Patent Description:
A combustor section in a modern turbine engine includes one or more fuel injectors. Each fuel injector is operable to inject fuel for combustion within a combustion chamber. Various types and configurations of fuel injectors are known in the art. While these known fuel injectors have various benefits, there is still room in the art for improvement. There is a need in the art, for example, for a fuel injector with reduced manufacturing costs, that facilitates reduced assembly time as well as provides precision fuel injection.

<CIT> discloses consumable support structures for additively manufactured combustor components.

According to the present invention, an apparatus for a turbine engine is provided in accordance with claim <NUM>.

The nozzle passage may have a non-annular, non-circular cross-sectional geometry at the nozzle orifice.

The nozzle passage may have a square or diamond shaped cross-sectional geometry at the nozzle orifice.

The solid polygonal cross-sectional geometry may have a diamond shape.

The solid polygonal cross-sectional geometry may have a square shape.

At least a first portion the nozzle passage may taper inward towards the centerline as the nozzle passage extends longitudinally along the centerline towards the nozzle orifice.

A slope of the taper may have a rise to run ratio of less than <NUM>, <NUM> or <NUM>.

At least a first portion of an exterior of the fuel nozzle may have a constant lateral width as the exterior of the fuel nozzle extends longitudinally along the centerline towards the nozzle orifice. The first portion of the exterior of the fuel nozzle may longitudinally overlap the first portion the nozzle passage along the centerline.

A second portion of the nozzle passage may be longitudinally between the first portion of the nozzle passage and the nozzle orifice along the centerline. The second portion of the nozzle passage may have a constant lateral width longitudinally along the centerline.

The first portion of the nozzle passage and the second portion of the nozzle passage may each have the solid polygonal cross-sectional geometry.

The first portion of the nozzle passage may be longitudinally between the nozzle orifice and a second portion of the nozzle passage along the centerline. The second portion of the nozzle passage may have a constant lateral width longitudinally along the centerline.

The first portion of the nozzle passage may have the solid polygonal cross-sectional geometry. The second portion of the nozzle passage may have a second cross-sectional geometry that is different than the solid polygonal cross-sectional geometry.

The solid polygonal cross-sectional geometry may extend along a longitudinal length of the nozzle passage.

The nozzle orifice may have a lateral width less than <NUM> inches (<NUM> centimeters).

The fuel nozzle may have a tubular sidewall forming the nozzle passage. The tubular sidewall may have a minimum lateral width that is less than <NUM> inches (<NUM> centimeters).

The turbine engine apparatus may also include a fuel conduit fluidly coupled with the fuel nozzle. The fuel nozzle may be configured to receive fuel from the fuel conduit within the nozzle passage. The fuel nozzle may also be configured to direct the fuel out of the nozzle passage through the nozzle orifice.

<FIG> illustrates a portion of an apparatus <NUM> for a turbine engine. This turbine engine apparatus <NUM> is configured as, or otherwise includes, a fuel injector assembly <NUM> for a combustor section of the turbine engine. The turbine engine apparatus <NUM> includes a fuel conduit <NUM> and a fuel nozzle <NUM>; e.g., a single and/or central orifice fuel nozzle. The turbine engine apparatus <NUM> of <FIG> may also include an apparatus base <NUM>, which apparatus base <NUM> may provide a structural support for the fuel conduit <NUM> and/or the fuel nozzle <NUM>.

The apparatus base <NUM> may be configured as any part of the turbine engine within the combustor section that is proximate the fuel injector assembly <NUM>. The apparatus base <NUM> of <FIG>, for example, may be configured as a turbine engine case such as, but not limited to, a combustor section case, a diffuser case and/or a combustor wall.

