A monolithic combustion liner for use in a gas turbine engine includes fuel channels integrated into the wall of the combustion liner. The integrated fuel channels can have an aerodynamic shape to reduce flow losses of cooling air flowing around the exterior of the combustion liner. The monolithic combustion liner allows more cooling air to flow around the combustion liner, increasing the cooling of the combustor region of the gas-turbine engine.

BACKGROUND

The present invention relates to combustion liners for use in gas turbine engines and, more particularly, to multi-function monolithic combustion liners for use in gas turbine engines.

Gas turbine engines are continuous combustion engines that can be used for various purposes, such as power generation and/or producing thrust in an aircraft. Gas turbine engines include one or more compressor sections, one or more combustor sections, and one or more turbine sections. The compressor section receives and compresses air to increase the pressure of the air before the air reaches the combustor section. The combustor section receives the high-pressure air, mixes the high-pressure air with a fuel, and ignites the fuel and air mixture to produce exhaust gases. The exhaust gases flow from the combustor section to the turbine section where energy is extracted from the exhaust gases for use by the gas turbine engine. Further, the combustor section of the gas turbine engine can include a combustion chamber, a combustion liner, and a fuel manifold. The combustion liner can be positioned within the combustion chamber and the fuel manifold can be utilized to inject fuel into the combustion liner. Three common configurations for the combustor section of the gas turbine engine are an annular combustor, a can combustor, and a can-annular combustor.

In one example in which the combustor section is an annular combustor, the fuel manifold can be a circular tube that is positioned outside the combustion liner and coupled to the combustion liner through a plurality of brackets secured to an outer surface of the combustion liner. As such, the fuel manifold is positioned within the volume between the outer surface of the combustion liner and the inner surface of the combustion chamber. This volume is utilized as a cooling flow path for air to flow and cool the combustion liner. The fuel manifold and the brackets are positioned within the cooling flow path and therefore disrupt cooling air as it traverses the outer surface of the combustion liner. This disruption of cooling air causes less surface cooling and increased temperatures of the combustion liner, potentially damaging the combustion liner during operation of the gas turbine engine. Thus, there is a need to achieve a suitable fuel distribution to the combustor while providing adequate cooling of the combustion liner to extend the parts life and improve performance of the gas turbine engine.

SUMMARY

According to one aspect of the disclosure, a combustion liner for use in a gas turbine engine is disclosed. The combustion liner includes an inner wall, an outer wall, a dome, a fuel channel, and a nozzle. The outer wall is spaced radially outward from the inner wall. The dome extends between and couples the inner wall to the outer wall. The fuel channel is formed as a single piece with the outer wall. A first distal end of the fuel channel receives fuel from a fuel source. The nozzle is positioned at a second distal end of the fuel channel and the nozzle is fluidly coupled to the fuel channel.

According to another aspect of the disclosure, a method of using a combustion liner in a gas turbine engine is disclosed. The method includes injecting fuel from a fuel source into a fuel channel, wherein the fuel channel is formed as a single piece with an outer wall of the combustion liner; flowing the fuel from a first distal end to a second distal end of the fuel channel, wherein the fuel is pre-heated as the fuel flows through the fuel channel; dispensing the pre-heated fuel through a nozzle into a dilution chute; and flowing the pre-heated fuel through the dilution chute into an interior of the combustion liner.

DETAILED DESCRIPTION

FIG.1Ais a schematic diagram of a portion of gas turbine engine100. As shown inFIG.1A, gas turbine engine100includes air input102, compressor section104, combustor section106, turbine section108, and exhaust nozzle110. Compressor section104is positioned on the upstream end of gas turbine engine100, with respect to airflow through gas turbine engine100. Compressor section104is configured to receive low-pressure air from air input102, compress the low-pressure air to increase the pressure of the low-pressure air, and then transfer the compressed high-pressure air to combustor section106. Combustor section106receives the high-pressure air, mixes the high-pressure air with a fuel, and ignites the fuel and air mixture to produce exhaust gases. The exhaust gases flow from combustor section106to turbine section108where energy is extracted from the exhaust gases for use by gas turbine engine100. The exhaust gases continue to flow downstream through gas turbine engine100to exhaust nozzle110, where the exhaust gases are accelerated and expelled from gas turbine engine100.

