EP4671612A1

EP4671612A1 - STAGED MIXER WITH MULTI-PREFILING COATINGS FOR A FUEL INJECTION NOZZLE

Info

Publication number: EP4671612A1
Application number: EP25172590.9A
Authority: EP
Inventors: Gregory Boardman; Brandon Williams
Original assignee: RTX Corp
Current assignee: RTX Corp
Priority date: 2024-06-28
Filing date: 2025-04-25
Publication date: 2025-12-31

Abstract

A gas turbine engine combustor includes a combustor liner dome, a fuel nozzle assembly (204) with a distal nozzle, and a fuel-air mixer situated between the distal nozzle and the dome. The fuel nozzle assembly (204) includes a generally axially-oriented primary fuel outlet (326), and stages of additional fuel outlets (328, 330) circumferentially distributed about the distal nozzle. The fuel air mixer includes a first mixer element (302), a second mixer element (310), and a floating connection between the first and second mixer elements (302, 310). The first and second mixers each include radial swirlers (304, 312) and, downstream of those radial swirlers (304, 312), prefilming surfaces (306, 314) situated in output paths of a respective stage additional fuel outlets (328, 330). The floating connection permits radial but not axial deflection of the first mixer element (302) relative to the second mixer element (310).

Description

BACKGROUND

The present invention relates generally to gas turbine engines, and more particularly to injectors for staged combustion.
Brayton cycle engines generate thrust and extract power from combustion of mixtures of fuel and compressed air. Nozzles for injection of fuel into gas turbine engine combustors are designed to atomize and disperse fuel to promote uniform and consistent mixing, avoid coking, and improve combustor dynamics. Some injector nozzles are constructed to facilitate staged use, i.e., with injection of fuel in different patterns and rates during different engine states.

SUMMARY

In one aspect, this disclosure presents a gas turbine engine combustor that includes a combustor liner dome, a fuel nozzle assembly with a distal nozzle, and a fuel-air mixer situated between the distal nozzle and the dome. The fuel nozzle assembly includes a generally axially-oriented primary fuel outlet, and stages of additional fuel outlets circumferentially distributed about the distal nozzle. The fuel air mixer includes a first mixer element, a second mixer element, and a floating connection between the first and second mixer elements. The first and second mixers each include radial swirlers and, downstream of those radial swirlers, prefilming surfaces situated in output paths of a respective stage additional fuel outlets. The floating connection permits radial but not axial deflection of the first mixer element relative to the second mixer element.
In another aspect, this disclosure presents an axial fuel-air mixer for a gas turbine engine. This axial fuel-air mixer includes a backplate defining an aperture for receiving a fuel nozzle, and first and second mixer elements. The first mixer element is axially forward of and rigidly anchored to the backplate. The second mixer element is axially forward of the first mixer element, and is secured to the first mixer element via a floating connection that permits radial but not axial movement of the second mixer element relative to the first mixer element. The first and second mixer elements each include radial swirlers and prefilming surface disposed downstream of their respective radial swirlers.
Features of embodiments are set forth in the dependent claims.
The present summary is provided only by way of example, and not limitation. Other aspects of the present disclosure will be appreciated in view of the entirety of the present disclosure, including the entire text, claims, and accompanying figures.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is schematic cross-sectional view of a simplified gas turbine engine.
FIG. 2 is a schematic cross-sectional view of a gas turbine engine combustor region.
FIG. 3 is a partial cross-sectional view of a staged injector fuel nozzle with multiple shrouds.

While the above-identified figures set forth one or more embodiments of the present disclosure, other embodiments are also contemplated, as noted in the discussion. In all cases, this disclosure presents the invention by way of representation and not limitation. It should be understood that numerous other modifications and embodiments can be devised by those skilled in the art, which fall within the scope and spirit of the principles of the invention. The figures may not be drawn to scale, and applications and embodiments of the present invention may include features and components not specifically shown in the drawings.