The fuel conduit <NUM> is configured as, or may be part of, a fuel supply for the fuel nozzle <NUM>. The fuel conduit <NUM>, for example, may be or may be part of a fuel supply tube, a fuel inlet manifold and/or a fuel distribution manifold. The fuel conduit <NUM> is arranged at and/or is connected to a first side <NUM> (e.g., an exterior and/or outer side) of the apparatus base <NUM>. The fuel conduit <NUM> is configured with an internal fuel supply passage <NUM> formed by an internal aperture (e.g., a bore, channel, etc.) within the fuel conduit <NUM>. The supply passage <NUM> and the associated aperture extend within and/or through the fuel conduit <NUM> along a (e.g., curved and/or straight) centerline <NUM> of the supply passage <NUM>, which may also be a centerline of the fuel conduit <NUM>.

Referring to <FIG>, the fuel nozzle <NUM> is configured to receive (e.g., liquid) fuel from the fuel conduit <NUM>, and inject that received fuel into a plenum <NUM> (e.g., a fluid passage such as an air passage) at a distal end <NUM> (e.g., tip) of the fuel nozzle <NUM>. The fuel nozzle <NUM> of <FIG> includes a nozzle body <NUM> and a nozzle passage <NUM>; e.g., a fuel passage.

The nozzle body <NUM> is arranged at and/or is connected to a second side <NUM> (e.g., an interior and/or inner side) of the apparatus base <NUM>, where the base second side <NUM> is opposite the base first side <NUM>. The nozzle body <NUM> of <FIG> includes a nozzle tube <NUM> and a nozzle support structure <NUM>; e.g., a web. A base end of the nozzle tube <NUM> is connected to the apparatus base <NUM>. The nozzle tube <NUM> projects longitudinally out from the apparatus base <NUM> along a (e.g., straight and/or curved) longitudinal centerline <NUM> of the nozzle passage <NUM> and/or the nozzle tube <NUM> to the fuel nozzle distal end <NUM>. The nozzle support structure <NUM> is connected to and extends between the apparatus base <NUM> and a (e.g., upstream) side of the nozzle tube <NUM>. The nozzle support structure <NUM> structurally ties the nozzle tube <NUM> to the apparatus base <NUM> and may thereby support the nozzle tube <NUM> within the plenum <NUM>. The nozzle support structure <NUM>, for example, may form a support gusset for the nozzle tube <NUM>.

An internal bore of the nozzle tube <NUM> at least partially (or completely) forms the nozzle passage <NUM>. The nozzle passage <NUM> extends longitudinally along the longitudinal centerline <NUM> within and/or through the apparatus base <NUM> and the nozzle tube <NUM> from the supply passage <NUM> to a downstream nozzle orifice <NUM> at the fuel nozzle distal end <NUM>. This nozzle orifice <NUM> provides an outlet from the nozzle passage <NUM> and, more generally, the fuel nozzle <NUM>.

Referring to <FIG>, the nozzle passage <NUM> includes one or more different flow portions (e.g., <NUM> and <NUM>) arranged longitudinally along the longitudinal centerline <NUM>. The nozzle passage <NUM> of <FIG>, for example, includes a (e.g., upstream) convergent portion <NUM> and a (e.g., downstream) throat portion <NUM>.

The convergent portion <NUM> is upstream of the throat portion <NUM>, for example at (e.g., on, adjacent or proximate) an upstream end <NUM> of the nozzle passage <NUM>. The convergent portion <NUM> of <FIG>, for example, is formed by one or more tapering convergent sidewall surfaces <NUM>; see also <FIG>. These convergent sidewall surfaces <NUM> and, thus, the convergent portion <NUM> extend longitudinally along the longitudinal centerline <NUM> from the supply passage <NUM> to the throat portion <NUM>, thereby defining a longitudinal length <NUM> of the convergent portion <NUM>.

A lateral width <NUM> (e.g., a diagonal axis) of the convergent portion <NUM> (e.g., continuously) decreases as the nozzle passage <NUM> extends longitudinally along the longitudinal centerline <NUM> towards the throat portion <NUM> / the nozzle orifice <NUM>. The convergent portion lateral width <NUM> at the nozzle passage upstream end <NUM> is greater than the convergent portion lateral width <NUM> at the throat portion <NUM>.