FIG.1Bis a front view of combustor section106taken along section line A-A ofFIG.1A. As shown inFIG.1B, combustor section106includes combustion liner10and combustion chamber12, which are used in gas turbine engine100. Combustion liner10is a component of a gas turbine engine that fits within and is secured to combustion chamber12of gas turbine engine100. Combustion liner10is configured to receive a fuel and air mixture and provide a location in which the fuel and air mixture is ignited within gas turbine engine100. As such, combustion liner10must be capable of withstanding high temperatures, without damage, during operation of gas turbine engine100. To prevent damage to combustion liner10, cooling air flows in the volume between combustion liner10and combustion chamber12, removing heat from combustion liner10as the cooling air traverses combustion liner10. The cooling air flows in axial direction AX, which is the central axis of combustion liner10, to remove heat from combustion liner10. Reducing cooling losses, by increasing cooling air flow through the volume between combustion liner10and combustion chamber12, extends the useful life of combustion liner10and improves the overall performance of gas turbine engine100.

Combustion liner10includes inner wall14, outer wall16, and dome17, which are the main body portions of combustion liner10. Inner wall14and outer wall16are coupled together through dome17, which extends between and couples inner wall14to outer wall16. Inner wall14is the innermost structure of combustion liner10. Inner wall14has a circular cross section when viewing in axial direction AX. Outer wall16is spaced radially outward from inner wall14, with respect to axial direction AX. Outer wall16has a circular cross section when viewing in axial direction AX, and outer wall16is concentric about inner wall14. The space between inner wall14and outer wall16is interior15of combustion liner10. Interior15is the portion of combustion liner10that receives high-pressure air and fuel and provides a location for the high-pressure air and fuel to mix before combustion, discussed further below.

FIG.2Ais a perspective view of a portion of combustion liner10taken along section line B-B ofFIG.1Bwith combustion chamber12removed.FIG.2Bis a cross-sectional view of fuel channel18of combustion liner10taken along section line C-C ofFIG.2A.FIG.2Cis a close-up cross-sectional view of a portion of fuel channel18.FIGS.2A-2Cwill be discussed together. Combustion liner10includes inner wall14, outer wall16, fuel channel18, nozzle20, splash plate22, dilution chute24, and support root26. Outer wall16includes outer surface28, first end face30, and second end face32(seeFIG.1). Outer surface28is the outermost surface along the outermost circumference of outer wall16. First end face30is the aft most surface of outer wall16and combustion liner10, with respect to flow direction FL of cooling air over combustion liner10. Second end face32is the forward most surface of outer wall16and combustion liner10, with respect to flow direction FL of cooling air over combustion liner10. Outer wall16is configured to receive and support other components/features of combustion liner10, discussed below.

Fuel channel18includes first distal end34, second distal end36, end face38, bore40, and internal geometry42. First distal end34is the aft most end of fuel channel18, with respect to flow direction FL of cooling air over combustion liner10. Second distal end36is the forward most end of fuel channel18, with respect to flow direction FL of cooling air over combustion liner10. End face38is an end surface of fuel channel18positioned adjacent second distal end36of fuel channel18. Bore40is an aperture extending through fuel channel18from first distal end34to second distal end36but bore40does not extend through second distal end36. In the embodiment shown bore40has a circular cross-sectional shape, but in another embodiment bore40can have any cross-sectional shape. Further, in the embodiment shown, bore40has a constant cross-section extending from first distal end34to second distal end36. In another embodiment, bore40can taper from first distal end34to second distal end36. In yet another embodiment, bore40can taper from second distal end36to first distal end34. Bore40is configured to receive fuel and provide a path for the fuel to flow through fuel channel18from first distal end34to second distal end36. In the embodiment shown inFIG.2B, fuel channel18includes internal geometry42extending from first distal end34to second distal end36. Internal geometry42is configured to induce swirling of the fuel flowing through bore40of fuel channel18. In some examples, internal geometry42can be a helical groove cut into bore40of fuel channel18, extending from first distal end34to second distal end36of fuel channel18. The helical groove causes the flowing fuel to swirl while it flows through bore40, which increases the atomization and burn performance of the fuel within combustion liner10. Although shown inFIG.2B, some embodiments of fuel channel18may not include internal geometry42.