DETAILED DESCRIPTION

This disclosure presents a fuel mixer with multiple nested mixer elements disposed coaxial with and partially concentrically about a fuel injector nozzle. In an illustrative example, the fuel injector nozzle includes both a primary axial fuel outlet and multiple stages of axially-aligned, circumferentially distributed, predominantly radially-oriented prefilming outlets. Each of these mixer elements includes a radial swirler, and in the aforementioned example at least two innermost of these mixer elements define frustoconical airblast prefilmer surfaces facing distinct stages of the prefilming outlets. These prefilmer surfaces receive air from upstream through respective radial swirlers, and promote fuel atomization and dispersion as filmed fuel leaves prefilmer surfaces at their downstream edges. The concentrically innermost of the mixer elements rides the fuel injector nozzle, while other mixer elements are directly or indirectly anchored to a combustor dome and interface with the innermost mixer element via a floating seal. This structure provides multiple stages of prefilming for improved fuel atomization and dispersion, forming a floating connection between the fuel injector and the dome while mounting allowing the prefilming surface closest to the fuel injector nozzle to remain at a fixed location relative to the fuel injector nozzle.
FIG. 1 is a schematic cross-sectional view providing a simplified illustration of gas turbine engine 10. Gas turbine engine 10 is a Brayton-cycle gas turbine engine such as an aircraft engine with air inlet 12, compressor section 14, combustor 16, turbine section 18, and exhaust section 20, all oriented generally sequentially along from forward to aft along engine axis A. Gas turbine engine 10 receives environmental air via air inlet 12 and compresses this air through multiple stages of compressor rotors and stators in compressor section 14. Compressed air from compressor section 14 is mixed with fuel in combustor 16, and the resulting fuel-air mixture is ignited. Multiple stages of airfoils in turbine section 18 extract energy (via torque) from resulting high pressure combustion gasses before these gasses are expelled via exhaust section 20.
Gas turbine engine 10 provides one simplified example of a gas turbine engine to contextualize further discussion, below. The illustrated features of gas turbine engine 10 should not be understood as limiting. FIG. 1 illustrates gas turbine engine 10 with a single-spool architecture, for example, whereby a single turbine section drives a single compressor section. In other examples, however, gas turbine engine 10 can be a multi-spool (e.g., 2-spool or 3-spool) system with multiple mechanically separate stages of compressor and/or turbine. Gas turbine engine 10 can, for example, be an aircraft or industrial gas turbine engine.
FIG. 2 is a schematic cross-sectional view of a region of combustor 16 of gas turbine engine 10. More specifically, although combustor 16 can be an annular section extending about circumferentially fully about engine axis A, FIG. 2 illustrates only a representative module of combustor 16 with axial mixer 200, liner 202, axial fuel nozzle assembly 204, outer case 206, outer mixers 208, outer fuel nozzle assembly 210, and dome 212.
FIG. 2 presents a simplified cartoon of a gas turbine engine combustor region. It should be noted that elements and dimensions shown in FIG. 2 are not drawn to scale. The illustrated section of combustor 16 extends from engine axis A to outer case 210, which can for example be a structural case of combustor 16. FIG. 2 illustrates combustor 16 as capable of receiving fuel through outer case 206 for injection both at axial mixer 200 and outer mixers 208, with axial mixer 200 being disposed at a distal end of axial fuel nozzle assembly 204, and outer mixers 208 being disposed at a distal end of outer fuel nozzle assembly 210. Although FIG. 2 illustrates both axial fuel nozzle assembly 204 and outer fuel nozzle assembly 210, some embodiments may eschew outer fuel nozzle assembly 210 (and accompanying outer mixer 208) at some circumferential locations, or altogether. In some embodiments, a circumferential region as shown in FIG. 2 can include multiple outer fuel nozzle assemblies, each with a separate outer mixer 208. Multiple outer fuel nozzle assemblies and corresponding mixers can, in some embodiments, be clustered together.
Fuel nozzle assemblies 204 and 210 carry fuel from a fuel source to combustor 16. In many embodiments fuel nozzle assemblies 204 and 210 can include fuel lines, supports, and seals. In other embodiments, fuel nozzle assemblies 204 and 210 can be deemed to only encompass flow-defining fuel outlets coinciding with mixers 200, 208, with other fluid routing and metering elements being considered separate from fuel nozzle assemblies 204 and 210. In the most general case, fuel injected into combustor 16 from axial fuel nozzle assembly 204 passes through mixer 200 to form combustion flow F-C1. Fuel nozzle assembly 204 can itself include fuel metering mechanisms, and/or can include multiple distinct fuel lines separately metered upstream of fuel nozzle assembly 204, such that fuel flow into combustor 16 can be controlled in both quantity and injection location relative to axial mixer 200. Fuel delivery to and/or through outer fuel nozzle assemblies 210 can be similarly metered, for example to separately control delivery of fuel to individual outer fuel nozzle assemblies 210.