A slope of a taper of the convergent portion <NUM> and its tapering convergent sidewall surfaces <NUM> has a rise to run ratio (Y/X; see <FIG>). This convergent portion rise to run ratio may be equal to or less than about (e.g., +/- <NUM>%) or exactly <NUM> (e.g., :S <NUM>), for example, to minimize head loss due to contraction. For example, referring to <FIG>, for every five (<NUM>) units the convergent portion <NUM> and its tapering convergent sidewall surfaces <NUM> extend longitudinally along the longitudinal centerline <NUM> (the run X), the convergent portion <NUM> and its tapering convergent sidewall surfaces <NUM> may extend laterally (e.g., in a direction perpendicular to the longitudinal centerline <NUM>) three (<NUM>) units (the rise Y). Such a convergent portion rise to run ratio may facilitate in the additive manufacturing of the fuel nozzle <NUM>, for example, by minimizing layer-to-layer overhangs and/or minimizing variation in lateral sidewall thickness <NUM> (see <FIG>) of the nozzle tube <NUM>. The present disclosure, however, is not limited to such an exemplary convergent portion rise to run ratio nor to any particular manufacturing techniques. For example, in some embodiments, the rise to run ratio may be equal to or less than <NUM>, <NUM>, <NUM>, etc. In other embodiments, the rise to run ratio may be greater than <NUM>, but less than <NUM> for example.

Referring to <FIG>, the throat portion <NUM> is downstream of the convergent portion <NUM>, for example at (e.g., on, adjacent or proximate) the fuel nozzle distal end <NUM>. A downstream most end of the throat portion <NUM> may also define the nozzle orifice <NUM>. The throat portion <NUM> of <FIG>, for example, is formed by one or more (e.g., non-tapered) throat sidewall surfaces <NUM> (see also <FIG>). These throat sidewall surfaces <NUM> and, thus, the throat portion <NUM> extend longitudinally along the longitudinal centerline <NUM> from the convergent portion <NUM> to (or towards) the nozzle orifice <NUM> in the fuel nozzle distal end <NUM>, thereby defining a longitudinal length <NUM> of the throat portion <NUM>.

The throat portion longitudinal length <NUM> may be different (e.g., less) than the convergent portion longitudinal length <NUM>. The convergent portion longitudinal length <NUM>, for example, may be more than two times (2x), five times (5x) or ten times (10x) the throat portion longitudinal length <NUM>. The present disclosure, however, is not limited to the foregoing dimensional relationship between the lengths.

A lateral width <NUM> (e.g., a diagonal axis <NUM> as shown in <FIG>) of the throat portion <NUM> may be about (e.g., +/- <NUM>%) or exactly constant as the nozzle passage <NUM> extends longitudinally along the longitudinal centerline <NUM> towards the nozzle orifice <NUM>. The throat portion lateral width <NUM> at the convergent portion <NUM> is equal to the throat portion lateral width <NUM> at the nozzle orifice <NUM>. Thus, the throat portion <NUM> is non-tapered.

Referring to <FIG> and <FIG>, one or more portion of the nozzle passage <NUM> may have a solid (e.g., non-annular) non-circular cross-sectional geometry (or other non-circular cross-sectional geometry), for example, when viewed in a plane perpendicular to the longitudinal centerline <NUM>. For example, each nozzle passage portion <NUM>, <NUM> and, thus, an entirety of the nozzle passage <NUM> of <FIG>, <FIG> and <FIG> has the (e.g., same) polygonal cross-sectional geometry. This polygonal cross-sectional geometry may be square shaped and/or diamond shaped as shown in <FIG> and <FIG>. The present disclosure, however, is not limited to such exemplary polygonal shapes. For example, in other embodiments, the polygonal cross-sectional geometry may have a triangular shape or any other polygonal shape.