Fuel channel18is a feature of combustion liner10that is positioned adjacent outer surface28of outer wall16. More specifically, as shown best inFIG.2A, fuel channel18is positioned partially within outer surface28of outer wall16and partially outside outer surface28of outer wall16. Fuel channel18is positioned partially within and partially outside outer surface28to pre-heat the fuel flowing through fuel channel18while not overheating the fuel flowing through fuel channel18. The heat produced during combustion of the fuel and air mixture within combustion liner10radiates outward and causes the fuel flowing through fuel channel18to increase to a desired temperature. If fuel channel18were positioned fully within outer wall16, the fuel flowing through fuel channel18could get too hot before entering an interior of combustion liner10, which could result in negative effects on the performance of gas turbine engine100(e.g., vaporization or coking in fuel channel18). If fuel channel18were positioned fully outside outer wall16, the fuel flowing through fuel channel18may not increase to the desired temperature before entering an interior of combustion liner10. Further, fuel channel18being positioned fully outside outer wall16could cause an increase in cooling air flow disruption (and a decrease in the cooling of combustion liner10) due to fuel channel18being positioned fully within the cooling air flow volume. As such, positioning fuel channel18partially within and partially outside outer surface28of outer wall16provides pre-heating of the fuel flowing through fuel channel18and decreases disruptions in cooling air flow over outer surface28.

Further, fuel channel18extends in axial direction AX of combustion liner10to reduce cooling air flow disruption. First distal end34of fuel channel18extends axially outward from first end face30of outer wall16and through a wall of combustion chamber12. First distal end34of fuel channel18is configured to be fluidly coupled to fuel source44of gas turbine engine100, such that first distal end34receives fuel from fuel source44. Fuel source44can be any fluid tight vessel that stores a fuel to be used by gas turbine engine100. In the embodiment shown, the portion of fuel channel18extending axially outward from first end face30of outer wall16is tubular in shape and has a circular cross section. In another embodiment, as shown inFIG.3, the portion of fuel channel18extending axially outward from first end face30of outer wall16can have a teardrop shaped cross-sectional area, when viewing in axial direction AX. The teardrop shape of the axially extending portion of fuel channel18can help decrease cooling air turbulence adjacent the bottom side of fuel channel18, as the cooling air flows over and past fuel channel18and combustion liner10. More specifically, the teardrop shaped cross section guides the cooling air flow along the outer surface of fuel channel18and away from fuel channel18, decreasing or preventing turbulence adjacent the bottom side of fuel channel18. In yet other embodiments, the portion of fuel channel18extending axially outward from first end face30of outer wall16can have any cross-sectional shape.

Fuel channel18is formed integral with outer surface28of combustion liner10, such that fuel channel18is formed as a single piece with outer wall16. In some examples, fuel channel18can be manufactured using an additive manufacturing process, such as powder bed fusion additive manufacturing. The additive manufacturing process can produce all features of fuel channel18, including internal geometry42within bore40of fuel channel18. Further, combustion liner10can include a plurality of fuel channels18that are equally spaced about a circumference of outer wall16of combustion liner10. In other examples, fuel channels18may not be equally spaced about a circumference of outer wall16of combustion liner10. In the example shown inFIG.1, combustion liner10includes twelve fuel channels18integrally formed with and spaced about a circumference of outer wall16of combustion liner10. In other examples, combustion liner10can include more or fewer than twelve fuel channels18integrally formed with and spaced about a circumference of outer wall16of combustion liner10. Fuel channel18is configured to receive fuel from fuel source44at first distal end34, flow the fuel through bore40from first distal end34to second distal end36, and then flow the fuel into nozzle20.