Axial mixer 200 is a fuel/air mixer with multiple radial swirlers and prefilming surfaces arranged to promote fuel atomization and dispersion, as discussed in greater detail below with respect to FIG. 3. Axial mixer 200 is distributed substantially symmetrically about a centerline (C, see FIG. 3) parallel or near parallel to engine axis A. By contrast, outer mixers 208 can be arranged about individual and/or cluster centerlines oriented generally radially towards engine axis A. Although FIG. 2 illustrates only one sectional region of combustor 16, combustor 16 can include multiple axial mixers 200 distributed circumferentially at locations on dome 212. Similarly, combustor 16 can include outer mixers 208 distributed circumferentially, or arranged in groupings distributed circumferentially, about an outer wall of liner 202.
As depicted in FIG. 2, combustor 16 receives compressed air from compressor section 14 as flow F-In. This flow F-In passes fuel nozzle assemblies 204 and is separated by liner 202 into outer flow F-O, inner flow F-I, and flow to axial mixer 200. Although liner 202 is schematically illustrated as a solid structure in FIG. 2, it should be understood that combustor liners such as liner 202 are perforated with holes allowing a degree of exchange between both inner flow F-I and outer flow F-O and the combustor interior defined by liner 202. Airflow from flow F-In through axial mixer 200 is directed to atomize and disperse fuel for consistently uniform fuel-air mixing, generating mixed flow F-M1. Outer mixers 208 are similarly arranged to receive outer flow F-O and mix this airflow with fuel to produce mixed flow F-M2. Mixed flows F-M1 and F-M2 are ignited (igniter not shown), and combustion gasses exit combustor 16 as flow F-Out to turbine section 18.
FIG. 3 is a partial cross-sectional view of axial mixer 200 illustrating subcomponents and geometry thereof, as well as fluid flow paths during operation. FIG. 3 depicts axial mixer 200, dome 212, and a distal-most end of fuel nozzle assembly 204. Axial mixer 200 is made up of first mixer element 302 (with first radial swirler 304, first airblast prefilmer surface 306, and outer radial flange 308), second mixer element 310 (with second radial swirler 312 and second airblast prefilmer surface 314), third mixer element 316 (with third radial swirler 318), backplate 320, retention collar 322, and lock ring 324. The illustrated distal end of fuel nozzle assembly 204 has a primary (pilot) fuel outlet(s) 326, second zone fuel outlets 328, and third zone fuel outlets 330.
First, second, and third mixer elements 302/310/316 are arranged generally coaxially and generally at least partly concentrically about centerline C, which can be parallel to engine axis A (not shown; see FIGs, 2, 3). Radial swirlers 304/312/318 of first, second, and third mixer elements 302/310/316, respectively, receive compressed airflow F-In from compressor 14 as flows F-1, F-2, and F-3, respectively, upstream of dome 212. Radial swirlers 304/312/318 impart centrifugal swirl on these airflows as they pass radially inward through axial mixer 200, as shown. In some embodiments all radial swirlers 304/312/318 can have the same orientation (i.e., clockwise or counterclockwise relative to centerline C). In other examples, orientations of swirlers 304/312/318 can vary, for example alternating such that a circumferential direction of flow imparted on F-2 by radial swirler 312 is opposite circumferential directions of flow imparted on F-1 and/or F-3 by radial swirlers 304/318, respectively. Swirler orientation and angle are selected to promote fuel atomization and dispersion within axial mixer 200 during operation of gas turbine engine 10.
Primary fuel outlet(s) 326 of fuel nozzle assembly 204 is or are oriented generally along or close to centerline C to eject fuel into first zone Z-1, past second and third mixer elements 310 and 316 into an interior of combustor 16. Second zone fuel outlets 328 and third zone fuel outlets 330, by contrast, are oriented towards airblast prefilming surfaces 306 and 314 of first and second mixer elements 302 and 310, defining second zone Z-2 and third zone Z-2, respectively. In the embodiment illustrated in FIG. 2, second zone fuel outlets 328 are distributed circumferentially about fuel nozzle assembly 204 at a first axial location, relative to centerline C, while third zone fuel outlets 330 are distributed circumferentially about fuel nozzle assembly 204 at a second axial location downstream of second zone fuel outlets 328. Second zone fuel outlets 328 are, in the illustrated embodiment, oriented primarily radially and slightly axially forward relative to centerline C, so as to direct fuel to impinge on first airblast prefilming surface 306. Third zone fuel outlets 330 are similarly oriented radially and axially forward to direct fuel past first mixer element 302 to second airblast prefilming surface 314. FIG. 3 depicts an embodiment with four second zone fuel outlets 328 and four third zone fuel outlets 330, with outlets of each zone being evenly circumferentially distributed about the distal end of fuel nozzle assembly 204. More generally, however, and number and size of fuel outlets selected and directed to distribute fuel across airblast prefilming surfaces 306 and 314 can be used, with larger scale systems benefitting from an increased number of circumferentially distributed fuel outlets for more uniform fuel distribution and dispersion. In at least some embodiments, fuel flow can be staged by metering fuel flow to or through primary fuel outlet 326 separately from fuel flow to or through second and third zone fuel outlets 328 and/or 330. Similarly, fuel flow to or through second zone fuel outlets 328 can, in some embodiments, be metered separately from flow to or through third zone fuel outlets 330. Flow metering for staging can, for example, direct different proportions of fuel through different fuel outlets depending on engine condition and/or flight stage (e.g., idle, cruise, climb).