Compared to a circular cross-sectional geometry for example, the polygonal cross-sectional geometry may aid in minimizing variation in as-formed surface finish (e.g., surface roughness and/or surface distortions) of the nozzle passage surfaces <NUM>, <NUM>, particularly where the fuel nozzle <NUM> is additively manufactured and/or the nozzle passage lateral width (e.g., <NUM>, <NUM>; see <FIG>) is relatively small. Configuring the nozzle passage <NUM> with the polygonal cross-sectional geometry may thereby reduce actual (e.g., additively manufactured, as-formed) dimensional and/or geometric deviation of the nozzle passage <NUM> and its nozzle orifice <NUM> from a (e.g., design) standard as schematically shown, for example, in <FIG>. By contrast, referring to <FIG>, layer-to-layer distortions produced during additive manufacturing may leave a nozzle passage <NUM> designed to have a circular cross-sectional geometry with a relative rough and/or otherwise distorted nozzle passage surface <NUM>. Such distortions may increase actual dimensional and/or geometric deviation of the circular nozzle passage <NUM> from its (e.g., design) standard. This increase in deviation particularly at a nozzle orifice <NUM> may reduce fuel metering precision through the circular nozzle orifice <NUM>. Furthermore, where a turbine engine includes multiple fuel nozzles with the circular nozzle orifice <NUM>, there may be a relatively significant deviation between the fuel injected by the fuel nozzles and, thus, a relatively high imbalance in fuel burn and hot streaks within the combustion chamber as well as downstream in the turbine section. However, by reducing the as-formed deviation as schematically shown in <FIG> by designing / providing the nozzle passage <NUM> and/or the nozzle orifice <NUM> with the polygonal cross-sectional geometry (or another non-circular cross-sectional geometry), fuel metering precision of the fuel nozzle <NUM> can be increased. Deviation between multiple fuel nozzles <NUM> can also be reduced and, thus, fuel burn and/or hot streak imbalance may also be reduced.

Referring to <FIG> and <FIG>, by reducing surface finish variation of the nozzle passage surfaces <NUM>, <NUM>, the fuel nozzle <NUM> may be designed with relatively small dimensions while still being producible via various manufacturing techniques including additive manufacturing. For example, the nozzle orifice <NUM> of <FIG> is configured with a (e.g., minimum or maximum) lateral width (e.g., the lateral width <NUM>) which may be equal to or less than about (e.g., +/- <NUM>%) or exactly <NUM> inches (<NUM> centimeters); e.g., ≤ <NUM> inches (<NUM> centimeters). In addition or alternatively, referring to <FIG>, a tubular sidewall <NUM> of the nozzle tube <NUM> may have a (e.g., minimum, smallest) lateral thickness <NUM> equal to or less than about (e.g., +/- <NUM>%) or exactly <NUM> inches (<NUM> centimeters). Note, at such relatively small dimensions for the nozzle orifice <NUM> and/or the tubular sidewall <NUM>, normally micro-issues in additive manufacturing may become macro-issues and poor melting exhibited by unsupported features (e.g., faces) may cause blockages. However, reducing the surface finish variation as described above may mitigate or prevent formation of such blockages. The present disclosure, of course, is not limited to the foregoing exemplary fuel nozzle dimensions nor to any particular manufacturing technique.

Referring to <FIG>, during turbine engine operation, (e.g., liquid) fuel is directed into the supply passage <NUM> from a fuel source (not shown). At least a portion (or all) of the fuel within the supply passage <NUM> is directed into the nozzle passage <NUM>. This fuel flows through the nozzle passage <NUM> and out of the fuel nozzle <NUM> through the nozzle orifice <NUM> and into the plenum <NUM>. The fuel within the plenum <NUM> may be mixed with air (e.g., compressed air) for subsequent combustion.

