <NUM> may be robustly fueled to establish a fuel-rich combustion environment and provide for ignition, combustion stability, and low emissions. When power increases, e.g. during cruise, climb and take-off, fuel flow to the downstream combustion zone <NUM> increases and fuel flow to the forward combustion zone <NUM> decreases whereby both combustion zones operate fuel lean in order to control NOx formation. The increased temperature of the combustion gases from the forward combustion zone <NUM> flow across the downstream combustion zone <NUM> to further facilitate stable combustion in the downstream zone and achievement of high combustion efficiency.

The pilot/main fuel schedule permits control of combustion dynamics in an axially-staged lean-lean combustor configuration. The fuel split between the pilot and main is optimized to mitigate dynamics while also meeting all other combustor performance metrics. For any given operating condition with required total fuel flow, the fuel split between pilot & main is optimized to mitigate combustor tones. This alters the axial heat release distribution and associated convective time delays. Application of this method to a liquid-fueled aeroengine axially-staged lean-lean combustor configuration is new.

The use of the terms "a" and "an" and "the" and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier "about" used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. It should be appreciated that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.

Although the different non-limiting embodiments have specific illustrated components, the embodiments of this invention are not limited to those particular combinations. It is possible to use some of the components or features from any of the non-limiting embodiments in combination with features or components from any of the other non-limiting embodiments.

It should be appreciated that like reference numerals identify corresponding or similar elements throughout the several drawings. It should also be appreciated that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.

Claim 1:
A method of controlling a pilot/main fuel schedule to an annular combustor (<NUM>) of a gas turbine engine (<NUM>) of an aircraft comprising:
scheduling a first percentage of a total fuel flow to an axial pilot fuel injection system (<NUM>) operable to supply the fuel into a forward combustion zone (<NUM>) of a combustion chamber (<NUM>) generally along an axis (F), and scheduling a second percentage of the total fuel flow to a radial main fuel injection system (<NUM>) that communicates with a downstream combustion zone of the combustion chamber (<NUM>) downstream of the axial pilot fuel injection system (<NUM>) and generally transverse to the axis (F),
wherein the scheduling comprises:
correlating, based on an engine operating parameter, a nominal fuel schedule comprising a range for the first and second percentages of the total fuel flow; characterized in that the scheduling further comprises
optimizing the nominal fuel schedule within the range, to mitigate combustor tones while also meeting other combustor metrics, to determine the first and second percentages of the total fuel flow.