Airblast prefilming surfaces 306 and 314 are inverse frustoconical surfaces tapering (i.e., narrowing) in an axial direction from upstream to downstream. Flows F-1 and F-2 through radial swirlers 304 and 312, respectively, assist in distributing and filming fuel from second and third zone fuel outlets 328 and 338 across airblast prefilming surfaces 306 and 314, respectively. Airblast prefilming surfaces 306 and 314 terminate at sharp downstream edges that promote atomization of filmed fuel, improving fuel dispersion and mixing.
First mixer element 302 is affixed to fuel nozzle assembly 204 via backplate 302, which also forms at least a partial fluid seal defining one dimension of flow F-1 after (and in some embodiments while) passing through radial swirler 304. The location of first mixer element 302 is therefore fixed radially with respect to fuel nozzle assembly 204, although fuel nozzle assembly 204 can translate axially within backplate 302 to accommodate tolerances and minor displacements. Because first mixer element 302 is radially fixed relative to fuel nozzle assembly 204 - i.e., because first mixer element and its airblast prefilming surface 306 float with fuel nozzle assembly 204 - fuel orientation and airflow from flow F-1 close to and partially defined by the outer surface of fuel nozzle assembly 204 are kept consistent, facilitating uniform first fuel filming across prefilming surface 306.
As depicted in FIG. 3, third mixer element 316 is secured to dome 212. In some examples, third mixer element 316 can be brazed directly to or otherwise installed and subsequently attached to dome 212. Second mixer element 310 rides within third mixer element 316, and is therefore also (indirectly) anchored relative to dome 212. First mixer element 302 thus floats radially relative to, but is axially fitted to, first and second mixer elements 310/316, collectively. To facilitating this floating fit, first mixer element 302 has outer radial flange (or grommet) 308 captured radially within retention collar 322. Retention collar 322 permits some radial freedom of movement of first mixer element 302, backplate 304, and fuel nozzle assembly 204 relative to second mixer element 310, third mixer element 316, and dome 212, while constraining the degree of that radial freedom and forming a fluid seal with outer radial flange 308. Lock ring 324 forms an upstream side of a U-shaped groove with retention collar 322 to secure first mixer element 302 in the aforementioned floating fit.
Although elements 302, 310, 316, 320, 322, and 324 are presented as separate elements for clarity of explanation, several of these elements can be directly joined (e.g., via brazing) or attached (e.g., via bolts or other fasteners) to each other, or formed integrally (e.g., as a single additively manufactured structure). In at least some examples, at least some of second mixer element 310, third mixer element 316, and retention collar 322 are formed as a single monolithic element and brazed to dome 212. Similarly, backplate 320 can be formed separately from first mixer element 302 and anchored to first mixer element 302 during assembly, or can be formed integrally with first mixer element 302. Lock ring 324 is attached to retention collar 322 after first mixer element 302 is installed, but can be attached or joined by any suitable means. As noted above, first mixer element 302 floats radially relative to second and third mixer elements 310/316, and is not fixedly attached to second or third mixer elements 310/316 directly or indirectly.
Axial mixer 200 provides a floating seal between fuel nozzle assembly 204 and elements fixedly attached to dome 212 (i.e., second and third mixer elements 310, 316), accommodating some radial and axial movement or displacement of fuel nozzle assembly 204 relative to dome 212. The inclusion of first mixer element 302 anchored to fuel nozzle assembly 204, however, allows for consistently uniform prefilming in a first prefilming stage very close (i.e., radially) to fuel nozzle assembly 204, without impairing the accommodation of relative displacement or movement of fuel nozzle assembly 204 and dome 212. This is accomplished via a floating fit between first and second mixer elements 302, 310 provided via outer radial flange 308, retention collar 322, and lock ring 324.
Although this disclosure presents mixer 200 as consisting qualitatively of three mixer elements each having a corresponding radial swirler, some embodiments consistent with this disclosure may include additional swirler/mixer stages, or as few as two stages (i.e., omitting third mixer element 316 and securing second mixer element 310 directly to dome 212. In general, this disclosure aims to describe and enable mixer structures with at least one upstream swirler and prefilming stage (e.g., via first mixer element 302) secured radially to a fuel nozzle assembly, and at least one downstream swirler and prefilming stage (e.g., via second mixer element 310) anchored to a dome, with a floating radial fit between these stages.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments of the present invention.