In some embodiments, referring to <FIG>, the nozzle passage <NUM> may also be configured with a flow channel portion <NUM>. This flow channel portion <NUM> is upstream of the convergent portion <NUM>, for example at (e.g., on, adjacent or proximate) the nozzle passage upstream end <NUM>. The flow channel portion <NUM> of <FIG>, for example, is formed by at least one (e.g., non-tapering, cylindrical) flow channel sidewall surface <NUM>. This flow channel sidewall surface <NUM> and, thus, the flow channel portion <NUM> extend longitudinally along the longitudinal centerline <NUM> from the supply passage <NUM> to the convergent portion <NUM>, thereby defining a longitudinal length <NUM> of the flow channel portion <NUM>.

The flow channel portion longitudinal length <NUM> may be different (e.g., greater) than the convergent portion longitudinal length <NUM>. The convergent portion longitudinal length <NUM>, for example, may be less than the flow channel portion longitudinal length <NUM> but greater than fifteen percent (<NUM>%) of the flow channel portion longitudinal length <NUM>. More particularly, the convergent portion longitudinal length <NUM> may be between twenty-five percent (<NUM>%) and seventy-five percent (<NUM>%) of the flow channel portion longitudinal length <NUM>. The present disclosure, however, is not limited to the foregoing dimensional relationship between the lengths <NUM> and <NUM>. For example, in other embodiments, the convergent portion longitudinal length <NUM> may be equal to or greater than the flow channel portion longitudinal length <NUM>.

A lateral width <NUM> (e.g., a diameter) of the flow channel portion <NUM> may be about (e.g., +/- <NUM>%) or exactly constant as the nozzle passage <NUM> extends longitudinally along the longitudinal centerline <NUM> towards the throat portion <NUM> / the nozzle orifice <NUM>. The flow channel portion lateral width <NUM> at the nozzle passage upstream end <NUM> is equal to the flow channel portion lateral width <NUM> at the convergent portion <NUM>. Thus, the flow channel portion <NUM> is non-tapered.

In some embodiments, referring to <FIG>, one portion of the nozzle passage <NUM> may have a different cross-sectional geometry than another portion of the nozzle passage <NUM>, for example, when viewed in respective planes perpendicular to the longitudinal centerline <NUM>. The throat portion <NUM> and at least an adjacent section of the convergent portion <NUM> of <FIG>, for example, may each be configured with the (e.g., same) solid polygonal cross-sectional geometry (see <FIG> and <FIG>). By contrast, the flow channel portion <NUM> and at least an adjacent section of the convergent portion <NUM> of <FIG> may be configured with a different cross-sectional geometry (see <FIG>); e.g., a solid (e.g., non-annular) circular cross-sectional geometry or another solid (e.g., non-annular) polygonal, elongated (e.g., oval) or other cross-sectional geometry. Of course, in other embodiments, each of the nozzle passage portions <NUM>, <NUM> and <NUM> may be configured with the (e.g., same) solid polygonal cross-sectional geometry.

In some embodiments, referring to <FIG>, at least a portion (or an entirety) of an exterior <NUM> of the fuel nozzle <NUM> and its nozzle tube <NUM> may have a constant lateral width <NUM> as the exterior <NUM> extends longitudinally along the longitudinal centerline <NUM>, for example, from the apparatus base <NUM> (see <FIG>) to (or towards) the fuel nozzle distal end <NUM> / the nozzle orifice <NUM>. This at least a portion (or the entirety) of the exterior <NUM> may (e.g., partially or completely) longitudinally overlap any one or more of the nozzle passage portions (e.g., <NUM>, <NUM> and/or <NUM>; <NUM> not shown in <FIG>) along the longitudinal centerline <NUM>.

In some embodiments, referring to <FIG>, at least a portion (or the entirety) of the exterior <NUM> of the fuel nozzle <NUM> and its nozzle tube <NUM> may have a variable lateral width <NUM>'. The exterior <NUM> of <FIG>, for example, laterally tapers inward towards the longitudinal centerline <NUM> as the exterior <NUM> extends longitudinally along the longitudinal centerline <NUM>, for example, from the apparatus base <NUM> (see <FIG>) to (or towards) the fuel nozzle distal end <NUM> / the nozzle orifice <NUM>. This at least a portion (or the entirety) of the exterior <NUM> may (e.g., partially or completely) longitudinally overlap any one or more of the nozzle passage portions (e.g., <NUM>, <NUM> and/or <NUM>; <NUM> not shown in <FIG>) along the longitudinal centerline <NUM>.