As best shown inFIGS.2B-2C, nozzle20is positioned at second distal end36of fuel channel18. More specifically, nozzle20is positioned within and extends through second distal end36of fuel channel18. Nozzle20is an aperture that extends through second distal end36of fuel channel18, such that nozzle20extends from bore40to end face38of fuel channel18. Nozzle20is configured to fluidly couple bore40of fuel channel18to an interior of combustion liner10. More specifically, nozzle20is configured to receive the fuel flowing through bore40, increase the pressure and flow rate of the flowing fuel, and then dispense the fuel into an interior of combustion liner10. In some examples, nozzle20can have a diameter of about 0.016 inches (0.4064 millimeters). In other examples, nozzle20can have a diameter of more or less than 0.016 inches. Further, in the described embodiment, nozzle20has a circular cross-sectional shape. In another embodiment, nozzle20can have any cross-sectional shape. In the embodiment shown, nozzle20includes a curved portion, but in another embodiment, nozzle20may not include the curved portion, such that nozzle20has a straight nozzle configuration. Nozzle20is formed integral with fuel channel18during the manufacturing process used to produce fuel channel18and combustion liner10. In some examples, combustion liner10can include a plurality of nozzles20equally spaced about a circumference of outer wall16of combustion liner10. In other examples, nozzles20may not be equally spaced about a circumference of outer wall16of combustion liner10. In the example shown inFIG.1, combustion liner10includes twelve nozzles20integrally formed with and positioned adjacent each fuel channel18of combustion liner10. In other examples, combustion liner10can include more or fewer than twelve fuel nozzles20integrally formed with and positioned adjacent each fuel channel18of combustion liner10, depending on the number of fuel channels18.

Splash plate22is positioned adjacent and coupled to fuel channel18. More specifically, splash plate22is spaced from end face38of fuel channel18by a distance, creating a gap between end face38of fuel channel18and splash plate22. Further, splash plate22is positioned parallel with end face38of fuel channel18. Splash plate22is also axially spaced from nozzle20by a distance, creating a gap between nozzle20and splash plate22. Splash plate22is a generally flat feature of combustion liner10that is offset from and positioned in close relation to nozzle20of fuel channel18. Splash plate22is configured to interrupt the flow of fuel dispensing from nozzle20before the fuel enters dilution chute24. More specifically, the fuel dispensing from nozzle20contacts splash plate22and the dispensing fuel is atomized before the fuel enters dilution chute24, increasing the burn performance of the fuel within combustion liner10. In some examples, combustion liner10can include a plurality of splash plates22equally spaced about a circumference of outer wall16of combustion liner10. In other examples, splash plates22may not be equally spaced about a circumference of outer wall16of combustion liner10. In the example shown inFIG.1, combustion liner10includes twelve splash plates22integrally formed with and positioned adjacent each fuel channel18of combustion liner10. In other examples, combustion liner10can include more or fewer than twelve splash plates22integrally formed with and positioned adjacent each fuel channel18of combustion liner10, depending on the number of fuel channels18.

As shown best inFIG.2A, dilution chute24is positioned adjacent second distal end36of fuel channel18and adjacent nozzle20, such that dilution chute24is offset from second distal end36and nozzle20by a distance or gap. Further, dilution chute24extends partially within and partially outside outer wall16of combustion liner10. As shown inFIG.2A, the portion of dilution chute24extending outside of outer wall16of combustion liner10is semi-circular in shape. As shown inFIG.1, the portion of dilution chute24extending within outer wall16of combustion liner10extends into interior15of combustion liner10at an acute angle. Dilution chute24is configured to receive atomized fuel after the fuel dispensing from nozzle20contacts splash plate22and then the atomized fuel flows into interior15of combustion liner10. The semi-circular shape of the portion of dilution chute24extending outside of outer wall16prevents the cooling airflow traversing outer surface28of outer wall16from interrupting the flow of fuel dispensing from nozzle20into dilution chute24. In other words, the semi-circular shape of dilution chute24allows the atomized fuel to enter dilution chute24by preventing the cooling airflow traversing outer surface28from blowing the atomized fuel spray away from dilution chute24.