A gas turbine engine combustor comprising: a combustor liner having a forward dome; a fuel nozzle assembly disposed to carry fuel to the combustor, the fuel nozzle assembly having a distal nozzle extending along a centerline axis toward the combustor liner, the fuel nozzle assembly comprising: a primary fuel outlet disposed at an end of the distal nozzle; a first stage of secondary fuel outlets distributed circumferentially about the distal nozzle; and a second stage of tertiary fuel outlets distributed circumferentially about the distal nozzle, between the secondary fuel outlets and the primary fuel outlet; and a fuel-air mixer disposed about a distal nozzle, the fuel-air mixer comprising: a first mixer element secured radially with respect to the distal nozzle and comprising: a first radial swirler; and a first prefilming surface disposed downstream of the first radial swirler and in an output path of the secondary fuel outlets; a second mixer element affixed to the forward dome and comprising: a second radial swirler axially forward of the first radial swirler; and a second prefilming surface disposed downstream of the second radial swirler and in an output path of the tertiary fuel outlets; and a floating connection between first mixer element and the second mixer element, the flowing (floating) connection permitting radial but not axial movement of the first mixer element relative to the second mixer element.
The gas turbine engine combustor of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A further embodiment of the foregoing gas turbine engine combustor, wherein the floating connection comprises a radially-extending flange extending from the first or second mixer elements, and a retention element constraining radial movement and preventing axial movement of the first mixer element relative to the second mixer element.
A further embodiment of the foregoing gas turbine engine combustor, further comprising a backplate rigidly attached to the first mixer element and abutting and surrounding the distal nozzle such that the backplate is radially fixed but axially free relative to the distal nozzle.
A further embodiment of the foregoing gas turbine engine combustor, further comprising a third mixer element disposed between the second mixer element and the forward dome, the third mixer element including a third radial swirler.
A further embodiment of the foregoing gas turbine engine combustor, wherein the third mixer element is brazed to the forward dome.
A further embodiment of the foregoing gas turbine engine combustor, wherein the second and third mixer elements are joined together or formed monolithically.
A further embodiment of the foregoing gas turbine engine combustor, wherein the secondary and tertiary fuel outlets are oriented generally axially forward, towards the dome, and at least partially radially outward relative to the axis.
A further embodiment of the foregoing gas turbine engine combustor, wherein the first stage of secondary fuel outlets and the second stage of tertiary fuel outlets each comprise at least four fuel outlets.
A further embodiment of the foregoing gas turbine engine combustor, wherein the secondary fuel outlets are evenly circumferentially distributed about the distal nozzle in the first stage, and the tertiary fuel outlets are evenly circumferentially distributed about the distal nozzle in the second stage, distally of the first stage.
A further embodiment of the foregoing gas turbine engine combustor, wherein fuel flow to the primary fuel outlet is configured to be metered separately from fuel flow to the secondary or tertiary fuel outlets.
A further embodiment of the foregoing gas turbine engine combustor, wherein the fuel nozzle assembly includes separate fuel lines to primary fuel outlet, the secondary fuel outlets, and the tertiary fuel outlets.
A further embodiment of the foregoing gas turbine engine combustor, wherein the first (and second) prefilming surfaces end at sharp atomizing edges, with an edge of the first prefilming surface terminating axially forward of the tertiary fuel nozzles (outlets).
An axial fuel-air mixer for a gas turbine engine combustor, the fuel-air mixer extending about a centerline axis and comprising: a backplate defining an aperture for receiving a fuel nozzle; a first mixer element axially forward of and rigidly anchored to the backplate, the first mixer element comprising: a first radial swirler; and a first prefilming surface disposed downstream of the first radial swirler; a second mixer element axially forward of the first mixer element and secured to the first mixer element via a floating connection permitting radial but not axial movement of the second mixer element relative to the first mixer element, the second mixer element comprising: a second radial swirler axially forward of the first radial swirler; and a second prefilming surface disposed downstream of the second radial swirler and radially outward of the first prefilming surface.
The axial fuel-air mixer of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:
A further embodiment of the foregoing axial fuel-air mixer, wherein the floating connection permits some constrains but does not prevent radial movement of the first mixer element relative to the second mixer element.
A further embodiment of the foregoing axial fuel-air mixer, further comprising a third mixer element axially forward of and rigidly attached to the second mixer element, the third mixer element comprising a third radial swirler axially forward of the second radial swirler.