In some embodiments, referring to <FIG>, the fuel nozzle <NUM> may be one of a plurality of fuel nozzles <NUM> connected to the apparatus base <NUM> and fluidly coupled with the fuel conduit <NUM>. These fuel nozzles <NUM> may be arranged circumferentially about a centerline / rotational axis <NUM> of the turbine engine in an annular array.

The turbine engine apparatus <NUM> also includes one or more fuel vaporizers <NUM>. Each fuel nozzle <NUM> is arranged with a respective one of the fuel vaporizers <NUM>. More particularly, each fuel nozzle <NUM> projects into a respective one of the fuel vaporizers <NUM> and is arranged within a fluid passage <NUM> (e.g., an air passage; the plenum <NUM> in <FIG>) of the respective fuel vaporizer <NUM>. With such an arrangement, each fuel nozzle <NUM> directs at least a portion of the fuel injected into the fluid passage <NUM> against a (e.g., tubular) surface <NUM> of the respective fuel vaporizer <NUM>. The fuel vaporizer <NUM> at least partially vaporizes the fuel impinging against its surface <NUM>.

In the specific embodiment of <FIG>, each fuel vaporizer <NUM> is configured as a structure such as a flow tube <NUM> (e.g., a fluid tube, an air tube) for a combustor <NUM> in the combustor section <NUM>. Note, the combustor <NUM> may also include at least one flow tube <NUM> in between, for example, each circumferentially neighboring set of the vaporizers <NUM>. Each of the flow tubes <NUM>, <NUM> is connected to and projects out from a wall <NUM> of the combustor <NUM> and into a (e.g., annular) combustion chamber <NUM> at least partially defined by the combustor wall <NUM>. The fluid passage <NUM> (e.g., air passage) of each flow tube <NUM> is configured to receive fluid and, more particularly, compressed air from a compressor section of the turbine engine (not visible in <FIG>) through another plenum <NUM>. This compressed air is directed through the respective fluid passage <NUM> and into the combustion chamber <NUM>. However, before reaching the combustion chamber <NUM>, the air within the respective fluid passage <NUM> is mixed with fuel injected by a respective one of the fuel nozzles <NUM>. By injecting the fuel within the flow tube <NUM>, the fuel may be more likely to vaporize within the respective fluid passage <NUM> upon impinging against the surface <NUM> (e.g., an inner side wall surface of the flow tube <NUM>).

In some embodiments, still referring to <FIG> (see also <FIG>), at least the apparatus base <NUM>, the fuel conduit <NUM> and each fuel nozzle <NUM> may be configured together in an integral, monolithic body. The turbine engine apparatus <NUM> and its elements <NUM>, <NUM> and <NUM>, for example, may be additively manufactured in a layer-by-layer build process. Referring to <FIG>, the additive manufacturing may be performed to (e.g., completely) form each nozzle passage <NUM> and its associated nozzle orifice <NUM>, for example, without any additional machining (e.g., drilling of the nozzle elements <NUM> and/or <NUM>). The present disclosure, however, is not limited to such an exemplary monolithic construction nor to additive manufacturing. For example, in other embodiments, one or more or all of the apparatus elements <NUM>, <NUM> and/or <NUM> and/or portions thereof may be individually formed (e.g., additively manufactured, cast, machined and/or formed via any other suitable technique) and subsequently connected (e.g., fastener and/or bonded) together.