Additionally, dilution chute24extends into interior15of combustion liner10to facilitate the swirling of the fuel and air mixture flowing within interior15of combustion liner10, increasing the burn performance within combustion liner10. Dilution chute24creates turbulence within interior15of combustion liner10, causing the fuel entering interior15to sufficiently mix with air entering interior15of combustion liner10. Dilution chute24is formed integral with outer wall16during the manufacturing process used to produce combustion liner10. In some examples, combustion liner10can include a plurality of dilution chutes24equally spaced about a circumference of outer surface28of outer wall16of combustion liner10. In other examples, dilution chutes24may not be equally spaced about a circumference of outer wall16. In the example shown inFIG.1, combustion liner10includes twelve dilution chutes24integrally formed with and positioned adjacent outer wall16of combustion liner10. In other examples, combustion liner10can include more or fewer than twelve dilution chutes24integrally formed with and positioned adjacent outer wall16of combustion liner10, depending on the number of fuel channels18within combustion liner10.

As shown best inFIG.2C, support root26is a feature of combustion liner10that extends from second distal end36of fuel channel18to a surface of dilution chute24, coupling second distal end36of fuel channel18to dilution chute24. More specifically, support root26extends from end face38of fuel channel18to splash plate22, coupling fuel channel18to splash plate22. Further, support root26extends from splash plate22to dilution chute24, coupling splash plate22to a surface of dilution chute24. Support root26extends generally in axial direction AX and, in some examples, support root26can have a generally circular cross-sectional shape when viewing in axial direction AX. In other examples, support root26can have any cross-sectional shape. Support root26is included as a feature of combustion liner10due to the manufacturing process used to produce combustion liner10. Support root26allows combustion liner10to be additively manufactured in axial direction AX, providing a contact surface to build upon layer by layer to produce complex geometries such as dilution chute24, splash plate22, and nozzle20. Removal of support root26would prevent the specific geometries and positioning of features of combustion liner10from being produced using an additive manufacturing process. As such, support root26is formed integral with fuel channel18, splash plate22, and dilution chute24during the manufacturing process used to produce combustion liner10.

Combustion liner10is an additively manufactured component of gas turbine engine100that is designed to improve performance of the combustion section of gas turbine engine100. Combustion liner10can be additively manufactured as a unitary single-piece construction, such that the combustion liner is constructed in a single, continuous manufacturing process. In one example, combustion liner10can be additively manufactured using a powder bed fusion additive manufacturing process. In other examples, combustion liner10can be manufactured using other additive manufacturing processes such as, but not limited to, direct metal laser sintering, electron beam melting, selective heat sintering, selective laser melting, and selective laser sintering. Further, combustion liner10can be additively manufactured in axial direction AX to produce complex geometries such as dilution chute24, splash plate22, and nozzle20, without the need for additional supports within combustion liner10(as described with reference to support root26). Additively manufacturing combustion liner10in axial direction AX ensures proper tolerancing and concentricity of combustion liner10is achieved. In contrast, traditionally manufactured combustion liners are perforated using laser-drilling and brazed or welded together, causing localized distortions due to the introduction of thermal energy into a low thermal mass system. As such, additively manufacturing combustion liner10in axial direction AX results in a combustion liner without localized distortions due to secondary joining processes and ensures proper tolerancing and concentricity are achieved. In some examples, combustion liner10can be additively manufactured from a nickel-based superalloy to ensure the mechanical and thermal properties of combustion liner10can withstand the harsh operating environment within gas turbine engine100. In some examples, the nickel-based superalloy can be one or more of Hastelloy X and Inconel.