Summation

Any relative terms or terms of degree used herein, such as "substantially", "essentially", "generally", "approximately" and the like, should be interpreted in accordance with and subject to any applicable definitions or limits expressly stated herein. In all instances, any relative terms or terms of degree used herein should be interpreted to broadly encompass any relevant disclosed embodiments as well as such ranges or variations as would be understood by a person of ordinary skill in the art in view of the entirety of the present disclosure, such as to encompass ordinary manufacturing tolerance variations, incidental alignment variations, alignment or shape variations induced by thermal, rotational or vibrational operational conditions, and the like.
While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims.

Claims

A gas turbine engine combustor (16) comprising:
a combustor liner (202) having a forward dome (212);

a fuel nozzle assembly (204) disposed to carry fuel to the combustor (16), the fuel nozzle assembly (204) having a distal nozzle extending along a centerline axis (C) toward the combustor liner (202), the fuel nozzle assembly (204) comprising:
a primary fuel outlet (326) disposed at an end of the distal nozzle;

a first stage of secondary fuel outlets (328) distributed circumferentially about the distal nozzle; and

a second stage of tertiary fuel outlets (330) distributed circumferentially about the distal nozzle, between the secondary fuel outlets (328) and the primary fuel outlet (326); and

a fuel-air mixer (200) disposed about a distal nozzle, the fuel-air mixer (200) comprising:
a first mixer element (302) secured radially with respect to the distal nozzle
and comprising:
a first radial swirler (304); and