The term additive manufacturing may describe a process where a component or components are formed by accumulating and/or fusing material together using an additive manufacturing device, typically in a layer-on-layer manner. Layers of powder material, for example, may be disposed and thereafter solidified sequentially onto one another to form the component(s). The term solidify may describe a process whereby material is sintered and/or otherwise melted thereby causing discrete particles or droplets of the sintered and/or melted material to fuse together. Examples of the additive manufacturing process include a laser powder bed fusion (LPBF) process and an electron beam powder bed fusion (EB-PBF) process. Examples of the additive manufacturing device include a laser powder bed fusion (LPBF) device and an electron beam powder bed fusion (EB-PBF) device. Of course, various other additive manufacturing processes and devices are known in the art, and the present disclosure is not limited to any particular ones thereof.

The turbine engine apparatus <NUM> of the present disclosure may be configured with various different types and configurations of turbine engines. <FIG> illustrates one such type and configuration of the turbine engine - a single spool, radial-flow turbojet turbine engine <NUM>. This gas turbine engine <NUM> is configured for propelling an aircraft such as, but not limited to, an unmanned aerial vehicle (UAV), a drone or any other manned or unmanned aircraft or self-propelled projectile. The present disclosure, however, is not limited to such an exemplary turbojet turbine engine configuration nor to an aircraft propulsion system application. For example, the gas turbine engine may alternatively be configured as an auxiliary power unit (APU) or an industrial gas turbine engine.

In the specific embodiment of <FIG>, the turbine engine <NUM> includes an upstream inlet <NUM>, a (e.g., radial) compressor section <NUM>, the combustor section <NUM>, a (e.g., radial) turbine section <NUM> and a downstream exhaust <NUM> fluidly coupled in series. A compressor rotor <NUM> in the compressor section <NUM> is coupled with a turbine rotor <NUM> in the turbine section <NUM> by a shaft <NUM>, which shaft <NUM> rotates about the centerline / rotational axis <NUM> of the turbine engine <NUM>.

The turbine engine apparatus <NUM> may be included in various turbine engines other than the one described above. The turbine engine apparatus <NUM>, for example, may be included in a geared turbine engine where a gear train connects one or more shafts to one or more rotors in a fan section, a compressor section and/or any other engine section. Alternatively, the turbine engine apparatus <NUM> may be included in a turbine engine configured without a gear train. The turbine engine apparatus <NUM> may be included in a geared or non-geared turbine engine configured with a single spool (e.g., see <FIG>), with two spools, or with more than two spools. The turbine engine may be configured as a turbofan engine, a turbojet engine, a propfan engine, a pusher fan engine or any other type of turbine engine. The present disclosure therefore is not limited to any particular types or configurations of turbine engines.

Claim 1:
An apparatus (<NUM>) for a turbine engine, comprising:
a fuel nozzle (<NUM>) comprising a nozzle passage (<NUM>) and a nozzle orifice (<NUM>);
the nozzle passage (<NUM>) extending longitudinally along a centerline (<NUM>) within the fuel nozzle (<NUM>) to the nozzle orifice (<NUM>);
an air tube (<NUM>) including an air passage (<NUM>), the fuel nozzle (<NUM>) projecting into the air passage (<NUM>); and
a combustor wall (<NUM>) at least partially forming a combustion chamber (<NUM>), the air tube (<NUM>) connected to the combustor wall (<NUM>) and projecting into the combustion chamber (<NUM>),
wherein the nozzle passage (<NUM>) has a solid polygonal cross-sectional geometry at the nozzle orifice (<NUM>) and the fuel nozzle (<NUM>) is configured to direct fuel out of the nozzle passage (<NUM>) through the nozzle orifice (<NUM>) into the air passage (<NUM>), wherein
the fuel nozzle (<NUM>) is configured to direct at least a portion of the fuel against a surface (<NUM>) of the air tube (<NUM>), wherein the air tube (<NUM>) is a fuel vaporizer (<NUM>) and wherein the fuel vaporizer (<NUM>) at least partially vaporizes the fuel impinging against its surface (<NUM>).