In addition, combustion liner10can be additively manufactured to include integral aerodynamic features such as fuel channel18. Fuel channels18are oriented in axial direction AX such that fuel channels18are axially parallel with the cooling airflow traversing outer surface28of combustion liner10. The parallel orientation of fuel channels18reduces disruptions to the cooling airflow, allowing the cooling air to remove heat from outer wall16of combustion liner10. Further, fuel channels18are integrally formed with outer wall16such that fuel channels18are partially within outer surface28and partially outside outer surface28to achieve desirable pre-heating of the fuel flowing through fuel channels18. The positioning and orientation of fuel channels18improves burn characteristics of the fuel within combustion liner10and improves cooling characteristics of outer surface28of combustion liner10, compared to traditional circular tube fuel manifolds positioned outside combustion liner10and coupled to combustion liner10through a plurality of brackets secured to outer surface28of combustion liner10. As such, integrally formed fuel channels18within combustion liner10improve the overall performance characteristics of gas turbine engine100.

Discussion of Possible Embodiments

A combustion liner for use in a gas turbine engine, the combustion liner comprising: an inner wall, an outer wall, and a dome, wherein the outer wall is spaced radially outward from the inner wall, and wherein the dome extends between and couples the inner wall to the outer wall; a fuel channel formed as a single piece with the outer wall, wherein a first distal end of the fuel channel receives fuel from a fuel source; and a nozzle positioned at a second distal end of the fuel channel, wherein the nozzle is fluidly coupled to the fuel channel.

The combustion liner of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components in any combination:

The fuel channel is positioned partially within and partially outside an outer surface of the outer wall, and wherein the fuel channel extends in an axial direction of the combustion liner.

The first distal end of the fuel channel extends axially outward from a first end face of the outer wall.

The portion of the fuel channel extending axially outward from the first end face of the outer wall comprises a teardrop shaped cross sectional area in an axial direction.

The fuel channel comprises internal geometry configured to induce swirling of the fuel flowing through a bore of the fuel channel.

The internal geometry of the fuel channel is a helical groove positioned within the bore of the fuel channel, and wherein the helical groove extends from the first distal end to the second distal end of the fuel channel.

The combustion liner comprises a plurality of fuel channels, and wherein the plurality of fuel channels are equally spaced about a circumference of the outer wall of the combustion liner.

A splash plate coupled to the fuel channel, wherein the splash plate is spaced from and positioned parallel with an end face of the fuel channel.

The splash plate is axially spaced from the nozzle, and wherein the splash plate is configured to interrupt a flow of fuel dispensing from the nozzle.

The combustion liner comprises a plurality of nozzles and a plurality of splash plates, and wherein the plurality of nozzles and the plurality of splash plates are equally spaced about a circumference of the outer wall of the combustion liner.

A dilution chute positioned adjacent the second distal end of the fuel channel and adjacent the nozzle, wherein the dilution chute extends partially within and partially outside the outer wall of the combustion liner.

The combustion liner comprises a plurality of dilution chutes, and wherein the plurality of dilution chutes are equally spaced about a circumference of the outer wall of the combustion liner.

The portion of dilution chute extending outside of the outer wall is semi-circular in shape; and the portion of dilution chute extending within the outer wall extends into an interior of the combustion liner at an acute angle.

A support root extends from the second distal end of the fuel channel to the dilution chute, coupling the second distal end of the fuel channel to the dilution chute.

The combustion liner is of unitary single-piece construction, such that the combustion liner is constructed in a single, continuous manufacturing process.

The combustion liner is constructed from a nickel-based superalloy.

The following are further non-exclusive descriptions of possible embodiments of the present invention.

A method of using a combustion liner in a gas turbine engine, the method comprising: injecting fuel from a fuel source into a fuel channel, wherein the fuel channel is formed as a single piece with an outer wall of the combustion liner; flowing the fuel from a first distal end to a second distal end of the fuel channel, wherein the fuel is pre-heated as the fuel flows through the fuel channel; dispensing the pre-heated fuel through a nozzle into a dilution chute; and flowing the pre-heated fuel through the dilution chute into an interior of the combustion liner.

Dispensing the pre-heated fuel through the nozzle onto a splash plate to atomize the pre-heated fuel before entering the dilution chute.

Inducing swirl, through internal geometry within a bore of the fuel channel, to the fuel as the fuel flows from the first distal end to the second distal end of the fuel channel.

The internal geometry of the fuel channel is a helical groove positioned within the bore of the fuel channel, and wherein the helical groove extends from the first distal end to the second distal end of the fuel channel.