a first prefilming surface (306) disposed downstream of the first radial swirler (304) and in an output path of the secondary fuel outlets (328);

a second mixer element (310) affixed to the forward dome (212) and
comprising:
a second radial swirler (312) axially forward of the first radial swirler (304); and

a second prefilming surface (314) disposed downstream of the second radial swirler (312) and in an output path of the tertiary fuel outlets (330); and

a floating connection between the first mixer element (302) and the second mixer element (310), the floating connection permitting radial but not axial movement of the first mixer element (302) relative to the second mixer element (310).
The gas turbine engine combustor of claim 1, wherein the floating connection comprises a radially-extending flange (308) extending from the first or second mixer elements (302, 310), and a retention element (322, 324) constraining radial movement and preventing axial movement of the first mixer element (302) relative to the second mixer element (310).
The gas turbine engine combustor of claim 1 or 2, further comprising a backplate (320) rigidly attached to the first mixer element (302) and abutting and surrounding the distal nozzle such that the backplate (320) is radially fixed but axially free relative to the distal nozzle.
The gas turbine engine combustor of any preceding claim, further comprising a third mixer element (316) disposed between the second mixer element (310) and the forward dome (212), the third mixer element (316) including a third radial swirler (318).
The gas turbine engine combustor of claim 4, wherein the third mixer element (316) is brazed to the forward dome (212).
The gas turbine engine combustor of claim 4 or 5, wherein the second and third mixer elements (310, 316) are joined together or formed monolithically.
The gas turbine engine combustor of any preceding claim, wherein the secondary and tertiary fuel outlets (328, 330) are oriented generally axially forward, towards the dome (212), and at least partially radially outward relative to the axis (C).
The gas turbine engine combustor of any preceding claim, wherein the first stage of secondary fuel outlets (328) and the second stage of tertiary fuel outlets (330) each comprise at least four fuel outlets (328, 330).
The gas turbine engine combustor of any preceding claim, wherein the secondary fuel outlets (328) are evenly circumferentially distributed about the distal nozzle in the first stage, and the tertiary fuel outlets (330) are evenly circumferentially distributed about the distal nozzle in the second stage, distally of the first stage.
The gas turbine engine combustor of any preceding claim, wherein fuel flow to the primary fuel outlet (326) is configured to be metered separately from fuel flow to the secondary or tertiary fuel outlets (328, 330).
The gas turbine engine combustor of claim 10, wherein the fuel nozzle assembly includes separate fuel lines to primary fuel outlet (326), the secondary fuel outlets (328), and the tertiary fuel outlets (330).
The gas turbine engine combustor of any preceding claim, wherein the first and second prefilming surfaces (306, 314) end at sharp atomizing edges, with an edge of the first prefilming surface (306) terminating axially forward of the tertiary fuel outlets (330).
An axial fuel-air mixer (200) for a gas turbine engine combustor (16), the fuel-air mixer (200) extending about a centerline axis (C) and comprising:
a backplate (320) defining an aperture for receiving a fuel nozzle;

a first mixer element (302) axially forward of and rigidly anchored to the backplate (320), the first mixer element (302) comprising:
a first radial swirler (304); and

a first prefilming surface (306) disposed downstream of the first radial swirler (304);

a second mixer element (310) axially forward of the first mixer element (302) and secured to the first mixer element (302) via a floating connection permitting radial but not axial movement of the second mixer element (310) relative to the first mixer element (302), the second mixer element (310) comprising:
a second radial swirler (312) axially forward of the first radial swirler (304); and

a second prefilming surface (314) disposed downstream of the second radial swirler (304) and radially outward of the first prefilming surface (306).
The axial fuel-air mixer of claim 13, wherein the floating connection permits some constrains but does not prevent radial movement of the first mixer element (302) relative to the second mixer element (310).
The axial fuel-air mixer of claim 13 or 14, further comprising a third mixer element (316) axially forward of and rigidly attached to the second mixer element (310), the third mixer element (316) comprising a third radial swirler (318) axially forward of the second radial swirler (312).